bluuebunny/crossref_metadata_2025 huggingface dataset:
Local Heat Transfer and Flow Distributions for Impinging Jet Arrays of Dense and Sparse Extent
Authors: Bailey J. C. , Bunker R. S.
Published: February 2009

Full-surface heat transfer coefficient distribution measurements have been made using a liquid crystal thermography technique for several cases of normally impinging jet arrays onto a flat, smooth surface within a region bounded on three sides. While the impingement target plate remains of a fixed size, the impingement jet array has been changed to cover a wide range of conditions, extending beyond the currently available literature data. Axial and lateral jet spacing values of x/D and y/D of 3, 6, and 9 have been used, all with square orientation and in-line jets. The jet plate-to-target surface distance z/D has been varied from 1.25 to 5.5. Jet Reynolds numbers ranged from 14,000 to 65,000. In the sparse array limiting case, the number of jet rows is four in the axial direction and three in the lateral direction. For the dense array limiting case, the number of jet rows is 26 in the axial direction and 20 in the lateral direction. Using both heat transfer and pressure distribution measurements, results are compared to the existing correlation of Florschuetz et al. [1], showing excellent agreement in regions of common parameters. In regions not previously reported in the literature, the present study extends the streamwsie row-averaged heat transfer coefficient correlation of [1] with a modified correlation for design use.
Heat Transfer Enhancement Using a Convex-Patterned Surface
Authors: Moon H. K. , O’Connell T. , Sharma R.
Published: February 2009

The heat transfer rate from a smooth wall in an internal cooling passage can be significantly enhanced by using a convex patterned surface on the opposite wall of the passage. This design is particularly effective for a design that requires the heat transfer surface to be free of any augmenting features (smooth). Heat transfer coefficients on the smooth wall in a rectangular channel, which had convexities on the opposite wall were experimentally investigated. Friction factors were also measured to assess the thermal performance. Relative clearances (δ/d) between the convexities and the smooth wall of 0, 0.024, and 0.055 were investigated in a Reynolds number (ReHD) range from 15,000 to 35,000. The heat transfer coefficients were measured in the thermally developed region using a transient thermochromic liquid crystal technique. The clearance gap between the convexities and the smooth wall adversely affected the heat transfer enhancement (NuHD). The friction factors (f), measured in the aerodynamically developed region, were largest for the cases of no clearance (δ/d = 0). The average heat transfer enhancement (NuHD) was also largest for the cases of no clearance (δ/d = 0), as high as 3.08 times at a Reynolds number of 11,456 in relative to that (Nuo) of an entirely smooth channel. The normalized Nusselt numbers (NuHD/Nuo), as well as the normalized friction factors (f/fo), for all three cases, decreased with Reynolds numbers. However, the decay rate of the friction factor ratios (f/fo) with Reynolds numbers was lower than that of the normalized Nusselt numbers. For all three cases investigated, the thermal performance ((NuHD/Nuo) /(f/fo)1/3) values were within 5% to each other. The heat transfer enhancement using a convex patterned surface was thermally more effective at a relatively low Reynolds numbers (less than 20,000 for δ/d = 0) than that of a smooth channel.
The Model Steam Turbine: More Details About the Heat Recovery Steam Generator Evaluating Method in a Combined Cycle Plant
Authors: Schmidt Dietmar , Arnold Michail
Published: February 2009

Turnkey and thermal island supply scopes present turbine suppliers with a perfect way to sell their rotating products. The popularity of these plant configurations, along with the recent availability of more holistic test codes, has led to the need for an accurate and reasonable method of determining the thermal performance of the externally-purchased HRSG component. To assess a multiple pressure HRSG, it is advantageous and convenient to have one single criterion for the evaluation of performance, especially when this criterion provides for the compensation of the different outlet energy streams. The so-called Model Steam Turbine method of HRSG evaluation was developed for these reasons. The result of the calculation, a lone performance criterion, is the shaft power of the fictitious Model Steam Turbine.
Effect of Inlet Air Cooling by Ice Storage on Unit Sizing of a Gas Turbine Cogeneration Plant
Authors: Yokoyama Ryohei , Ito Koichi
Published: February 2009

In the commercial sector, heat and power demands peak in the summer daytime because of high space cooling demands, and cogeneration plants are required to produce maximum heat and power to meet their demands. However, gas turbine cogeneration plants have the disadvantage of decreases in maximum power output in the summer daytime, which reduces the availability of gas turbines. One of the ways to avoid the aforementioned disadvantage is to cool inlet air and augment maximum power output. In addition, one of the ways for inlet air cooling is to make ice by driving electric compression refrigerators using off-peak power generated during the nighttime, store it in ice banks, and use its heat for inlet air cooling during the on-peak period. The objective of this paper is to investigate the effect of inlet air cooling by ice storage on the unit sizing and cost of a gas turbine cogeneration plant. An optimal unit sizing method based on the mixed-integer linear programming is used to rationally determine equipment capacities and operational strategies of the plant. A numerical study is conducted, in which the gas turbine cogeneration plants with and without inlet air cooled by ice storage are compared with each other, and the effect of inlet air cooling on the equipment capacities as well as the annual total cost and its items is clarified.
Surface Generation and Machining With Rational B-Splines for Turbomachinery Components
Authors: Tsay Der Min , Chen Chien Wen , Chen Jiun Ren
Published: February 2009

To five-axis point mill turbomachinery components, a least squares approximation procedure for generating tool paths in terms of parametric rational B-splines is presented. The detailed processes of using the appropriate parameters of rational B-splines that will reduce the amount of geometric deviations between the designed and machined part surfaces are proposed. Compared to a traditional linear tool path, a rational B-spline tool path can substantially reduce the number of blocks on machine controllers without compromising the accuracy. As a result, the machining efficiency and machine dynamic behavior can be improved. Numerical and experimental results are given to show such improvements. A practical application example for machining a centrifugal compressor impeller with 12 blades is also illustrated to demonstrate the reliability and usefulness of the developed procedure.
Aerodynamic-Rotordynamic Interaction in Axial Compression Systems: Part I — Modeling and Analysis of Fluid-Induced Forces
Authors: Al-Nahwi Ammar A. , Paduano James D. , Nayfeh Samir A.
Published: February 2009

This paper presents a first principles-based model of the fluid-induced forces acting on the rotor of an axial compressor. These forces are primarily associated with the presence of a nonuniform flow field around the rotor, such as that produced by a rotor tip clearance asymmetry. Simple, analytical expressions for the forces as functions of basic flow field quantities are obtained. These expressions allow an intuitive understanding of the nature of the forces and—when combined with a rudimentary model of an axial compressor flow field (the Moore-Greitzer model)—enable computation of the forces as a function of compressor geometry, torque and pressure-rise characteristics, and operating point. The forces predicted by the model are also compared to recently published measurements and more complex analytical models, and are found to be in reasonable agreement. The model elucidates that the fluid-induced forces comprise three main contributions: fluid turning in the rotor blades, pressure distribution around the rotor, and unsteady momentum storage within the rotor. The model also confirms recent efforts in that the orientation of fluid-induced forces is locked to the flow nonuniformity, not to tip clearance asymmetry as is traditionally assumed. The turning and pressure force contributions are shown to be of comparable magnitudes—and therefore of equal importance—for operating points between the design point and the peak of the compressor characteristic. Within this operating range, both “forward” and “backward” rotor whirl tendencies are shown to be possible. This work extends recent efforts by developing a more complete, yet compact, description of fluid-induced forces in that it accounts for all relevant force contributions, both tangential and radial, that may influence the dynamics of the rotor. Hence it constitutes an essential element of a consistent treatment of rotordynamic stability under the action of fluid-induced forces, which is the subject of Part II of this paper.
Aerodynamic-Rotordynamic Interaction in Axial Compression Systems: Part II — Impact of Interaction on Overall System Stability
Authors: Al-Nahwi Ammar A. , Paduano James D. , Nayfeh Samir A.
Published: February 2009

This paper presents an integrated treatment of the dynamic coupling between the flow field (aerodynamics) and rotor structural vibration (rotordynamics) in axial compression systems. This work is motivated by documented observations of tip clearance effects on axial compressor flow field stability, the destabilizing effect of fluid-induced aerodynamic forces on rotordynamics, and their potential interaction. This investigation is aimed at identifying the main nondimensional design parameters governing this interaction, and assessing its impact on overall stability of the coupled system. The model developed in this work employs a reduced-order Moore-Greitzer model for the flow field, and a Jeffcott-type model for the rotordynamics. The coupling between the fluid and structural dynamics is captured by incorporating a compressor pressure rise sensitivity to tip clearance, together with a momentum based model for the aerodynamic forces on the rotor (presented in Part I of this paper). The resulting dynamic model suggests that the interaction is largely governed by two nondimensional parameters: the sensitivity of the compressor to tip clearance and the ratio of fluid mass to rotor mass. The aerodynamic-rotordynamic coupling is shown to generally have an adverse effect on system stability. For a supercritical rotor and a typical value of the coupling parameter, the stability margin to the left of the design point is shown to decrease by about 5% in flow coefficient (from 20% for the uncoupled case). Doubling the value of the coupling parameter not only produces a reduction of about 8% in the stability margin at low flow coefficients, but also gives rise to a rotordynamic instability at flow coefficients 7% higher than the design point.
Stator-Rotor Interactions in a Transonic Compressor: Part 2 — Description of a Loss Producing Mechanism
Authors: Gorrell Steven E. , Okiishi Theodore H. , Copenhaver William W.
Published: February 2009

A previously unidentified loss producing mechanism resulting from the interaction of a transonic rotor blade-row with an upstream stator blade-row is described. This additional loss occurs only when the two blade rows are spaced closer together axially. Time-accurate simulations of the flow and high-response static pressure measurements acquired on the stator blade surface reveal important aspects of the fluid dynamics of the production of this additional loss. At close spacing the rotor bow shock is chopped by the stator trailing edge. The chopped bow shock becomes a pressure wave on the upper surface of the stator that is nearly normal to the flow and that propagates upstream. In the reference frame relative to this pressure wave, the flow is supersonic and thus a moving shock wave that produces an entropy rise and loss is experienced. The effect of this outcome of blade-row interaction is to lower the efficiency, pressure ratio, and mass flow rate observed as blade-row axial spacing is reduced from far to close. The magnitude of loss production is affected by the strength of the bow shock and how much it turns as it interacts with the trailing edge of the stator. At far spacing the rotor bow shock degenerates into a bow wave before it interacts with the stator trailing edge and no significant pressure wave forms on the stator upper surface. For this condition, no additional loss is produced.
Preliminary Gas Turbine Design Using the Multidisciplinary Design System MOPEDS
Authors: Jeschke Peter , Kurzke Joachim , Schaber Reinhold , Riegler Claus
Published: February 2009

A prototype preliminary design task for gas turbines is set up to outline the four major requirements a preliminary design program must typically meet: assessment of all major engine components and their interrelations; inclusion of all relevant disciplines; designing over several operating points; and model fidelity zooming at least for individual components. It is described how the “MOdular Performance and Engine Design System” (MOPEDS) — MTU Aero Engines’ software package for the preliminary design of airborne and stationary gas turbines — fulfills these requirements. The program structure, the graphical user interface, and the physical models are briefly presented. A typical design example is carried out emphasizing the necessity for a numerical procedure to find a solution to the many variables and constraints that the design problem comprises. Finally, some dominating multidisciplinary effects are discussed.
Performance Modeling Methodology: Efficiency Definitions for Cooled Single and Multistage Turbines
Authors: Kurzke Joachim
Published: February 2009

Quoting numbers for the efficiency of a turbine is ambiguous if it is not known how this efficiency is defined. This is especially true for a heavily cooled turbine where for the same machine the efficiency may be quoted as 88% or 91%, for example. In aero-engine industry, several different turbine efficiency bookkeeping systems are in use. Since nearly always a consortium of two or more companies is involved into any new engine project it is important to understand the various bookkeeping systems. Turbine designers and performance specialists should not quote a number for the efficiency of a cooled turbine without clearly stating on which methodology and control volume it is based. In this paper, the commonly used turbine performance bookkeeping methods are compared. The pros and cons for the different methodologies are discussed and correlations between the definitions are shown for single and multistage turbines with various amounts of cooling air. Furthermore it is pointed out, that simulating a cooled multistage turbine with an equivalent single-stage model requires the use of an equivalent Stator Outlet Temperature SOTeqv which differs from the true SOT. It depends on the bookkeeping system used what the calculated impact on turbine performance is when the amount of cooling air changes. Without carefully adhering to a unique bookkeeping system with a clearly defined control volume the probability of misunderstandings in collaborative engine development projects is not to be underestimated.
Importance of Inlet Total Pressure Conditions in Evaluating Performance of Non-Symmetric Gas Turbine Exhaust Ducts
Authors: Cunningham M. H. , Birk A. M. , Di Bartolomeo W.
Published: February 2009

When highly non-symmetric exhaust ducts are installed on a gas turbine engine, the asymmetries result in a non-uniform circumferential total pressure condition at the inlet of the duct. When testing these ducts experimentally or computationally the correct inlet conditions are often not known or cannot be reproduced. To study the sensitivity of duct performance to inlet conditions, an experimental and computational study of a non-symmetric gas turbine exhaust duct that includes a 160° turn with an annular to rectangular transition, has been carried out over a range of inlet conditions. The inlet conditions varied include circumferential total pressure profiles and swirl. The experimental studies have been carried out in cold flow with several non-uniform total pressure inlet conditions. Computational fluid dynamic (CFD) techniques validated against the experimental results, have been used to extend the range of inlet conditions beyond the range that could be obtained experimentally to those typical of an engine installation. Results show that the total pressure inlet conditions have a significant effect on the flow structure in the exhaust duct and that the performance of the exhaust duct degrades as the level of circumferential non-uniformities increase. However, trends in geometric optimization identified experimentally using cold flow and uniform total pressure inlet conditions are confirmed computationally with circumferential non-uniformities typical of actual engine operations. This suggests that although inlet conditions are important for determining the level of performance, the configuration of the optimized geometry is somewhat independent of the inlet conditions.
A Computer Code for Gas Turbine Engine Weight and Disk Life Estimation
Authors: Tong Michael T. , Halliwell Ian , Ghosn Louis J.
Published: February 2009

Reliable engine-weight estimation at the conceptual design stage is critical to the development of new aircraft engines. It helps to identify the best engine concept amongst several candidates. In this paper, the major enhancements to NASA’s engine-weight estimate computer code (WATE) are described. These enhancements include the incorporation of improved weight-calculation routines for the compressor and turbine disks using the finite-difference technique. Furthermore, the stress distribution for various disk geometries was also incorporated, for a life-prediction module to calculate disk life. A material database, consisting of the material data of most of the commonly-used aerospace materials, has also been incorporated into WATE. Collectively, these enhancements provide a more realistic and systematic way to calculate the engine weight. They also provide additional insight into the design trade-off between engine life and engine weight. To demonstrate the new capabilities, the enhanced WATE code is used to perform an engine weight/life trade-off assessment on a production aircraft engine.
Confidence Intervals for Weibull Fast-Fracture Parameters
Authors: Chati Mandar , Johnson Curtis , Kaya Ahmet , Schenk Bjoern
Published: February 2009

Practical limits on number of specimens that can be tested lead to uncertainty in the estimated Weibull parameters. This paper presents an evaluation of four techniques for estimating confidence intervals for size-scaled Weibull parameters of monolithic ceramics. The techniques include normal approximation method, likelihood ratio technique, nonparametric bootstrap, and parametric bootstrap methods. For uncensored fast-fracture data, the confidence intervals for Weibull parameters are compared to the method used in ASTM Standard C1239. A simulation fracture experiment is conducted to evaluate the statistical characteristics, in particular coverage probability, of the four methods. For fast-fracture data with multiple failure modes, the statistical assessment of the confidence interval techniques for size-scaled Weibull parameters complement the existing literature. Overall, it was observed that the likelihood ratio technique and parametric bootstrap method perform very well. These techniques can also be extended for confidence interval estimation using fast-fracture data obtained from various geometry’s of test specimens and/or loading conditions (pooled data).
Ceramic Gas Turbine Materials Impact Evaluation
Authors: van Roode Mark , Jimenez Oscar , McClain John , Price Jeff , Parthasarathy Vijay , Poormon Kevin L. , Ferber Mattison K. , Lin Hua-Tay
Published: February 2009

Impact of foreign or domestic material on components in the hot section of gas turbines with ceramic components is a common cause of catastrophic failure. Several such occurrences were observed during engine testing under the Ceramic Stationary Gas Turbine program sponsored by the U.S. Department of Energy. A limited analysis was carried out at Solar Turbines Incorporated (Solar), which involved modeling of the impact in the hot section. Based on the results of this study an experimental investigation was carried out at the University of Dayton Research Institute Impact Physics Laboratory to establish the conditions leading to significant impact damage in silicon-based ceramics. The experimental set up involved impacting ceramic flexure bars with spherical metal particulates under conditions of elevated temperature and controlled velocity. The results of the study showed a better correlation of impact damage with momentum than with kinetic energy. Increased test specimen mass and fracture toughness were found to improve impact resistance. Continuous fiber-reinforced ceramic composite (CFCC) materials have better impact resistance than monolithics. A threshold velocity was established for impacting particles of a defined mass. Post-impact metallography was carried out at Oak Ridge National Laboratory to further establish the impact mechanism.
Power Law Versus Exponential Form of Slow Crack Growth of Advanced Structural Ceramics: Dynamic Fatigue
Authors: Choi Sung R. , Gyekenyesi John P.
Published: February 2009

The life prediction analysis based on an exponential crack velocity formulation was examined using a variety of experimental data on glass and advanced structural ceramics in constant stress-rate (“dynamic fatigue”) and preload testing at ambient and elevated temperatures. The data fit to the strength versus ln (stress rate) relation was to found be very reasonable for most of the materials. It was also found that preloading technique was equally applicable for the case of slow crack growth (SCG) parameter n>30. The major limitation in the exponential crack velocity formulation, however, was that an inert strength of a material must be known priori to evaluate the important SCG parameter n, a significant drawback as compared to the conventional power-law crack velocity formulation.
A Novel Technique for Estimation of Critical Frontal Process Zone Size in Ceramics
Authors: Choi Seong-Min , Okamoto Masato , Awaji Hideo
Published: February 2009

A novel technique for estimating the critical size of the frontal process zone in ceramics was proposed using a single-edge V-notched beam (SEVNB) test method. A three-point flexure test was carried out on a silicon nitride and two kinds of graphite test specimens with sharp V-shaped notches whose depths varied from 20 to 1,500 μm. The critical local stress was calculated at a critical distance from the notch tip. The critical size of the frontal process zone was determined as the distance between the notch tip and the point where the critical local stress had the same value as the flexural strength of test specimens without notches. The relationship among the critical size of the frontal process zone, the fracture toughness, and the flexural strength was discussed.
Nuclear Gas Turbines: Small-Scale Inherently Safe, Well-Proven Nuclear Power
Authors: van Dam H. , van der Hagen T. H. J. J.
Published: February 2009

The paper discusses uranium as a new fuel for gas turbines used as energy conversion installations for the markets of: stand-alone heat production, combined heat and power generation, stand-alone electricity production and as prime mover on board ships. This development is a logical step in a historical trend in energy conversion. The paper discusses the availability of the fuel, uranium and the construction of the fuel which makes this combination of gas turbine and uranium suitable for the non-utility markets.
Degradation Effects on Combined Cycle Power Plant Performance: Part 3 — Gas and Steam Turbine Degradation Effects
Authors: Zwebek A. I. , Pilidis P.
Published: February 2009

This paper presents an investigation of the degradation effects that gas and steam turbine cycles components have on combined cycle (CCGT) power plant performance. Gas turbine component degradation effects were assessed with TurboMatch, the Cranfield Gas Turbine simulation code. A new code was developed to assess bottoming cycle performance deterioration. The two codes were then joined to simulate the combined cycle performance deterioration as a whole unit. Areas examined were gas turbine compressor and turbine degradation, HRSG degradation, steam turbine degradation, condenser degradation, and increased gas turbine back-pressure due to HRSG degradation. The procedure, assumptions made, and the results obtained are presented and discussed. The parameters that appear to have the greatest influence on degradation are the effects on the gas generator.
Experimental Results of Hot Section Metal Temperature Measurements of Rolls-Royce Allison 501KH Under Massive Steam Injection
Authors: Cheng Dah Yu , Nelson Albert L. C.
Published: February 2009

It has always been thought by the gas turbine industry that steam injection will shorten the effective life of certain gas turbine parts. Recently it was shown that a number of steam injected Cheng Cycle Rolls-Royce Allison 501KH gas turbines, accumulated more than 2.5 million logged hours of operation and with a prolonged parts life. The “hot parts” of a Rolls-Royce Allison 501KH gas turbine engine that are of concern, are the first stage nozzle, the first stage blade, and the second stage nozzle. These parts are all air cooled through the first stages internal passages. (The second stage blades and on down are not internally cooled.) The concern raised in many gas turbine institutions is that the metal temperatures of these hot parts, due to the heat conductivity properties of injected steam, will make them deteriorate faster. An experiment was completed using a steam injected Cheng Cycle, on an Allison 501KH gas turbine engine. In the experiment, a substantial number of thermocouples were attached to the surfaces of the turbines hot parts. This engine had a steam injection rate of up to 18% airflow. The experimental results showed that if steam could be properly mixed with the cooling air before the air enters into the cooling passages of the hot parts, the metal temperatures did not increase. During the operation of the engines, it was recorded that the hot parts lifetime increased from 25,000 hours before the hot parts section had to be overhauled, to 42,000 hours (on average) before they needed to be overhauled. This paper will report the measurement installation in detail. The results before and after steam injection in the hot parts sections of the Rolls-Royce Allison 501KH engine will also be discussed.
Cycle Optimisation Using an Advanced, Real Engine Performance Prediction Model
Authors: Coppinger Miles , Cox Graham , Hannis John , Cox Nigel
Published: February 2009

A whole gas-turbine engine model has been developed incorporating all of the key turbomachinery aerothermal relationships. The aim of the model has been to predict trends in gas-turbine performance with a high degree of confidence that they reflect real engine design limitations. Simple cycles, recuperated, inter-cooled, and inter-cooled recuperated cycles can be assessed across a wide of range of operating parameters. The model is spreadsheet-based with additional macro programming. The major part of it is concerned with establishing representative overall turbine characteristics. A non-integer number of stages is determined as a function of technology level inputs. Individual stage geometry features are derived allowing the calculation of the coolant requirements and efficiencies. The results of various studies are presented for a number of cycle types. The resulting trends are believed to be sensible because of the realistic turbine features. Confidence in the method is established by the modelling of a number of existing industrial gas turbines.
Gas Properties as a Limit to Gas Turbine Performance
Authors: Wilcock R. C. , Young J. B. , Horlock J. H.
Published: February 2009

Although increasing the turbine inlet temperature has traditionally proved the surest way to increase cycle efficiency, recent work suggests that the performance of future gas turbines may be limited by increased cooling flows and losses. Another limiting scenario concerns the effect on cycle performance of real gas properties at high temperatures. Cycle calculations of uncooled gas turbines show that when gas properties are modelled accurately, the variation of cycle efficiency with turbine inlet temperature at constant pressure ratio exhibits a maximum at temperatures well below the stoichiometric limit. Furthermore, the temperature at the maximum decreases with increasing compressor and turbine polytropic efficiency. This behaviour is examined in the context of a two-component model of the working fluid. The dominant influences come from the change of composition of the combustion products with varying air/fuel ratio (particularly the contribution from the water vapour) together with the temperature variation of the specific heat capacity of air. There are implications for future industrial development programmes, particularly in the context of advanced mixed gas-steam cycles.
Film Cooling on a Modern HP Turbine Blade: Part IV — Compound-Angle Shaped Holes
Authors: McGrath E. Lee , Leylek James H. , Buck Frederick A.
Published: February 2009

The performance and physics of film cooling with compound-angle shaped holes on a modern high-pressure turbine airfoil is studied in detail using state-of-the-art computational simulations. Computations model high-speed single-airfoil-passage cascade experiments, and computational results show good agreement with experimental data. Evaluation of physics includes examination of flow features and adiabatic effectiveness. The blowing ratios (M) simulated on the pressure surface (PS) of the blade are 1.5, 3.0, and 4.5, with a single density ratio of 1.52. On the pressure surface the dominant mechanism affecting coolant behavior is vorticity, which increasingly tucks hot crossflow under the coolant as the blowing ratio increases. Thus at high blowing ratios, a lower percentage of the coolant provides thermal protection for the blade until the vortices dissipate far downstream. Also, the vortex structures cause large lateral temperature gradients despite the lateral motion of the flow induced by the compound-angle injection. The dominance of vorticity can be attributed to poor diffusion of the coolant inside the diffuser of the film hole. On the suction surface (SS), the simulated blowing ratios are 1.0, 1.5, and 2.0, with a single density ratio of 1.52. Pressure gradients normal to the SS result in the flow pushing the coolant onto the blade. Also, vorticity is less dominant since diffusion of coolant inside the film hole is better due to low blowing ratios and due to a hole metering section that is almost 3 times longer than that of the PS hole. Hot crossflow ingestion into the film hole is observed at M = 2.0. Ingested crossflow causes heating of the surface inside the hole that extends down to the end of the hole metering section, where the surface temperatures are approximately equal to an average of the crossflow and coolant temperatures. These results demonstrate the inadequacy of 1-D, empirical design tools and demonstrate the need for a validated CFD-based film cooling methodology.
Effects of Aeroderivative Combustor Turbulence on Endwall Heat Transfer Distributions Acquired in a Linear Vane Cascade
Authors: Ames Forrest E. , Barbot Pierre A. , Wang Chao
Published: February 2009

Vane endwall heat transfer distributions are documented for a mock aeroderivative combustion system and for a low turbulence condition in a large-scale low speed linear cascade facility. Inlet turbulence levels range from below 0.7 percent for the low turbulence condition to 14 percent for the mock combustor system. Stanton number contours are presented at both turbulence conditions for Reynolds numbers based on true chord length and exit conditions ranging from 500,000 to 2,000,000. Low turbulence endwall heat transfer shows the influence of the complex three-dimensional flow field, while the effects of individual vortex systems are less evident for the high turbulence cases. Turbulent scale has been documented for the high turbulence case. Inlet boundary layers are relatively thin for the low turbulence case while inlet flow approximates a nonequilibrium or high turbulence channel flow for the mock combustor case. Inlet boundary layer parameters are presented across the inlet passage for the three Reynolds numbers and both the low turbulence and mock combustor inlet cases. Both midspan and 95 percent span pressure contours are included. This research provides a well-documented database taken across a range of Reynolds numbers and turbulence conditions for assessment of endwall heat transfer predictive capabilities.
Combustor Turbine Interface Studies: Part 1 — Endwall Effectiveness Measurements
Authors: Colban W. F. , Thole K. A. , Zess G.
Published: February 2009

Improved durability of gas turbine engines is an objective for both military and commercial aeroengines as well as for power generation engines. One region susceptible to degradation in an engine is the junction between the combustor and first vane given that the main gas path temperatures at this location are the highest. The platform at this junction is quite complex in that secondary flow effects, such as the leading edge vortex, are dominant. Past computational studies have shown that the total pressure profile exiting the combustor dictates the development of the secondary flows that are formed. This study examines the effect of varying the combustor liner film-cooling and junction slot flows on the adiabatic wall temperatures measured on the platform of the first vane. The experiments were performed using large-scale models of a combustor and nozzle guide vane in a wind tunnel facility. The results show that varying the coolant injection from the upstream combustor liner leads to differing total pressure profiles entering the turbine vane passage. Endwall adiabatic effectiveness measurements indicate that the coolant does not exit the upstream combustor slot uniformly but instead accumulates along the suction side of the vane and endwall. Increasing the liner cooling continued to reduce endwall temperatures, which was not found to be true with increasing the film-cooling from the liner.
Combustor Turbine Interface Studies: Part 2 — Flow and Thermal Field Measurements
Authors: Colban W. F. , Lethander A. T. , Thole K. A. , Zess G.
Published: February 2009

Most turbine inlet flows resulting from the combustor exit are non-uniform in the near-platform region as a result of cooling methods used for the combustor liner. These cooling methods include injection through film-cooling holes and injection through a slot that connects the combustor and turbine. This paper presents thermal and flow field measurements in the turbine vane passage for a combustor exit flow representative of what occurs in a gas turbine engine. The experiments were performed in a large-scale wind tunnel facility that incorporates combustor and turbine vane models. The measured results for the thermal and flow fields indicate a secondary flow pattern in the vane passage that can be explained by the total pressure profile exiting the combustor. This secondary flow field is quite different than that presented for past studies with an approaching flat plate turbulent boundary layer along the upstream platform. A counter-rotating vortex that is positioned above the passage vortex was identifed from the measurements. Highly turbulent and highly unsteady flow velocities occur at flow impingment locations along the stagnation line.
Controlling Secondary-Flow Structure by Leading-Edge Airfoil Fillet and Inlet Swirl to Reduce Aerodynamic Loss and Surface Heat Transfer
Authors: Shih T. I.-P. , Lin Y.-L.
Published: February 2009

Computations, based on the ensemble-averaged compressible Navier-Stokes equations closed by the shear-stress transport (SST) turbulence model, were performed to investigate the effects of leading-edge airfoil fillet and inlet-swirl angle on the flow and heat transfer in a turbine-nozzle guide vane. Three fillet configurations were simulated: no fillet (baseline), a fillet whose thickness fades on the airfoil, and a fillet whose thickness fades on the endwall. For both fillets, the maximum height above the endwall is positioned along the stagnation zone/line on the airfoil under the condition of no swirl. For each configuration, three inlet swirls were investigated: no swirl (baseline) and two linearly varying swirl angle from one endwall to the other (+30° to −30° and −30° to +30°). Results obtained show that both leading-edge fillet and inlet swirl can reduce aerodynamic loss and surface heat transfer. For the conditions of this study, the difference in stagnation pressure from the nozzle’s inlet to its exit were reduced by more than 40% with swirl or with fillet without swirl. Surface heat transfer was reduced by more than 10% on the airfoil and by more than 30% on the endwalls. When there is swirl, leading-edge fillets became less effective in reducing aerodynamic loss and surface heat transfer, because the fillets were not optimized for swirl angles imposed. Since the intensity and size of the cross flow were found to increase instead of decrease by inlet swirl and by the type of fillet geometries investigated, the results of this study indicate that the mechanisms responsible for aerodynamic loss and surface heat transfer are more complex than just the intensity and the magnitude of the secondary flows. This study shows their location and interaction with the main flow to be more important, and this could be exploited for positive results.
A Novel Transient Liquid Crystal Technique to Determine Heat Transfer Coefficient Distributions and Adiabatic Wall Temperature in a Three Temperature Problem
Authors: Chambers Andrew C. , Gillespie David R. H. , Ireland Peter T. , Dailey Geoffrey M.
Published: February 2009

Transient liquid crystal techniques are widely used for experimental heat transfer measurements. In many instances it is necessary to model the heat transfer resulting from the temperature difference between a mixture of two gas streams and a solid surface. To non-dimensionally characterise the heat transfer from scale models it is necessary to know both the heat transfer coefficient and adiabatic wall temperature of the model. Traditional techniques rely on deducing both parameters from a single test. This is a poorly conditioned problem. A novel strategy is proposed in which both parameters are deduced from a well conditioned three test strategy. The heat transfer coefficient is first calculated in a single test; the contribution from each driving gas stream is then deduced using additional tests. Analytical techniques are developed to deal with variations in the temperature profile and transient start time of each flow. The technique is applied to the analysis of the heat transfer within a low aspect ratio impingement channel with initial cross flow.
Liburdi Power Metallurgy: New Compositions for High Strength Repairs of Turbine Components
Authors: Sparling Rob , Liburdi Joseph
Published: February 2009

Over the past five years the Liburdi Powder Metallurgy (LPM™) process for repair, overlay or joining of nickel and cobalt based alloys has been improved through the addition of new alloy combinations. In addition to the original IN738™ based LPM™, two MarM247™ based compositions have been used for repair and manufacture of turbine blades and vanes. These new compositions show improved properties over the original IN738™ material. The new LPM™ compositions can fill gaps up to 30 mm wide and 6.3mm deep per application. These new compositions can also be used for surface build-ups up to 0.25m2 at 6.3mm thick per application. This ability to fill large gaps and create large surface build-ups makes these new alloy compositions very flexible tools for manufacture and repair of turbine components. The structure of these new allow combinations along with stress rupture, and tensile properties are presented in this paper. Examples of typical LPM™ applications and performance are presented.
A Novel Method for the Performance Control of a Gas Transmission Compressor
Authors: Pearson W. N. , Armitage A. F. , Henderson D. S.
Published: February 2009

This paper presents the application of feed forward neural networks to the performance control of a gas transmission compressor. It is estimated that a global saving in compressor fuel gas of 1% could prevent the production of 6 million tonnes of CO2 per year, [1]. Results of compressor model testing suggest that compressor speed can be estimated to within ± 2.5%. The neural network property of function approximation is used to predict compressor speed for given process constraints and instrument input sets. The effects of training set size, instrument noise, reduced input sets and extrapolation from the training domain, are quantified. Various neural network architectures and training schema were examined. The embedding of a neural network into an expert system is also discussed.
Flow Physics and Vortex Evolution in Annular Turbine Cascades
Authors: Treiber M. , Abhari R. S. , Sell M.
Published: February 2009

The evolution of flow vortices downstream of annular cascades with a specific focus upon the association between vorticity and the mechanisms of aerodynamic loss generation has been experimentally investigated. Spatially-resolved experimental results on an unstructured measurement pattern is used to identify small scale flow structures on the surface of the investigated blades and downstream of two different cascades. The test case comparison is based upon two moderately loaded turbine blade profiles, one stacked prismatically and the other on a bowed stacking line. The design operating conditions for both cascade is at a nominal exit Mach number of 0.5. The measurements were made using an ultra miniature (0.9 mm Dia) 5-hole probe. The measurement were made on 4 planes behind and one plane in front of the trailing and leading edge of the cascade respectively. The experimental results clearly show the influence of the flow structure on the evolution of the stream wise vorticities behind the blading. Detailed measurements are used to compare and evaluate the mixing processes downstream of these two cascades. It is shown that the interaction of the different vortices with the endwall boundary layer has a substantial impact on the overall total pressure loss generation.
A Turbomachinery Blade Design and Optimization Procedure
Authors: Xu C. , Amano R. S.
Published: February 2009

With the development of the advanced technology, the combustion temperature is raised for increased efficiencies. At the same time, the turbine and compressor pressure ratio and the mass flow rate rise; thus causing turbine and compressor blades turning and blade lengths increase. Moreover, the high efficiency requirements had made the turbine and compressor blade design difficult. A turbine airfoil has been custom designed for many years, but an optimization for the section design in a three-dimensional consideration is still a challenge. For a compressor blade design, standard section cannot meet the modern compressor requirements. Modern compressor design has not only needs a custom designed section according to flow situation, but also needs three-dimensional optimizations. Therefore, a good blade design process is critical to the turbines and compressors. A blade design of the turbomachines is one of the important steps for a good turbomachine design. A blade design process not only directly influences the overall machine efficiency but also dramatically impact the design time and cost. In this study, a blade design and optimization procedure was proposed for both turbine and compressor blade design. A compressor blade design was used as a test case. It was shown that the current design process had more advantages than conventional design methodology.
Low Pressure Turbine Lapse Rate Study: CFD Model vs. Rig Data
Authors: Liu J.-S. , Celestina M. L. , Heitland G. B. , Bush D. B. , Mansour M. L. , Adamczyk J. J.
Published: February 2009

As an aircraft engine operates from sea level take-off (SLTO) to altitude cruise, the low pressure (LP) turbine Reynolds number decreases. As Reynolds number is reduced the condition of the airfoil boundary layer shifts from bypass transition to separated flow transition. This can result in a significant loss. The LP turbine performance fall-off from SLTO to altitude cruise, due to the loss increase with reduction in Reynolds number, is referred to as a lapse rate. A considerable amount of research in recent years has been focused on understanding and reducing the loss associated with the low Reynolds number operation. A recent 3-1/2 stage LP turbine design completed a component rig test program at Honeywell. The turbine rig test included Reynolds number variation from SLTO to altitude cruise conditions. While the rig test provides detailed inlet and exit condition measurements, the individual blade row effects are not available. Multi-blade row computational fluid dynamics (CFD) analysis is used to complement the rig data by providing detailed flow field information through each blade row. A multi-blade row APNASA model was developed and solutions were obtained at the SLTO and altitude cruise rig conditions. The APNASA model predicts the SLTO to altitude lapse rate within 0.2 point compared to the rig data. The global agreement verifies the modeling approach and provides a high confidence level in the blade row flow field predictions. Additional Reynolds number investigation with APNASA will provide guidance in the LP turbine Reynolds number research areas to reduce lapse rate. To accurately predict the low Reynolds number flow in the LP turbine is a challenging task for any computational fluid dynamic (CFD) code. The purpose of this study is to evaluate the capability of a CFD code, APNASA, to predict the sensitivity of the Reynolds number in LP turbines.