bluuebunny/crossref_metadata_2025_split huggingface dataset:
Conjugate Heat Transfer Analysis of an Engine Internal Cavity
Authors: Montenay A. , Paté L. , Duboué J. M. 
Published: August 2014

The analysis of heat transfer in engine cavities or blade internal cooling systems is one of the most challenging work for aircraft engines designers for two main reasons. Firstly, the efficiency of such systems has a direct influence on both life and performance of these engines. Secondly, the available tools to predict heat transfer in both solid parts and surrounding cooling gases, i.e. Navier Stokes and conduction codes, are often used independently. An interaction model between the fluid and solid media is generally required and remains a difficult issue in engine configurations. A coupling procedure between a Navier-Stokes code and a conduction solver is therefore the only way to achieve heat transfer predictions in all flow situations. The objective of this work is to present such a procedure, which has been developed at Snecma and based on a Finite Volume Navier-Stokes code and a commercial Finite Element solver.
The first application showed in the paper demontrates, with an uncoupled calculation that the Navier-Stokes code MSD, from ONERA, is able to predict heat transfer with an acceptable accuracy. The discretization used in the solid to predict heat conduction is briefly presented. Then the steady state coupling procedure is exposed and validated with an analytical solution. Finally, a conjugate heat transfer computation in a rotor/rotor cavity of a real engine, with rotating solid disks, is described in detail.
Remote Monitoring and Diagnostics System for GE Heavy Duty Gas Turbines
Authors: Ozgur Dincer , Lakshminarasimha Arkalgud N. , Rucigay Richard , Morjaria Mahesh , Sanborn S. 
Published: August 2014

The paper describes GE’s Remote Monitoring and Diagnostic (RM&D) system, its operation, its unique features and our experiences in applying it to improve performance and availability of heavy-duty gas turbine fleet worldwide. A key success factor of the system is that it relies upon an effective combination of advanced computer automation complemented by technical experts who have an in-depth understanding of the power generation equipment to achieve its objectives.
The RM&D system enables GE experts to remotely access operational data of power plants operating world wide, and exercise sophisticated algorithms, which can detect abnormal operating conditions. The experts, with ready access to the design information, operation and maintenance information and in-depth knowledge about the turbines, identify and track operational signatures, which may indicate potential problems. The real key is to be able to distinguish false indications from the ones that are true early indicators of potential problems. In this paper we describe a design approach for developing a successful RM&D system. We illustrate our approach through various examples of performance, vibration and combustion diagnostics scenarios.
Degradation in Gas Turbine Systems
Authors: Kurz Rainer , Brun Klaus 
Published: August 2014

Any prime mover exhibits the effects of wear and tear over time. The problem of predicting the effects of wear and tear on the performance of any engine is still a matter of discussion. Because the function of a gas turbine is the result of the fine tuned cooperation of many different components, the emphasis of this paper is on the gas turbine and its driven equipment (compressor or pump) as a system, rather than on isolated components. We will discuss the effect of degradation on the package as part of a complex system (e.g. a pipeline, a reinjection station, etc.).
Treating the gas turbine package as a system reveals the effects of degradation on the match of the components as well as on the match with the driven equipment. This article will contribute insights into the problem of gas turbine system degradation. Based on some detailed studies on the mechanisms that cause engine degradation, namely
- changes in blade surfaces due to erosion or fouling, and the effect on the blade aerodynamics;
- changes in seal geometries and clearances, and the effect on parasitic flows,
- changes in the combustion system (e.g. which result in different pattern factors),
the effects of degradation will be discussed.
The study includes a methodology to simulate the effects of engine and driven equipment degradation. With a relatively simple set of equations that describe the engine behavior, and a number of linear deviation factors, which can easily be obtained from engine maps or test data, the equipment behavior for various degrees of degradation will be studied. A second model, using a stage by stage model for the engine compressor, is used to model the compressor deterioration. The authors have avoided to present figures about the speed of degradation, because it is subject to a variety of operational and design factors, that typically cannot be controlled entirely.
Effect of the Decaying Overshoot of the Rankine Vortex on the Axial Flow Field Behind the Inlet Guide Vane of Centrifugal Compressors
Authors: Chen Y. N. , Hagelstein D. , Kassens I. , Haupt U. , Rautenberg M. 
Published: August 2014

The present analysis of the experimental results obtained by Kassens (1997), and Kassens and Rautenberg (1998) about the variation of the axial velocity profile of the rotating flow travelling downstream the inlet guide vane revealed the controlling effect of the decaying overshoot of the Rankine vortex in this process.
The disappearance of the overshoot indicates the stop of its excessive drive on the rotating flow. It will be shown that this decay process is directly connected to the braking effect of the impeller blades on the rotating flow generated by the IGV. For the case of an IGV setting angle of 60°, the flow with a very strong swirl component generated by the IGV outlet cannot be admitted by the inlet of the impeller without shock. This shock acts as a braking force of the impeller blades on the swirl component of the approaching flow, a process that is expressed as the spin-down of the swirling flow. A secondary toroidal ring vortex in the frontal zone of the impeller will be generated by the fundamental flow. The recirculation velocity of this toroidal ring vortex has an axial component, which directs towards the impeller in the outer zone of the suction duct, but turns away from the impeller in its central zone.
At the same time, the field of the uniform tangential velocity of the IGV outlet is getting into the influence sphere of the spin-down mentioned. As a consequence, it will be converted into the flow field of a potential vortex, which is accompanied by a field of axial velocity having a normal value uniformly distributed in its outer potential zone and a strongly peaked value over its central core. This process is similar to that of the flow field in a bathtub vortex during draining its fluid out of the central pipe. The new flow field consists of a secondary toroidal ring vortex superimposed on the fundamental flow. The rotating sense of this new toroidal ring vortex is just opposite to that formed in front of the impeller.
We then have a pair of toroidal ring vortices staying one behind another between the IGV outlet and the impeller inlet. Each member of the pairs serves as the border for the other one. It can thus be concluded that the pair is induced simultaneously by the spin-down of the swirling flow of the IGV outlet brought about by the braking effect of the impeller blades.
Effects of Blade Number on the Efficiency of Centrifugal Compressor Impellers
Authors: Rodgers Colin 
Published: August 2014

The results of a study directed towards identifying optimum blade numbers for aviation and turbocharger type centrifugal compressor impellers is presented. The study focused upon a simplified one dimensional frictional model supported by impeller test data over a wide specific speed range, plus limited correlation with computational fluid dynamic models.
The study results indicated that optimum blade numbers were a function of specific speed, blade backsweep angle, and DeLaval number. With a backsweep angle of 45 degrees and DeLaval number of unity the optimum blade numbers were ≈ 30 and ≈ 15 at low and high specific speeds respectively. It was shown that efficiency decreased with low blade number primarily due to low momentum exchange i.e., low slip factor.
In order to condense the analysis only the impeller component of the compressor stage was studied, wherein the geometry and operating conditions were characterized in terms of five parameters, with particular emphasis on specific speed.
Mixed Flow Turbines: Inlet and Exit Flow Under Steady and Pulsating Conditions
Authors: Karamanis N. , Martinez-Botas R. F. , Su C. C. 
Published: August 2014

The performance and detailed flow characteristics of a high pressure ratio mixed flow turbine has been investigated under steady and pulsating flow conditions. The rotor has been designed to have a nominal constant incidence (based on free vortex flow in the volute) and it is for use in an automotive high speed diesel turbocharger. The results indicated a departure from the quasi-steady analysis commonly used in turbocharger turbine design. The pulsations from the engine have been followed through the inlet pipe and around the volute; the pulse has been shown to propagate close to the speed of sound and not according to the bulk flow velocity as stated by some researchers.
The flow entering and exiting the blades has been quantified by a laser Doppler velocimetry system. The measurements were performed at a plane 3.0 mm ahead of the rotor leading edge and 9.5 mm behind the rotor trailing edge. The turbine test conditions corresponded to the peak efficiency point at 29,400 and 41,300 rpm. The results were resolved in a blade-to-blade sense to examine in greater detail the nature of the flow at turbocharger representative conditions. A correlation between the combined effects of incidence and exit flow angle with the isentropic efficiency has been shown.
The unsteady flow characteristics have been investigated at two flow pulse frequencies, corresponding to internal combustion engine speeds of 1600 and 2400 rpm. Four measurement planes have been investigated: one in the pipe feeding the volute, two in the volute (40° and 130° downstream of the tongue) and one at the exit of the turbine. The pulse propagation at these planes has been investigated; the effect of the different planes on the evaluation of the unsteady isentropic efficiency is shown to be significant. Overall, the unsteady performance efficiency results indicated a significant departure from the corresponding steady performance, in accordance with the inlet and exit flow measurements.
Design of Low-Speed Cascades for Investigating Viscous Effects in High-Speed Axial Turbines
Authors: Zavitz K. , Sjolander S. A. 
Published: August 2014

For turbine flow phenomena which are dominated by viscous effects, many valuable insights into the flow physics can be gained through low-speed cascade measurements. For example, for low-pressure turbines unsteady wake-blade interactions can be investigated in cascade provided that the Reynolds number, freestream turbulence conditions and the pressure coefficient distributions are the same in the cascade as in the high-speed counterpart.
This paper describes an iterative procedure for inversely designing low-speed linear cascades with prescribed blade pressure-coefficient distributions. The inverse-design problem is treated as an optimization problem. The optimization strategy features the use of a genetic algorithm and a gradient-type algorithm. At the end of each global iteration of the design procedure a Navier-Stokes analysis is used to see if the final cascade geometry gives the specified pressure-coefficient distribution to the desired degree of accuracy.
Although the resulting cascade may be designed to the level of accuracy afforded by the Navier-Stokes analysis, the method takes advantage of the fact that the pressure distribution in the low-speed cascade can be predicted with good accuracy and very rapidly using a panel method solution for the potential flow through the cascade. A panel method flow solver is used to minimize the number of Navier-Stokes evaluations to three or four for a given inverse-design problem. As a result, the present procedure is very efficient.
Aerodynamic Performance Measurements of a Fully-Scaled Turbine in a Short Duration Facility
Authors: Keogh R. C. , Guenette G. R. , Sommer T. P. 
Published: August 2014

This paper describes the implementation and results of an experiment to measure the isentropic efficiency of a fully scaled transonic turbine stage in a short-duration blowdown type facility. Short duration test facilities offer the potential for low cost, high accuracy testing of turbomachinery components under fully scaled conditions. The method developed employs the ‘brake’ approach for turbine efficiency measurement. This paper describes newly developed techniques to measure the turbine torque and mass flow in this type of facility.
The isentropic efficiency was measured over a test window of 400 to 500 milliseconds. The estimated measurement uncertainty (U95) is 0.45% for absolute measurements, and 0.25% for relative measurements. Corrected mass flow and isentropic efficiency measurements are presented for a range of corrected speed (97–112%) and compared with CFD predictions. Close agreement was observed both in overall magnitudes and trends between the measured data and predictions.
The New PGT5B/2: A State-of-the-Art 6MW Two-Shaft Gas Turbine
Authors: Bianchi Dino , Noccioni Paolo , Silvestri Catherine J. 
Published: August 2014

The new PGT5B is a compact, state-of-the-art, 6 MW industrial gas turbine. The unit is available in two configurations: a single shaft for power generation and a two-shaft for mechanical drive applications. Maximum commonality has been maintained between the single and two shaft models. Both units share a common gas generator, with evident benefits in all the installations where mixed operations are required. The two-shaft engine is simply obtained by removing the second stage turbine from the single shaft unit and adding a two-stage power turbine.
The unit is an evolution of the existing PGT5, building on the experience of the successful PGT10 model. The synergy between GE Aircraft Engines and GE Power System Nuovo Pignone, with the benefit of GE’s Six Sigma Total Quality Methodology, has permitted the application of the most advanced technology in the development of the major components of the engine.
The high efficiency axial compressor is a scaled down version of the successful PGT10B, chosen as the base compressor for the new PGT’s product line. It has been obtained by the combination of advanced fully 3D aero-design, custom airfoil shape and blade stacking optimization.
The annular DLE, dual fuel combustion system, developed according to the proven aircraft engines design, is able to satisfy the most stringent environmental regulations.
The axial turbine, with sophisticated air cooling system, is designed to reach high values of firing temperature consistent with future uprates. The application of advanced CFD analysis using 3D viscous multi-rows codes and unsteady simulation has resulted in a very high efficiency turbine.
The two stage uncooled power turbine capitalizes on Nuovo Pignone’s experience in the mechanical drive market. While the single shaft version is particularly suitable for power generation and cogeneration due to the high exhaust temperature, the two-shaft version, with wide operating speed range, is designed to be a reliable and efficient mechanical drive for compressors and pumps.
The Power Turbine speed for optimum application efficiency (up to 34.%) is 12,500 rpm, with a capability range from 50% to 105% of nominal speed.
The flexibility of the standard package satisfies the different installation requirements of oil and gas market.
The simplicity of construction and the modular design concept of both the engine and package facilitate maintenance and reduce downtime and labor during servicing.
Unsteady Flow and Whirl-Inducing Forces in Axial-Flow Compressors: Part I — Experiment
Authors: Storace A. F. , Wisler D. C. , Shin H.-W. , Beacher B. F. , Ehrich F. F. , Spakovszky Z. S. , Martinez-Sanchez M. , Song S. J. 
Published: August 2014

An experimental and theoretical investigation has been conducted to evaluate the effects seen in axial-flow compressors when the centerline of the rotor is displaced from the centerline of the static structure of the engine. This creates circumferentially non-uniform rotor-tip clearances, unsteady flow, and potentially increased clearances if the rotating and stationary parts come in contact. The result not only adversely affects compressor stall margin, pressure rise capability and efficiency but also generates an unsteady, destabilizing, aerodynamic force, called the Thomas/Alford force, which contributes significantly to rotor whirl instabilities in turbomachinery. Determining both the direction and magnitude of this force in compressors, relative to those in turbines, is especially important for the design of mechanically-stable turbomachinery components. Part I of this two-part paper addresses these issues experimentally and Part II presents analyses from relevant computational models.
Our results clearly show that the Thomas/Alford force can promote significant backward rotor whirl over much of the operating range of modern compressors, although some regions of zero and forward whirl were found near the design point. This is the first time that definitive measurements, coupled with compelling analyses, have been reported in the literature to resolve the long-standing disparity in findings concerning the direction and magnitude of whirl-inducing forces important in the design of modern axial-flow compressors.
Some Aspects of Modelling Compressor Behavior in Gas Turbine Performance Calculations
Authors: Riegler Claus , Bauer Michael , Kurzke Joachim 
Published: August 2014

Performance calculation procedures for gas turbine engines are usually based on the performance characteristics of the engine components, and especially the turbo components are of major interest. In this paper methods of modelling compressors in gas turbine performance calculations are discussed. The basic methodologies based on Mach number similarity are summarized briefly including some second order effects.
Under extreme enginepartload conditions, as for example subidle or windmilling, the operating points in the compressor map are located in a region which is usually not covered by rig tests. In addition the parameters usually used in compressor maps are no longer appropriate. For these operating conditions a method is presented to extrapolate compressor maps towards very low spool speed down to the locked rotor. Instead of the efficiency more appropriate parameters as for example specific work or specific torque are suggested. A compressor map prepared with the proposed methods is presented and discussed.
As another relevant topic the performance modelling of fans for low bypass ratio turbofans is covered. Due to the flow splitter downstream of such a fan the core and bypass stream may be throttled independently during engine operation and bypass ratio becomes a third independent parameter in the map. Because testing a fan on the rig for various bypass ratios is a very costly task, a simplified method has been developed which accounts for the effects of bypass ratio.
Viscous-Flow 2D-Analysis Including Secondary Flow Effects
Authors: Mönig Reinhard , Mildner Frank , Röper Ralf 
Published: August 2014

During the last few decades extremely powerful Quasi-3D codes and fully 3D Navier-Stokes solvers have been developed and successfully utilized in the design process and optimization of multistage axial-flow compressors.
However, most of these methods proved to be difficult in handling and extremely time consuming. Due to these disadvantages, the primary stage design and stage matching as well as the off-design analysis is nowadays still based on fast 2D methods incorporating loss-, deviation- and end wall modeling. Only the detailed 3D optimization is normally performed by means of advanced 3D methods.
In this paper a fast and efficient 2D calculation method is presented, which already in the initial design phase of multistage axial flow compressors considers the influence of hub leakage flows, tip clearance effects and other end wall flow phenomena. The method is generally based on the fundamental approach by Howard and Gallimore (1992). In order to allow a more accurate prediction of skewed and non-developed boundary layers in turbomachines an improved theoretical approach was implemented. Particularly the splitting of the boundary layers into an axial and tangential component proved to be necessary in order to account for the change between rotating and stationary end walls. Additionally, a new approach is used for the prediction of the viscous end wall zones including hub leakage effects and strongly skewed boundary layers. As a result, empirical correlations for secondary flow effects are no longer required.
The results of the improved method are compared with conventional 2D-results including 3D loss- and deviation-models, with, experimental data of a 3-stage research compressor of the Institute for Jet Propulsion and Turbomachinery of the Technical University of Aachen and with 3D Navier-Stokes solutions of the V84.3A compressor and of a multi-stage Siemens research compressor. The results obtained using the new method show a remarkable improvement in comparison with conventional 2D-methods. Due to the high quality and the extremely short computation time the new method allows an overall viscous design of multistage compressors for heavy duty gas turbines and aeroengine applications.
Limitations on Gas Turbine Performance Imposed by Large Turbine Cooling Flows
Authors: Horlock J. H. , Watson D. T. , Jones T. V. 
Published: August 2014

Calculations of the performance of modern gas turbines usually include allowance for cooling air flow rate; assumptions are made for the amount of the cooling air bled from the compressor, as a fraction of the mainstream flow, but this fractional figure is often set in relatively arbitrary fashion.
There are two essential effects of turbine blade cooling:
[i] the reduction of the gas stagnation temperature at exit from the combustion chamber [entry to the first nozzle row] to a lower stagnation temperature at entry to the first rotor and
[ii] a pressure loss resulting from mixing the cooling air with the mainstream.
Similar effects occur in the following cooled blade rows.
The paper reviews established methods for determining the amount of cooling air required and semi-empirical relations, for film cooled blading with thermal barrier coatings, are derived. Similarly, the pressure losses related to elements of cooling air leaving at various points round the blade surface are integrated over the whole blade. This gives another semi-empirical expression, this time for the complete mixing pressure loss in the blade row, as a function of the total cooling air used.
These two relationships are then used in comprehensive calculations of the performance of a simple open-cycle gas turbine, for varying combustion temperature and pressure ratio. These calculations suggest that for maximum plant efficiency there may be a limiting combustion temperature [below that which would be set by stoichiometric combustion]. For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.