bluuebunny/crossref_metadata_2025_split huggingface dataset:
Prediction of the Becalmed Region for LP Turbine Profile Design
Authors: Schulte Volker , Hodson Howard P. 
Published: December 2014

Recent attention has focused on the so called ‘becalmed region’ that is observed inside the boundary layers of turbomachinery blading and is associated with the process of wake-induced transition. Significant reductions of profile loss have been shown for high lift LP turbine blades at low Reynolds-numbers due the effects of the becalmed region on the diffusing flow at the rear of the suction surface.
In this paper the nature and the significance of the becalmed region are examined using experimental observations and computational studies. It is shown that the becalmed region may be modelled using the unsteady laminar boundary layer equations. Therefore, it is predictable independently of the transition or turbulence models employed. The effect of the becalmed region on the transition process is modelled using a spot-based intermittency transition model. An unsteady differential boundary layer code was used to numerically simulate a deterministic experiment involving an isolated turbulent spot.
The predictability of the becalmed region means that the rate of entropy production can be calculated in that region. It is found to be of the order of that in a laminar boundary layer. It is for this reason and because the becalmed region may be encroached upon by pursuing turbulent flows that for attached boundary layers, wake-induced transition cannot significantly reduce the profile loss. However, the becalmed region is less prone to separation than a conventional laminar boundary layer. Therefore, the becalmed region may be exploited in order to prevent boundary layer separation and the increase in loss that this entails. It is shown that it should now be possible to design efficient high lift LP turbine blades.
Intermittent Flow and Thermal Structures of Accelerating Transitional Boundary Layers: Part 1 — Mean Quantities
Authors: Wang Ting , Keller F. Jeffrey 
Published: December 2014

A conditional sampling technique was employed to separate the turbulent and non-turbulent parts of accelerated boundary layers undergoing laminar-turbulent transition on a uniformly heated flat plate. Tests were conducted with zero pressure gradient and two levels of streamwise acceleration parameter: K = 0.07 ×10−6 and 0.16 × 10−6.
The conditionally sampled distribution of the skin friction coefficients revealed that the values for Cf in the non-turbulent and turbulent portions significantly deviated from the respective laminar and turbulent correlations. These deviations increased as acceleration increased. Reconstructing the local average Cf values using the laminar and fully turbulent correlations consistently overestimated the unconditioned Cf values. Using the conditionally sampled data for reconstructing Cf values provided better results, but does not necessarily result in the same unconditioned Cf values.
The mean velocity profiles from the turbulent portions had the appearance of a low-Reynolds-number turbulent boundary layer with a large wake region. In the late transition region, as acceleration increased, the wake region in the turbulent portion was suppressed relative to the unconditioned result. The integral parameters, δ*, θ, and shape factor, H were conditionally sampled and analyzed.
Design of New Three Stage Low Pressure Turbine for the BMW Rolls-Royce BR715 Turbofan Engine
Authors: Cobley Keith , Coleman Neil , Siden Gunnar , Arndt Norbert 
Published: December 2014

In 1990, BMW and Rolls Royce plc (RR) joined to form a new company BWW-Rolls-Royce GmbH (BRR), to develop the BR700 family of engines aimed at the 12K and 25K lbs thrust range, using advanced technology and a modern organisation working in integrated teams to minimise the engine development timescales. After a successful development programme the BR710 engine rated at 14K lbs thrust, will shortly enter service in Gulfstream and Canadair Executive Jets.
The recent launch of the BR715 engine at 21K lbs thrust, builds on the high pressure core developed for the BR710, plus a low pressure system with an increased diameter fan and 2 stage booster driven by a three stage turbine.
This paper will describe, the advanced design technology incorporated, including the latest three dimensional aerodynamic philosophy using advanced high lift aerofoils for reduced parts count, plus the mechanical design issues addressed to optimise the LP turbine module configuration and the simultaneous design/make process employed to achieve the required parts delivery timescales.
Turbine Airfoil Manufacturing Technology
Authors: Kortovich Charles S. , Garlock Robert M. , Hayes Craig R. 
Published: December 2014

The Department of Energy’s Turbine Airfoil Manufacturing Technology Program was initiated at PCC Airfoils, Inc./General Electric Power Generation Group. The specific program goal is to define manufacturing methods for single crystal (SX) technology to be applied to airfoil components for power generation applications. A number of technical issues are being addressed and these form the task structure of the program and will be discussed including:
• Alloy Melt Practice to Reduce Sulfur Content in Alloys
• Modification/Improvement of SX Casting Process
• Core Materials and Design
• Grain Orientation Control
Melt desulfurization has successfully reduced sulfur to below 1 part/million with accompanying improvements in oxidation resistance compared to undesulfurized material. Casting variables have been identified which can optimize casting quality. Core material evaluations to date indicate the initial core compositions need to be optimized to eliminate core break during processing operations. Grain orientation control addresses the formation of a database to enable decisions to be made concerning the establishment of grain limit defect criteria. Liberalizing the defect criteria will have a positive effect on the producibility and yields associated with the SX castings. The planned testing has been formulated.
Land Based Turbine Casting Initiative
Authors: Mueller Boyd A. , Price Allen R. , Murphy Kenneth S. , Bell Gregory B. 
Published: December 2014

Howmet is currently conducting a Department of Energy funded Advanced Turbine Systems program, the Land Based Turbine Casting Initiative, to scale single crystal casting technological advances developed for aircraft gas turbine engines up to land based gas turbine sized components. The program encompasses technical activities in the areas of: low sulfur alloys (low sulfur melt processing treatments), casting process development (casting process, mold, core and wax), post-cast process development (mold, core, and gating removal, heat treatment, HIP, and inspection), and casting defect tolerance level establishment (freckles, off-orientation, low angle boundaries, and recrystallized grains).
This presentation will highlight recent progress in the areas of low sulfur alloy, large single crystal casting process development, and 2.5D reconstruction X-ray inspection system development. Howmet has developed a melt treatment to reduce the sulfur content of the nickel-based superalloy ingot it manufactures. Sulfur contents achieved with this technique will be compared to sulfur levels present in currently available ingot. The effect of the reduced sulfur contents on oxidation performance is being investigated and will be reported. The casting process development activities are foundry experiments to determine the effect of different process factors on both grain and dimensional quality. The status of casting activities to scale aeroengine processes up to land based sized components will be presented. The 2.5D reconstruction uses X-ray metrology and multiple 2D X-ray views of the casting to reconstruct the 3D geometry of selected features. The large cross sections and long path lengths inherent in large utility land based components will be particularly challenging.
Fundamental Heat Transfer Experiments of Heat Pipes for Turbine Cooling
Authors: Yamawaki Shigemichi , Yoshida Toyoaki , Taki Masanobu , Mimura Fujio 
Published: December 2014

Fundamental heat transfer experiments were carried out for three kinds of heat pipes which may be applied to turbine cooling in future aero-engines. In the turbine cooling system with a heat pipe, heat transfer rate and start-up time of the heat pipe are the most important performance criteria to evaluate and compare with conventional cooling methods. Three heat pipes are considered, called heat pipe A, B and C, respectively. All heat pipes have a stainless steel shell and nickel sintered powder metal wick. Sodium(Na) was the working fluid for heat pipes A and B; heat pipe C used eutectic sodium-potassium(NaK). Heat pipes B and C included non-condensible gas for rapid start-up. There were fins on the cooling section of heat pipes. In the experiments, an infrared image furnace supplied heat to the heat pipe simulating turbine blade surface conditions.
In the results, heat pipe B demonstrated the highest heat flux of 17 to 20 W/cm2. The start-up time was about 6 minutes for heat pipe B and about 16 minutes for heat pipe A. Thus adding non-condensible gas effectively reduced start-up time. Although NaK is a liquid phase at room temperature, the start-up time of heat pipe C (about 7 to 8 minutes) was not shorter than the heat pipe B. The effect of a gravitational force on heat pipe performance was also estimated by inclining the heat pipe at an angle of 90 degrees. There was no significant gravitational dependence on heat transport for heat pipes including non-condensible gas.
Life Assessment of Gas Turbine Blades and Vanes
Authors: Hepworth John K. , Wilson John D. , Allen James M. , Quentin George H. , Touchton George 
Published: December 2014

The hot stage components of a large industrial gas turbine engine operate under extremely arduous conditions and as a result have a finite life. In particular, the first stage blades and vanes are expensive to replace and there are strong financial incentives to maximise their useful life and to refurbish them, where possible, to extend that life. In order to achieve this life extension, it is essential to have a detailed knowledge of the conditions under which the blades operate and of their responses to those conditions.
Detailed thermal and stress analyses have been carried out on the first stage blades and vanes of a large gas turbine, operating in a combined cycle power generation plant. The boundary conditions for the analysis were taken from an aerothermal analysis of the engine. From these, the finite element method was used to compute temperature distributions and then stress distributions in both the blades and vanes. Consideration was given to the effect of thermal transients and also to the effect of creep in allowing some redistribution of the thermal stresses.
The resulting stress and temperature distributions were used to estimate damage accumulation in the material due to fatigue, creep and, where appropriate, oxidation. These damage estimates have then been applied to predict crack initiation. Estimates have also been made of likely crack growth behaviour. Damage predictions have been compared with observations made on blades removed from service. The damage accumulation has implications for the operation, maintenance and blade replacement strategies for the engine.
Computation of Separated-Flow Transition Using a Two-Layer Model of Turbulence
Authors: Papanicolaou Elias L. , Rodi Wolfgang 
Published: December 2014

A model for predicting transition in flows with separation is presented in this study. The two-layer model of turbulence is employed, along with a suitably defined intermittency function, which takes appropriate values in the laminar-, transitional- and turbulent-flow regions. Correlations derived from measurements are used for this purpose. Two test cases were selected: the flow over a long horizontal body with semi-circular leading edge and the flow over the backward-facing step of small height (expansion ratio of 1:1.01). In the former, oncoming flows with a freestream turbulence level encountered in practical applications was considered (0.2%–5.6%), whereas in the latter the corresponding level was much lower. The Reynolds numbers, based on the diameter in the first case and on the step height in the second, lie in the range of 1600–6600, in which limited numerical investigations were previously available and where bubbles with laminar separation and turbulent reattachment are expected.
The predictions were found to compare well with the corresponding measurements, both in terms of the lengths of the separation and the transition regions and of velocity and turbulence intensity profiles at various stream wise locations. The results show that, for the transition criterion chosen, in all cases transition is completed downstream of the reattachment point and that the rate at which it is completed increases with the freestream turbulence level.
Turbine Tip Clearance Measurement System Evaluation on an Industrial Gas Turbine
Authors: Gill S. J. , Ingallinera M. D. , Sheard A. G. 
Published: December 2014

The continuing development of industrial gas turbines is resulting in machines of increasing power and efficiency. The need to continue this trend is focusing attention on minimizing all loss mechanisms within the machine, including those associated with turbine blade tip clearance.
In order to study tip clearance in the turbine, real time measurement is required of clearance between turbine blades and the casing in which they run. This measurement is not routinely performed, due to the harsh nature of the turbine environment. On those occasions when turbine tip clearance is measured, it is typically in development vehicles, often using cooled probes that are somewhat unsuitable for use in production gas turbines.
In this paper a program of work is reported that was undertaken with the purpose of identifying a promising turbine tip clearance measurement system that used the capacitive gap measurement technique. Issues surrounding the application of three systems to the turbine section of a GE MS6001FA gas turbine are identified and reported. Performance of the three evaluated systems is analyzed.
Bird Strike Investigations in the Development Process of a Transonic Fan Blisk
Authors: Frischbier Jörg D. 
Published: December 2014

Initiated in 1994, a transonic fan in blisk configuration is being developed at MTU with the intention to replace a conventionally bladed titanium fan rotor stage. A major challenge in the structural development process of this blisk stage was to find a blade geometry with a sufficient bird strike resistance. This paper gives an overview of the combined analytical and experimental design process towards the final blisk blade geometry capable of resisting a worst-case 1 lb low speed bird strike at a typical aircraft take-off operating situation without blade loss.
Starting with a first blade geometry a rig ingestion test in 1995 revealed an insufficient bird strike potential of this original blisk blade standard. The intensive analysis of the test result also showed that the analytical model established at MTU had to be adapted with respect to some decisive aspects of the analytical approach. This led to a revised method of meshing the blade leading edge, to a revised failure criterion of the material model and to a revised concept in modelling the bird slicing effect. This was done on the basis of several ‘static’ shooting tests on cantilever steel specimens and on single blade segments of the original blisk. The intention of these tests was to verify the analytical model with respect to quantitatively reliable predictions of local large strain (up to 60% plastic strain) at high rates of strain.
With these revised analytical methods the blisk geometry was redesigned (increased blade hub cross sections) and with a rig ingestion test of a 1 lb pigeon at take-off rotational speed the required capability of the final blisk geometry was verified in 1996.
Failure Mechanisms of Coating Systems Applied to Advanced Turbine Components
Authors: Daleo Joseph A. , Boone Donald H. 
Published: December 2014

The maintenance period of critical gas turbine components operating in advanced D, E and F technology industrial turbines is often determined by the life of the coating system applied to protect the base alloy substrate. The progressively higher firing temperatures used in all of the advanced engine designs results not only in very high metal surface temperatures but also in very high temperature gradients and concomitant thermal stresses Induced in part by the complex cooling systems. In order to develop optimum component life strategies, it is important to establish the actual operating conditions of each component, and, to define the predominant degradation modes.
Metallurgical life assessment of these advanced component designs has identified several distinct coating/base metal failure mechanisms not generally encountered in earlier generation turbines, in addition to the more usual degradation modes, accelerated by the increases in temperature. A review and examples of these degradation mechanisms encountered on service exposed coating systems currently used by some of today’s major manufactures are presented.
Performance Improvements of a Natural Gas Injection Station Using Gas Turbine Inlet Air Cooling
Authors: De Lucia Maurizio , Carnevale Ennio , Falchetti Massimo , Tesei Alberto 
Published: December 2014

Gas Turbine (GT) performance seriously deteriorates at increased ambient temperature. This study analyses the possibility of improving GT power output and efficiency by installing a gas turbine inlet air cooling system.
Different cooling systems were analyzed and preliminary cost evaluations for each system were carried out.
The following three cooling systems were considered in detail:
a) Traditional compression cooling system;
b) Absorption single-acting cooling system using a solution of lithium bromide;
c) Absorption double-acting cooling system using a solution of lithium bromide.
Results clearly indicate that there is a great potential for GT performance enhancement by application of an Inlet Air Cooling (IAC). Technical and economical analyses lead to selection of a particular type of IAC for significant savings in capital outlay, operational and maintenance costs and other additional advantages.
Analysis of Vortices in Three-Dimensional Jets Introduced in a Cross-Flow Boundary-Layer
Authors: Sgarzi O. , Leboeuf F. 
Published: December 2014

The aim of this work is to numerically investigate the different vortical structures present in the flow generated by a jet in crossflow. The test case which is relevant to an hydraulic experiment consists of a single jet ejecting normally into a laminar main stream.
The computation is performed using a stationary three-dimensional Navier-Stokes code. A multi-block technique is used to compute the flow in the injection pipe. In addition high resolution is achieved in the region of jet-mainstream interaction.
The flow analysis relies on the visualization of particle trajectories. These particles are introduced into vortex cores that are located by the secondary velocity field they induce. The three-dimensional behavior of these structures is enlightened. This provides us with the origin of the fluid of which they are comprised. The mixing between the jet and the incoming viscous layer appears to begin on the windward side of the jet boundary.
Up to five types of vortices are identified including the well known counter rotating vortices that dominate the downstream flow. They clearly result from the stretching and warping of the annular vorticity rings issuing from the pipe. At the jet exit, they each split into two sub-structures with different growth and downstream development. Less prominent structures are also seen. The “horse shoe” vortex has been captured which is typical of the near wall effects due to blockage induced by the jet in the main stream. Its downstream legs are sucked through the jet boundary. Another weak structure located on the upstream jet boundary is the “lip” vortex which results from the upstream part of the flow specific topology. It is shown that viscous effects play an important role in both the generation and interaction between vortical structures.
The Influence of Film Cooling on the Efficiency of an Annular Nozzle Guide Vane Cascade
Authors: Day C. R. B. , Oldfield M. L. G. , Lock G. D. 
Published: December 2014

This paper examines the effect of aerofoil surface film cooling on the aerodynamic efficiency of an annular cascade of transonic nozzle guide vanes. A dense foreign gas (SF6/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios under ambient conditions. The flowfield measurements have been obtained using a four-hole pyramid probe in a short duration blowdown facility which correctly models engine Reynolds and Mach numbers, as well as the inlet turbulence intensity.
The use of foreign gas coolant poses specific challenges not present in an air-cooled cascade, and this paper addresses two. Firstly, a novel method for the determination of mass flow from pneumatic probe data in a heterogeneous gas environment is presented which eliminates the need to measure concentration in order to determine loss. Secondly, the authors argue on the grounds of dimensionless similarity that momentum flux ratio is to be preferred to blowing rate for the correct parametrisation of film cooling studies with varying coolant densities.
Experimental results are presented as area traverse maps, from which values for loss have been calculated. It is shown that air and foreign gas at the same momentum flux ratio give very similar results, and that the main difference between cooled and uncooled configurations is an increase in wake width. Interestingly, it is shown that an increase in the momentum flux ratio above the design value with foreign gas coolant reduces the overall loss compared with the design value. The data has been obtained both for purposes of design and for CFD code validation.
Wake Recovery Performance Benefit in a High-Speed Axial Compressor
Authors: Van Zante Dale E. , Adamczyk John J. , Strazisar Anthony J. , Okiishi Theodore H. 
Published: December 2014

This paper addresses the significant differences in compressor rotor wake mixing loss which exist in a stage environment relative to a rotor in isolation. The wake decay for a rotor in isolation is due solely to viscous dissipation which is an irreversible process and thus leads to a loss in both total pressure and efficiency. Rotor wake decay in the stage environment is due to both viscous mixing and the inviscid strain imposed on the wake fluid particles by the stator velocity field. This straining process, referred to by Smith (1993) as recovery, is reversible and for a 2D rotor wake leads to an inviscid reduction of the velocity deficit of the wake. It will be shown that for the rotor/stator spacing typical of core compressors, wake stretching is the dominant wake decay process within the stator with viscous mixing playing a secondary role.
A model for the rotor wake decay process is developed and used to quantify the viscous dissipation effects relative to those of inviscid wake stretching. The model is verified using laser anemometer measurements acquired in the wake of a transonic rotor operated in isolation and in a stage configuration at near peak efficiency and near stall operating conditions. Additional insight is provided by a time-accurate 3D Navier Stokes simulation of the compressor stator flow field at the corresponding stage loading levels. Results from the wake decay model exhibit good agreement with the experimental data. Data from the model, laser anemometer measurements, and simulations indicate that wake straining (stretching) is the primary decay process in the stator passage. The implications of these results on compressor stage design are discussed.