Showing papers in "Journal of Turbomachinery-transactions of The Asme in 2003"

Large Eddy Simulation of Film Cooling Flow From an Inclined Cylindrical Jet

Abstract: Predictions of turbine blade film cooling have traditionally employed Reynolds-averaged Navier-Stokes solvers and two-equation models for turbulence. Evaluation of several versions of such models have revealed that the existing two-equation models fail to resolve the anisotropy and the dynamics of the highly complex flow field created by the jet-crossflow interaction. A more accurate prediction of the flow field can be obtained from large eddy simulations (LES) where the dynamics of the larger scales in the flow are directly resolved. In the present paper, such an approach has been used, and results are presented for a row of inclined cylindrical holes at blowing ratios of 0.5 and I and Reynolds numbers of 11,100 and 22,200, respectively, based on the jet velocity and hole diameter. Comparison of the time-averaged LES predictions with the flow measurements of Lavrich and Chiappetta (UTRC Report No. 90-04) shows that LES is able to predict the flow field with reasonable accuracy. The unsteady three-dimensional flow field is shown to be dominated by packets of hairpin-shaped vortices. The dynamics of the hairpin vortices in the wake region of the injected jet and their influence on the unsteady wall heat transfer are presented. Generation of hot spots and their migration on the film-cooled surface are associated with the entrainment induced by the hairpin structures. Several geometric properties of a mixing interface around hairpin coherent structures are presented to illustrate and quantify their impact on the entrainment rates and mixing processes in the wake region.

Mistuning Identification of Bladed Disks Using a Fundamental Mistuning Model—Part I: Theory

Abstract: This paper is the first in a two-part study of identifying mistuning in bladed disks. It develops a new method of mistuning identification based on measurements of the vibratory response of the system as a whole. As a system-based method, this approach is particularly suited to integrally bladed rotors, whose blades cannot be removed for individual measurements. The method is based on a recently developed reduced order model of mistuning called the Fundamental Mistuning Model, FMM, and is applicable to isolated families of modes. Two versions of FMM system identification are presented: a basic version that requires some prior knowledge of the system’s properties, and a somewhat more complex version that determines the mistuning completely from experimental data.Copyright © 2003 by ASME

Predictions of Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions Using an Intermittency Transport Equation

Abstract: A new transport equation for the intermittency factor was proposed to predict separated and transitional boundary layers under low-pressure turbine airfoil conditions. The intermittent behavior of the transitional flows is taken into account and incorporated into computations by modifying the eddy viscosity, t , with the intermittency factor, y. Turbulent quantities are predicted by using Menter s two-equation turbulence model (SST). The intermittency factor is obtained from a transport equation model, which not only can reproduce the experimentally observed streamwise variation of the intermittency in the transition zone, but also can provide a realistic cross-stream variation of the intermittency profile. In this paper, the intermittency model is used to predict a recent separated and transitional boundary layer experiment under low pressure turbine airfoil conditions. The experiment provides detailed measurements of velocity, turbulent kinetic energy and intermittency profiles for a number of Reynolds numbers and freestream turbulent intensity conditions and is suitable for validation purposes. Detailed comparisons of computational results with experimental data are presented and good agreements between the experiments and predictions are obtained.

Active Flow Separation Control of a Stator Vane Using Embedded Injection in a Multistage Compressor Experiment

Abstract: Micro-flow control actuation embedded in a stator vane was used to successfully control separation and improve near stall performance in a multistage compressor rig at NASA Glenn. Using specially designed stator vanes configured with internal actuation to deliver pulsating air through slots along the suction surface, a research study was performed to identify performance benefits using this microflow control approach. Pressure profiles and unsteady pressure measurements along the blade surface and at the shroud provided a dynamic look at the compressor during microflow air injection. These pressure measurements lead to a tracking algorithm to identify the onset of separation. The testing included steady air injection at various slot locations along the vane. The research also examined the benefit of pulsed injection and actively controlled air injection along the stator vane. Two types of actuation schemes were studied, including an embedded actuator for on-blade control. Successful application of an online detection and flow control scheme will be discussed. Testing showed dramatic performance benefit for flow reattachment and subsequent improvement in diffusion through the use of pulsed controlled injection. The paper will discuss the experimental setup, the blade configurations, and preliminary CFD results which guided the slot location along the blade. The paper will also show the pressure profiles and unsteady pressure measurements used to track flow control enhancement, and will conclude with the tracking algorithm for adjusting the control.

Separation Control on Low-Pressure Turbine Airfoils Using Synthetic Vortex Generator Jets

Abstract: Oscillating vortex generator jets have been used to control boundary layer separation from the suction side of a low-pressure turbine airfoil. A low Reynolds number (Re = 25,000) case with low free-stream turbulence has been investigated with detailed measurements including profiles of mean and fluctuating velocity and turbulent shear stress. Ensemble averaged profiles are computed for times within the jet pulsing cycle, and integral parameters and local skin friction coefficients are computed from these profiles. The jets are injected into the mainflow at a compound angle through a spanwise row of holes in the suction surface. Preliminary tests showed that the jets were effective over a wide range of frequencies and amplitudes. Detailed tests were conducted with a maximum blowing ratio of 4.7 and a dimensionless oscillation frequency of 0.65. The outward pulse from the jets in each oscillation cycle causes a disturbance to move down the airfoil surface. The leading and trailing edge celerities for the disturbance match those expected for a turbulent spot. The disturbance is followed by a calmed region. Following the calmed region, the boundary layer does separate, but the separation bubble remains very thin. Results are compared to an uncontrolled baseline case in which the boundary layer separated and did not reattach, and a case controlled passively with a rectangular bar on the suction surface. The comparison indicates that losses will be substantially lower with the jets than in the baseline or passively controlled cases.Copyright © 2003 by ASME

A Method for Use of Cyclic Symmetry Properties in Analysis of Nonlinear Multiharmonic Vibrations of Bladed Discs

Abstract: An effective method for analysis of periodic forced response of nonlinear cyclically symmetric structures has been developed. The method allows multiharmonic forced response to be calculated for a whole bladed disk using a periodic sector model without any loss of accuracy in calculations and modeling. A rigorous proof of the validity, of the reduction of the whole nonlinear structure to a sector is provided. Types of bladed disk forcing for which the method may be applied are formulated. A multiharmonic formulation and a solution technique for equations of motion have been derived for two cases of description for a linear part of the bladed disk model: (i) using sector finite element matrices and (ii) using sector mode shapes and frequencies. Calculations validating the developed method and a numerical investigation of a realistic high-pressure turbine bladed disk with shrouds have demonstrated the high efficiency of the method

Heat Transfer Coefficients on the Squealer Tip and Near-Tip Regions of a Gas Turbine Blade With Single or Double Squealer

Abstract: Detailed heat transfer coefficient distributions on a gas turbine squealer tip blade were measured using a hue detection based transient liquid crystals technique. The heat transfer coefficients on the shroud and near tip regions of the pressure and suction sides of a blade were also measured. Squealer rims were located along (a) the camber line, (b) the pressure side, (c) the suction side, (d) the pressure and suction sides, (e) the camber line and the pressure side, and (f) the camber line and the suction side, respectively. Tests were performed on a five-bladed linear cascade with a blow down facility. The Reynolds number based on the cascade exit velocity and the axial chord length of a blade was 1.1×106 and the overall pressure ratio was 1.2. Heat transfer measurements were taken at the three tip gap clearances of 1.0%, 1.5% and 2.5% of blade span. Results show that the heat transfer coefficients on the blade tip and the shroud were significantly reduced by using a squealer tip blade. Results also showed that a different squealer geometry arrangement changed the leakage flow path and resulted in different heat transfer coefficient distributions. The suction side squealer tip provided the lowest heat transfer coefficient on the blade tip and near tip regions compared to the other squealer geometry arrangements.Copyright © 2003 by ASME

Impact of Geometric Variability on Axial Compressor Performance

Abstract: A probabilistic methodology to quantify the impact of geometric variability on compressor aerodynamic performance is presented. High-fidelity probabilistic models of geometric variability are derived using a Principal-Component-Analysis (PCA) of blade surface measurements. This probabilistic blade geometry model is then combined with a compressible, viscous blade-passage analysis to estimate the impact on the passage loss and turning using a Monte Carlo simulation. Finally, a mean-line multi-stage compressor model, with probabilistic loss and turning models from the blade-passage analysis, is developed to quantify the impact of the blade variability on overall compressor efficiency and pressure ratio. The methodology is applied to a flank-milled Integrally-Bladed Rotor (IBR). Results demonstrate that overall compressor efficiency can be reduced by approximately 1% due to blade-passage effects arising from representative manufacturing variability.Copyright © 2003 by ASME

Review of Centrifugal Compressor’s Application and Development

Abstract: The paper historically describes the application and development of the centrifugal compressor from the very beginning of its introduction until today. It focuses on selected practical and theoretical examples that — to the author’s opinion — pushed the centrifugal’s standard from simple, low efficiency designs to its current high level status. The main events related with this development like the impact of the industrial revolution and the introduction of jet propulsion are pointed out. The implication of improved theoretical tools becoming available with raising computer capacity and the impetus of advanced measurement techniques on the centrifugal’s improvement are described. A considerable number of references offers the possibility to engross the thoughts.Copyright © 2003 by ASME

Active Control of Tip Clearance Flow in Axial Compressors

Abstract: Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.

Effect of Internal Coolant Crossflow on the Effectiveness of Shaped Film-Cooling Holes

Abstract: Film-cooling was the subject of numerous studies during the past decades. However, the flow conditions on the entry side of the film-cooling hole have not received much attention up to now. A stagnant plenum which is widely used in experimental and numerical studies to feed the holes is not necessarily a right means to represent real engine conditions. For this reason, the present paper reports on an experimental study investigating the effect of a coolant crossflow feeding the holes that is oriented perpendicular to the hot gas flow direction to model a flow situation that is, for instance, of common use in the mid-portion of turbine blades. A comprehensive set of experiments was performed to evaluate the effect of perpendicular coolant supply direction on film-cooling effectiveness over a wide range of blowing ratios (M = 0.5{hor{underscore}ellipsis}2.0) and coolant crossflow Mach numbers (Ma{sub c} = 0{hor{underscore}ellipsis}0.6). The coolant to hot gas density ratio, however, was kept constant at 1.85 which can be assumed to be representative for typical gas turbine applications. Three different hole geometries, including a cylindrical hole as well as two holes with expanded exits, were considered. Particularly, two-dimensional distributions of local film-cooling effectiveness acquired by means of anmore » infrared camera system were used to give detailed insight into the governing flow phenomena. The results of the present investigation show that there is a profound effect of how the coolant is supplied to the hole on the film-cooling performance in the near hole region. Therefore, crossflow at the hole entry side must be taken into account when modeling film-cooling at engine representative conditions.« less

Passive Flow Control on Low-Pressure Turbine Airfoils

Abstract: Two-dimensional rectangular bars have been used in an experimental study to control boundary layer transition and reattachment under low-pressure turbine conditions. Cases with Reynolds numbers (Re) ranging from 25,000 to 300,000 (based on suction surface length and exit velocity) have been considered at low (0.5%) and high (8.5% inlet) free-stream turbulence levels. Three different bars were considered, with heights ranging from 0.2% to 0.7% of suction surface length. Mean and fluctuating velocity and intermittency profiles are presented and compared to results of baseline cases from a previous study. Bar performance depends on the bar height and the location of the bar trailing edge. Bars located near the suction surface velocity maximum are most effective. Large bars trip the boundary layer to turbulent and prevent separation, but create unnecessarily high losses. Somewhat smaller bars had no immediate detectable effect on the boundary layer, but introduced small disturbances that caused transition and reattachment to move upstream from their locations in the corresponding baseline case. The smaller bars were effective under both high and low free-stream turbulence conditions, indicating that the high free-stream turbulence transition is not simply a bypass transition induced by the free stream. Losses appear to be minimized when a small separation bubble is present, so long as reattachment begins far enough upstream for the boundary layer to recover from the separation. Correlations for determining optimal bar height are presented. The bars appear to provide a simple and effective means of passive flow control. Bars that are large enough to induce reattachment at low Re, however, cause higher losses at the highest Re. Some compromise would, therefore, be needed when choosing a bar height for best overall performance. @DOI: 10.1115/1.1626685#

Generic Friction Models for Time-Domain Vibration Analysis of Bladed Disks

Abstract: New efficient models have been developed to describe dynamic friction effects in order to facilitate analysis of the vibration of bladed discs in the time domain. These friction models describe friction forces occurring at contact interfaces under time-varying normal load variations, including cases of separation. The friction models developed allow one to take into account time-varying friction contact parameters, such as friction coefficient and contact stiffness coefficients. Anisotropy and variation of the friction characteristics over the contact surfaces are included in the proposed models. The capabilities of the new friction models are demonstrated. Analysis of the friction forces is performed for different motion trajectories and different time variations of the normal load, and the effects of anisotropy, variation in time of the friction characteristics and normal load variation are discussed. A numerical analysis of transient vibrations of shrouded blades using the new models is presented.Copyright © 2003 by ASME

Experimental and Numerical Investigation of Impingement on a Rib-Roughened Leading-Edge Wall

Abstract: Effective cooling of the airfoil leading-edge is imperative in gas turbine designs. Amongst several methods of cooling the leading edge, impingement cooling has been utilized in many modern designs. In this method, the cooling air enters the leading edge cavity from the adjacent cavity through a series of crossover holes on the partition wall between the two cavities. The crossover jets impinge on a smooth leading-edge wall and exit through the film holes, and, in some cases, form a crossflow in the leading-edge cavity and move toward the end of the cavity. It was the main objective of this investigation to measure the heat transfer coefficient on a smooth as well as rib-roughened leading-edge wall. Experimental data for impingement on a leading edge surface roughened with different conical bumps and radial ribs are reported by the same authors, previously. This investigation, however, deals with impingement on different horseshoe ribs and makes a comparison between the experimental and numerical results. Three geometries representing the leading-edge cooling cavity of a modern gas turbine airfoil with crossover jets impinging on 1) a smooth wall, 2) a wall roughened with horseshoe ribs, and 3) a wall roughened with notched-horseshoe ribs were investigated. The tests were run for a range of flow arrangements and jet Reynolds numbers. The major conclusions of this study were: a) Impingement on the smooth target surface produced the highest overall heat transfer coefficients followed by the notched-horseshoe and horseshoe geometries. b) There is, however, a heat transfer enhancement benefit in roughening the target surface. Amongst the three target surface geometries, the notched-horseshoe ribs produced the highest heat removal from the target surface which was attributed entirely to the area increase of the target surface. c) CFD could be considered as a viable tool for the prediction of impingement heat transfer coefficients on an airfoil leading-edge wall.Copyright © 2003 by ASME

Mistuning Identification of Bladed Disks Using a Fundamental Mistuning Model: Part 2 — Application

Abstract: This paper is the second in a two-part study of identifying mistuning in bladed disks. It presents experimental validation of a new method of mistuning identification based on measurements of the vibratory response of the system as a whole. As a system based method, this approach is particularly suited to integrally bladed rotors, whose blades cannot be removed for individual measurements. The method is based on a recently developed reduced order model of mistuning called the Fundamental Mistuning Model and is applicable to isolated families of modes. Two versions of FMM system identification are applied to the experimental data: a basic version that requires some prior knowledge of the system’s properties, and a somewhat more complex version that determines the mistuning completely from experimental data.Copyright © 2003 by ASME

Unsteady Surface Pressures Due to Wake-Induced Transition in a Laminar Separation Bubble on a Low-Pressure Cascade

Abstract: This paper presents unsteady surface pressures measured on the suction surface of a LP turbine cascade that was subject to wake passing from a moving bar wake generator. The surface pressures measured under the laminar boundary layer upstream of the steady flow separation point were found to respond to the wake passing as expected from the kinematics of wake convection. In the region where a separation bubble formed in steady flow, the arrival of the convecting wake produced high frequency, short wavelength, fluctuations in the ensemble averaged blade surface pressure. The peak-to-peak magnitude was 30% of the exit dynamic head. The existence of fluctuations in the ensemble averaged pressure traces indicates that they are deterministic and that they are produced by coherent structures. The onset of the pressure fluctuations was found to lie beneath the convecting wake and the fluctuations were found to convect along the blade surface at half of the local freestream velocity. Measurements performed with the boundary layer tripped ahead of the separation point showed no oscillations in the ensemble average pressure traces indicating that a separating boundary layer is necessary for the generation of the pressure fluctuations. The coherent structures responsible for the large amplitude pressure fluctuations were identified using PIV to be vortices embedded in the boundary layer. It is proposed that these vortices form in the boundary layer as the wake passes over the inflexional velocity profiles of the separating boundary layer and that the rollup of the separated shear layer occurs by an inviscid Kelvin-Helmholtz mechanism.Copyright © 2003 by ASME

Effect of Endwall Motion on Blade Tip Heat Transfer

Abstract: Local mass transfer measurements were conducted on the tip of a turbine blade in a five-blade linear cascade with a blade-centered configuration. The tip clearance levels ranged from 0.6 to 6.9% of blade chord. The effect of relative motion between the casing and the blade tip was simulated using a moving endwall made of neoprene mounted on the top of the wind tunnel. Data were obtained for a single Reynolds number of 2.7 ×10 5 based on cascade exit velocity and blade chord. Pressure measurements indicate that the effect of endwall motion on blade loading at a clearance of 0.6% of blade chord is to reduce the pressure gradients driving the tip leakage flow. With the introduction of endwall motion, there is a reduction of about 9% in mass transfer levels at a clearance of 0.6% of chord. This is presumably due to the tip leakage vortex coming closer to the suction side of the blade and 'blocking the flow,' leading to reduced tip gap velocities and hence lower mass transfer.

Turbine Blade Trailing Edge Flow Characteristics at High Subsonic Outlet Mach Number

Abstract: The paper presents an experimental investigationof the effect of the trailing edge vortex shedding on the steady and unsteady trailing blade pressure distribution of a turbine blade at high subsonic Mach number (M 2,is =0.79) and high Reynolds number (RE =2.8×10 6 ). The vortex formation and shedding process is visualized using a high-speed schlieren camera and a holographic interferometric density measuring technique. The blade is equipped with a rotatable trailing edge cylinder instrumented side-by-side with a pneumatic pressure tap and a fast response pressure sensor for detailed measurements of the trailing edge pressure distribution. The experiments demonstrate that contrary to the isobaric dead air region demonstrated at low subsonic Mach numbers the data reveal the existence of a highly nonuniform trailing edge pressure distribution with a strong pressure minimum at the center of the trailing edge. This finding is significant for the determination of the base pressure coefficient that is in general measured with a single pressure-sensing hole at the trailing edge center. The paper investigates further the effect of the vortex shedding on the blade rear suction side and discusses the superposition of unsteady effects emanating from the trailing edge and from the neighboring blade. The experimental data are a unique source for the validation of unsteady Navier-Stokes codes.

Advanced High Turning Compressor Airfoils for Low Reynolds Number Condition: Part 1 — Design and Optimization

Abstract: High performance compressor airfoils at a low Reynolds number condition at ~Re51.3310 5 ! have been developed using evolutionary algorithms in order to improve the performance of the outlet guide vane (OGV), used in a single low pressure turbine (LPT) of a small turbofan engine for business jet aircrafts. Two different numerical optimization methods, the evolution strategy (ES) and the multi-objective genetic algorithm (MOGA), were adopted for the design process to minimize the total pressure loss and the deviation angle at the design point at low Reynolds number condition. Especially, with respect to the MOGA, robustness against changes of the incidence angle is considered. The optimization process includes the representation of the blade geometry, the generation of a numerical grid and a blade-to-blade analysis using a quasi-three-dimensional Navier-Stokes solver with a k-v turbulence model including a newly implemented transition model to evaluate the performance. Overall aerodynamic performance and boundary layer properties for the two optimized blades are discussed numerically. The superior performance of the two optimized airfoils is demonstrated by a comparison with conventional controlled diffusion airfoils (CDA). The advantage in performance has been confirmed by detailed experimental investigations, which are presented in Part II of this paper. @DOI: 10.1115/1.1737780#

Time-Resolved Vane-Rotor Interaction in a High-Pressure Turbine Stage

Abstract: This paper describes the time-varying aerodynamic interaction mechanisms that have been observed within a transonic high-pressure turbine stage; these are inferred from the time-resolved behavior of the rotor exit flow field. It contains results both from an experimental program in a turbine test facility and from numerical predictions. Experimental data was acquired using a fast-response aerodynamic probe capable of measuring Mach number, whirl angle, pitch angle, total pressure, and static pressure. A 3-D time-unsteady viscous Navier-Stokes solver was used for the numerical predictions. The unsteady rotor exit flowfield is formed from a combination of four flow phenomena: the rotor wake, the rotor trailing edge recompression shock, the tip-leakage flow, and the hub secondary flow. This paper describes the time-resolved behavior of each phenomenon and discusses the interaction mechanisms from which each originates. Two significant vane periodic changes (equivalent to a time-varying flow in the frame of reference of the rotor) in the rotor exit flowfield are identified. The first is a severe vane periodic fluctuation in flow conditions close to the hub wall and the second is a smaller vane periodic fluctuation occurring at equal strength over the entire blade span. These two regions of periodically varying flow are shown to be caused by two groups of interaction mechanisms. The first is thought to be caused by the interaction between the wake and secondary flow of the vane with the downstream rotor; and the second is thought to be caused by a combination of the interaction of the vane trailing edge recompression shock with the rotor, and the interaction between the vane and rotor potential fields.

Performance of Strongly Bowed Stators in a 4-Stage High Speed Compressor

Abstract: The FVV-sponsored-Project “Bow Blading” (c.f. acknowledgments) at the Turbomachinery Laboratory of the University of Hannover addresses the effect of strongly bowed stator vanes on the flow field in an 4-stage high speed axial compressor with controlled diffusion airfoil (CDA) blading. The compressor is equipped with more strongly bowed vanes than have previously been reported in the literature. The performance map of the present compressor is being investigated experimentally and numerically. The results show that the pressure ratio and the efficiency at the design point and at the choke limit are reduced by the increase in friction losses on the surface of the bowed vanes, whose surface area is greater than that of the reference (CDA) vanes. The mass flow at the choke limit decreases for the same reason. Because of the change in the radial distribution of axial velocity, pressure rise shifts from stage 3 to stage 4 between the choke limit and maximum pressure ratio. Beyond the point of maximum pressure ratio, this effect is not distinguishable from the reduction of separation by the bow of the vanes. Experimental results show that in cases of high aerodynamic loading, i.e. between maximum pressure ratio and the stall limit, separation is reduced in the bowed stator vanes so that the stagnation pressure ratio and efficiency are increased by the change to bowed stators. It is shown that the reduction of separation with bowed vanes leads to a increase of static pressure rise towards lower mass flow so that the present bow bladed compressor achieves higher static pressure ratios at the stall limit.Copyright © 2003 by ASME

Experimental and Numerical Investigation of Trailing Edge Film Cooling by Circular Coolant Wall Jets Ejected From a Slot With Internal Rib Arrays

Abstract: The present study concentrates on the experimental and computational investigation of a cooled trailing edge in a modern turbine blade. The trailing edge features a pressure side cutback and a slot, stiffened by two rows of evenly spaced ribs in an inline configuration. Cooling air is ejected through the slot and forms a cooling film on the trailing edge cutback region. In the present configuration the lateral spacing of the ribs equals two times their width. The height of the ribs, i.e. the height of the slot equals their width. Since the ribs are provided with fillet radii of half the slot height in size, circular coolant jets are exiting the slot tangentially to the trailing edge cutback. The adiabatic wall temperature mappings on the trailing edge cutback indicate that strong three-dimensional flow interaction between the coolant jets and the hot main flow takes place in such a way that two or more coolant jets coalesce depending on the blowing ratio. Experimental and numerical data to be presented in the present study include adiabatic film cooling effectiveness on the trailing edge cutback, the pressure distribution along the internal ribbed passage as well as slot discharge coefficients for different blowing ratios ranging from M = 0.35 to 1.1.Copyright © 2003 by ASME

Experimental and Numerical Study of Stall Flutter in a Transonic Low-Aspect Ratio Fan Blisk

Abstract: Experiments are performed on a modern design transonic shroudless low-aspect ratio fan blisk that experienced both subsonic/transonic and supersonic stall-side flutter. High-response flush mounted miniature pressure transducers are utilized to measure the unsteady aerodynamic loading distribution in the tip region of the fan for both flutter regimes, with strain gages utilized to measure the vibratory response at incipient and deep flutter operating conditions. Numerical simulations are performed and compared with the benchmark data using an unsteady three-dimensional nonlinear viscous computational fluid dynamic (CFD) analysis, with the effects of tip clearance, vibration amplitude, and the number of time steps-per-cycle investigated. The benchmark data are used to guide the validation of the code and establish best practices that ensure accurate flutter predictions.Copyright © 2003 by ASME

Heat Transfer and Friction Factors for Flows Inside Circular Tubes With Concavity Surfaces

Abstract: Heat transfer and friction coefficients measurements have been obtained for fully developed, turbulent internal flows in circular tubes with six different concavity (dimple) surface array geometries. Two different concavity depths and three different concavity array densities were tested using tube bulk flow Reynolds numbers from 20,000 to 90,000. Liquid Crystal Thermography was used to measure the temperature distributions on the outside of the concavity tubes. Using the average heat transfer coefficient for the fully developed region, the overall heat transfer enhancements are compared to baseline smooth tube results. Friction coefficients are also compared to values for a smooth tube. Dimple depths of 0.2 to 0.4 relative to the dimple surface diameter were used, with surface area densities ranging from 0.3 to 0.7. Dimple arrays were all in-line geometries. The results showed that heat transfer enhancements for dimpled internal surfaces of circular passages can reach factors of 2 or more when the relative dimple depth is greater than 0.3 and the dimple array density is about 0.5 or higher. The associated friction factor multipliers for such configurations are in the range of 4 to 6. The present study provides a first insight into the heat transfer and friction effects of various concavity arrays for turbulent flows.Copyright © 2003 by ASME

Aerodynamic Blade Row Interactions in an Axial Compressor: Part I — Unsteady Boundary Layer Development

Abstract: This two-part paper presents experimental investigations of unsteady aerodynamic blade row interactions in the first stage of the four-stage Low-Speed Research Compressor of Dresden. Both the unsteady boundary layer development and the unsteady pressure distribution of the stator blades are investigated for several operating points. The measurements were carried out on pressure side and suction side at midspan. In part I of the paper the investigations of the unsteady boundary layer behaviour are presented. The experiments were carried out using surface-mounted hot-film sensors. Additional information on the time-resolved flow between the blade rows were obtained with a hot-wire probe. The unsteady boundary layer development is strongly influenced by the incoming wakes. Within the predominantly laminar boundary layer in the front part of the blade a clear response of the boundary layer to the velocity and turbulence structure of the incoming wakes can be observed. The time-resolved structure of the boundary layer for several operating points of the compressor is analyzed in detail. The topic “calmed regions”, which can be coupled to the wake passing, is discussed. As a result an improved description of the complex boundary layer structure is given.Copyright © 2003 by ASME

Unsteady Turbine Blade Wake Characteristics

Abstract: The paper presents an experimental investigation of large coherent structures, commonly referred to as “von Karman vortex street”, in the wake of a turbine blade at high subsonic Mach number (M2,is = 0.79) and high Reynolds number (RE = 2.8×106 and their effect on the steady and unsteady pressure and temperature distribution in the wake. Ultra short smoke visualizations and two interferometric measurement techniques, holographic interferometry and white light differential interferometry provide insight into the vortex formation and shedding process. In addition, the interferometric measurement provides quantitative information on the stream wise evolution of the minimum density associated with the vortices and on their lateral spreading. Wake traverses are performed with a four-head fork probe carrying a Kiel probe and a fast response Kulite pressure probe for pressure measurements and a thermo-couple probe and a cold wire resistance probe for temperature measurements. The results confirm the observation of energy separation in the wake as found by other researchers. The experimental data are a unique source for the validation of unsteady Navier-Stokes codes.© 2003 ASME

Design of Industrial Axial Compressor Blade Sections for Optimal Range and Performance

Abstract: A design system for the blade sections of industrial axial compressors has been developed. The method combines a parametric geometry definition method, a powerful blade-to-blade flow solver (MISES) and an optimization technique (breeder genetic algorithm) with an appropriate fitness function. Particular effort has been devoted to the design of the fitness function for this application which includes non-dimensional terms related to the required performance at design and off-design operating points. It has been found that essential aspects of the design (such as the required flow turning, or mechanical constraints) should not be part of the fitness function, but need to be treated as so-called “killer” criteria in the genetic algorithm. Finally, it has been found worthwhile to examine the effect of the weighting factors of the fitness function to identify how these affect the performance of the sections. The system has been tested on the design of a repeating stage for the middle stages of an industrial axial compressor. The resulting profiles show an increased operating range compared to an earlier design using NACA65 profiles.Copyright © 2003 by ASME

Effect of Unsteady Stator Wake—Rotor Double-Leakage Tip Clearance Flow Interaction on Time-Average Compressor Performance

Abstract: A study has been conducted, using unsteady three-dimensional Reynolds-averaged Navier-Stokes simulations to determine the impact on rotor performance of the interaction between upstream (steady defect and time-varying defect) stator wakes and rotor tip clearance flow. The key effects of the interaction between steady stator wakes and rotor tip clearance flow are: 1) a decrease in loss and blockage associated with tip clearance flow; 2) an increase in passage static pressure rise. Performance benefit is seen in the operability range from near design to high loading. The benefit is modest near design and increases with loading. Significant beneficial changes due to the stator-rotor interaction occur when the phenomenon of tip clearance flow double-leakage is present. Double-leakage occurs when the tip clearance flow passes through the tip gap of the adjacent blade. It is detrimental for compressor performance. The effect of strong stator-rotor interaction is to suppress double-leakage on a time-average basis. Double-leakage typically takes place at high loading but can be present at design condition as well, for modern highly loaded compressor. A benefit due to unsteady interaction is also observed in the operability range of the rotor. A new generic causal mechanism is proposed to explain the observed changes in performance. It identifies the interaction between the trip clearance flow and the pressure pulses, induced on the rotor blade pressure surface by the upstream wakes, as the cause for the observed effects. The direct effect of the interaction is a decrease in the time-average double-leakage flow through the tip clearance gap so that the stream-wise defect of the exiting tip flow is lower with respect to the main flow. A lower defect leads to a decrease in loss and blockage generation and hence an enhanced performance compared to that in the steady situation. The performance benefits increase monotonically with loading and scale linearly with upstream wake velocity defect. With oscillating defect stator wakes, rotor performance shows dependence on oscillation frequency. Changes in the tip region occur at a particular reduced frequency leading to (1) decrease in blockage, and (2) increase in passage loss. The changes in rotor performance at a particular reduced frequency are hypothesized to be associated with the inherent unsteadiness of the tip clearance vortex and its resonance behavior excited by the oscillating wakes.

A Technique for Processing Transient Heat Transfer, Liquid Crystal Experiments in the Presence of Lateral Conduction

Abstract: New techniques for processing transient liquid crystal heat transfer experiment have been developed. The methods are able to measure detailed local heat transfer coefficient and adiabatic wall temperature in a three temperature system from a single transient test using the full intensity history recorded. Transient liquid crystal processing methods invariably assume that lateral conduction is negligible and so the heat conduction process can be considered one dimensional into the substrate. However, in regions with high temperature variation such as immediately downstream of a film-cooling hole, it is found that lateral conduction can become significant. For this reason, a procedure which allows for conduction in three dimensions was developed by the authors. The paper is the first report of a means of correcting data from the transient heat transfer liquid crystal experiments for the effects of significant lateral conduction. The technique was applied to a film cooling system as an example and a detailed uncertainty analysis performed.Copyright © 2003 by ASME

Interaction of Shroud Leakage Flow and Main Flow in a Three-Stage LP Turbine

Abstract: Endwall losses significantly contribute to the overall losses in modern turbomachinery, especially when aerodynamic airfoil load and pressure ratios are increased In turbines with shrouded airfoils a large portion of these losses are generated by the leakage flow across the shroud clearance Generally the related losses can be grouped into losses of the leakage flow itself and losses caused by the interaction with the main flow in subsequent airfoil rows In order to reduce the impact of the leakage flow and shroud design related losses a thorough understanding of the leakage losses and especially of the losses connected to enhancing secondary flows and other main flow interactions has to be understood Therefore, a three stage LP turbine typical for jet engines is being investigated For the three-stage test turbine 3D Navier-Stokes computations are performed simulating the turbine including the entire shroud cavity geometry in comparison with computations in the ideal flow path Numerical results compare favorably against measurements carried out at the high altitude test facility at Stuttgart University The differences of the simulations with and without shroud cavities are analyzed for several points of operation and a very detailed quantitative loss breakdown is presented