bluuebunny/crossref_metadata_2025 huggingface dataset:
Effects of Using Hydrogen on Aero Gas Turbine Pollutant Emissions, Performance and Design
Authors: Svensson Fredrik , Singh Riti
Published: November 2008

The majority of scientists today are in agreement that discharging greenhouse gases into the atmosphere has an impact on the global climate. Furthermore, the dwindling fossil oil resources raise concerns. Finding an alternative to kerosene for civil aviation is thus essential. Taking a completely new approach for civil aviation by fuelling with liquid hydrogen would enable compatibility with the environment. This paper covers the main effects on aero engines when changing to hydrogen fuel. Particularly, emission and performance issues are discussed, but some design matters are also briefly covered. The study shows that burning hydrogen in an aero gas turbine seems to be feasible from a technical point of view. By simply changing to hydrogen, small performance gains may be obtained. The results of the calculations suggest that there is the potential to design a combustion system using hydrogen that produces less NOx emissions than any system using kerosene.
Flowfield Measurements in a Ribbed Channel Relevant to Internal Turbine Blade Cooling
Authors: Graham A. , Sewall E. , Thole K. A.
Published: November 2008

The internal cooling of turbine blades is an important part of overall blade cooling that is necessary to survive current turbine inlet temperatures. This internal cooling is often performed by means of rib-turbulated passages designed to provide maximum heat transfer coefficients. One determining factor in the cooling effectiveness is the flow pattern created by the ribbed channel. This paper presents a comparison of the effect of two different channel blockage ratios on the fully developed flow pattern in a stationary ribbed channel. For the experimental investigation, a scaled-up ribbed passage was tested using of a closed-loop channel facility. The facility included a passage consisting of a developing region, a fully developed region, a 180° bend, and a second developing region following the bend. Friction factors were determined through the use of pressure taps throughout the channel, and flow fields were measured using a two-component laser Doppler velocimeter. A majority of the testing performed was done under conditions of in-line (i.e. non-staggered) ribs with ratios of rib height to channel hydraulic diameter of 0.10 and 0.17. All LDV measurements were performed at a Reynolds number of 20,000. Results indicate that both blockage ratios create significant recirculation regions immediately before and immediately following the rib. While the smaller rib results indicate a reattachment downstream of the rib, the larger rib results indicate that the flow is separated along the entire streamwise distance between the ribs. The smaller blockage ratio also indicated a small recirculation region directly on top of the rib, whereas the larger blockage ratio showed no such region. Additionally, the increased friction factor created by the larger ribs can be attributed to the larger recirculation regions between ribs.
Laser Based Investigations of Periodic Combustion Instabilities in a Gas Turbine Model Combustor
Authors: Giezendanner R. , Weigand P. , Duan X. R. , Meier W. , Meier U. , Aigner M. , Lehmann B.
Published: November 2008

The driving mechanism of pulsations in gas turbine combustors depends on a complex interaction between flow field, chemistry, heat release, and acoustics. Experimental data on all these factors are therefore required to obtain insight into the coupling mechanisms during a pulsation period. In order to develop a comprehensive experimental data base to support a phenomenological understanding and to provide validation data for numerical simulation, a standard burner for optical investigations was established that exhibits strong self-excited oscillations. The burner was a swirl-stabilized non-premixed model combustor designed for gas turbine applications and operated using methane as fuel at atmospheric pressure. It was mounted in a combustion chamber which provides almost unobstructed optical access. The periodic combustion instabilities were studied by a variety of phase-resolved laser based diagnostic techniques, locked to the frequency of the dominant pressure oscillation. Measurement techniques used were LDV for velocity measurements, planar laser-induced fluorescence for imaging of CH and OH radicals, and laser Raman scattering for the determination of the major species concentrations, temperature, and mixture fraction. The phase-resolved measurements revealed significant variations of all measured quantities in the vicinity of the nozzle exit, which trailed off quickly with increasing distance. A strong correlation of heat release rate and axial velocity at the nozzle was observed, while the mean mixture fraction as well as the temperature in the periphery of the flame is phase-shifted with respect to axial velocity oscillations. A qualitative interpretation of the experimental observations is given, which will help to form a better understanding of the interaction between flow field, mixing, heat release, and temperature in pulsating reacting flows, particularly when accompanied by corresponding CFD simulations which are currently under way.
On the Use of 5-Hole Probes in the Testing of Industrial Centrifugal Compressors
Authors: Gilarranz Jose´ L. , Ranz Andrew J. , Kopko Jason A. , Sorokes James M.
Published: November 2008

This paper addresses the use of 5-hole probes in the testing of industrial centrifugal compressors. The 5-hole probes utilized for this work are of the conical-tip type and were used in a non-nulling configuration (i.e. the probes do not need to be rotated or moved in any way during the tests). These 5-hole probes proved to be fairly robust, making them practical for a non-laboratory setting such as an industrial multistage compressor test stand. A discussion of 5-hole probes and how they function is provided, including an overview of the mathematical formulations and calibrations required to translate the pressure data gathered from the 5 holes into static and total pressures, velocities and flow angles. A method to transform these variables from a probe-based coordinate system to a machine-based coordinate system is also presented and schematics of this process are provided to aid the reader’s understanding. The testing performed on a prototype multistage centrifugal compressor using 5-hole probes is also discussed showing that the probes provided valuable insight into the flowfield exiting the impellers and at the return bend. The hub-to-shroud velocity profile exiting an impeller was found to be more skewed than expected and was contributing to poor performance in the downstream stationary components. The measured flowfield from one of the tests is also compared against 3-D CFD results and comments are offered regarding the agreement between the analytical and measured results. Advantages and disadvantages of 5-hole probes as compared to more conventional instrumentation are presented. Finally, conclusions are drawn regarding the value of 5-hole probe data in the development and/or troubleshooting of high performance turbomachinery and in the validation / calibration of design and analysis tools.
Set Up and Operational Experience With a Micro-Turbine Engine for Research and Education
Authors: Davison Craig R. , Birk A. M.
Published: November 2008

Set up and operation of a mid to large size gas turbine requires a significant investment in the engine and the test cell. The continuing cost of operation and maintenance is also substantial. Both the capital and operating costs are well outside the budget of many educational institutions, and small research centres. Micro-turbines, in particular engines used for model aircraft, are a viable alternative. Their capital costs are low and existing facilities can often be modified to support them. Micro-turbines are complete turbojets often utilizing a centrifugal compressor from a turbo-charger and an axial turbine. This work utilized a micro-jet engine commonly used to propel remote control aircraft. The engine has been set up for research into component degradation and as a laboratory for an upper year engineering course in turbomachinery. It is rated by the manufacturer to produce 150 N of thrust at 132,000 RPM. This paper examines some of the problems encountered in installing small total pressure and temperature probes in a micro engine. The entire test rig, including measurement of thrust and mass flow is presented. As well, the addition of a PC based control system is discussed. Operating data is presented and compared to larger engines to demonstrate the viability of this engine as a test bed. Some problems encountered in using an engine such as this beyond its normal operating envelope, along with some solutions, are presented.
Steady State and Transient Modeling of a Micro-Turbine With Comparison to Operating Engine
Authors: Davison Craig R. , Birk A. M.
Published: November 2008

Steady state and transient computer models of a micro turbine were produced. The engine under study was a micro-jet engine that when tested at 126,000 RPM provided 95 N thrust. The aero-thermal model uses generic performance maps for the compressor and turbine which were modified, based on operating data, to represent the components in the engine under study. The model also includes the inlet ducting connected to the engine. It simulates engine operation from idle to full power over the expected operating range of ambient temperature, pressure and humidity. A comparison of steady state model results to actual engine operating data is presented over the full range of speeds. The effect of ambient humidity on the engine operating point is examined for a micro-engine, in particular at temperatures above 30° Celsius. The techniques for introducing component faults are given and their effect on the engine operation is presented. The degraded components are the turbine and inlet flow passages. The methods for modeling the transient behavior of the engine are also presented. Results are presented for both acceleration and deceleration of the engine between steady state operating point. These results are also compared to the operating engine.
Probabilistic System Reliability for a Turbine Engine Airfoil
Authors: Beachkofski Brian K. , Grandhi Ramana V.
Published: November 2008

A probabilistic assessment for an integrally bladed disk design is a system reliability problem where a failure in at least one blade constitutes a system failure. Turbine engine fan and compressor blade life is dominated by High Cycle Fatigue (HCF) which is initiated either by pure HCF or Foreign Object Damage (FOD). The system model includes frequency prediction, model stress variation, mistuning amplification, FOD effect, and random material capability. Although the analysis time for these submodels would have ruled out any sampling method as recently as two years ago, novel reduced order models have made this process tractable. A large system problem has more random variables than some probability integration methods can handle; this system has nearly 700 variables, 29 for each of the 24 blades. Sampling methods, such as the Latin Hypercube Sampling or Monte Carlo used in this analysis, can handle such large numbers of variables easily. Additionally, both methods construct a confidence interval around the most likely failure rate estimate. The results converge to the same statistic with equal confidence intervals for all analysis methods.
Analytical Approach to Steam Turbine Heat Transfer in a Combined Cycle Power Plant
Authors: Brilliant Howard M. , Tolpadi Anil K.
Published: November 2008

Combined cycle units have become very popular in recent years as a source of power generation. Such units have a gas turbine as the topping cycle and a steam turbine as the bottoming cycle and can reach combined cycle efficiencies as high as 60%. The exhaust from the gas turbine is passed through a heat exchanger in which steam is generated for the steam turbine. This combined arrangement makes it less polluting as well. An important element of a combined cycle power plant is the steam turbine, which is the subject of this paper. Improvements to the design of advanced steam turbines require an improved understanding of the heat transfer within the various components of the unit. Physics-based ANSYS models for typical GE high pressure and intermediate pressure units have been developed. Components such as the rotor, diaphragm, and shells have been analyzed. The boundary conditions were derived from full-load, steady state flow analyses, steam turbine performance code outputs and computational fluid dynamics (CFD) analyses to develop normalized (non-dimensional) local flow conditions, with the normalizing parameters based on key cycle parameters. These normalized local flow conditions and cycle parameters were then used to define local transient boundary temperatures and heat transfer coefficients for input to the thermal ANSYS models. Transient analyses of components were performed. The results were compared with temperature measurements taken during the complete cycle of an operational steam turbine to validate and improve the methodology, and were applied to structural models of the components to predict their thermal growth and the net impact on the clearance between the rotor and diaphragms and other secondary flow paths in the steam turbine, including the packing seals. This paper will focus on the thermal modeling of a typical steam turbine. It will discuss the process used (summarized above) and the basic equations employed in the analyses. Results will be compared with shell temperature measurements obtained during the start up of a steam turbine in the field. Implications of the thermal results on power systems operation will be discussed. Plans for future improvements will be presented.
A Combustion Control and Diagnostics Sensor for Gas Turbines
Authors: Thornton Jimmy D. , Straub Douglas L. , Chorpening Benjamin T. , Huckaby E. David , Richards Geo. A. , Benson Kelly
Published: November 2008

The implementation of sophisticated combustion control schemes in modern gas turbines is motivated by the desire to maximize thermodynamic efficiency while meeting NOx emission restrictions. To achieve target NOx levels, modern turbine combustors must operate with a finely controlled fuel-air ratio near the fuel-lean flame extinction limit, where the combustor is most susceptible to instabilities. In turbine configurations with multiple combustors arranged around the annulus, differences in flow splits caused by manufacturing variations or engine wear can compromise engine performance. Optimal combustion control is also complicated by changes in environmental conditions, fuel-quality, or fuel-type. As a consequence, engines must be commissioned in the field with adequate stability margin such that manufacturing tolerances, normally expected component wear, fuel-quality, and environmental conditions will not cause unstable combustion. A lack of robust combustion in-situ monitoring has limited the ability of modern turbines to achieve stable ultra-low emission performance over the entire load range. This paper describes a combustion control and diagnostics sensor (CCADS) that can potentially revolutionize the manner in which modern gas turbines are controlled. This robust sensor uses the electrical properties of the flame to detect key events and monitor critical operating parameters within the combustor. The CCADS is integrated into the fuel nozzle such that low cost and long life are achieved. Tests conducted at turbine conditions in laboratory combustors instrumented with CCADS have demonstrated the following potential capabilities: 1) detection of incipient flashback and autoignition 2) detection of incipient lean blowout 3) detection of dynamic pressure oscillations 4) and a qualitative measure of equivalence ratio within the combustor. Many of these capabilities have been reported in other publications with data from an atmospheric combustion rig. This paper will summarize each of the capabilities with recent data at turbine conditions. The expectation is that CCADS will provide the key in-situ monitoring for diagnostics and control of modern gas turbines, allowing them to achieve stable ultra-low emissions performance.
Application of CFD-Based Analysis Technique for Design and Optimization of Gas Turbine Combustors
Authors: Koutsenko I. G. , Onegin S. F. , Sipatov A. M.
Published: November 2008

The design and operational development of gas turbine combustors is a complex process, involving a great volume of design and experimental work. The application of computational fluid dynamics (CFD) methods allows to lower the volume of experimental works on operational development of combustors and to make changes to the design of combustion chambers on early design stages. In this paper the application of commercial CFD package CFX-TASCflow for calculation of flow structure and analysis of nitric oxide formation process in the combustion chamber of the PS-90A gas turbine and its modifications is considered. The results of the analysis show, that the basic determinative criterion of a nitric oxide emission level is the residence time of a combustion products in high-temperature zones. With help of this criterion, an optimization of the PS-90A combustion chamber was performed. A design of an optimized combustion chamber allows to achieve a low level of nitric oxide emissions.
Suppression of Unstable Flow at Small Flow Rates in a Centrifugal Blower by Controlling Tip Leakage Flow and Reverse Flow
Authors: Ishida Masahiro , Surana Taufan , Ueki Hironobu , Sakaguchi Daisaku
Published: November 2008

The effects of the inlet recirculation arrangement on inducer stall and the diffuser width on diffuser stall in a high specific speed type centrifugal impeller with inducer were analyzed by a numerical simulation and also verified experimentally. It is found that the incipient unstable flow occurs due to a rolling-up vortex flow resulting from an interaction between the tip leakage flow and the reverse flow accumulated at the pressure side immediately downstream of the inducer tip throat in which a strong streamwise component of vorticity is included. By forming the inlet recirculation flow, the tip leakage vortex is effectively sucked into the suction ring groove, and the flow incidence is decreased simultaneously. The unstable flow range of the test blower was reduced significantly by about 45% without deteriorating the impeller characteristics by implementing optimally both the ring groove arrangement and the narrowed diffuser width.
Dynamics of a Spinning Hollow Rayleigh Beam in General Boundary Conditions
Authors: Sheu Geeng Jen
Published: November 2008

The analytical solutions of the whirl speed, the critical speed and the mode shape of a spinning hollow beam in six general boundary conditions are presented in this paper. The beam is represented by a Rayleigh model with the rotatory moment inertia and the gyroscopic effects. It is shown that the critical speed be written in a function of the whirl ratio (λ) defined by the spinning speed over the whirl speed, the hollowness ratio (α) defined by the hollow area over the total area of the cross section, and the slenderness ratio (l) defined by the beam’s length over its outer radius. Only a finite number of critical speeds exist, and hence finite precessional modes can be found during overall operation ranges when λ > 1/2, and the number is independent of the boundary conditions. The steady state unbalanced response can be therefore expressed analytically in terms of the finite precessional modes in resonance.
Performance Assessment of Turbulence Models for the Quantitative Prediction of Tip Leakage Flow in Turbomachines
Authors: Lee Gong Hee , Park Jong Il , Baek Je Hyun
Published: November 2008

It is experimentally well-known that high anisotropy of turbulent flow field, which results from the complex shear and rotation effect, is dominant inside tip leakage vortex (TLV). With all this fact, the greater part of the numerical studies to deal with TLV have used an isotropic eddy viscosity model (EVM), and their predictions showed some disagreement with the measurement data. The main objective of the present study is to show the superior ability of the Reynolds stress model (RSM), which can naturally consider the effect of system rotation on turbulence via the rotational production term, over the isotropic EVM for predicting appropriately the complex tip leakage flow in axial type of turbomachines. To achieve this aim, the results obtained from steady-state Reynolds averaged Navier-Stokes simulations based on the Spalart-Allmaras model, Renormalization Group (RNG) k-ε model and RSM are compared with the experimental data for two test configurations: a linear compressor cascade and a forward-swept axial-flow fan. This comparative study of turbulence models suggests that the RSM should be used to predict reasonably the complex tip leakage flow, especially in a rotating environment.
Lean Blowout Limit and NOx-Production of a Premixed Sub-ppm NOx Burner With Periodic Flue Gas Recirculation
Authors: Kalb Jochen R. , Sattelmayer Thomas
Published: November 2008

The technological objective of this work is the development of a lean-premixed burner for natural gas. Sub-ppm NOx emissions can be accomplished by shifting the lean blowout limit (LBO) to slightly lower adiabatic flame temperatures than the LBO of current standard burners. This can be achieved with a novel burner concept utilizing periodic flue gas recirculation: Hot flue gas is admixed to the injected premixed fresh mixture with a mass flow rate of comparable magnitude, in order to achieve self-ignition. The subsequent combustion of the diluted mixture again delivers flue gas. A fraction of the combustion products is then admixed to the next stream of fresh mixture. This process pattern is to be continued in a cyclically closed topology, in order to achieve stable combustion of e.g. natural gas in a temperature regime of very low NOx production. The principal ignition behavior and NOx production characteristics of one sequence of the periodic process was modeled by an idealized adiabatic system with instantaneous admixture of partially or completely burnt flue gas to one stream of fresh reactants. With the CHEMKIN-II package a reactor network consisting of one perfectly stirred reactor (PSR, providing ignition in the first place) and two plug flow reactors (PFR) has been used. The effect of varying burnout and the influence of the fraction of admixed flue gas have been evaluated. The simulations have been conducted with the reaction mechanism of Miller and Bowman and the GRI-Mech 3.0 mechanism. The results show that the high radical content of partially combusted products leads to a massive decrease of the time required for the formation of the radical pool. As a consequence, self-ignition times of 1 ms are achieved even at adiabatic flame temperatures of 1600 K and less, if the flue gas content is about 50%–60% of the reacting flow after mixing is complete. Interestingly, the effect of radicals on ignition is strong, outweighs the temperature deficiency and thus allows stable operation at very low NOx emissions.
A Review of L20A Engine Design and Field Operating Experience
Authors: Sugimoto Takao , Nagai Katsushi , Ryu Masanori , Tanaka Ryozo , Kimura Takeshi
Published: November 2008

Kawasaki Heavy Industries launched the L20A gas turbine, rated at 18MW, in 2001. The design philosophy adopted for the turbine includes a high efficiency transonic axial-flow compressor with eight can-type combustors and a high turbine inlet temperature of 1250 Deg C. This results in a thermal efficiency of 35%, an electric efficiency of 50% for combined cycle power plants and an overall thermal efficiency of 81% for cogeneration systems. In addition, the NOx emissions from the combustor are lower than 23 ppm and the engine has a long service life. These features permit long-term continuous operations under various environmental limitations. Details are presented in 2002-GT-30255.(1) The first commercial unit has been in operation as a daily start and stop cogeneration plant in Kawasaki’s Akashi Works in Japan since October 2001. Accumulated operation hours are 5500 hours and 410 starts as of March 2004. Reliability higher than 99% has been demonstrated during this period. During the shop test and the commercial operation, some additional improvements have been developed for the compressor, turbine and combustor.
Studies on Endwall Treatment for the Generation of High Incidence Flow and Its Influence on Flow Structures in an Annular Compressor Cascade
Authors: Weidenfeller Jo¨rg , Lawerenz Martin
Published: November 2008

The flow in the endwall region of axial turbo compressors is of particular interest with respect to the development of losses and vortices. The relative movement between the rotating hub and the stator causes high incidence flow in the hub region, which leads to an overturning and distinctive vortex systems. In order to analyze the influence of the inlet flow on three dimensional flow effects, experimental investigations were performed in an isolated compressor stator for a Mach number of Ma = 0.35. The annular cascade consists of prismatic blades with a small hub clearance (0.8% of blade chord) and a solidity of unity at the mean radius. The cylindrical flow path has a hub to tip ratio of 0.67. In contrast to investigations in former published work, where the inlet flow angle near the endwall was influenced by a rotating hub, in this work active and passive methods were used: The flow angle distribution in the lower half of the passage can be altered by endwall treatment (passive) and a bleed-air extraction at the casing upstream of the investigated stator. Extensive flow field measurements at the stator inlet and outlet were performed with pneumatic five hole and three hole probes. In addition the distribution of the static wall pressure at the stator passage was measured. To identify the blade loading both the flow fields and the pitch averaged flow quantities are analyzed. The results indicate the influence of different inlet swirl flows on the development of loss and vortex regions.
Optimization of Microturbine Aerodynamics Using CFD, Inverse Design and FEM Structural Analysis: 1st Report — Compressor Design
Authors: Ashihara Kosuke , Goto Akira , Guo Shijie , Okamoto Hidenobu
Published: November 2008

In this paper, a new aerodynamic design procedure is presented for a centrifugal compressor stage of a microturbine system. To optimize the three-dimensional (3-D) flows and the performance, an inverse design method, which numerically generates the 3-D blade geometry for specified blade loading distribution, has been applied together with the numerical validation using CFD (Computational Fluid Dynamics) and FEM (Finite Element Method). The blade profile along the shroud surface of the impeller was optimized based on the 3-D inverse design and CFD. However, the blade profile towards the hub surface was modified geometrically to achieve a nearly radial blade element especially at the inducer part of the impeller, in order to meet the required structural strength. The modified impeller successfully kept similar aerodynamic performance as that of a blade with a fully 3-D shape, whilst showing improved structural reliability. So, the proposed method to adopt the blade profile designed by the inverse method along the shroud, and to geometrically modify the blade profile towards the hub, was confirmed to be effective to design a high-speed compressor impeller. The vaned diffuser has also been re-designed using the inverse design method. The corner separation in the conventional wedge-type diffuser channel was suppressed in the new design. The stage performance improvements were confirmed by stage calculations using CFD.
Accelerating the Calibration of Multi-Hole Pressure Probes by Applying Advanced Computational Methods
Authors: Franken Arnoud R. C. , Ivey Paul C.
Published: November 2008

The miniature multi-hole pneumatic pressure probe is widely regarded as a cost-effective, easy-to-use and accurate method for performing two- or three-dimensional flow field measurements in turbomachinery. The major downside to the use of these probes is that the influence of fabrication imperfections on probe characteristics necessitates an extensive and highly time-intensive and, therefore, costly calibration of each individual probe. Unless these probes can be fabricated to such standards that make individual probe calibrations superfluous, the only way to significantly reduce the time and costs associated with probe calibration is to shorten the calibration process. The latter is only possible if all essential information can be obtained from less calibration data. This paper describes a novel approach to the calibration of a series of multi-hole pressure probes in which advanced computational methods are used to make this possible. By exploiting the key features of a probe’s characteristic this approach requires only a fraction of the size of a conventional calibration database for the accurate modeling of the relationships between port pressures and flow conditions. As a result, calibration time and costs can be reduced without the sacrifice of quality.
Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller
Authors: Higashimori Hirotaka , Hasagawa Kiyoshi , Sumida Kunio , Suita Tooru
Published: November 2008

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and CFD. This report presents the measured results of the high transonic flow at the impeller inlet using LDV and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460m/s and the relative Mach number reaches about 1.6. Using an LDV, about 500m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.
Time-Averaged and Time-Resolved Heat Flux Measurements on a Turbine Stator Blade Using Two-Layered Thin-Film Gauges
Authors: Iliopoulou V. , De´nos R. , Billiard N. , Arts T.
Published: November 2008

This paper describes the steps undertaken to measure heat flux in a turbine tested in a blowdown windtunnel when using a two-layered thin film gauge array. The sensor consists of a nickel thermo resistor deposited onto a flexible polymide sheet that can be easily bounded on a substrate using double sided adhesive. The assembly constitutes a two-layered system. First, a numerical algorithm is proposed to extract the wall heat flux from the surface temperature history measured by the thin film gauge. It is very flexible and handles multi-layered systems. Then, an original procedure is proposed to determine the thermal properties and the thickness of the different layers. It uses the above numerical algorithm coupled with a minimization routine. The repeatability of the procedure is assessed. Finally, tests are processed according to the proposed method. The results are successfully compared with measurements performed with single-layered thin film gauges.
Squeeze-Film Damper Predictions for Simulation of Aircraft Engine Rotordynamics
Authors: Defaye Cyril , Laurant Franck , Carpentier Philippe , Arghir Mihai , Bonneau Olivier , Colboc Samuel
Published: November 2008

On aircraft engines, a common recurring problem is excessive vibration levels generated by unbalance. With rotors mounted on usual undamped ball bearings, an amount of damping is required to limit peak amplitudes at traversed critical speeds: a solution is to introduce external damping with squeeze-film dampers. Such dampers can be added with minor modifications of the rotor system design. This paper presents experimental and theoretical work in progress focused on the analysis of squeeze film dampers (SFD) based on serial aircraft engines design. Several squeeze-film geometries were tested to measure the influence of different design parameters as the fluid clearance and the groove feeding system. Next, a damper model based on the numerical solution of the Reynolds equation is correlated with the experimental data to obtain predictive global forces. It is shown that the theoretical model is a good predictive tool if it is correctly adjusted and if temporal inertia forces are negligible. The present damper model is further compared with analytical models taken from the literature which are obviously more appropriate to be used in whole engine rotordynamic analysis. The limits of the models are then underlined by comparisons with experimental results.
Resultant Benefits of Standardized Overhaul Packages for LM2500 Propulsion Gas Turbine Engines in U.S. Navy Applications
Authors: Driscoll Matthew J. , Picozzi Joseph
Published: November 2008

This paper discusses the Unites States Navy’s program to standardize repair and overhaul packages/workscopes for their LM2500 propulsion gas turbine engines. The General Electric LM2500 gas turbine engine is utilized for main propulsion aboard the Navy’s newest surface combatants including the FFG 7, DD 963, CG 47 and DDG 51 class ships. The Navy employs a condition based maintenance philosophy for its fleet of 450 LM2500 engines; removing engines from ships only when in place corrective actions can no longer be effected. Consequently, most LM2500 gas generators and power turbines have exhausted much of their useful life once they arrive at the depot for overhaul. Beginning in 1999, NAVSEA implemented a standardized workscope for these engines to ensure post repair life and performance goals were achieved. This paper discusses the contents of the standardize repair package and the resultant benefits and metrics associated with its execution at both military and commercial facilities.
Effect of the Hub Endwall Cavity Flow on the Flow-Field of a Transonic High-Pressure Turbine
Authors: Paniagua G. , De´nos R. , Almeida S.
Published: November 2008

In high-pressure turbines, a small amount of cold flow is ejected at the hub from the cavity that exists between the stator and the rotor disk. This prevents the ingestion of hot gases into the wheel-space cavity, thus avoiding possible damage. This paper analyses the interaction between the hub-endwall cavity flow and the mainstream in a high-pressure transonic turbine stage. Several cooling flow ratios are investigated under engine representative conditions. Both time-averaged and time-resolved data are presented. The experimental data is successfully compared with the results of a 3D steady Navier-Stokes computation. Despite the small amount of gas ejected, the hub-endwall cavity flow has a significant influence on the mainstream flow. The Navier-Stokes predictions show how the ejected cold flow is entrained by the rotor hub vortex. The time-resolved static pressure field around the rotor is greatly affected when traversing the non-uniform vane exit flow field. When the cavity flow rate is increased, the unsteady forces on the rotor airfoil are reduced. This is linked to the decrease of vane exit Mach number caused by the blockage of the ejected flow.
A Low-Re RSTM Model for Computations of Heat Transfer and Fluid Flow for Impingement and Convective Cooling
Authors: Jia Rongguang , Sunde´n Bengt
Published: November 2008

A new Reynolds stress transport model (RSTM) aimed for engineering applications is proposed with consideration of near-wall turbulence. This model employs the SSG pressure strain term, the ω equation, and the SST model for the shear stresses at the near-wall region (say y+ < 30). The models are selected based on the following merits: The SSG RSTM model performs well in the fully turbulent region and does not need the wall normal vectors; the ω equation can be integrated down to the wall without damping functions; The SST model is a proper two-equation model that performs well for flows with adverse pressure gradient, while most two-equation models can have a good prediction of the shear stresses. A function is selected for the blending of the RSTM and SST. Three cases are presented to show the performance of the present model: (1) fully developed channel flow with Reτ = 395, (2) backward-facing step with an expansion ratio of 1.2 and Re = 5,200 base on the step height, (3) circular impingement with the nozzle-to-wall distance H = 4D and Re = 20,000.
Metallurgical Analysis of Rainbow Rotor Coatings: Analysis of Fleet Blades
Authors: Neff Robert A. , Katz Gerald B. , Nagaraj B. , Tarvin Rachel
Published: November 2008

In a continuing effort to increase component lifespan and decrease overhaul cost, the US Navy has completed a 2nd phase of the Rainbow Rotor project. This project, initiated in the early 1990’s, consists of three LM2500 main propulsion engine, high pressure turbines (HPT) built up with refurbished blade pairs protected by various coatings. This turbine was operated for over 7,000 hours on a Cruiser-class ship where it was subjected to a typical operating profile. Six coatings were examined ranging from differing chemical compositions to application processes. The coating compositions were of four types, CoCrAlHf, PtAl, Physical Vapor Deposition (PVD) Zirconia thermal barrier coating (TBC) with a PtAl bond coat and a silicon aluminide type coating. The BC-22 (CoCrAlHf) overlay coatings were applied by either a plasma spray process or an electroplating process. The PtAl coatings, supplied by two vendors, and the TBC were applied by standard commercial processes. The goal behind this study is to find a coating that has the best balance between cost and performance. With the already realized cost savings in using refurbished components to overhaul the gas turbine engine, the emphasis is now placed on delaying the deterioration of the reprocessed blade pairs. The following discussion covers all aspects of this completed phase.
Effect of Clocking on the Second Stator Pressure Field of a One and a Half Stage Transonic Turbine
Authors: Gadea J. , De´nos R. , Paniagua G. , Billiard N. , Sieverding C. H.
Published: November 2008

This paper focuses on the experimental investigation of the time-averaged and time-resolved pressure field of a second stator tested in a one and a half stage high-pressure transonic turbine. The effect of clocking and its influence on the aerodynamic and mechanical behaviour are investigated. The test program includes four different clocking positions, i.e. relative pitch-wise positions between the first and the second stator. Pneumatic probes located upstream and downstream of the second stator provide the time-averaged component of the pressure field. For the second stator airfoil, both time-averaged and time-resolved surface static pressure fields are measured at 15, 50 and 85% span with fast response pressure transducers. Regarding the time-averaged results, the effect of clocking is mostly observed in the leading edge region of the second stator, the largest effects being observed at 15% span. The surface static pressure distribution is changed locally, which is likely to affect the overall performance of the airfoil. The phase-locked averaging technique allows to process the time-resolved component of the data. The pressure fluctuations are attributed to the passage of pressure gradients linked to the traversing of the upstream rotor. The pattern of these fluctuations changes noticeably as a function of clocking. Finally, the time-resolved pressure distribution is integrated along the second stator surface to determine the unsteady forces applied on the vane. The magnitude of the unsteady force is very dependent on the clocking position.
Experimental and Computational Investigation of Cross-Flow Fan Propulsion for Lightweight VTOL Aircraft
Authors: Hobson Garth V. , Cheng W. T. , Seaton M. Scot , Gannon Anthony , Platzer Max F.
Published: November 2008

Cross-flow fan propulsion has not been seriously considered for aircraft use since an Vought Systems Division (VSD) study for the U.S. Navy in 1975. A recent conceptual design study of light-weight, single seat VTOL aircraft suggest that rotary-engine powered cross-flow fans may constitute a promising alternative to the conventional lift-fan vertical thrust augmentation systems for VTOL aircraft. The cross-flow fan performance data obtained by VSD supported the hypothesis that they could be improved to the point where their thrust augmentation could be used in a VTOL aircraft. In this paper we report results of a NASA Glenn supported experimental and computational cross-flow fan investigation which is currently in progress and we provide an assessment of the potential suitability of crossflow fans for VTOL aircraft propulsion. The tests are carried out in the Turbopropulsion Laboratory of the Naval Postgraduate School, using an existing Turbine Test Rig as a power source to drive the cross-flow fan. A 0.305 m (12-inch) diameter, 38.1 mm (1.5-inch) span cross-flow fan test article was constructed to duplicate as closely as possible the VSD fan so that baseline comparison performance data could be obtained. Performance measurements were taken over a speed range of 1,000 to 7,000 RPM and results comparable to those measured by Vought Systems Division were obtained. At 3,000 RPM a 2:1 thrust-to-power ratio was measured which dropped to one as the speed was increased to 6,000 RPM. Performance maps were experimentally determined for the baseline configuration as well as one with both cavities blanked off, for the speed range from 2,000 to 6,000 rpm. Using Flo++, a commercial PC-based computational fluid dynamics software package by Softflo, 2-D numerical simulations of the flow through the cross-flow fan were also obtained. Based on the performance measurements it was concluded that the optimum speed range for this rotor configuration was in the 3,000 to 5,000 rpm range. The lower speed producing the best thrust-to-power ratio and the upper speed range producing the highest efficiency over sizeable throttling range.
A Numerical Simulation Capability for Analysis of Aircraft Inlet – Engine Compatibility
Authors: Hale Alan , Davis Milt , Sirbaugh Jim
Published: November 2008

Two primary aircraft propulsion subsystems are the inlet and the engine. Traditionally these subsystems have been designed, analyzed, and tested as isolated systems. The interaction between the subsystems is modeled primarily through evaluating inlet distortion in an inlet test and then simulating this distortion in engine tests via screens or similar devices. Recently, it has been recognized that significant improvements in both performance and operability can be realized when both the inlet and the engine are designed with full knowledge of the other. In this paper, a computational tool called Turbine Engine Analysis Compressor Code (TEACC) is used to evaluate the effect of inlet distortion on a three-stage military fan. This three-stage military fan is further connected to an F-16 inlet and forebody operating at an angle of attack and sideslip to demonstrate the effect of inlet distortion generated by flight maneuvers. The computational approach of simulating an integrated inlet-engine system is expected to provide additional insight over evaluating the components separately.
Experimental and Computational Evaluation of an Ejector System Utilizing a Primary S-Bend Transition Duct
Authors: Mateer J. C. , Birk A. M. , Poirier D.
Published: November 2008

An experimental and computational study of an air-air ejector system applicable to aerospace applications was conducted. The geometry of the ejector consists of a primary S-bend duct that expels into an oblong mixing tube. The focus of this study was to determine the performance and practicality of an ejector implemented downstream of an S-bend duct. Experimental and computational results were compared to determine computational adequacy. The experimental apparatus utilized a cold flow blower delivering air through an annulus to an S-bend duct that incorporated a downstream mixing tube. Flow measurements were made at the inlet and outlet of the duct as well as at the outlet of the mixing tube. This allowed losses of the primary duct as well as ejector pumping performance to be evaluated. The computational work involved using a commercial CFD package implementing k-ε turbulence modeling and non-equilibrium wall functions.
Test and Evaluation of a Gas Turbine Electric Starter System
Authors: Quin˜ones Martin , Gustafson James R. , Marks Kenneth A. , Parsons Douglas , Lapointe Richard
Published: November 2008

The Naval Surface Warfare Center, Carderock Division, Philadelphia Site successfully completed test and evaluation of the first gas turbine electric starter targeted for Navy use at the Land Based Engineering Site (LBES) of the Naval Surface Warfare Center, Carderock Division (NSWCCD). The Electric Start System (ESS) was developed and manufactured by Hamilton Sundstrand Corporation and installed on an LM2500 main propulsion engine. The ESS was successfully demonstrated to start an LM2500 marine gas turbine with the added benefits of reduced start system configuration (fewer components), ability to control engine acceleration during the early phase of the start cycle, functional transparency to the engine start logic, and speed synchronization between the ESS electric motor and the accessory gearbox of the LM2500 engine with no impact torque. To date it has accumulated over 30 successful starts on the engine. This paper will discuss the results of the aforementioned test, along with advantages and challenges of installing such a system aboard ship.