bluuebunny/crossref_metadata_2025 huggingface dataset:
Surface Roughness Measurements on Gas Turbine Blades
Authors: Taylor Robert P.
Published: March 2015

Results are presented from profilometer measurements of the surface roughness on inservice turbine engine blades from F-100 and TF-39 aeroengines. On each blade, one roughness profile is taken in the region of the leading edge, the mid-chord and the trailing edge on both the pressure and suction sides for a total of 6 profiles. Thirty 1st stage turbine blades are measured from each engine. Statistical computations are performed on these profiles and the root mean square height, skewness and kurtosis of the roughness height distribution are presented along with the correlation length of the autocorrelation function.

The purpose of this work is to provide insight into the nature of surface roughness characteristics of inservice turbine blades which can be used in the development of scaled laboratory experiments of boundary layer flow and heat transfer on turbine engine blades.
Prediction of Heat Transfer in a Rotating Cavity With a Radial Outflow
Authors: Ong C. L. , Owen J. M.
Published: March 2015

Solutions of the differential boundary-layer equations, using the Keller-box scheme and the Cebeci-Smith eddy-viscosity model for turbulent flow, have been used to predict the Nusselt numbers on the discs of a heated rotating cavity with a radial outflow of cooling air.

Computed Nusselt numbers were in satisfactory agreement with analytical solutions of the elliptic equations for laminar flow and with solutions of the integral equations for turbulent flow. For a wide range of flow rates, rotational speeds and disc-temperature profiles, the computed Nusselt numbers were in mainly good agreement with measurements obtained from an air-cooled rotating cavity. It is concluded that the boundary-layer equations should provide solutions accurate enough for application to air-cooled gas-turbine discs.
User Experience: Operating a 300 MW Base Load Cogeneration Plant With High Water Injection Rates to Control NOx Emissions
Authors: Hauhe William E. , Haub Gary L. , Myers Charles O. , Guthan Donald C. , Fitts David O.
Published: March 2015

This paper describes user experience with the operation and maintenance of a gas turbine based cogeneration plant operating at base load while injecting up to 80 gpm (303 l/min) of water to control NOx emissions to 42 ppmv (at 15% O2). The plant, located in the Kern River Oil Field, near Bakersfield, California, has produced an average of 294.6 MWe and 1.903 million lbs/hr (0.863 million kg/hr) of steam since achieving commercial operation in August, 1985. To date, the plant has achieved an operational reliability and availability of 98.9% and 95.4%, respectively. The effects of water injection on combustion hardware, as well as, overall gas turbine reliability and availability and equipment enhancements will be discussed.
Annular Cascade Study of Low Back-Pressure Supersonic Fan Blade Flutter
Authors: Kobayashi Hiroshi
Published: March 2015

Low back-pressure supersonic fan blade flutter in the torsional mode was examined using a controlled-oscillating annular cascade test facility. Precise data of unsteady aerodynamic forces generated by shock wave movement due to blade oscillation and the previously measured data of chordwise distributions of unsteady aerodynamic forces acting, on an oscillating blade were joined, and then the nature of cascade flutter was evaluated. These unsteady aerodynamic forces were measured by direct and indirect pressure measuring methods. Our experiments covered a range of reduced frequencies based on a semi-chord from 0.0375 to 0.547, 6 interblade phase angles and inlet flow velocities from subsonic to supersonic flow.

The occurrence of unstalled cascade flutter in relation to reduced frequency, interblade phase angle and inlet flow velocity was clarified including the role of unsteady aerodynamic blade surface forces on flutter. Reduced frequency of the flutter boundary increased greatly when the blade suction surface flow became transonic flow. Interblade phase angles which caused flutter were in the range from 40° to 160° for flow fields ranging from high subsonic to supersonic. Shock wave movement due to blade oscillation generated markedly large unsteady aerodynamic forces which stimulated blade oscillation.
Optical Measurement of Unducted Fan Blade Deflections
Authors: Kurkov Anatole P.
Published: March 2015

A nonintrusive optical method for measuring unducted fan (or propeller) blade deflections is described and evaluated. The measurement does not depend on blade surface reflectivity. Deflection of a point at the leading edge and a point at the trailing edge in a plane nearly perpendicular to the pitch axis is obtained with a single light beam generated by a low-power, helium-neon laser. Quantitative analyses are performed from taped signals on a digital computer. Averaging techniques are employed to reduce random errors. Measured static deflections from a series of high-speed wind tunnel tests of a counterrotating unducted fan model are compared with available predicted deflections, which are also used to evaluate systematic errors.
Through Flow Theory for Nonaxisymmetric Turbomachinery Flow: Part I — Formulation
Authors: Dring Robert P. , Oates Gordon C.
Published: March 2015

Through flow theory has been limited in its applicability and in its accuracy by the fact that it has not historically been cast in a form which can account for the nonaxisymmetries that naturally arise in turbomachinery flow due to the presence of finite numbers of rotor and stator airfoils. The attempt to circumvent this limitation by the introduction of an aerodynamic blockage factor has been demonstrated in earlier work to produce fundamental inconsistencies in the calculation which lead to significant errors in the regions of the flow where the nonaxisymmetries are severe. The formulation in Part I of the present work is a derivation of a system of through flow equations for nonaxisymmetric flow.

A benchmark data base is used in Part II to provide input to the calculation and to help identify the dominant terms. It is demonstrated that the dominant effect of nonaxisymmetry is contained in two terms that relate the total pressure of the averaged flow to the mass averaged total pressure. It is also demonstrated that the present formulation produces a result which is more accurate than that obtained with the historical blockage-based formulation.
Practical Design Considerations for Integrating the Propulsion System With the Aircraft for Jetborne Flight
Authors: Groll D. B. , Bossart T. R.
Published: March 2015

Results of a pilot-in-the-loop simulator evaluation for engine and configuration variations to the AV-8B Harrier II aircraft are presented. Configurations were tested for various center of gravity locations, weights, pitching moments of inertia, and thrust center locations in jetborne flight. The effects of the propulsion system’s characteristics in relation to aircraft parameters are generalized with regard to pitch handling qualities such as control power, pitch sensitivity, disturbance susceptibility, and control margin. A proposed modification to improve pitch handling qualities was also evaluated and results for this modification are discussed.
Part-Circumference Casing Treatment and the Effect on Compressor Stall
Authors: Cumpsty N. A.
Published: March 2015

Results are presented and discussed from an axial compressor rotor operated with an axial skewed slot casing treatment over part of the circumference. The compressor was one for which stall was initiated in the tip region and for this type there is some potential for stall margin improvement with lower loss using this. The main significance of the experiments is, however, the possibility of looking at aspects of stall inception. Normally stall inception is a brief transient with an unknown start time and is difficult to study but with the partial casing treatment it was possible to make the untreated section operate continuously in such a way that it underwent the processes normally leading to stall. For a tip stalling rotor the experiments identify the annulus boundary layer as the crucial region of the flow and spillage of the tip-clearance flow forward of the blades as a process leading to the rapid build up of blockage prior to instability and stall.
Advanced Oxidation Protection System for Carbon-Carbon Composites
Authors: Eitman D. A. , Kidd R. W. , Dirling R. B.
Published: March 2015

Carbon-carbon composites possess a number of desirable attributes including low density, high strength and stiffness at temperatures well beyond the capabilities of refractory alloys, low thermal expansion coefficient, tailorable orthotropic properties, absence of strategic materials, and resistance to thermal shock, fatigue, and brittle failures. However, for many applications of interest (such as aircraft and aerospace vehicle structure and engines) resistance to oxidation in high-temperature air or engine exhaust streams is a requirement which is not satisfied by unprotected carbon-carbon composites. The elements of an advanced oxidation protection system for carbon-carbon composites are described in this paper. The system is comprised of both an oxidation resistant coating intended to provide the primary barrier to oxygen ingress and inhibitors added to the matrix of the carbon-carbon composite to increase its oxidation resistance without significant losses in mechanical properties. The composite inhibition system is designed to be complementary to the coating and to enhance its long-term performance. A description of the principal elements of the system is presented along with recent test data and current research directions.
A Computer Aided Method for the Design of Centrifugal Impellers Using Optimization Techniques
Authors: Al-Zubaidy Sarim Jawad
Published: March 2015

The paper proposes a design method which may be used to evaluate the impeller detailed geometry for a prescribed mean stream velocity distribution. The analysis makes extensive use of numerical optimization, firstly in the step by step channel geometry calculation, and secondly in controlling the manner in which the relative velocity components are permitted to vary along the mean streamline for the condition of minimum entropy generation along the flow path. It should be mentioned that the components of the relative velocity were not allowed to take arbitrary values. The variation of the radial, axial and tangential components was governed in such a way that the prescribed diffusion schedule is satisfied at every point.

An example impeller is shown, which was designed using the proposed method. The general configuration of the example design shows it to be not difficult to manufacture.
Boundary Layer Transition of Strongly Concave Surfaces
Authors: Riley Stephen , Johnson Mark W. , Gibbings John C.
Published: March 2015

Boundary layer transition has been studied on two blades of constant 0.5 and 1 metre radius of curvature with free stream turbulence levels of 0.7%, 2.6% and 7.2%. Zero pressure gradients were used throughout. Strong Gortler vortices developed in the boundary layer which led to growth rates of up to ten times the flat plate rate. The boundary layer profile was also highly distorted by the vortices. Transition correlation techniques for flat plates proved totally inadequate for the concave surface data, but a method of obtaining correlations for these surfaces was suggested by considering the inner critical region of the boundary layer alone.
STOVL Hot Gas Ingestion Control Technology
Authors: Amuedo K. C. , Williams B. R. , Flood J. D. , Johns A. L.
Published: March 2015

A comprehensive wind tunnel test program was conducted to evaluate control of Hot Gas Ingestion (HGI) on a 9.2% scale model of the McDonnell Aircraft Company model 279-3C advanced Short Takeoff and Vertical Landing (STOVL) configuration. The test was conducted in the NASA-Lewis Research Center 9 foot by 15 foot Low Speed Wind Tunnel during the summer of 1987. Initial tests defined baseline HGI levels as determined by engine face temperature rise and temperature distortion. Subsequent testing was conducted to parametrically evaluate HGI control using: Lift Improvement Devices (LIDs), forward nozzle splay angle, combination of LIDs and forward nozzle splay angle, and main inlet blocking. The results from this test program demonstrate that HGI can be effectively controlled and that HGI is not a barrier to STOVL aircraft development.
A New Low-Thermal-Expansion, High-Strength Alloy for Gas Turbines
Authors: Srivastava S. K. , Lai George Y.
Published: March 2015

HAYNES® alloy No. 242, based on Ni-Mo-Cr system, was developed primarily for turbine seal applications. The key characteristics of the alloy are derived from critical control of Ni, Mo, Cr, and several minor elements. The high strength — about twice that of alloy B — is achieved from a long-range ordering reaction. The paper reviews comparative evaluation of 242™ alloy versus several currently used gas-turbine alloys such as alloys B, S, N and INCOLOY® alloy 909. Environmental resistance is particularly emphasized. It is shown that 242 alloy possesses an excellent combination of low-thermal-expansion, high strength, and environmental resistance for gas turbine service up to 1400°F (760°C).
Influences of Splitter Blades on the Centrifugal Fan Performances
Authors: Gui Lichuan , Gu Chuangang , Chang Hongshou
Published: March 2015

Centrifugal fan and compressors with splitter blades are widely utilized in engineering practices. A detail investigation of the influences of splitter blades on a forward-curved centrifugal fan performances is presented in the paper. The study includes two parts: experiment and numerical calculation. The experiments were produced in a specific impeller with adjustable splitter blades in order to get the performances in different conditions. The internal flow field in the impeller was calculated by means of FEASM (Finite Element Approximate Solution Method, Gu, 1984).

The experimental result show that changing the circumferential positions of the splitter blades has a noticeble influence on the fan performance, the incidence of splitter blades also has a certain effect on it, and properly lengthened splitter blades can raise the total pressure coefficient.

The velocity and load distributions on the blade surfaces calculated can be used to analyse the phenomena above satisfactorily.
Blade Mistuning Coupled With Shaft Flexibility Effects in Rotor Aeroelasticity
Authors: Khader Naim , Loewy Robert G.
Published: March 2015

The effect of bladed-disk polar dissymmetry, resulting from variations in mass from one blade to another, on aeroelastic stability boundaries for a fan stage is presented. In addition to both in-plane and out-of-plane deformations of the bladed-disk, bending of the supporting shaft in two planes is considered, and the resulting Coriolis forces and gyroscopic moments are included in the analysis. A quasi-steady aerodynamics approach is combined with the Lagrangian method to develop the governing equations of motion for the flexible bladed-disk-shaft assembly. Calculations are performed for an actual fan stage.
Condition Monitoring of Aeroderivative Gas Turbines
Authors: Boyce M. P. , Meher-Homji Cyrus B. , Wooldridge Byron
Published: March 2015

Aeroderivative gas turbines play a critically important role in the gas distribution and cogeneration industry. The need for maintaining high efficiency and availability has created the need for condition monitoring systems. This paper covers aspects relating to online condition monitoring techniques.

The subject matter of this paper applies to all aeroderivative gas turbines. Owners and operators of these machines can reduce operating costs and enhance availability and performance by paying close attention to the operating condition of their equipment. By doing so, performance tuning for optimum operating costs can be accomplished. Likewise, preventive maintenance efforts can be enhanced.
Gas Turbine Compressor Operating Environment and Material Evaluation
Authors: Haskell R. W.
Published: March 2015

The reliability and performance of a gas turbine compressor is strongly dependent upon the environment in which it operates, the materials which are used, and the filtration system. Erosion and to a certain extent fouling can be controlled by the filtration system, but corrosion is largely controlled through site and material selection. The factors which determine the corrosivity of a site are humidity, the concentration of acid-forming gases, and the composition of particulates. The interrelationships of these factors are discussed with an aim of reducing their impact on compressor operation. A necessary condition for corrosion is the presence of moisture. The acidity of the moisture results from its interaction with the gases and particulates of the environment. The details of these interactions which are important to turbine operators are discussed. A considerable amount of corrosion testing of base materials and coatings has been performed and this is reviewed. A table is presented for selection of compressor materials based on the nature of the site environment and the type of compressor filtration.
Viscous Flow Calculations in Turbomachinery Channels
Authors: Martelli Francesco , Michelassi Vittorio
Published: March 2015

An implicit procedure based on the artificial compressibility formulation is presented for the numerical solution of the two-dimensional incompressible steady Navier-Stokes equations in the presence of large separated regions. Turbulence effects are accounted for by the Chien low Reynolds number form of the K-ε turbulence model and the Baldwin-Lomax algebraic expression for turbulent viscosity. The governing equations are written in conservative form and implicitly solved in fully coupled form using the approximate factorization technique. Preliminary tests were carried out in a laminar flow regime to check the accuracy and stability of the method in two-dimensional and cylindrical axisymmetric flow configurations. After testing in laminar and turbulent flow regimes and comparing the two turbulence models, the code was successfully applied to an actual gas turbine diffuser at low Mach numbers.
Compressor Blade Boundary Layers: Part 1 — Test Facility and Measurements With No Incident Wakes
Authors: Dong Y. , Cumpsty N. A.
Published: March 2015

The boundary layers on compressor blades are sensitive to the conditions at which transition occurs and transition can be affected by the convection of wakes from upstream blade rows. This paper and its companion, Part 2 by the same authors, describes an experiment to study the effect of the moving wakes on the boundary layer of a compressor blade. This paper describes the background and facility devised to introduce wakes together with results obtained on the blades in tests without the wakes present. Part II describes the measurements made with the wakes present and presents conclusions for the whole project. Further details of all aspects of the work can be found in Dong (1988).
Compressor Blade Boundary Layers: Part 2 — Measurements With Incident Wakes
Authors: Dong Y. , Cumpsty N. A.
Published: March 2015

This paper follows directly from Part I** by the same authors and describes measurements of the boundary layer on a supercritical-type compressor blade with wakes from a simulated moving upstream blade row convected through the passage. (The blades and the test facilities togehter with the background are described in Part I.) The results obtained with the wakes are compared to those with none for both low and high levels of inlet turbulence. The transition process and boundary layer development is very different in each case though the overall momentum thickness at the trailing edge is fairly similar. None of the models for transition is satisfactory when this is initiated by moving wakes.
Blade Sweep for Low-Speed Axial Fans
Authors: Wright Terry , Simmons William E.
Published: March 2015

The available literature on aerodynamic and acoustic properties of axial fans with swept blades is presented and discussed with particular emphasis on noise mechanisms and the influence of high-intensity inlet turbulence on “excess” noise. The acoustic theory of Kerschen and Envia for swept cascades is applied to the problem of axial fan design. These results are compared to available data and a provisional model for specifying sweep angles is presented. The aerodynamic performance theory for swept-bladed rotors of Smith and Yeh is adapted for use in designing low speed axial fans.

Three prototype fans were designed using the resultant computer codes. One is a baseline fan with blade stacking lines radially oriented, and two are fans having swept blades of increasingly greater forward sweep. Aerodynamic testing shows that performance of the fans lie within a band width of about ± two percent of volume flow rate and pressure rise predictions in the region of design performance, effectively validating the design procedure for selection of the blading parameters.

Noise testing of the fans was carried out and the results show an average noise reduction for the swept-bladed fans of about 7 dBA overall, and a reduction of pure tone noise at blade-pass frequency of about 10 dB compared to the zero-sweep baseline model in close agreement with the theory of Kerschen and Envia.
The Influence of the Recirculation Region: A Comparison of the Convective Heat Transfer Downstream of a Backward-Facing Step and Behind a Jet in a Cross Flow
Authors: Scherer V. , Wittig S.
Published: March 2015

Convective heat transfer is examined in two typical examples of separated flows, namely: the flow over a backward-facing step and a two-dimensional jet entering a cross flow. Local Nusselt numbers were determined in and behind the recirculation region. The main parameters influencing the heat transfer, the Reynolds number and the momentum flux ratio of the jet and the cross flow, have been varied in a wide range. In addition to heat transfer measurements, the flow field has been documented using a LDA-system and oil film techniques. The static pressure distribution at the wall within the separated flow is also given. The measurements are compared with the results of a numerical code, based on a finite volume method, where the well known k-ε-model is employed. The differences in Nusselt numbers predicted with a one- and a two-layer model are shown to demonstrate the influence of wall functions on heat transfer. The numerical and experimental results are compared with available data, and the differences and similarities in the heat transfer behaviour of separated flows are discussed.
Measurement and Calculation of the Three-Dimensional Flow in Axial Compressor Stators, With and Without End-Bends
Authors: Robinson C. J. , Northall J. D. , McFarlane C. W. R.
Published: March 2015

This paper presents results from tests of two stators, one conventional and one with end-bends, operating at low-speed in the buried stage environment of the Cranfield, 4-stage research compressor (LSRC). The aerofoil velocity distributions are modelled on those of a high speed, 10 stage, high pressure compressor with ‘supercritical’ blading, and the stators were cantilevered with clearances of 1.8% of annulus height. The test results are compared with predictions from the Moore Elliptic Flow Program (MEFP) which solves the full, three-dimensional, Navier-Stokes equations with a pressure correction algorithm. The calculation results capture the essential physics of the viscous flow in these two bladerows. The calculated deviations agree well with experimental data across the blade spans, including the near hub-region, which is dominated by the clearance flow. The calculated, radial profiles of loss are in reasonable agreement with experiment, although the magnitude of loss is over-predicted.
Stall Inception in Axial Compressors
Authors: McDougall N. M. , Cumpsty N. A. , Hynes T. P.
Published: March 2015

Detailed measurements have been made of the transient stalling process in an axial compressor stage. The stage is of high hub-casing ratio and stall is initiated in the rotor. If the rotor tip clearance is small stall inception occurs at the hub, but at clearances typical for a multistage compressor the inception is at the tip. The crucial quantity in both cases is the blockage caused by the endwall boundary layer. Prior to stall disturbances rotate around the inlet flow in sympathy with rotating variations in the endwall blockage; these can persist for some time prior to stall, rising and falling in amplitude before the final increase which occurs as the compressor stalls.
A Comparison Between the Design Point and Near Stall Performance of an Axial Compressor
Authors: McDougall N. M.
Published: March 2015

Detailed measurements have been made within an axial compressor operating both at design point and near stall. Rotor tip clearance was found to control the performance of the machine by influencing the flow within the rotor blade passages. This was not found to be the case in the stator blade row, where hub clearance was introduced beneath the blade tips. Although the passage flow was observed to be altered dramatically, no significant changes were apparent in the overall pressure rise or stall point.

Small tip clearances in the rotor blade row resulted in the formation of corner separations at the hub, where the blade loading was highest. More representative clearances resulted in blockage at the tip due to the increased tip clearance flow. The effects which have been observed emphasize both the three dimensional nature of the flow within compressor blade passages, and the importance of the flow in the endwall regions in determining the overall compressor performance.
Optimum Weight Design of a Rotor Bearing System With Dynamic Behavior Constraints
Authors: Shiau Ting Nung , Hwang Jon Li
Published: March 2015

An efficient design algorithm for optimum weight design of a rotor bearing system with dynamic behavior constraints is investigated. The constraints include the restrictions on stresses, unbalance response, and/or critical speeds. The system dynamic behaviors are analyzed by the finite element method. And the exterior penalty function method is used as the optimization technique to minimize the system weight.

The system design variables are the cross-sectional areas of the shaft and the stiffnesses of the bearings. The sensitivity analysis of the system parameters is also investigated. The example of a single spool rotor bearing system is employeed to demonstrate the merits of the design algorithm with different combination of dynamic behavior constraints. At the optimum stage, it is shown that the weight of rotor system can be significantly reduced. Moreover, the optimum design weights are quite difference for various combinations of dynamic behavior constraints.
Unbalance Response of a Jeffcott Rotor Incorporating Short Squeeze Film Dampers
Authors: El-Shafei A.
Published: March 2015

The steady state unbalance response of a Jeffcott rotor incorporating short squeeze film dampers executing circular centered whirl is obtained by a fast algorithm. Savings in execution time of the order of fifty times are gained over numerical integration. Fluid inertia forces are included in the model of the squeeze film dampers. The fast algorithm allows parametric studies to be performed. It is shown that fluid inertia results in the excitation of a second mode for the Jeffcott rotor, decreases the possibility of jump resonance, and decreases the useful range of vibration isolation of the dampers. It is also shown that a squeeze film damper with no centering spring (or a very soft spring) may be advantageous with regards to the unbalance response and the vibration isolation capability of the dampers.
Life Extension and Performance Enhancement of an Industrial Gas Turbine Through Upgrading
Authors: Hawthorne Paul R. , Davis Robert A.
Published: March 2015

This paper describes the upgrade of a GE Frame 3 industrial gas turbine with new advanced technology parts and a new microprocessor-based control system. Based on pre- and post-upgrade performance tests, the gas turbine heat rate was reduced by 9.4%. In addition, with the new triple redundant control system and upgraded hardware, unit availability is anticipated to improve by 18% over the extended life of the gas turbine. The paper describes the mechanical and control changes and the method used to measure the enhancement of unit performance.
Secondary Flow due to the Tip Clearance at the Exit of Centrifugal Impellers
Authors: Ishida Masahiro , Senoo Yasutoshi , Ueki Hironobu
Published: March 2015

The velocity distribution was measured at the exit of two different types of unshrouded centrifugal impellers under four different tip clearance conditions each; one with twenty radial blades and inducers and the other with sixteen backward-leaning blades. And the effect of tip clearance on input power was also measured. By increasing the tip clearance, the input power was hardly changed in the radial blade impeller and was reduced in the backward-leaning blade impeller. The velocity distribution normalized by the passage width between hub and shroud wall was hardly changed at the exit of the radial blade impeller by varying the tip clearance, on the other hand, the relative flow angle was reduced significantly and monotonously by an increase of tip clearance in the backward-leaning blade impeller. The change in input power due to the tip clearance was clearly related to the change of flow pattern at the exit of impeller due to the secondary flow, which is most likely caused by the component, normal to the blade, of the shear force to support the fluid in the clearance space against the pressure gradient in the meridional plane without blades.
Analysis of Three-Dimensional Turbomachinery Flows on C-Type Grids Using an Implicit Euler Solver
Authors: Weber K. F. , Thoe D. W. , Delaney R. A.
Published: March 2015

A 3-D Euler analysis for turbomachinery flows on a C-type grid is presented. The analysis is based on the Beam and Warming implicit algorithm for solution of the unsteady Euler equations and is derived from the ARC3D code developed by Pulliam at NASA Ames Research Center. Modifications made to convert this code from external flow applications to internal turbomachinery flows are given in detail. These changes include the addition of inflow, outflow, and periodic boundary point calculation procedures. Also presented are the C-grid construction procedures. Finally, results of code experimental verification studies for 3-D compressor cascade and rotor flows are presented.
Ten Years’ Operating History of a Large Gas Turbine/Recompression Train
Authors: Conquergood C. Peter , Blauser Dave , Willbourn Peter
Published: March 2015

In 1978, Shell Canada Limited commissioned a large aero-derivative gas turbine driven compressor unit in its Waterton Straddle Plant. This unspared unit provides the primary recompression service in the “Deep Cut” ethane extraction facility. Significant operating features of this unit include flat rating and three stages of waste heat recovery from the turbine exhaust.

Throughout its history, this unit has demonstrated over 99% reliability and has operated for long periods without significant maintenance. All routine turbine maintenance has been accomplished on-site.

This paper describes the features of the installation, the operating and maintenance philosophy, and the experience obtained from ten years’ service, thus providing the reader with insight in regard to features and practices which can provide for a successful installation.
Preliminary Performance Estimates for a GE Steam Injected LM1600 Gas Turbine
Authors: Thames J. M. , Coleman R. P.
Published: March 2015

Preliminary performance modeling and conceptual designs have been completed for the addition of a steam injection system to a natural gas fired General Electric (GE) LM1600 gas turbine. The steam injected LM1600 gas turbine (STIG™ system) is an attractive size, 13 to 17 mw, for both cogeneration and small power applications. The STIG™ system offers numerous benefits including 1) increased power output in a compact single engine system, 2) improved efficiency, 3) reduced exhaust emissions, 4) steam-to-power flexibility, and 5) attractive system economics.

This paper describes the LM1600 STIGTM system conceptual design, performance estimates and projected emissions.
Heat Transfer and Pressure Drop for Short Pin-Fin Arrays With Pin-Endwall Fillet
Authors: Chyu M. K.
Published: March 2015

The effects of array configuration and pin-endwall fillet on the heat transfer and pressure drop of short pin-fin arrays are investigated experimentally. The pin-fin element with endwall-fillet, typical in actual turbine cooling applications is modeled by a spool-like cylinder. The arrays studied include an in-line and a staggered array, each having 7 rows of 5 pins. These arrays have the same geometric parameters, i.e. H/D = 1, S/D = X/D = 2.5, and the Reynolds number ranging from 5 × 103 to 3 × 10. One of the present results shows that the staggered array always has a higher array-averaged mass transfer coefficient than its in-line counterpart. However, the pressure drop for the staggered array is higher compared to the in-line configuration. These trends are unaffected by the existence of the pin-endwall fillet. Another significant finding is that an array with pin-endwall fillet generally produces lower heat transfer coefficient and higher pressure drop than that without endwall-fillet. This leads to the conclusion that pin-endwall fillet is undesirable for heat transfer augmentation. In addition, naive use of the heat transfer results obtained with perfectly circular cylinders tends to overestimate the pin-fin cooling capability in the actual turbine. The effects of endwall-fillet on the array heat transfer and pressure drop are much more pronounced for the staggered array than for the in-line array; however, they diminish as the Reynolds number increases.
The Influence of Air/Fuel Ratio and Swirl Number on the Combustion Characteristics of a Model Combustor
Authors: Bicen A. F. , Tse D. G. N. , Whitelaw J. H.
Published: March 2015

Combustion characteristics of a model can-type combustor are reported for air/fuel ratios encompassing take-off and ground-idle conditions and for two swirlers with gaseous fuel and at atmospheric pressure. Temperatures were obtained with fine-wire thermocouples, and concentrations of UHC, H2, CO, CO2, O2 sampled through a water-cooled probe with a flame ionization detector, a gas chromotograph and infrared and paramagnetic analyzers.

The results indicate combustion efficiencies greater than 97%. For the air/fuel ratio corresponding to that of take-off, 45% of the mass flow of the primary jets turned upstream after impingement, combustion occurred in the wakes of the swirler and of the primary jets resulting in a pattern factor of 0.46. At ground-idle condition, 58% of the primary jet flow turned upstream reducing the pattern factor to 0.43. An 18% reduction in the swirl number at the ground-idle condition led to a pattern factor to 0.35 and, for takeoff, to a pattern factor of around 0.37 with combustion occurred mainly in the intermediate zone.
Vibration Analysis of Cracked Turbine and Compressor Blades
Authors: Ostachowicz W. , Krawczuk M.
Published: March 2015

The object of the paper is to present a method of vibration analysis of cracked turbine and compressor blades. A transverse crack in a blade introduces local flexibility, which changes dynamic characteristics of a structure.

The assumption of the open and closed crack leads to a model with point finite elements which are introduced to the crack model. The blade is modelled by spatial finite elements.

The method described in the paper can be used for the study of the effects of crack locations and sizes on the vibrational behaviour of the structure. The analysis is provided along with an algorithm and an example of calculations.

The authors have compared the calculation results of the beam model with a crack (regardless of friction in the crack) with the investigation results of paper [2] and have secceeded in obtaining a high compatibility of the results. Hovever, the comparison of the results relevant to the model taking into account friction in the crack was impossible due to inavailabity of similar papers in literature.
Influence of High Mainstream Turbulence on Leading Edge Film Cooling Heat Transfer
Authors: Mehendale A. B. , Han J. C.
Published: March 2015

The influence of high mainstream turbulence on leading edge film effectiveness and heat transfer coefficient was studied. High mainstream turbulence was produced by a passive grid and a jet grid. Experiments were performed using a blunt body with a semi-cylinder leading edge with a flat afterbody. The mainstream Reynolds number based on leading edge diameter was about 100,000. Spanwise and streamwise distributions of film effectiveness and heat transfer in the leading edge and on the flat sidewall were obtained for three blowing ratios, through rows of holes located at ±15° and ±40° from stagnation. The holes in each row were spaced three hole-diameters apart and were angled 30° and 90° to the surface in the spanwise and streamwise directions respectively. The results indicate that the film effectiveness decreases with increasing blowing ratio, but the reverse is true for the heat transfer coefficient. The leading edge film effectiveness for low blowing ratio (B = 0.4) is significantly reduced by high mainstream turbulence (Tu = 9.67% and 12.9%). The mainstream turbulence effect is diminished in the leading edge for higher blowing ratios (B = 0.8 and 1.2) but still exists on the flat sidewall region. Also, the leading edge heat transfer coefficient for blowing ratio of 0.8 increases with increasing mainstream turbulence; but the effect for other blowing ratios (B = 0.4 and 1.2) is not so systematic as for B = 0.8. Surface heat load is significantly reduced with leading edge film cooling.
The Influence of Curvature on Film Cooling Performance
Authors: Schwarz S. G. , Goldstein R. J. , Eckert E. R. G.
Published: March 2015

The effects of injection rate and strength of curvature on film cooling performance of gas injected through a row of holes on a convex surface is studied. Comparisons are made to film cooling of concave and flat surfaces. Three different relative strengths of curvature (ratio of radius of curvature to radius of injection hole), two density ratios (0.95 and 2.0), and a wide range of blowing rates (0.3 to 2.7) are considered. A foreign gas injection technique (mass transfer analogy) is used. The strength of curvature was controlled by varying the injection hole diameter. At low blowing rates, film cooling is more effective on the convex surface than on a flat or a concave surface. The cross stream pressure gradient present in curved flows tends to push the jet into the convex wall. As the injection rate is increased, normal and tangential jet momentum promote lift-off from the convex surface, thereby lowering performance. In contrast, previous studies show that on a concave surface, tangential jet momentum, flow instabilities, and blockage improve performance on a concave surface as blowing rate is increased.
Advances in the Numerical Analysis of Linearized Unsteady Cascade Flows
Authors: Usab W. J. , Verdon J. M.
Published: March 2015

This paper describes two new developments in the numerical analysis of linearized unsteady cascade flows, that have been motivated by the need for an accurate analytical procedure for predicting the onset of flutter in highly loaded compressors. In previous work, results were determined using a two-step or single-pass procedure in which a solution was first determined on a rectilinear-type cascade mesh to determine the unsteady flow over an extended blade-passage solution domain and then on a polar-type local mesh to resolve the unsteady flow in high-gradient regions. In the present effort a composite procedure has been developed in which the cascade- and local-mesh equations are solved simultaneously. This allows the detailed features of the flow within the local-mesh region to impact the unsteady solution over the entire domain. In addition, a new transfinite local mesh has been introduced to permit a more accurate modeling of unsteady shock phenomena. Numerical results are presented for a two-dimensional compressor-type cascade operating at high subsonic inlet Mach number and high mean incidence to demonstrate the impact of the new composite- and local-mesh analyses on unsteady flow predictions.
Three-Dimensional Rotational Flow in Transonic Turbomachines: Part I — Solution Obtained by Using a Number of S1 Stream Filaments of Revolution and a Central S2 Stream Filament
Authors: Chung-Hua Wu , Zhengming Wang , Hongji Chen
Published: March 2015

The general theory of three-dimensional flow in subsonic and supersonic turbomachines (Wu, 1952) is extended to the three-dimensional rotational flow in transonic turbomachines. In Part I of the paper, an approximation that the S1 stream filaments are filaments of revolution is made. Then, the three-dimensional solution is obtained by an iterative solution between a number of S1 stream filaments and a single S2 stream filament. Recently developed relatively simple and quick method of solving the transonic S1 flow is utilized. The complete procedure is illustrated with the solution of the 3D flow in the DFVLR rotor operating at the design point. The solution is presented in detail, special emphasis being placed on the fulfillment of the convergence requirement. The character of the three-dimensional field obtained is examined with the three-dimensional structure of the passage shock, the relative Mach number contours on a number of S1 surfaces, S2 surfaces, and cross surfaces, and the variations of the thickness of S1 and S2 filaments. Comparison between the calculated three-dimensional field with the DFVLR measured data shows: the character of the flow field and the streamwise variation of the flow variables in the middle of the flow channel are in good agreement. It is recommended that the method presented herein can be used for three-dimensional design of transonic turbomachines.
Inviscid-Viscous Interaction Analysis of Compressor Cascade Performance
Authors: Barnett M. , Hobbs D. E. , Edwards D. E.
Published: March 2015

An inviscid-viscous interaction technique for the analysis of quasi-three-dimensional turbomachinery cascades has been developed. The inviscid flow is calculated using a time-marching, multiple-grid Euler analysis. An inverse, finite-difference viscous-layer analysis, which includes the wake, is employed so that boundary-layer separation can be modeled. This analysis has been used to predict the performance of a transonic compressor cascade over the entire incidence range. The results of the numerical investigation in the form of cascade total pressure loss, exit gas angle and blade pressure distributions are compared with existing experimental data and Navier-Stokes solutions for this cascade, and show that this inviscid-viscous interaction procedure is able to accurately predict cascade loss and airfoil pressure distributions. Several other aspects of the present interaction analysis are examined, including transition and wake modeling, through comparisons with data.
Rotating Cavity With Axial Throughflow of Cooling Air: Flow Structure
Authors: Farthing P. R. , Long C. A. , Owen J. M. , Pincombe J. R.
Published: March 2015

A rotating cavity with an axial throughflow of cooling air is used to provide a simplified model for the flow that occurs between adjacent corotating compressor discs inside a gas-turbine engine. Flow visualization and laser-Doppler anemometry are employed to study the flow structure inside isothermal and heated rotating cavities for a wide range of axial-gap ratios. G. rotational. Reynolds numbers, Reφ, axial Reynolds numbers, Rez, and temperature distributions.

For the isothermal case, the superposed axial flow of air generates a powerful toroidal vortex inside cavities with large gap ratios (G > 0.400) and weak counter-rotating toroidal vortices for cavities with small gap ratios. Depending on the gap ratio and the Rossby number, ε (where ε ∝ Rez/Reφ), axisymmetric and nonaxisymmetric vortex breakdown can occur, but circulation inside the cavity becomes weaker as ε is reduced.

For the case where one or both discs of the cavity are heated, the flow becomes nonaxisymmetric: cold air enters the cavity in a “radial arm” on either side of which is a vortex. The cyclonic and anti-cyclonic circulations inside the two vortices are presumed to create the circumferential pressure gradient necessary for the air to enter the cavity (in the radial arm) and to leave (in Ekman layers on the discs). The core of fluid between the Ekman layers precesses with an angular speed close to that of the discs, and vortex breakdown appears to reduce the relative speed of precession.
Three-Dimensional Wake Decay Inside of a Compressor Cascade and its Influence on the Downstream Unsteady Flow Field: Part I — Wake Decay Characteristics in the Flow Passage
Authors: Poensgen C. , Gallus H. E.
Published: March 2015

A measuring technique based on multisensor 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.

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.
Service Temperature Estimation of Turbine Blades Based on Microstructural Observations
Authors: Aurrecoechea Jose M. , Brentnall William D. , Gast Joseph R.
Published: March 2015

Optical and electron metallographic (SEM) examination was performed on MAR-M-421 samples subjected to controlled furnace exposures, to quantify the microstructural changes associated with the prolonged high temperature exposures. Gamma prime size measurements were used to generate a mathematical model, based on diffusion controlled kinetics, designed to estimate temperatures. This computational technique was utilized to estimate exposure temperatures of turbine blades which had seen service in land based gas turbine engines. The engines had accumulated from 1,200 to more than 98,000 hours, operating under a variety of conditions. The procedure is generally applicable to commonly used gamma prime strengthened nickel-base superalloys.
Mechanics on the Tip Clearance Loss of Impeller Blades
Authors: Senoo Yasutoshi
Published: March 2015

For predicting the tip clearance loss of turbomachines, different equations are published in the literatures based on differnt principles. In 1986 the present author posturated a new theory where the pressure loss consisted of two parts, one was the pressure loss induced by the drag force of the leaked flow and the other was the pressure loss to support the axial pressure difference without blades in the tip clearance zone. There were comments such as the two losses were the same loss looked from two different view points, or at least a part of the former was included in the latter or vice versa.

In this paper the pressure loss due to the tip clearance is examined based on a macroscopic balance of forces and the two kinds of loss are derived. Furthermore, it is made clear that the former comes from the induced drag which is parallel to the blade while the latter comes from the missing blade force normal to the blade in the clearance zone. Because these two forces are mutually perpendicular, the two losses are entirely different in nature and they do not even partially overlap to each other. It is also made clear quantitatively, how the loss of the kinetic energy of leaked flow is related to the induced drag of the clearance flow.