bluuebunny/crossref_metadata_2025_split huggingface dataset:
The Calculation of the Three-Dimensional Flow Through a Transonic Fan Including the Effects of Blade Surface Boundary Layers, Part-Span Shroud, Engine Splitter and Adjacent Blade Rows
Authors: Cedar R. D. , Holmes D. G. 
Published: March 2015

In the past, predictions of the three-dimensional flow through turbomachinery blades have often assumed that:-
a) The blade passage forms a single channel.
b) The effects of adjacent blade rows and engine structures are negligible.
c) The flow is inviscid.
Unfortunately, for a transonic fan in a high bypass ratio aero-engine, these assumptions are poor because:-
a) The blade passage is split into a number of channels by the part-span shroud and the engine splitter.
b) The fan aerodynamics may be influenced by the presence of the engine splitter and downstream blade rows.
c) The blockage due to the blade surface boundary layers can have a significant effect on the shock structure.
This paper describes the extensions that have been made to the three-dimensional Euler solver developed by Holmes and Tong (1985) in order to remove the assumptions listed above. The problem of modeling the part-span shroud and the engine splitter has been generalized to that of solving a flow containing any number of internal solid bodies. This has been achieved in a novel manner that allows the finite volume scheme to be fully vectorized and the multigrid acceleration technique to be implemented efficiently. The effect of adjacent blade rows has been modeled using passage averaged mass, momentum and energy source terms to represent the neighboring blades in a similar manner to that described by Adamczyk, Mulac and Celestina (1986). The blade surface boundary layers have been modeled using a series of two-dimensional boundary layer calculations coupled to the Euler solver using a transpiration model to account for the boundary layer displacement thickness. To demonstrate the capabilities of this program, results from the analysis of a General Electric advanced fan blade design, with a part-span shroud, engine splitter and down-stream stator and rotor blade are presented. This is followed by an investigation and discussion of the individual effect of not modeling the viscous blockage, the part-span shroud engine, splitter and down-stream blade rows. It is concluded that if an accurate prediction of the fan aerodynamics is to be obtained all these features must be modeled.
A Method for the Prediction of Supersonic Compressor Blade Performance
Authors: Freeman C. , Cumpsty N. A. 
Published: March 2015

A simple model is used to calculate the flow in compressor blades with supersonic relative inlet flow. The one-dimensional model utilises the conservation of stagnation enthalpy, mass flow and momentum in the inlet region. The momentum equation is applied in the direction parallel to the blade surface at inlet and one of the fundamental simplifications adopted is that the projected area of the blades gives such a small force in this direction that a very simple approximation for it suffices.
The model is able to predict the loss creation in the inlet region. The level of predicted loss agree well with measured values when the Mach numbers are sufficiently high for the inlet loss to dominate. Furthermore the correct trends of loss with incidence and blade speed are predicted. The pressure rise and flow can also be predicted when the correct deviation and meridional streamtube convergence are given. The method predicts trends usually seen in measurments: the narrowing of the useful operating range between choke and flow instability (surge) as blade speed is increased and a steepening of the surge line at the higher speeds.
Directional Solidification of Land-Based Gas Turbine Buckets
Authors: Peterson Lance G. 
Published: March 2015

Improvements in land-based gas turbine performance require materials capable of operating at increasingly higher temperatures. Innovative approaches to alloy and process improvement, such as directional solidification, can meet the more demanding conditions of new gas turbine designs. This paper outlines GE TBO’s multi-faceted research program for the development of directionally solidified GTD-111 first stage turbine buckets. Included are discussions of the theoretical background, casting program, mechanical tests and machining studies.
The GTD-111 directional solidification program successfully met the technological challenges encountered in scaling up directional solidification experience from relatively small aircraft parts. Directionally solidified buckets are in production for several first stage product lines. Improvements in mold technology coupled with experience gained in the on-going research program indicate that more complex parts as well as single crystal parts are now feasible.
Modified Fuel Control for a Large Heavy-Duty Combustion Turbine-Generator
Authors: Quentin G. , Brawley G. , Loft A. , Perez-Daple W. 
Published: March 2015

Thermal transients during start-up and shutdown of a heavy-duty gas turbine stress the hot gas path components and reduce their effective life due to low-cycle fatigue. Experience has shown that stress-related cracking and wear is more extensive on gas turbines in cyclic service that start and stop more frequently, than machines in baseload service that operate more continuously.
The start-up schedules of most engines in peaking duty are generally designed to minimize the time to reach full load. Rapid loading is not a critical element for many gas turbine users, who would prefer alternate start-up schedules that could extend component life, by as much as a factor of three.
It is the fuel controls that dictate thermal energy input to a combustion turbine during start-up and shutdown. Therefore, experiments performed on the rates of acceleration and deceleration led to modified fuel controls for a natural gas-fired combustion turbine (General Electric Model MS7001B), Unit 41 at the T. H. Wharton Plant of Houston Lighting & Power Co. The results may be generalized and applied to other gas turbines and control systems, using different fuels.
Specific findings relate to quicker ignition, more gradual initial warm-up, a lower acceleration for a soft start, and lower deceleration during fired shutdown. Additionally, new minimum fuel settings enhanced running reliability during load-shedding operation with water injection for NOx control. The results are described and illustrated in detail to demonstrate the new start-up and shutdown sequence. Subsequently, one HL&P unit was permanently modified for check-out, and across-the-board updating of other units is planned.
A Methodology Proposal to Design and Analyse Counterrotating High Speed Propellers
Authors: Nicoud D. , Brochet J. , Goutines M. 
Published: March 2015

Contrarotating high speed propellers are able to significantly reduce fuel consumption of high subsonic aircrafts. The achievement of this goal requires the optimization of the transonic flowfield on the blades in order to obtain high efficiency. For several years, 2D and 3D aerodynamic computational methods have been used to design high performance turbofans. A similar methodology can be developed for high speed propeller design, and this paper presents a typical application of such methods.
We first present an application of the through-flow method. An outer fictitious casing is chosen in order to simulate undisturbed flow far from the propellers, and the mesh is adapted to the high swept blades. Radial distribution of loading is selected using aerodynamic criteria.
Then, a quasi geometrical method supplies the bidimensional profiles accounting for structural specifications such as chord length, maximum thickness and root attachment. Suction side incidence and downstream deviation are also specified. After the profile stacking operations, which use conformal application on the axisymetric stream surfaces, the tridimensional transonic flowfield is drawn by a 3D Euler solver on an appropriate domain. This code uses a multi-domains technique and includes the energy equation for non-constant rothalpy cases. Particular interest is focused on the Mach number distributions and on the shock strength. The final loss prediction is made by means of a shock loss model and a bidimensional boundary layer calculation based on the Euler static pressure distributions.
The profile shapes are modified and the above process is repeated until the required deflection, a convenient throat margin, and sufficiently thin and well attached boundary layers are obtained. Finally, the global performances are issued from 3D Euler and boundary layer computations, completed by the calculation of secondary flow effects.
Boundary Layer Calculation Including the Prediction of Transition and Curvature Effects
Authors: Guyon B. , Arts T. 
Published: March 2015

The calculation of surface temperature on gas turbine blades in severe operating conditions requires a detailed knowledge of boundary layers behaviour. The prediction of laminar to turbulent transition as to existence and location, as well as the evaluation of heat transfer rates are major concerns.
The program developed by SNECMA for this purpose is presented, in which models are introduced to take into account the main effects occuring on blades without film-cooling. The algorithm and discretisation scheme for boundary layer equations is Patankar and Spalding’s, with profiles initialization by Pohlhausen’s method. The turbulence and transition model, after Mc Donald and Fish, was improved in search for more stability and to have a better detection of the beginning of the transition. Adams and Johnston’s model for curvature, including propagation effects, was adapted to a transitional boundary layer.
The validation tests of this program are described, which are based on numerous experimental data taken from a bibliography of tests over flat plates and blades. Other tests use heat transfer rate measurements conducted by SNECMA, together with VKI, on vanes and blades in non-rotating grids. The calculation results are further compared to the STAN5 program results; they show a superiority in predicting the transfer rates on a convex surface and for transitional boundary layers.
Fiber Optic Probe for Flame Monitoring in Gas Turbine Combustors
Authors: Morey W. W. , Quentin G. H. , Angello L. A. 
Published: March 2015

A new probe, utilizing fiber optics, has been developed that allows one to view the flame inside an industrial gas turbine combustor. The probe was tested on a GE 7001 gas turbine at Houston Lighting & Power’s T. H. Wharton plant. Views of the flame were recorded at the nozzle and from in front of the nozzle to the combustor exit. A small portion of the flame in the adjoining combustor could also be viewed through the connecting cross fire port.
During field testing with the probe, we observed such events as delayed ignitions, start ups, and load changes. Different flame patterns could be associated with different turbine operating conditions in start up and loading sequences. An interesting flame instability was observed during engine loading where the flame attaches and detaches from the fuel nozzle. At full power the flame is fully detached from the nozzle and extends from the cross fire port to the combustor exit. Plots of flame centroid position can be correlated with the turbine load. Operators, familiar with the flame image, could diagnose certain engine malfunctions such as fuel nozzle fouling, loss of water injection, and liner damage. A viewing system is currently under development that will monitor all the combustor flames simultaneously on one video screen. This system will significantly enhance operator diagnostics by allowing comparisons between combustors and detecting imbalances or differences between combustor flames.
Flow Field Investigation in a Low Solidity Inducer by Laser Doppler Velocimeter
Authors: Boccazzi A. , Perdichizzi A. , Tabacco U. 
Published: March 2015

The results of an experimental investigation of the flow-field within a low solidity inducer at design and off design flow rate are presented and discussed; particular attention is devoted to the analysis of the flow-field, at the tip in front of the leading edge, for the flow rate close to the back-flow onset.
The flow-field was measured by means of a laser doppler velocimeter at four different axial positions: upstream, within and downstream of the inducer. Axial, tangential and relative flow angle distributions, in the measuring planes, are presented for three different flow coefficients.
At the lower flow rate, the plots show the presence of reverse flow in the region close to the hub downstream of the trailing edge. For the same flow rate, quite low axial velocities are detected at the tip. This is in agreement with pressure probe traverses carried out in a slightly downstream section; these measurements also show radial inward velocities of the same order of magnitude as the axial velocities.
Circumferentially averaged losses were evaluated from specific work and total head rise given by pressure probes.
An Experimental Study of Turbine Vane Heat Transfer With Leading Edge and Downstream Film Cooling
Authors: Nirmalan V. , Hylton L. D. 
Published: March 2015

This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine vane external heat transfer. Steady state experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters — Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio — were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio.
The vane external heat transfer data obtained in this program indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil. The data presented in this paper illustrate the interaction of these variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.
Experimental Examination of the Aerothermal Performance of a Gas Turbine Engine Test Facility
Authors: Cooper E. E. , Kodres C. A. 
Published: March 2015

The U.S. Navy standard jet engine test cell was experimentally examined to determine aerodynamic and thermodynamic characteristics. Pressures and temperatures were measured throughout the facility while testing different engines at different power settings from idle through afterburning. The data include gas prescures and temperatures measured at both the test cell intakes and exhaust from which velocities were calculated and, subsequently, mass flow rates calculated and compared.
The highest test cell wall temperature recorded was about 300 degrees Centigrade reached during A/B testing of the TF30 engine. The vacuum drawn within the facility was within structural limits, reaching a minimum of only 2.5 cm of water. Airflow approaching the engine inlet was always steady and free of eddies or recirculation.
Although the facility performed well, some intuitively puzzling flow characteristics were observed. For example, the total test cell airflow increases with engine power setting from idle up to military power (100% rpm) but remains constant as afterburning is initiated. Another characteristic to be pondered is the skewing of the jet as it mixes with the entrained ambient air.
Three-Dimensional Wake Decay Inside of a Compressor Cascade and its Influence on the Downstream Unsteady Flow Field: Part II — Unsteady Flow Field Downstream of the Stator
Authors: Poensgen C. , Gallus H. E. 
Published: March 2015

A measuring technique based on multi-sensor hot-wire anemometry has been developed to determine the unsteady three-dimensional velocity vector and the structure of turbulent flows. It then has been applied to the passage and the exit flow of an annular compressor cascade, which is periodically disturbed by the wakes of a cylinder rotor, located about 50 percent of blade chord upstream. In Part I of this paper the decay of the rotor wakes will be described first without stator and secondly through a stator passage. The time-dependent turbulent flow field downstream of this stator is discussed in Part II of this paper.
The rotor wakes have a major influence on the development of three-dimensional separated regions inside the compressor cascade, and this interaction will be addressed in both parts of this paper.
Heat Transfer and Aerodynamics of a High Rim Speed Turbine Nozzle Guide Vane Tested in the RAE Isentropic Light Piston Cascade (ILPC)
Authors: Harasgama S. P. , Wedlake E. T. 
Published: March 2015

Detailed heat transfer and aerodynamic measurements have been made on an annular cascade of highly loaded nozzle guide vanes. The tests were carried out in an Isentropic Light Piston test facility at engine representative Reynolds number, Mach number and gas-to-wall temperature ratio.
The aerodynamics indicate that the vane has a weak shock at 65–70% axial chord (mid span) with a peak Mach number of 1.14. The influence of Reynolds number and Mach number on the Nusselt number distributions on the vane and endwall surfaces are shown to be significant. Computational techniques are used for the interpretation of test data.
Solar Dynamic CBC Power for Space Station Freedom
Authors: Harper William B. , Boyle Robert V. , Kudija Charles T. 
Published: March 2015

The National Aeronautics and Space Administration (NASA) has selected a closed Brayton cycle (CBC) solar dynamic (SD) power conversion system to provide electric power for the growth phase of Space Station Freedom. This selection was based in part on the comparitive reduction in solar radiation collection made possible by the high efficiency, closed Brayton cycle system when compared to the efficiency of static conversion systems. The smaller required collector area results in a significant reduction in drag in the low earth orbit which is planned for the Space Station.
Under contract to the Rocketdyne Division of Rockwell International, the Allied-Signal Aerospace Company, Garrett Fluid Systems Division is developing the Solar Dynamic Power Generating Subsystem (PGS). This subsystem consists of a solar receiver, which accepts sunlight from the multifaceted collector mirror (the concentrator), and a CBC heat engine that converts the solar heat into electricity. This paper will review the design requirements for the Space Station Solar Dynamic Power Module and present results of the PGS cycle analysis and preliminary design.
Aerodynamic Analysis of a Two Stage Transonic Compressor With Variable Stator Vanes
Authors: Broichhausen K.-D. , Harster P. 
Published: March 2015

Within the scope of a compressor research program a highly loaded two stage transonic compressor was investigated.
The submitted paper presents global information on the performance characteristics of the compressor as a whole and of the two stages as a function of the variable Stator and inlet guide vane settings. For the investigated two stage compressor a variable stator 1 is able to improve the performance to the same extent as a variable inlet guide vane.
Subsequently on the basis of L2F velocimetry, hotfilm gauges and probe traverses a more detailed view of the flow inside of the compressor is given. The main results can be summarized as follows:
In spite of high supersonic relative Mach numbers in front of rotor 1 the incidence angles change evenly from hub to tip depending on the back pressure. This is in contrast to plane cascade experience and is attributed to the bow shocks.
In the downstream rotor the shock pattern is strongly influenced by the relative position between rotor 1 and rotor 2 blades. As the glue on hotfilms on the stator 2 suction side show, the influence of the rotor 1 wakes can still be recognized in the laminar separation bubble and the turbulent boundary layer. Downstream of the compressor the mixing of the rotor wakes causes strong pitchwise gradients of total temperature.
Viscous Flow Computations in Turbomachine Cascades
Authors: Chevrin P.-A. , Vuillez C. 
Published: March 2015

Accurate prediction of the flow in turbomachinery requires numerical solution of the Navier-Stokes equations. A two-dimensional Navier-Stokes solver developed at ONERA for the calculation of the flow in turbine and compressor cascades was adapted at SNECMA to run on different types of grid.
The solver uses an explicit, time-marching, finite-volume technique, with a multigrid acceleration scheme. A multi-domain approach is used to handle difficulties due to the geometry of the flow. An H-C grid was used in the calculations. Two turbulence models, based on the mixing length approach, were used.
The flow in a transonic compressor cascade, a subsonic and a transonic turbine cascade were computed. Comparison with experiments is presented.
Development of Materials and Coating Processes for High Temperature Heavy Duty Gas Turbines
Authors: Nakamura S. , Fukui Y. , Kojima Y. , Siga M. , Yamaguchi S. 
Published: March 2015

This paper describes materials developed for the first stage nozzle and all stages turbine disc, and the new coating processes of the first stage bucket for high temperature, heavy duty gas turbines.
A cobalt base casting superalloy (30Cr-10Ni-7W-0.2Nb-0.15Zr-0.2Ti-bal.Co), having high creep rupture strength and ductility, was used for the first stage nozzle and 12Cr forging material, having high creep rupture strength, was used for all stages turbine disc of the high temperature, heavy duty gas turbine, the H25. To increase hot corrosion properties of the first stage bucket, MCrAlY was applied on the bucket external surfaces by a low-pressure plasma-spray coating technique. A procedure of chemical vapor deposition for Al coating on the bucket internal cooling paths was also established for future high temperature gas turbines.
Experimental Evaluation of a Low Emissions, Variable Geometry, Small Gas Turbine Combustor
Authors: Smith K. O. , Samii M. H. , Mak H. K. 
Published: March 2015

The results of an on-engine evaluation of an ultra-low NOx, natural gas-fired combustor for a 200 kW gas turbine are presented. The combustor evaluated used lean-premixed combustion to reduce NOx emissions and variable geometry to extend the range over which low emissions were obtained.
Test results showed that ultra-low NOx emissions could be achieved from full load down to approximately 50% load through the combination of lean-premixed combustion and variable primary zone airflow.
SNECMA M88 Engine Development Status
Authors: Corde Jean Christophe 
Published: March 2015

After describing the historical origins of the SNECMA M88 engine development program, the paper presents the general development planning until the production engine flight qualification at the end of 1995.
Then the technical aspects of the program are emphasized: the main new technologies involved in the engine, the general arrangement and the performance are described, and the engine development status is fully presented (development philosophy, main component tests, and complete engine tests).