bluuebunny/crossref_metadata_2025 huggingface dataset:
The Effect of Compound Angle on Nozzle Suction Side Film Cooling
Authors: Zhang Luzeng , Yin Juan , Moon Hee Koo
Published: August 2013

The film cooling hole compound angle effect on nozzle suction side film cooling effectiveness was experimentally investigated using a single row of shaped holes. Engine operating conditions were simulated in a scaled warm cascade, which was built based on industrial gas turbine nozzle vanes. Local film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Three separate nozzle test models were fabricated, which have the same cooling supply plenum configurations. The baseline configuration has a row of shaped holes on the suction surface without a compound angle. The other two test models have same size film holes at the same location, but one with a compound angle in co-flow and the other in counter-flow direction to the cooling supply. Four cooling mass flow ratios (MFR) were studied for each of the nozzle test models and two-dimensional film effectiveness distributions were measured. Then the film effectiveness distributions were spanwise averaged for comparison. The compound angle injections prevent jets lift-off for higher MFR cases, but for a lower MFR, the baseline configuration results in higher film effectiveness. The co-flow or counter-flow compound angle results in film shifting to upper or lower endwall, but the effect of the direction of the compound angle is less significant compared to the compound angle effect.
Pressure Losses for Jet Array Impingement With Crossflow
Authors: Levy Yeshayahou , Rao Arvind G. , Erenburg V. , Sherbaum V. , Gaissinski I. , Krapp V.
Published: August 2013

Jet impingement is an efficient heat transfer method and has been used successfully in cooling of turbine blades in gas turbine engines. Although many studies have been conducted on the heat transfer characteristics of jet impingement array, there is a lack of knowledge in pressure drop characteristics of large jet impingement arrays. The pressure losses encountered are becoming increasingly important when applied to micro gas turbines, cooling concentrated solar panels and high density electronic chips. The present work focuses on experimental and theoretical investigation of pressure losses in low Re impingement arrays, 200< Re <3000.

Experiments were carried out on jet impingement array with nozzle diameters of 200 to 800 μm. Numerical simulations were also performed with available commercial CFD tools. Reasonable comparisons between experimental results and numerical simulations were obtained. Detailed flow structure, mass flow rate distribution, jet velocity profiles, and pressure drop within the array in the streamwise direction were obtained from the CFD simulations. These simulations enhance the understanding of the physics within multiple jet impingement system.

Additionally a semi empirical–analytical method is developed for calculating the total pressure loss within a multi jet impingement system. This simple methodology can provide a quick estimate of the total pressure drop and hence is suited for first order optimization. The methodology is validated by results obtained from experiments and from CFD simulations.
Evaporation of Volatile Liquid Pools Under Forced Convection: Part 2—Integration of a Vaporization Model in a CFD Software
Authors: Doué Nicolas , Loueslati Karim , Alonso Dominique , Genin Ghislain , Molière Michel
Published: August 2013

The transfer under dynamic conditions of volatile species from a liquid pool to the surrounding air is gaining interest in the engineering community.

In particular, increasingly stringent regulations and standards apply to all types of flammable substances. This is especially the case in stationary gas turbine applications, where the vaporization of accidentally occurring pools of liquid fuels attracts increasing attention.

Since the flame-to-explosion transition cases are insufficiently controlled, the current, only practicable approach to assess the explosion risks arising from a fuel pool in an enclosure consists in quantifying the amount of vapor that leaves the pool and minimizing, by means of a proper dilution strategy, the potential damages entailed by the ignition of the resulting cloud. This approach requires two steps:

(1) The accurate assessment of vaporization rates under given ventilation conditions, a task that calls for skills in thermal and mass transfers.

(2) The reliable prediction of the transport of the vapors in the ventilation stream, a task specifically focused on fluid dynamics.

The three teams involved in this paper have joined their efforts to achieve this multidisciplinary objective. As a first task, the LRGP team (Laboratoire des Reactions et Génie des Procédés) and GE Energy have experimentally validated a vaporization model initially devised for water pools. This work has been reported in a recent paper. Concurrently, EURO/CFD and GE Energy have developed a CFD approach devoted to the mixing/dilution processes in defined enclosure geometries and under specified ventilation conditions. Finally both approaches havebeen coupled by EURO/CFD to produce predictive isopleth pictures of the vapor clouds generated under given temperature and velocity conditions.

The present paper covers the integration of the liquid pool vaporization model in thecommercial CFD software ANSYS Fluent and sets out the results obtained.

This dual, concerted approach is a first of the kind to the authors’ knowledge and proves fruitful for the prediction of the spatial distributions of the volatile species developed when a volatile pool vaporizes in a ventilated enclosure. It fills a gap in the analysis of safety scenarios arising from spillages of liquid fuels and provides a rational tool in zone classification studies.
Measurement of Thermal Contact Conductance of a Heat Shield for Gas Turbine Combustors in a Realistic Test Rig Set-Up
Authors: Facchini B. , Simonetti F. , Tarchi L. , Abram R. , Maritano M.
Published: August 2013

This paper presents the results of the experimental campaign on the measurement of thermal contact conductance (TCC) in a realistic test rig replicating the contact between the ceramic heat shield and the liner for an ANSALDO Energia s.p.a. heavy-duty gas turbine. The designed test rig is made up of a single sector consisting of a combustor ceramic brick and a liner sector with a non-rigid ceramic fiber textile insulating layer in between, to better replicate the operating conditions of the combustor components.

The experimental campaign was carried out with a steady-state technique applying an uniform thermal load on the heat shield surface. The effect of brick-holders, that in the actual combustor anchor the bricks on the liner walls, was replicated with static loads.

The liner sector was tested without the brick-holder grooves first and afterwards their effect was assessed in the same test conditions; the test campaign also aimed at measuring the load cycling effect on TCC. Validation of the measured TCC was then performed with a finite element code to compare the predicted temperature profiles with the measured ones.
3D Coupled Fluid-Solid Thermal Simulation of a Turbine Disc Through a Transient Cycle
Authors: Sun Zixiang , Chew John W. , Hills Nicholas J. , Barnes Christopher J. , Valencia Antonio Guijarro
Published: August 2013

Thermal analysis of a turbine disc through a transient test cycle is demonstrated using 3D computational fluid dynamics (CFD) modeling for the cooling flow and 3D finite element analysis (FEA) for the disc. The test case is a 3D angular sector of the high pressure (HP) turbine assembly of a civil jet engine and includes details of the coolant flow around the blade roots. Proprietary FEA and CFD solvers are used to simulate the metal and fluid domains, respectively. Coupling is achieved through an iterative loop with smooth exchange of information between the FEA and CFD simulations at each time step, ensuring consistency of temperature and heat flux on the coupled interfaces between the metal and fluid domains. The coupled simulation can be completed within a few weeks using a PC cluster with multiple parallel CFD executions. The FEA/CFD coupled result agrees well with corresponding rig test data and the baseline 3D and 2D FEA solutions, which have been calibrated using test data. Provision of upstream boundary conditions and modeling of rapid transients are identified as areas of uncertainty. Averaging of CFD solutions and relaxation is used to overcome difficulties caused by CFD oscillations associated with flow unsteadiness. The present work supports the continued use and development of the FEA/CFD coupling method for industrial applications.
The Optimal Distribution of Chordwise Rib Fin Arrays for Blade Tip Cap Underside Cooling
Authors: Ledezma Gustavo A. , Bunker Ronald S.
Published: August 2013

A fundamental question in the design of fin-augmented heat transfer surfaces is how to determine the optimal spacing between the fins. It has already been demonstrated that considerable heat transfer augmentation in the underside of an HPT turbine blade tip cap can be achieved using arrays of discrete shaped pins [5]. However, it is desirable to predict the maximum heat transfer augmentation that can be achieved by installing the array of fins and the geometric arrangement (fin-to-fin spacing) that has to be used to achieve such augmentation. In this paper chordwise parallel ribs installed on the underside of a blade tip cap are studied. The objective is to maximize the overall thermal conductance between the fin array and the surrounding fluid. The optimization is performed numerically in the range 25,000<Re<100,000 and Pr = 0.72. The behavior of the optimal spacing data is explained and correlated analytically using the method of intersecting the two asymptotes: small spacing and large spacing heat transfer [7].
Influence of Passage Flow on Tip Film Cooling Characteristics
Authors: Wang Jin , Yu Yong , Zeng M. , Wang Q. W.
Published: August 2013

Three-dimensional simulations of the squealer tip on the GE-E3 blade with eight film cooling holes were carried out. The effect of different blade spans and different blowing ratios on the tip flow and cooling performance was revealed with the k-ε model. For the squealer tip, the depth of the cavity and the height of the tip clearance were fixed, the influence of different spans (10%, 25%, 50%, 75% and 100% span) on the tip heat transfer was investigated. It was found that the velocity field above the blade tip and the heat transfer distribution on the groove floor for the 10% span (cut-back span) model had no difference from that for the 100% span (whole span) model obviously. However, the leakage flow for the 10% span model showed larger interaction with the passage flow. With different spans, the effect of different blowing ratios, i.e., M = 0.4, 0.8 and 1.2, was investigated. Increasing the blowing ratio (from M = 0.4 to 1.2) increased the film cooling effectiveness and made the heat transfer coefficients of all the models smaller. Because the cut-back model for the 10% span had similar tip flow field with the 100% span model, the simulation for the 10% span model could be used to find out the tip flow and heat transfer for the 100% span model.
Full Coverage Effusion Cooling With a Narrow Duct Backside Coolant Supply
Authors: Andrews G. E.
Published: August 2013

Full coverage effusion cooling was studied for a square array of 90° effusion cooling holes with backside cooling using a 5 mm duct air supply to the coolant holes. Active cooling was used with metal walls and 300K effusion cooling into a 27 m/s mean velocity duct flow at 770K crossflow temperature. The 152 mm square test section had 15 rows of holes and the hole diameter, D, was varied for constant hole pitch, X. The X/D values studied were 11.0, 7.0 and 4.6. At a constant coolant mass flow rate the wall pressure loss was reduced as X/D was reduced and there was an associated reduction in the film blowing ratio, M. The duct air feed to the holes enhanced the backside cooling of the wall. These results were compared with previous work using a plenum chamber air feed. The increased duct air feed velocity relative to the plenum low velocity air feed resulted in an increase in the overall cooling effectiveness due to the additional heat transfer by the duct crossflow velocity. However, the trailing edge cooling effectiveness improvement was small as there was no residual cross flow here and the greatest effect was at the leading edge of the test wall. The decrease in X/D was the most effective way of increasing the overall cooling effectiveness as this reduced the blowing rate without decreasing the coolant mass flow rate. This was more effective than using 30° inclined holes with an X/D of 11, as the hole exit velocity was much lower for the same coolant mass flow rate with 90°holes at an X/D of 4.8 than with 30°holes at an X/D of 11.0.
Leading Edge Shielding Concept in Gas Turbines With Can Combustors
Authors: Aslanidou Ioanna , Rosic Budimir , Kanjirakkad Vasudevan , Uchida Sumiu
Published: August 2013

The remarkable developments in gas turbine materials and cooling technologies have allowed a steady increase in combustor outlet temperature and hence in gas turbine efficiency over the last half century. However, the efficiency benefits of higher gas temperature, even at the current levels, are significantly offset by the increased losses associated with the required cooling. Additionally, the advancements in gas turbine cooling technology have introduced considerable complexities into turbine design and manufacture. Therefore, a reduction in coolant requirements for the current gas temperature levels is one possible way for gas turbine designers to achieve even higher efficiency levels.

The leading edges of the first turbine vane row are exposed to high heat loads. The high coolant requirements and geometry constraints limit the possible arrangement of the multiple rows of film cooling holes in the so called showerhead region. In the past, investigators have tested many different showerhead configurations, varying the number of rows, inclination angle and shape of the cooling holes. However the current leading edge cooling strategies using showerheads have not been shown to allow further increase in turbine temperature without excessive use of coolant air. Therefore new cooling strategies for the first vane have to be explored.

In gas turbines with multiple combustor chambers around the annulus, the transition duct walls can be used to shield, i.e. to protect the first vane leading edges from the high heat loads. In this way the stagnation region at the leading edge and the shower-head of film cooling holes can be completely removed, resulting in a significant reduction in the total amount of cooling air that is otherwise required. By eliminating the showerhead the shielding concept significantly simplifies the design and lowers the manufacturing costs.

This paper numerically analyses the potential of the leading edge shielding concept for cooling air reduction. The vane shape was modified to allow for the implementation of the concept and non-restrictive relative movement between the combustor and the vane. It has been demonstrated that the coolant flow that was originally used for cooling the combustor wall trailing edge and a fraction of the coolant air used for the vane showerhead cooling can be used to effectively cool both the suction and the pressure surfaces of the vane.
Prediction of De-Swirled Radial Inflow in Rotating Cavities With Hysteresis
Authors: May David , Chew John W. , Scanlon Timothy J.
Published: August 2013

De-swirl nozzles are sometimes used in turbomachinery to reduce the pressure drop when air is drawn radially inwards through a rotating cavity. However, this can lead to non-unique steady state solutions with operating conditions achieved depending on how the steady point is approached. In the present study, a transient, 1D model of flow in a rotating cavity has been created. The model allows the vortex profile to change with through flow rate, and links this to estimates of disk windage, tangential velocity and, consequently, the vortex pressure gradient. The model was applied to the simulation of de-swirl nozzle fed, rotating cavities with radial inflow. The steady vortex flow characteristics (non-dimensional flow versus pressure ratio) predicted by the model were validated for 2 distinct cases. For a smooth rectangular cavity the flow characteristic was predicted using the model’s default parameters. For an engine-representative case with non-axisymmetric geometric features, the flow characteristic of the cavity was reproduced with some alignment of the model.

The transient model reproduced experimentally observed hysteresis, discontinuity in operating characteristics, and regions where no steady-state solution could be found. A transient model is required as a steady state model would choose one of the possible solutions without physical justification. In the case of the engine-representative rig, part of the flow characteristic could not be obtained in testing. This is determined to be due to the interaction of the negative resistance region of the vortex and the flow modulating valve characteristic. Measures that allow the full capture of the flow characteristic in rig testing are identified.
Direct Numerical Simulations of Transitional Separation–Bubble Development in Swept–Blade Flow Conditions
Authors: Brinkerhoff Joshua R. , Yaras Metin I.
Published: August 2013

This paper describes numerical simulations of the instability mechanisms in a separation bubble subjected to a three-dimensional freestream pressure distribution. Two direct numerical simulations are performed of a separation bubble with laminar separation and turbulent reattachment under low freestream turbulence at flow Reynolds numbers and streamwise pressure distributions that approximate the conditions encountered on the suction side of typical low-pressure gas-turbine blades with blade sweep angles of 0° and 45°. The three-dimensional pressure field in the swept configuration produces a crossflow-velocity component in the laminar boundary layer upstream of the separation point that is unstable to a crossflow instability mode. The simulation results show that crossflow instability does not play a role in the development of the boundary layer upstream of separation. An increase in the amplification rate and most amplified disturbance frequency is observed in the separated-flow region of the swept configuration, and is attributed to boundary-layer conditions at the point of separation that are modified by the spanwise pressure gradient. This results in a slight upstream movement of the location where the shear layer breaks down to small-scale turbulence and modifies the turbulent mixing of the separated shear layer to yield a downstream shift in the time-averaged reattachment location. The results demonstrate that although crossflow instability does not appear to have a noticeable effect on the development of the transitional separation bubble, the 3D pressure field does indirectly alter the separation-bubble development by modifying the flow conditions at separation.
Experimental Investigation Into Thermal Behaviour of Steam Turbine Components: Part 1—Temperature Measurements With Optical Probes
Authors: Mohr W. F. , Ruffino P.
Published: August 2013

The first-in-time application of intensity pyrometry to measure in-situ the hot rotor surface temperature of a standard, combined cycle, intermediate pressure steam turbine is presented. The data cover a cold-start and cooling from base load. The pyrometric temperatures are compared to standard temperature measurements on static turbine parts and an upstream steam temperature measured on a thermo well.

It is reported, how the applicability of pyrometry in steam turbines was assessed. Details are given about a newly developed USC autoclave, which was used to measure steam transmittance, and about the measurement of the emissivity of the rotor metal.

Further the steps taken towards a steam-pyrometer are shown; how it was developed, validated in terms of its precision and lifetime in hot steam environment, and how its integration to a standard turbine was prepared.
Experimental Investigation Into Thermal Behavior of Steam Turbine Components: Part 2—Natural Cooling of Steam Turbines and the Impact on LCF Life
Authors: Marinescu Gabriel , Ehrsam Andreas
Published: August 2013

Steam turbine cool-down has a significant impact on the cyclic fatigue life. A lower initial metal temperature after standstill results in a higher temperature difference to be overcome during the next start-up. Generally, lower initial metal temperatures result in higher start-up stress. In order to optimize steam turbines for cyclic operation, it is essential to fully understand natural cooling, which is especially challenging for rotors.

A two-dimensional numerical procedure is described for the assessment of the thermal regime during natural cooling including the rotors, casings, valves and main pipes. The concept of the cooling calculation is to replace the steam gross buoyancy during the gland steam ingestion phase by an equivalent fluid conductivity, that gives the same thermal effect on the metal parts. The fluid equivalent conductivity is calculated based on measurements. The approach is calibrated with experimental data. Finally, the highly sensitive nature of the cyclic lifetime to the predicted cooling evolution is demonstrated. This paper is complementary with the paper [1].
Validation of the Immersed Mesh Block (IMB) Approach Against a Cooled Transonic Turbine Stage
Authors: Lad B. , He L. , Romero E.
Published: August 2013

Due to difficulties associated with directly meshing and solving cooled turbine geometries, evaluation of cooling designs at an early design stage is not practical. This results in a severely disjointed aerodynamic and heat transfer design process and affects turbine optimisation. In order to reduce the time cost of evaluating cooled geometries, the Immersed Mesh Block (IMB) approach was previously proposed. The IMB is an independent, high density cooling hole mesh that can be mapped to the surface of a turbine and solved with two-way coupling within hours, thus allowing cooling holes to be consistently and reliably modelled at an early design stage. The main focus of this paper is to validate the IMB approach by comparing its results of a cooled transonic NGV against experimental data. Prior to this, a series of mesh dependency tests are conducted on the base solver for a turbine stage. These highlight the need for greater rigour in mesh generation for RANS predictions of heat transfer over aerodynamics. The effect of mesh density on cooling hole predictions is then considered; a range of meshes are employed to model a row of holes on a flat plate. Hence the need for high resolution meshes for cooled geometries is emphasised. Subsequently, it is demonstrated that grid dependency studies of cooling holes on the turbine surface can easily be conducted with the IMB approach. Finally, the IMB method is validated by comparison of its results to experimental data for a cooled turbine stage. The results show that the IMB predicted Nusselt number distribution agrees well with experimental data.
Experimental and Numerical Analysis of Multiple Impingement Jet Arrays for an Active Clearance Control System
Authors: Andreini Antonio , Da Soghe Riccardo , Facchini Bruno , Maiuolo Francesco , Tarchi Lorenzo , Coutandin Daniele
Published: August 2013

The turbine blade tip clearances control in large aeroengines is currently performed by means of impinging fan air on the outer case flanges. The aim of the present study is to evaluate both the heat transfer coefficient and the adiabatic thermal effectiveness characteristics of an engine-like ACC system, and in particular, to comprehend the effects of the undercowl flow on the impingement jets. The considered geometry replicates the impingement tubes and the by-pass duct used in active control clearance systems. The tube’s internal diameter is D = 12mm, the cooling holes’ diameter is d = 1mm and the span-wise pitch is Sy/d = 12. In order to simulate the undercowl flow, the impingement arrays are inserted inside a tunnel that replicates the typical shape of a real engine by-pass duct. Tests were conducted varying both the mainstream Reynolds number and the jets Reynolds number in a range typical of real-engine operative conditions (Rej = 2000–10000, β = 1.05–1.15). Numerical calculations are finally proposed to point out if CFD is able to confidently reproduce the experimental evidences.
Preliminary Experimental Investigation of the Effects of Particulate Deposition on IGCC Turbine Film-Cooling in a High-Pressure Combustion Facility
Authors: Murphy Robert G. , Nix Andrew C. , Lawson Seth A. , Straub Douglas , Beer Stephen K.
Published: August 2013

Researchers at West Virginia University are working with the U.S. Department of Energy, National Energy Technology Laboratory (NETL) to study the effects of particulate deposition on turbine film cooling in a high pressure and high temperature environment. To simulate deposition on the pressure side of an Integrated Gasification Combined Cycle (IGCC) turbine first stage vane, angled film-cooled test articles with thermal barrier coatings (TBC) are subjected to accelerated deposition at a pressure of approximately 4 atm and a gas temperature of 1100°C. Two different test article geometries were designed, with angles of 10° and 20° to the mainstream flow. Both geometries have straight-cooling holes oriented at a 30° angle to the hot-side surface. A high pressure seeding system was used to generate a particulate concentration of approximately 33.3 ppmw. Particle concentrations of 0.02 ppmw exist in the IGCC hot gas path. An accelerated simulation method was developed to simulate deposition that would occur in 10000 hr of engine operation. Preliminary tests were performed at 4 atm and 1100 °C to validate the deposition process. The results showed more deposition on the 20° test article than the 10° test articles; however no substantial deposition developed on either test article. A lumped mass analysis showed that the fly ash particles dropped below the theoretical sticking temperature as they approached the test article. Deposition was analyzed non-destructively through visual observation and scanning with a scanning laser microscope. Based on the initial test run results, a detailed plan was created to increase the operating temperature of the rig and allow two 3-hour tests to be performed on each of the test articles. Non-destructive testing will be used before, in between and after the runs to examine the evolution of the deposition growth. Following the final run, destructive testing will be used to examine the chemical composition of the deposits and their potential interaction with the TBC. Preliminary work will lead to a future study the would enhance the understanding of particle deposition evolution and examine the effects of deposition on film cooling by performing the tests in a high-pressure and high-temperature environment that is similar to the high-pressure combustion exhaust gas environment of the first stage region in IGCC turbines.
Characterizing Compressor Rotor Unclamp in a Gas Turbine Engine
Authors: Alban Cory , Tolstykh Masha , Hilty Devin , Bollman Andrew
Published: August 2013

In the aerospace industry, many gas turbine compressors rely on a tie bolt to mechanically hold together all rotating components in the compressor rotor. Maintaining this clamp load is essential to the performance of the engine. In the event of an unclamp, the engine will experience a reduction in tip clearance due to a change in rotational dynamics; increased temperatures and pressures in secondary air systems; and a decrease in critical component life. Accordingly, designers must be aware of the variables effecting compressor rotor clamp loads observed for component assembly and operational missions. During testing, an axial gas turbine engine unexpectedly experienced a compressor rotor unclamp which led to an increase in turbine temperature and front sump buffer air temperature and pressure. Further investigation revealed a thermal expansion mismatch between the tie bolt and inner gas path rim during a specific transient condition. Because of this thermal effect, the rotor will experience an unclamped condition which will result in ingesting compressor discharge air into the drum. The compressor rotor will remain unclamped until the engine is operated at a lower power setting or shut down for an extended period. This paper documents and explains the transient condition at which the engine experienced unclamp through review of test data, characterizes the design space around the tie bolt by using heat transfer and structural finite element analysis codes, and shows how robust design tools were used to find an optimized solution that eliminated the risk of thermally driven unclamp through robust design and assembly choices.
The Development of a Large Radial Movement Brush Seal
Authors: Franceschini Gervas , Jones Terry V. , Gillespie David R. H.
Published: August 2013

Brush seals provide an advantage over traditional seals such as the labyrinth seal by tolerating a small amount of contact between the seal and the rotor without performance degradation. The bristles are compliant and are able to accommodate a rotor movement towards the seal by mechanical deflection of the bristles. Blow-down, an aerodynamic phenomenon whereby the bristles move towards the shaft under the application of a differential pressure, allows the bristles to close a gap between the seal and the rotor. These mechanisms ensure that the brush seal gives good performance with small rotor growths and rotor translations. In a conventional seal the need to accommodate large eccentric movements may result in excessive seal wear as the axial stiffness of the seals in the unsupported region either becomes insufficient to avoid serious seal deflection, or is increased to such an extent that wear is excessive in normal operation. Rolls-Royce has developed a large radial movement seal using a conventional brush seal as its primary element. The seal housing is designed to accommodate a much larger range of rotor translations, through the use of a tight tolerance, essentially static-sliding, secondary seal. Continuous stiffness, torque and flow measurements were made during a series of eccentric shaft movements at engine scale and differential pressures with the shaft rotating at slow speed. Results show that at small deflections the seal acts identically to a conventional brush seal, to which the results are compared. At higher eccentric displacements the floating seal limits the maximum bristle tip contact load and the load on the rotor. Here the stiffness characteristics are determined by the frictional properties of the seal housing, the precise geometry of which is also the limit to the range of radial movement. The experimental campaign has shown that the large radial movement seal is a viable solution for engine locations where large eccentric movements need to be accommodated, but small radial growths are expected.
Numerical Investigation of Ash Deposition on Nozzle Guide Vane Endwalls
Authors: Casaday B. , Ameri A. , Bons J. P.
Published: August 2013

A computational study was performed to determine the factors that affect ash deposition rates on the endwalls in a nozzle guide vane passage. Deposition tests were simulated in flow around a flat plate with a cylindrical leading edge, as well as through a modern, high performance turbine vane passage. The flow solution was first obtained independent of the presence of particulates, and individual ash particles were subsequently tracked using a Langrangian tracking model. Two turbulence models were applied and their differences were discussed. The critical viscosity model was used to determine particle deposition. Features that contribute to endwall deposition, such as secondary flows, turbulent dispersion, or ballistic trajectories, were discussed and deposition was quantified. Particle sizes were varied, to reflect Stokes numbers ranging from 0.01 to 1.0, to determine the effect on endwall deposition. Results showed that endwall deposition rates can be as high as deposition on the leading edge for particles with a Stokes number less than 0.1, but endwall deposition rates for a Stokes numbers of 1.0 were less than 25% of the deposition rates on the leading edge or pressure surface of the turbine vane. Deposition rates on endwalls were largest near the leading edge stagnation region on both the cylinder and vane geometries, with significant deposition rates downstream showing a strong correlation to the secondary flows.
Measurement Versus Predictions of Rotordynamic Coefficients of a Hole-Pattern Gas Seal With Negative Preswirl
Authors: Brown Philip D. , Childs Dara W.
Published: August 2013

Test results are presented for rotordynamic coefficients of a hole-pattern annular gas seals at supply pressures to 84 bar and running speeds to 20,200 RPM. The principal test variable of interest was negative preswirl. Preswirl signifies the circumferential fluid flow entering a seal, and negative preswirl indicates a fluid swirl in a direction opposite to rotor rotation. The influences of pressure ratio and rotor speed were also investigated. Measured results produce direct and cross-coupled stiffness and damping coefficients that are a function of excitation frequency Ω.

Changes in pressure ratio had only small effects on most rotordynamic coefficients. Cross-coupled stiffness showed slightly different profiles through the mid-range of Ω values.

Increasing rotor speed significantly increased cross-coupled stiffness and cross-coupled damping. At 10,200 RPM, high negative inlet preswirl produced negative cross-coupled stiffness over an excitation frequency range of 200–250 Hz. Negative preswirl did not affect direct stiffness and damping coefficients.

Effective damping combines the stabilizing effect of direct damping and the destabilizing effect of cross-coupled stiffness. The cross-over frequency is the precession frequency where effective damping transitions from a negative value to a positive value with increasing frequency. At 20,200 RPM with a pressure ratio of 50%, peak effective damping was increased by 50%, and the cross-over frequency was reduced by 50% for high-negative preswirl versus zero preswirl. Hence, reverse swirl can greatly enhance the stabilizing capacity of hole-pattern balance-piston or division-wall seals for compressors.

A two-control-volume model that uses the ideal gas law at constant temperature was used to predict rotordynamic coefficients. The model predicted direct rotordynamic coefficients well, but substantially under predicted cross-coupled rotordynamic coefficients especially at high negative preswirls.
Impact of Swirl Flow on the Cooling Performance of an Effusion Cooled Combustor Liner
Authors: Wurm B. , Schulz A. , Bauer H.-J. , Gerendas M.
Published: August 2013

An experimental study on combustor liner cooling of modern direct lean injection (DLI) combustion chambers using coolant ejection from both effusion cooling holes and a starter film has been conducted. The experimental setup consists of a generic scaled three sector planar rig in an open loop hot gas wind tunnel, which has been described earlier in Wurm et al. [1]. Experiments are performed without combustion. Realistic engine conditions are achieved by applying engine-realistic Reynolds numbers, Mach numbers, and density ratios. A Particle Image Velocimetry (PIV) measurement technique is employed, which has been adjusted to allow for high resolution near wall velocity measurements with and without coolant ejection. As the main focus of the present study is a deeper understanding of the interaction of swirl flows and near wall cooling flows, wall pressure measurements are performed for the definition of local blowing ratios and to identify the impact on the local cooling performance. For thermal investigations an infrared thermography measurement technique is employed that allows high resolution thermal studies on the effusion cooled liner surface. The effects of different heat shield geometry on the flow field and performance of the cooling films are investigated in terms of near wall velocity distributions and film cooling effectiveness. Two different heat shield configurations are investigated which differ in shape and inclination angle of the so called heat shield lip. Operating conditions for the hot gas main flow are kept constant. The pressure drop across the effusion cooled liner is varied between 1% and 3% of the total pressure. Results show the impact of the swirled main flow on the stability of the starter film and on the effusion cooling performance. Stagnation areas which could be identified by wall pressure measurements are confirmed by PIV measurements. Thermal investigations reveal reduced cooling performance in the respective stagnation areas.
Radiative Models for Hot Gas With Various H2O-CO2 Ratios
Authors: Sun Peng , Zhou Yuji , Ren Jing , Jiang Hongde
Published: August 2013

Radiation is the principal mode of heat transfer in many engineering problems, such as in the performance calculation of the gas turbine. Models for gaseous radiative properties have been well-established for air combustion. The weighted-sum-of-gray-gases-model (WSGGM) is the most widely used global model for the computation of the gaseous radiative properties. It is a reasonable compromise between the oversimplified gray gas model and a complete model that takes in account the particular absorption bands. However, the WSGGM coefficients have generally been determined only at the H2O-CO2 ratios of 1 and 2 or other special ratios. In the advanced gas turbines, the H2O-CO2 ratios sometimes fall outside the range of the weighted-sum-of-gray-gases-model parameters reported in the literature. In order to obtain the radiative parameters for various H2O-CO2 ratios, a computer code is developed based on the HITRAN 2008 database. The total emissivity of CO2 and H2O are evaluated under different temperature, pressure and path length utilizing the statistic narrow band (SNB) model. On the basis of the SNB model, the parameters of WSGGM for various H2O-CO2 mixtures are computed with the Flether-Powell technique. Comparisons are executed with the other WSGGM.
The Effect of Particle Size and Film Cooling on Nozzle Guide Vane Deposition
Authors: Bonilla C. , Webb J. , Clum C. , Casaday B. , Brewer E. , Bons J. P.
Published: August 2013

An accelerated deposition test facility is used to study the effect of particle size and film cooling on deposit roughness, spatial distribution and thickness. Tests were run at gas turbine representative inlet Mach numbers (0.08) and temperatures (1080°C). Deposits were created from a sub-bituminous coal fly ash with mass median diameters from 4 to 16 microns (Stokes numbers ranging from 0.1 to 1.9. Two CFM56-5B nozzle guide vane doublets comprising three full passages and two half passages of flow were utilized as the test articles. Tests were run with three levels of film cooling. The addition of film cooling to the vanes was shown to decrease deposit capture efficiency by as much as a factor of 3 and shift the primary location of deposit buildup to the leading edge coincident with an increased region of positive cooling backflow margin. Video taken during tests noted that film cooling holes with negative backflow margin were primary areas of deposit formation regardless of film cooling percentage. Stokes number was shown to have a marked effect on vane capture efficiency, with the largest Stokes number ash (St = 1.9) approximately 3 times as likely to stick to the vane as the smallest Stokes number ash (St = 0.1). Post test observations on deposit thickness were made using a coordinate measurement machine. Deposit thickness was noted to be reduced with decreasing Stokes number and increased film cooling percentage. Deposit surface roughness falls with particle size but is only weakly dependent on cooling level.
A Comparative Study of Film Cooling Effectiveness on a Flat Plate With Adiabatic and Conjugate Conditions for Different Hole Shapes
Authors: Chandran P. M. Dileep , Halder Paresh , Panda Rajesh Kumar , Prasad B. V. S. S. S.
Published: August 2013

A computational study is carried out for comparison of the effectiveness of film cooling through holes of different shapes: cylindrical, shaped, trenched cylindrical and trenched shaped hole. Both adiabatic and conjugate wall thermal condition with various blowing ratios (0.6 to 1.4) are considered for the investigation. The coolant-to-mainstream density ratio and temperature ratio are maintained at 1.6 and 0.625 respectively. The κ-ω SST turbulence model is used for computation. The complex flow structures within the film hole, the exit and in the interaction zone of jet with cross flow are reported. Among the configurations studied, the shaped and trenched shaped holes have shown smaller jet lift off and uniform coolant coverage in both lateral and streamwise directions. Further, the results showed a significant difference in the indicated values of centerline surface temperature and effectiveness with the adiabatic and the conjugate conditions; the conjugate values being much lower but more uniform.
Numerical Study of Transitional Rough Wall Boundary Layer
Authors: Elsner Witold , Warzecha Piotr
Published: August 2013

The paper presents the verification of boundary layer modeling approach, which relies on a γ-Reθt model proposed by Menter et al. [1]. This model was extended by laminar-turbulent transition correlations proposed by Piotrowski et al. [2] as well as Stripf et al. [3] correlations, which take into account the effects of surface roughness. To blend between the laminar and fully turbulent boundary layer over rough wall the modified intermittency equation is used.

To verify the model a flat plate with zero and non-zero pressure gradients test cases as well as the high pressure turbine blade case were chosen. Further on, the model was applied for unsteady calculations of turbine blade profile as well as the Lou and Hourmouziadis [4] flat plate test case, with induced pressure profile typical for suction side of highly-loaded turbine airfoil. The combined effect of roughness and wake passing were studied.

The studies proved that the proposed modeling approach (ITMR hereinafter) appeared to be sufficiently precise and enabled for a qualitatively correct prediction of the boundary layer development for the tested simple flow configurations. The results of unsteady calculations indicated that the combined impact of wakes and the surface roughness could be beneficial for the efficiency of the blade rows, but mainly in the case of strong separation occurring on highly-loaded blade profiles. It was also demonstrated that the roughness hardly influences the location of wake induced transition, but has an impact on the flow in between the wakes.
Heat Transfer Measurements in an Internal Cooling System Using a Transient Technique With Infrared Thermography
Authors: Egger Christian , von Wolfersdorf Jens , Schnieder Martin
Published: August 2013

In this paper a transient method for measuring heat transfer coefficients in internal cooling systems using infrared thermography is applied. The experiments are performed with a two-pass internal cooling channel connected by a 180° bend. The leading edge and the trailing edge consist of trapezoidal and nearly rectangular cross sections, respectively, to achieve an engine-similar configuration. Within the channels rib arrangements are considered for heat transfer enhancement. The test model is made of metallic material. During the experiment the cooling channels are heated by the internal flow. The surface temperature response of the cooling channel walls is measured on the outer surface by infrared thermography. Additionally, fluid temperatures as well as fluid and solid properties are determined for the data analysis. The method for determining the distribution of internal heat transfer coefficients is based on a lumped capacitance approach which considers lateral conduction in the cooling system walls as well as natural convection and radiation heat transfer on the outer surface. Because of time-dependent effects a sensitivity analysis is performed to identify optimal time periods for data analysis. Results are compared with available literature data.
Disk Heat Transfer Analysis in a Heated Rotating Cavity With an Axial Throughflow
Authors: Tian Shuqing , Zhu Yatao
Published: August 2013

In the rotating disk cavities of aero-engine compressors, buoyancy-induced flow and heat transfer can occur due to thermal gradients between cooling air and hot surfaces. The simplified rotating cavity with two plane discs, a shaft and a cylindrical rim has been investigated numerically and compared with the available measurements. Two models have been solved using a commercial CFD code, Fluent, with the RNG k-ε turbulence model. The first one is the conventional model with only fluid region solved, a temperature profile with the linear radial gradient imposed at the disk walls, and an isothermal boundary condition imposed at the shroud wall. The second one is the model with thick-walled disks and shroud, an adiabatic boundary condition imposed at the outer walls of the disks, and an isothermal boundary condition imposed at the outer wall of the shroud. The fluid and solid are coupled solved simultaneously. The disk temperatures are computed.

In the present work, the numerical results are in reasonable agreement with the measurements. The computed disk temperatures in the second model have approximately linear radial gradients over the first three-quarters of the disks, and in the last quarter of the disks the temperature radial gradients are obviously non-linear. The different disk temperature profiles in these two models do not lead to obviously different disk heat transfers.

The heat transfer in the rotating cavity leads to a considerable temperature increase of the cavity core fluid, therefore a corresponding increase of the outlet temperature. These two temperature increases are critical for the cooling design in aero-engines.
Cooling Effectiveness of Impingement/Effusion Cooling With and Without Turbulence Promoter Ribs
Authors: Yang Qian , Lin Yuzhen , Xu Quanhong , Zhang Chi , Sung Chih-Jen
Published: August 2013

In order to solve the cooling problem of aeroengine, a variety of highly efficient cooling methods came into application, such as effusion wall, impingement/effusion wall, etc. The impingement/effusion cooling methods with and without turbulence promoter ribs are introduced here. In the impingement/promoter ribs/effusion walls, turbulence promoter ribs are on the coolant side of effusion wall so as to enhance the internal cooling.

Numerical and experimental approaches were taken to study cooling effectiveness of impingement/effusion liners cooling with and without turbulence promoter ribs. Cooling effectiveness of the two types cooling configurations were investigated on the same hot and cold air condition of inlet. Infrared camera was used to measure the temperature of effusion cooled surfaces on the side of hot main stream. The results showed double wall cooling with turbulence promoter ribs achieves higher cooling effectiveness than impingement/effusion cooling configurations.

In addition, the two test configurations were investigated numerically to get clearer characterization of flow field in the double walls cooling configurations. High velocity vectors were found in the flow field within the gap of the double walls. A heat transfer enhancement was resulted in complex flows and increased area caused by the inclined holes and turbulence promoter ribs.