bluuebunny/crossref_metadata_2025 huggingface dataset:
Solar Turbines Incorporated “Taurus 60” Gas Turbine Development
Authors: Van Leuven Vern
Published: February 2015

The Taurus gas turbine was first introduced in 1989 with ratings of 6200 HP for single shaft and 6500 HP for twin shaft configurations. A new design of the single shaft third stage turbine rotor and exhaust diffuser brought its power to 6500 HP in 1991. A program was initiated early in 1992 to identify opportunities to further optimize performance of the Taurus. Thorough investigation of performance sensitivity to thermodynamic cycle parameters has resulted in significant improvement over the original design with no change in firing temperature. Aerodynamic and mechanical design changes were implemented in 1993 which raised Taurus performance to 7000 HP and 32% thermal efficiency. Selection of the final design configuration was the outcome of performance maximization versus cost increase, durability risk and loss of commonality with previous engines. This paper details these changes and the design selection process.
Testing an APU for Potential Service Aboard a U.S. Naval Destroyer
Authors: Francis George W.
Published: February 2015

Incentives exist for replacing ships high pressure gas turbine emergency start air systems with auxiliary power units (APUs). The Allied Signal, Model GTCP 100-82 is one option. It is currently used in Naval aircraft start carts. Interest has been kindled in a shipboard application primarily for emergency starting Ship Service Gas Turbine Generator Sets. This APU is tested at the Naval Surface Warfare Center Carderock Division facility in Philadelphia. The target ships for this application is the future addition to the DDG-51 Class, AEGIS Destroyers. Advantages from both financial and life cycle management perspectives are expected from standardized air and sea service. This APU application concept, and variations of it, are overtly suited to a broad array of similar installations.
An Operation and Maintenance Strategy to Maximize Performance of a 70-MW Combined-Cycle Power Plant
Authors: Anuskiewicz Robert J. , Hickerson Thomas A.
Published: February 2015

A 70-MW combined-cycle power plant was constructed in January 1989 at a manufacturing plant located in York, Pennsylvania. The purpose of the power plant was to provide the manufacturing plant with a reliable source of lower-cost process steam and electricity and to produce excess electricity for sale to the local utility company. This paper is an overview of the planning and implementation of the operation and maintenance (O&M) practices which have resulted in superior cash flow from plant operations. A primary element of the power plant project was to include the O&M needs in plant design, construction, commissioning, and start-up.
300 kW Class Ceramic Gas Turbine Development (CGT 301)
Authors: Sakakida Masaru , Sasa Tadashi , Akiyama Kazuho , Tanaka Shinya
Published: February 2015

CGT 301 is a recuperated, single-shaft, ceramic gas turbine for cogeneration capable of continuous full load application. In order to reduce its size, thermal stress, and deformations, ceramic parts are designed axi-symmetrically. The combustor is located on a shaft axis just before the turbine, therefore it does not have a large scroll. The turbine is a two-stage axial flow-type with ceramic blades.

For the first phase of the program, the primary-type gas turbine with all-metallic parts was fabricated and tested under various conditions. The test results confirmed the rotation stability of the gas turbine. After the test of preliminary metallic gas turbine, all-ceramic parts were fabricated and various tests were carried out to confirm their reliability.

The configuration and structure of the ceramic turbine were improved based on the data obtained from the tests of the primary-type gas turbine and the fundamental tests for ceramic components. The primary-type ceramic gas turbine of TIT 1200°C was designed and fabricated for the second phase of the program. This paper outlines the concept of the ceramic component design, test results of ceramic parts in the hot section, and the engine test.
Digital Control Brings Large Turbofan Benefits to the Regional Jetliner Turbofan Market
Authors: Sisson Patterson B. , Faymon David K.
Published: February 2015

In keeping with the general industry trend of applying Full Authority Digital Electronic Control (FADEC) technology to small gas turbine engines, Textron Lycoming and Chandler Evans Division of Coltec Industries have developed and qualified a single channel control system for use on the Textron Lycoming LF507-1F turbofan engine. The LF507-1F is the world’s smallest FADEC-equipped airline turbofan engine and is the only FADEC-equipped turbofan developed and certified for the regional jetliner market. The application for this powerplant, the four-engine AVRO International Aerospace RJ Avroliner series of aircraft, began airline service in April of 1993.

The FADEC employs modern control algorithms to achieve surge-free operation over the flight envelope while providing rapid transient performance and crisp handling qualities. The control interfaces with the aircraft via an ARINC 429 data link to control each engine automatically to the desired power setting with or without N1 synchronization. A simple hydromechanical backup control provides full dispatch capability in the event of a critical FADEC system failure. In addition, the FADEC includes advanced diagnostics for fault identification to the line replaceable unit (LRU) level without specialized test equipment.

This paper describes the architecture, primary features, and development process of the engine control system. Emphasis is placed on the design characteristics and technical challenges unique to the development of an inexpensive control system for the low thrust turbofan market.
Computational Modeling of Three-Dimensional Endwall Flow Through a Turbine Rotor Cascade With Strong Secondary Flows
Authors: Ho Y.-H. , Lakshminarayana B.
Published: February 2015

A steady, three-dimensional Navier-Stokes solver which utilizes a pressure-based technique for incompressible flows is used to simulate the three-dimensional flow field in a turbine cascade. A new feature of the numerical scheme is the implementation of a second-order plus fourth-order artificial dissipation formulation, which provides a precise control of the numerical dissipation. A low-Reynolds-number form of a two-equation turbulence model is used to account for the turbulence effects. Comparison between the numerical predictions and the experimental data indicates that the numerical model is able to capture most of the complex flow phenomena in the endwall region of a turbine cascade, except the high gradient region in the secondary vortex core. The effects of inlet turbulence intensity and turbulence length scale on secondary vortices, total pressure loss, and turbulence kinetic energy inside the passage are presented and interpreted. It is found that higher turbulence intensity energizes the vortical motions and tends to move the passage vortex away from the endwall. With a larger turbulence length scale the secondary flow inside the passage is reduced. However, the total pressure loss increases due to higher turbulence kinetic energy production.
Computation and Simulation of Wake-Generated Unsteady Pressure and Boundary Layers in Cascades: Part 1 — Description of the Approach and Validation
Authors: Fan S. , Lakshminarayana B.
Published: February 2015

The unsteady pressure and boundary layers on a turbomachinery blade row arising from periodic wakes due to upstream blade rows are investigated in this paper. A time accurate Euler solver has been developed using an explicit four-stage Runge-Kutta scheme. Two dimensional unsteady non-reflecting boundary conditions are used at the inlet and the outlet of the computational domain. The unsteady Euler solver captures the wake propagation and the resulting unsteady pressure field, which is then used as the input for a 2-D unsteady boundary layer procedure to predict the unsteady response of blade boundary layers. The boundary layer code includes an advanced k-ε model developed for unsteady turbulent boundary layers. The present computational procedure has been validated against analytic solutions and experimental measurements. The validation cases include unsteady inviscid flows in a flat plate cascade and a compressor exit guide vane (EGV) cascade, unsteady turbulent boundary layer on a flat plate subject to a traveling wave, unsteady transitional boundary layer due to wake passing and unsteady flow at the mid-span section of an axial compressor stator. The present numerical procedure is both efficient and accurate in predicting the unsteady flow physics resulting from wake/blade-row interaction, including wake induced unsteady transition of blade boundary layers.
Endwall and Unsteady Flow Phenomena in an Axial Turbine Stage
Authors: Gallus H. E. , Zeschky J. , Hah C.
Published: February 2015

Detailed experimental and numerical studies have been performed in a subsonic, axial-flow turbine stage to investigate the secondary flow field, the aerodynamic loss generation, and the spanwise mixing under a stage environment. The experimental study includes measurements of the static pressure distribution on the rotor blade surface and the rotor exit flow field using three-dimensional hot-wire and pneumatic probes. The rotor exit flow field was measured with an unsteady hot-wire probe which has high temporal and spatial resolution. Both steady and unsteady numerical analyses were performed with a three-dimensional Navier-Stokes code for the multiple blade rows. Special attention was focused on how well the steady multiple-blade-row calculation predicts the rotor exit flow field and how much the blade interaction affects the radial distribution of flow properties at the stage exit. Detailed comparisons between the measurement and the steady calculation indicate that the steady multiple-blade-row calculation predicts the overall time-averaged flow field very well. However, the steady calculation does not predict the secondary flow at the stage exit accurately. The current study indicates that the passage vortex near the hub of the rotor is transported toward the mid-span due to the blade interaction effects. And, the structure of the secondary flow field at the exit of the rotor is significantly modified by the unsteady effects. The time-averaged secondary flow field and the radial distribution of the flow properties, which are uses for the design of the following stage, can be predicted more accurately with the unsteady flow calculation.
Evaluation of an Euler Code With Non-Reflecting Boundary Conditions for the Analysis of 3D Flow in Radial Impellers
Authors: Fatsis Antonios , Van den Braembussche René
Published: February 2015

The aim of the present paper is the investigation of the influence of non-reflecting boundary conditions on the calculation of the three-dimensional inviscid steady rotational flow inside a high pressure ratio centrifugal impeller.

The space discretization of the Euler equations was done using the finite volume technique and the multistep Runge-Kutta scheme was used for the time integration. Nonreflecting boundary conditions using Fourier decomposition are applied at inflow and outflow boundaries.

Computations show that convergence to steady state is reached faster when non-reflecting boundary conditions are used compared to the reflecting boundary conditions. Studying the influence of the non-reflecting boundary conditions on the accuracy of the flowfield prediction, it is concluded that the location of the upstream boundary has a negligible influence on the solution even in the case where a shock is crossing the boundary.
A Rapid Aerodynamic Loading Procedure for Centrifugal Impeller Design
Authors: Howard J. H. G. , Osborne Colin , Japikse David
Published: February 2015

A crucial aspect of the design process for centrifugal impellers is the establishment of specific blade shapes. A rapid inviscid flow analysis procedure was developed for incorporation within a geometry manipulation code. Using a single streamtube model, a single-pass computation technique was generated. A two-zone model ensures that key features of the passage flow physics are incorporated.

Several examples of industrial design problems are employed to demonstrate the capabilities of the rapid loading method and its use in a geometry design procedure (used by some 20 industrial design groups worldwide). Comparisons with a quasi-three-dimensional method are included. The rapid loading method is most accurate when the meridional stream paths have similar shapes to those for the hub and shroud contours. The technique is useful within a geometry generation program since rapid aerodynamic screening of candidate configurations is allowed with sufficient accuracy to avoid the need for quasi-three-dimensional approaches. If required, the final design may be analyzed using three-dimensional viscous flow calculation methods.
A Numerical Investigation of Blade Lean Angle Effects on Flow in a Centrifugal Impeller
Authors: Howard J. H. G. , Ashrafizaadeh M.
Published: February 2015

An established 3-D viscous code (TASCflow3d) for compressible turbulent flow has been used to numerically investigate the effects of lean angle modifications to an existing high performance centrifugal compressor. Numerical results show that an appropriate compound lean can reduce the leakage flow, unload the blade tip and increase the total pressure ratio at a given flow rate while keeping its efficiency at a constant level. They also indicate that a simple lean can marginally improve the impeller efficiency while it lowers the impeller head versus flow curve.
A New Procedure for Simulating Unsteady Flows Through Turbomachinery Blade Passages
Authors: Chen Jen Ping , Celestina M. L. , Adamczyk J. J.
Published: February 2015

This paper outlines the development of two new unsteady wake-blade row aerodynamic interaction models and a rotor-stator unsteady aerodynamic interaction model. These flow models take, as input, solutions from Adamczyk’s average-passage flow model to provide an initial guess as to the transport of vortical, entropic and potential disturbances through a blade row. The response to these disturbances is calculated using an unsteady code known as MSUTC. This code solves the full Reynolds-averaged form of the Navier-Stokes equations using a modified Baldwin-Lomax turbulence model. The code can run with and without the use of wall functions. The solver is an implicit finite volume scheme with flux Jacobians evaluated by flux-vector splitting and residual fluxes by Roe’s flux-difference splitting. A series of simulations will show that the CPU time for all three models is nearly two orders of magnitude faster than current rotor-stator numerical simulation models. In addition, because of the numerical efficiency of the models, routine execution of numerical experiments for gaining insight into the flow processes controlling turbomachinery blade row performance become practical.
Experimental Investigation of the Flow Field of Deep Rotating Stall in a Centrifugal Compressor
Authors: Chen Y. N. , Seidel U. , Chen J. , Haupt U. , Rautenberg M.
Published: February 2015

The pressure field of deep rotating stall of a centrifugal compressor with two stall cells is analysed by means of the two-dimensional pressure pattern in the impeller determined by Chen et al. (1993). These authors transferred the pressure pattern measured on the shroud surface (i.e. in the absolute frame) to that related to the rotating blade channels. The transferred pressure pattern is thus a two-dimensional one.

The existence of the low and high pressure vortices according to the Rossby wave theory is confirmed by this experiment. The development stages of the two vortices, in combination with the Rossby wave that steers the rotating stall, can be evaluated very well.

The vortex low is developed from the front between the reverse flow (with high temperature and entropy) and the forward flow (with low temperature and entropy) due to baroclinic instability. Its center is situated within the channel of the splitter blade. This front is accompanied by a squall line of small-scaled eddies. This is the same phenomenon as can be observed on the meteorological polar front.

The vortex high is induced by the vortex low. Its embryo starts on the pressure surface. Its center is situated behind the inlet edge of the splitter blade.

It can be further verified that the stall cell is caused by the backflows of the induction fields of the two vortices (low and high).
Surface Heat Transfer From a Three-Pass Blade Cooling Passage Simulator
Authors: Chyu M. K. , Natarajan V.
Published: February 2015

Using an analog mass transfer system based on naphthalene sublimation, the present research focuses on investigating the local heat transfer characteristics from three-pass smooth and turbulated blade cooling passages. To simulate the actual passage geometry, the test model is incorporated with trapezoidal cross-sections including variable passage sizes. Measured local mass transfer results reveal strong evidence of velocity re-distribution over the trapezoidal flow area. Elevated mass transfer always exists in the vicinity of a sharp turn. However, in the present study, one of the most notable mass transfer increases is perceived in the third pass, downstream to the second turn, where the flow area is reduced severely. Overall, the combined effects of the three-pass and two sharp turns virtually doubles the mass transfer as compared to its straight counterpart with fully developed, turbulent flow. With a pitch-to-height ratio equal to 10 and 90-degree orientation, the rib turbulators produce approximately an additional 30% of overall mass transfer enhancement in comparison to the smooth case. Locally, rib-induced enhancement varies with different surfaces and passes. The greatest enhancement lies on the first pass, about 40%; the other two passes are comparable, less than 20%.
The Streamwise Development of Görtler Vortices in a Favorable Pressure Gradient
Authors: Finnis M. V. , Brown A.
Published: February 2015

Measurements are presented of the streamwise velocity variation within a laminar boundary layer on a concave surface of 4 m radius of curvature for which the free-stream velocity gradient factor ν/U02dU0/dx was approximately 1 × 10−6. The velocity variation was consistent with the presence of counter-rotating vortices resulting from the Görtler instability. The vortices exhibited exponential growth over the streamwise extent of the measurements to a disturbance amplitude of approximately 13% of the local freestream velocity. The vortex growth rates were found to be less than those for a zero velocity gradient factor indicating that a favorable pressure gradient stabilizes the flow with respect to the Görtler instability.

Boundary layer profiles at local upwash and downwash positions are compared with the linear theory for which the mean flow was modelled using the Pohlhausen approximation to the solution of the boundary layer equations. The agreement between the measured and predicted profiles indicates that the linear stability theory can provide a fair approximation to the small amplitude growth of the Görtler instability.
Component and Procedural Improvements for the T-62T-40-7 Gas Turbine in the LCAC Fleet
Authors: Ehrhardt Jerome P. , Spring Soren K.
Published: February 2015

The Sundstrand Power Systems T-62T-40-7 Gas Turbine Auxiliary Power Unit (APU) was adapted from an aircraft-borne APU to a marine application on the U.S. Navy’s Landing Craft Air Cushion (LCAC).

Although the LCAC APU experienced less operating time than its aircraft version, the environmental conditions that exist cause unusual wear and component failures. Component and procedural improvements have been developed to extend the reliability of the T-62T-40-7 on board the LCAC.
Heat Transfer Measurements in an Annular Cascade of Transonic Gas Turbine Blades Using the Transient Liquid Crystal Technique
Authors: Martinez-Botas R. F. , Lock G. D. , Jones T. V.
Published: February 2015

Heat transfer measurements have been made in the Oxford University Cold Heat Transfer Tunnel employing the transient liquid crystal technique. Complete contours of the heat transfer coefficient have been obtained on the aerofoil surfaces of a large annular cascade of high pressure nozzle guide vanes (mean blade diameter of 1.11 m and axial chord of 0.0664 m). The measurements are made at engine representative Mach and Reynolds numbers (exit Mach number 0.96 and Reynolds number 2.0 × 106).

A novel mechanism is used to isolate five preheated blades in the annulus before an unheated flow of air passes over the vanes, creating a step change in heat transfer. The surfaces of interest are coated with narrow-band thermochromic liquid crystals and the colour crystal change is recorded during the run with a miniature CCD video camera. The heat transfer coefficient is obtained by solving the one dimensional heat transfer equation for all the points of interest.

This paper will describe the experimental technique and present results of heat transfer and flow visualisation.
System Development Test Program for the WR-21 Intercooled Recuperated (ICR) Gas Turbine Engine System
Authors: Hawkins William J. , Mathieson Douglas , Bruce Chris J. , Socoloski Paul
Published: February 2015

Westinghouse Electric Corporation has teamed with Rolls-Royce to develop an affordable, commercially based Intercooled/Recuperated Gas Turbine Engine System (ICR) for the United States Navy. This engine system known as WR-21 will become the next prime mover on Navy new construction surface combatants. The system development test program for the WR-21 engine system will be carried out at two test sites in geographically different locations. These are the US Navy’s Test Site at the Carderock Division Naval Surface Warfare Center in Philadelphia, Pa. and the Royal Navy’s Admiralty Test House at the Test and Evaluation Establishment, Pyestock in the United Kingdom.

This paper will briefly describe the WR-21 engine system with a more detailed discussion of the system development test program itself. This will include descriptions of the system development testing to be performed and the test facilities and data acquisition systems at each test site location. Also discussed are the methods used to establish the required design commonality between each test site to establish test bed cross-calibration and provide test program flexibility and interchangeability of testing at each site.
CADDMAS: A Real-Time Parallel System for Dynamic Data Analysis
Authors: Tibbals Thomas F. , Bapty Theodore A. , Abbott Ben A.
Published: February 2015

Arnold Engineering Development Center (AEDC) has designed and built a high-speed data acquisition and processing system for real-time online dynamic data monitoring and analysis. The Computer Assisted Dynamic Data Monitoring and Analysis System (CADDMAS) provides 24 channels at high frequency and another 24 channels at low frequency for online real-time aeromechanical, vibration, and performance analysis of advanced turbo-engines and other systems. The system is primarily built around two different parallel processors and several PCs to demonstrate hardware independence and architecture scalability. These processors provide the computational power to display online and in real-time what can take from days to weeks using existing offline techniques. The CADDMAS provides online test direction and immediate hardcopy plots for critical parameters, all the while providing continuous health monitoring through parameter limit checking. Special in-house developed Front End Processors (FEP) sample the dynamic signals, perform anti-aliasing, signal transfer function correction, and bandlimit filtering to improve the accuracy of the time domain signal. A second in-house developed Numeric Processing Element (NPE) performs the FFT, threshold monitoring, and packetizes the data for rapid asynchronous access by the parallel network. Finally, the data are then formatted for display, hardcopy plotting, and cross-channel processing within the parallel network utilizing off-the-shelf hardware. The parallel network is a heterogeneous message-passing parallel pipeline configuration which permits easy scaling of the system. Advanced parallel processing scheduler/controller software has been adapted specifically for CADDMAS to allow quasi-dynamic instantiation of a variety of simultaneous data processing tasks concurrent with display and alarm monitoring functions without gapping the data. Although many applications of CADDMAS exist, this paper describes the features of CADDMAS, the development approach, and the application of CADDMAS for turbine engine aeromechanical testing.
Turbine Cascade Losses in the Presence of Progressive Particle Deposition
Authors: Kladas D. D. , Georgiou D. P.
Published: February 2015

Gas turbine efficiency deterioration due to particulate deposition is a well known phenomenon especially in industrial machines which operate on “heavy” fuels. The result is the need for frequent blade surface cleaning after certain hours of operation. The present study investigates the flowfield modification and the corresponding aerodynamic losses due to the progressive particle deposition on the aerofoil surfaces of a representative turbine stator cascade.

The flowfield is calculated with the Katsanis (TSONIC) code coupled to the STAN5 one for the boundary layer. The losses are calculated by momentum scanning - Stewart’s method. The problem is quasi-steady since the time scale in turbine blade flows is of the order of milliseconds, while in the deposition the data indicate slow growth rates with time scales of the ordr of hours. The particle deposition calculation is based on a turbulent diffusion model proposed by Yau and Young (18) and the particle characteristics refer to ash. The main parameters calculated are the new thickness distribution, the throat blockage, the aerodynamic losses and the transition point. The paper comments on the cleaning process based on these results.
Integration of Magnetic Bearings in the Design of Advanced Gas Turbine Engines
Authors: Storace A. F. , Sood D. , Lyons J. P. , Preston M. A.
Published: February 2015

Active magnetic bearings provide revolutionary advantages for gas turbine engine rotor support. These advantages include tremendously improved vibration and stability characteristics, reduced power loss, improved reliability, fault-tolerance, and greatly extended bearing service life. The marriage of these advantages with innovative structural network design and advanced materials utilization will permit major increases in thrust to weight performance and structural efficiency for future gas turbine engines.

However, obtaining the maximum payoff requires two key ingredients. The first key ingredient is the use of modern magnetic bearing technologies such as innovative digital control techniques, high-density power electronics, high-density magnetic actuators, fault-tolerant system architecture, and electronic (sensorless) position estimation. This paper describes these technologies and the test hardware currently in place for verifying the performance of advanced magnetic actuators, power electronics and digital controls. The second key ingredient is to go beyond the simple replacement of rolling element bearings with magnetic bearings by incorporating magnetic bearings as an integral part of the overall engine design. This is analogous to the proper approach to designing with composites, whereby the designer tailors the geometry and load carrying function of the structural system or component for the composite instead of simply substituting composites in a design originally intended for metal material. This paper describes methodologies for the design integration of magnetic bearings in gas turbine engines.
Design and Development of an Advanced Two-Stage Centrifugal Compressor
Authors: Palmer D. L. , Waterman W. F.
Published: February 2015

This paper describes the aero-mechanical design and development of a 3.3 kg/sec (7.3 lb/sec), 14:1 pressure ratio two-stage centrifugal compressor which is used in the T800-LHT-800 helicopter engine. The design employs highly nonradial, splitter bladed impellers with swept leading edges and compact vaned diffusers to achieve high performance in a small and robust configuration. The development effort quantified the effects of impeller diffusion and passive inducer shroud bleed on surge margin as well as the effects of impeller loading on tip clearance sensitivity and the impact of sand erosion and shroud roughness on performance. The developed compressor exceeded its performance objectives with a minimum of 23-percent surge margin without variable geometry. The compressor provides a high performance, rugged, low-cost configuration ideally suited for helicopter applications.
Forced Harmonic Response of Grouped Blade Systems: Part II — Application
Authors: Wagner L. F. , Griffin J. H.
Published: February 2015

The vibration of grouped blades on a flexible disk should, for purposes of economy and clarity of modal identification, be analyzed using procedures developed for cyclically-symmetric structures. In this paper, a numerical model, based on the theory of cyclically-symmetric structures, is applied to the vibration analysis, and in particular, the harmonic response, of a flexible disk supporting a number of groups, or packets, of turbine blades. Results are presented to show variations in the modal participation factors as a function of such parameters as disk flexibility, blade density, and the total number of assembled groups. It is also shown that many characteristics of the system spectra of natural frequencies are strongly dependent on the number of blade groups.
A Review of Droplet Dynamics and Vaporization Modeling for Engineering Calculations
Authors: Peng F. , Aggarwal S. K.
Published: February 2015

The present paper reviews the methodologies for representing the droplet motion and vaporization history in two-phase flow computations. The focus is on the use of droplet models that are realistic in terms of their efficient implementation in comprehensive spray simulations, representation of important physical processes, and applicability under a broad range of conditions. The methodologies available at present to simulate droplet motion in complex two-phase flows may be broadly classified into two categories. First one is based on the modified BBO equation. This approach is more comprehensive, but requires modifications and/or correlations at higher droplet Reynolds number. Second approach aims at developing correlations, using detailed numerical simulations or laboratory experiments, for the effects of flow nonuniformity and droplet relative acceleration on the instantaneous drag and lift coefficients. Recent advances made in the droplet vaporization models are also discussed. The advanced vaporization models include the effects of transient liquid heating, gas-phase convection, and variable thermophysical properties. All of these models are discussed, and recommendations are made for their inclusion in comprehensive two-phase computations.
Studies of Lean Blowout in a Step Swirl Combustor
Authors: Durbin Mark D. , Ballal Dilip R.
Published: February 2015

The design requirements of a modem gas turbine combustor are increasingly dictated by wide stability limits, short flame length, and uniform mixing. To achieve the best trade-off between the above three factors, flame characteristics (length, shape, mixedness), lean blowout (LBO), and optimum combustor configuration should be investigated over a wide range of inner and outer air velocities, inner and outer vane angles, and co- vs. counter-swirl arrangements. Such an investigation was performed in a step swirl combustor (SSC) designed to simulate the fuel-air mixing pattern in a gas turbine combustor dome fitted with an airblast atomizer.

It was found that an increase in the outer vane angle and a decrease in inner air velocity decreased the flame length. LBO was improved when outer flow swirl intensity was increased. An optimum hardware and velocity configuration for the SSC was found for inner swirl = 45°, outer swirl = 60°, co-swirl direction, and inner air velocity = outer air velocity = 16 m/s. This optimum SSC configuration yielded: (i) low values of LBO, (ii) short flame length, (iii) uniformly mixed stable flame, and (iv) little or no variation in these characteristics over the range of operation of SSC. Finally, the co- vs. counter-swirl arrangements and the operation of the optimized combustor configuration were discussed.
The Development of an Aviation Fuel Thermal Stability Test Unit
Authors: Daggett David L. , Veninger Al , Lewis Chris , Bullock Stuart , Kamin Rick
Published: February 2015

A test unit has been developed by Rolls-Royce for the U.S. Navy to evaluate fuel thermal deposition typically found in various aircraft engine components. Although the current Jet Fuel Thermal Oxidation Tester (JFTOT) provides a qualitative thermal stability evaluation, it may not be able to predict in service problems. Conditioning and testing of the fuel under realistic conditions is crucial if one is to accurately predict deposit formations. Engine fuel deposit evaluations and evidence from unpublished Rolls-Royce laboratory rig tests were used to help design a test unit that would address fuel stability problems in current or future aircraft. The Aviation Fuel Thermal Stability (AFTS) test unit embodies test modules that were selected with extensive fuel systems experience to enable the evaluation of thermal deposition in various fuel components using properly conditioned fuel. The test modules are controlled and results are recorded by a computer. This paper includes a review of the AFTS test unit design and preliminary test results thereof.
Effects of Rotating Inlet Distortion on Multisage Compressor Stability
Authors: Longley J. P. , Shin H.-W. , Plumley R. E. , Silkowski P. D. , Day I. J. , Greitzer E. M. , Tan C. S. , Wisler D. C.
Published: February 2015

In multi-spool engines, rotating stall in an upstream compressor will impose a rotating distortion on the downstream compressor, thereby affecting its stability margin. In this paper experiments are described in which this effect was simulated by a rotating screen upstream of several multistage low-speed compressors. The measurements are complemented by, and compared with, a theoretical model of multistage compressor response to speed and direction of rotation of an inlet distortion.

For co-rotating distortions (i.e., distortions rotating in the same direction as rotor rotation), experiments show that the compressors exhibited significant loss in stability margin and that they could be divided into two groups according to their response. The first group exhibited a single peak in stall margin degradation when the distortion speed corresponded to roughly 50% of rotor speed. The second group showed two peaks in stall margin degradation corresponding to distortion speeds of approximately 25–35% and 70–75% of rotor speed. These new results demonstrate that multistage compressors can have more than a single resonant response. Detailed measurements suggest that the two types of behavior are linked to differences between the stall inception processes observed for the two groups of compressors and that a direct connection thus exists between the observed forced response and the unsteady flow phenomena at stall onset.

For counter-rotational distortions, all the compressors tested showed minimal loss of stability margin. The results imply that counter-rotation of the fan and core compressor, or LP and HP compressors, could be a worthwhile design choice.

Calculations based on the two-dimensional theoretical model show excellent agreement for the compressors which had a single peak for stall margin degradation. We take this first-of-a-kind comparison as showing that the model, though simplified, captures the essential fluid dynamic features of the phenomena. Agreement is not good for compressors which had two peaks in the curve of stall margin shift versus distortion rotation speed. The discrepancy is attributed to the three-dimensional and short length scale nature of the stall inception process in these machines; this includes phenomena that have not yet been addressed in any model.
An Improved Aspirating Probe for Total-Temperature and Total-Pressure Measurements in Compressor Flows
Authors: Van Zante Dale E. , Suder Kenneth L. , Strazisar Anthony J. , Okiishi Theodore H.
Published: February 2015

The aspirating probe originally designed by Epstein and Ng at MIT was modified by replacing the two platinum coated tungsten hot wires normally used with platinum iridium alloy wires. The resulting improved unsteady total pressure and total temperature resolution of the modified probe is demonstrated. Flowfield measurements were made downstream of NASA Rotor 37 for a part speed operating condition to test the performance of the probe. Time resolved blade-to-blade total temperature and total pressure as calculated from the two platinum iridium hot wire voltages are shown. The flowfield measurements are compared with independent measurements of total pressure with high response transducers and total temperature calculated from laser anemometer measurements. Limitations of a more often used unsteady temperature data reduction method which involves only one aspirating probe hot wire voltage and a high-response pressure measurement are discussed.
LDA Investigation of the Flow Development Through Rotating U-DUCTS
Authors: Cheah S. C. , Iacovides H. , Jackson D. C. , Ji H. , Launder B. E.
Published: February 2015

This paper reports results from the use of laser Doppler anemometry (LDA) to measure the mean and the fluctuating flow field in a U-bend of strong curvature, Rc/D = 0.65, that is either stationary or rotating in orthogonal mode (the axis of rotation being parallel to the axis of curvature). The data acquisition system enables a stationary optical fibre probe to collect flow data from a rotating U-bend sweeping past it. Three cases have been examined all concerning a flow Reynolds number of 100,000; a stationary case, a case of positive rotation (the pressure side of the duct coincides with the outer side of the U-bend) at a Rotational number (ΩD/Um) of 0.2 and a case of negative rotation at a Rotational number of −0.2. Measurements have been obtained along the symmetry plane of the duct and also along a plane near top wall. The most important influence on the development of the mean and the turbulence flow fields is exerted by the streamwise pressure gradients that occur over the entry and exit regions of the U-bend. In the stationary case a 3-dimensional separation bubble is formed along the inner wall at the 90° location and it extends to about 2 diameters downstream of the bend causing the generation of high turbulence levels. Along the outer side, opposite the separation bubble, turbulence levels are suppressed due to streamwise flow acceleration. For the Rotation numbers examined, the Coriolis force also has a significant effect on the flow development. Positive rotation doubles the length of the separation bubble and generally suppresses turbulence levels. Negative rotation causes an extra separation bubble at the bend entry, raises turbulence levels within and downstream of the bend, increases velocity fluctuations in the cross-duct direction within the bend and generates strong secondary motion after the bend exit. It is hoped that the detailed information produced in this study will assist in the development of turbulence models suitable for the numerical computation of flow and heat transfer inside blade-cooling passages.
Plasticity Considerations in Probabilistic Ceramic-to-Metal Joint Design
Authors: O’Neil D. A. , Selverian J. H. , Kim K. S.
Published: February 2015

A new probabilistic failure criterion was developed for the design of high-temperature ceramic-to-metal joints. The essential feature of the theory is the inclusion of the energy dissipated during plastic deformation of the adjacent braze layer in the joint. A large number of bi-material interface fracture simulations were performed for different crack positions and orientations near the bimaterial interface to determine the effect on stresses in the ceramic near the interface. The effective stress values were then ported to a probabilistic failure analysis code, which permitted simple inclusion of the new failure criterion. Brazed joints were made and failure tested in torsion to verify the failure criterion. Results show that the new failure criterion more closely approximates the failure of the ceramic-to-metal joints over the entire range of ultimate loads, an is a significant improvement in the failures criteria previously used for this type of joint design. Aspects of the failure criterion, material systems, residual stresses, mechanical behavior, and strength predictions will be presented.
Optimization of Regenerated Gas Turbines
Authors: Beck Douglas Stephen
Published: February 2015

An algorithm for the optimization of regenerated gas turbines is given. For sets of inputs that are typical for automotive applications, the optimum cycle pressure ratio and a set of optimized regenerator parameters that maximize thermal efficiency are given. A second algorithm, an algorithm for sizing regenerators based on outputs of the optimization algorithm, is given. With this sizing algorithm, unique regenerator designs can be determined for many applications based on the presented optimization data. Results of example sizings are given. The data indicate that one core (instead of two cores) should be used to maximize thermal efficiency. The data also indicate that thermal efficiencies of over 50% should be achievable for automotive applications if ceramic turbines are used.
The Aerodynamic Mixing Effect of Discrete Cooling Jets With Mainstream Flow on a Highly Loaded Turbine Blade
Authors: Wilfert Günter , Fottner Leonhard
Published: February 2015

For the application of film cooling to turbine blades experimental investigations were performed on the mixing processes in the near hole region with a row of holes on the suction side of a turbine cascade. Data were obtained using pneumatic probes, pressure tappings, and a three-dimensional subminiature hot wire probe, as well as surface flow visualization techniques.

It was found that at low blowing rates a cooling jet behaves very much like a normal obstacle and the mixing mainly takes place in the boundary layer. With increasing blowing rates the jet penetrates deeper into the mainstream. The variation of the turbulence level at the inlet of the turbine cascade and the Reynolds number showed a strong influence on the mixing behavior. The kidney-shaped vortex and as an important achievement the individual horse-shoe vortex of each single jet were detected and their exact positions were obtained. This way it was found that the position of the horse-shoe vortex is strongly dependent on the blowing rate and this influences the aerodynamic mixing mechanisms. A two-dimensional code for the calculation of boundary layer flows called GRAFTUS was used, however the comparison with the measurements showed only limited agreement for cascade flow with blowing due to the strong three-dimensional flow pattern.
Design of a Low Aspect Ratio Transonic Compressor Stage Using CFD Techniques
Authors: Sanger Nelson L.
Published: February 2015

A transonic compressor stage has been designed for the Naval Postgraduate School Turbopropulsion Laboratory. The design relied heavily on CFD techniques while minimizing conventional empirical design methods. The low aspect ratio (1.2) rotor has been designed for a specific head ratio of .25 and a tip relative inlet Mach number of 1.3. Overall stage pressure ratio is 1.56. The rotor was designed using an Euler code augmented by a distributed body force model to account for viscous effects. This provided a relatively quick-running design tool, and was used for both rotor and stator calculations. The initial stator sections were sized using a compressible, cascade panel code. In addition to being used as a case study for teaching purposes, the compressor stage will be used as a research stage. Detailed measurements, including non-intrusive LDV, will be compared with the design computations, and with the results of other CFD codes, as a means of assessing and improving the computational codes as design tools.
In Flight Heat Transfer Measurements on an Aero-Engine Nacelle
Authors: Bown Nicholas W. , Cain Terrence M. , Jones Terence V. , Shipley Philip P. , Barry Brian
Published: February 2015

This paper describes heat transfer measurements on the external surface of a Natural Laminar Flow (NLF) nacelle. The measurement technique employed temperature sensitive liquid crystals and platinum resistance thermometers (PRTs) to measure the surface temperature over an electrically heated pad. This gave an immediate visual indication of the transition location. The heat transfer distribution along the length of the pad has been determined and is compared with a simple theoretical model. Results are presented for the cruise condition of Mach 0.56 at an altitude of 6400m.
Aerodynamic Performance of a Transonic Turbine Guide Vane With Trailing Edge Coolant Ejection: Part II — Numerical Approach
Authors: Michelassi V. , Martelli F. , Amecke J.
Published: February 2015

The effect of coolant ejection from the trailing edge of a turbine blade is studied by means of numerical simulation codes. Two solvers based on explicit and implicit algorithms with and without multigridding are used in connection with H, HOH and HC grid types. A simple procedure to model the coolant jet at the trailing edge of the blade is presented which does not require excessive grid refinement. The features of the different approaches are discussed in terms of isentropic Mach number distribution on the blade, base pressures and wake profiles at various exit Mach numbers and coolant flow rates which match the experimental conditions. The coolant ejection effect on the shock pattern is discussed. The trailing edge ejection is found to have a marginal effect on the wake shape and on the measured and computed base pressures.
Efficient Computations of Navier-Stokes Equations With Nonequilibrium Chemistry
Authors: Groenner J. , Hilgenstock M. , von Lavante E.
Published: February 2015

An efficient computer program was developed for the computation of supersonic combustion units. Several test cases show the capabilities of implementations of different flux splitting schemes in calculating chemical reactions. It can be shown that for more efficient supersonic combustion enhancement of the mixing of fuel and air is necessary.
A Rapid Method for Predicting Global and Local Performance of Cascades With Special Emphasis on the Calculation of the Transition Region
Authors: Kus Umit , Chauvin Jacques
Published: February 2015

The principle of a rapid method of cascade performance prediction is presented. It is based on a strong coupling between a potential flow approach and an integral boundary layer method, and takes into account laminar, transitional and turbulent regimes as well as the separated regions. Special emphasis is given to the transitional region. In addition to a pre-existing bubble separation model and following an extensive literature survey, two transition onset and three transition length criteria have been incorporated in the method as well as a Görtler type approach. Comparison of the criteria on typical cascades show that they are relatively coherent between themselves and lead to a satisfactory prediction of the global performance for a large number of different compressor and turbine cascades among which the NACA and DCA cascades discussed in this paper. Local boundary layer performance is compared with Deutsch and Zierke’s data, where, unhappily, transition is chiefly through separation bubbles. The comparison shows the validity of Roberts ‘bubble transition model. Applying the Görtler curvature approach gives similar results to the other criteria. The global performance prediction for this case is also satisfactory. Additionally, parametric studies are carried out on the effect of external turbulence level.