bluuebunny/crossref_metadata_2025_split huggingface dataset:
The Periodical Interaction of the Tip Clearance Flow in the Blade Rows of Axial Compressors
Authors: Mailach Ronald , Sauer Helmut , Vogeler Konrad 
Published: July 2014

Current models on the tip clearance flow in turbomachines only describe the time-averaged behaviour of the flow structures. However, the real tip clearance flow is periodically fluctuating in time. This fact has to be considered for the design of turbomachine bladings especially with regard to blade vibrations and tip clearance noise.
Detailed experimental investigations on the time-resolved behaviour of the flow in the rotor blade tip region were carried out in a four-stage low-speed research compressor. A strong time-periodic interaction of the blade tip vortices of adjacent blades can be shown for relatively large tip clearance of the rotor blades for operating points near the stability limit of the compressor. The resulting flow pattern, which frequency is not related to the rotor frequency, moves along the blade row. It can be described as a multicell configuration with strongly fluctuating cell number and size. The structure and propagation of the flow instability can be summarized in a model of the periodic fluctuating tip clearance flow (Mailach et al., 2000).
Additional experiments were carried out in a straight cascade to improve the understanding of this flow phenomenon. It can be shown by means of time-resolved measurements that the same disturbance exists for comparable inlet flow conditions in the blade tip region of the cascade. Flow visualizations show that the blade tip vortex is strongly fluctuating and moves sometimes ahead of the leading edge of the adjacent blade. The result of this is a short-lengthscale flow pattern, which is propagating along the blade row. These experiments confirm the model of the time-periodic tip clearance flow proposed for compressors. A Strouhal-number for the estimation of the frequency of the flow fluctuation will be presented, which includes both design and aerodynamic parameters.
Analysis of Two Automotive Turbocharger Compressor Radial Vaneless Diffusers for Use in Improving a Performance Prediction Technique
Authors: Meese H. M. , Swain E. 
Published: July 2014

An experimental and analytical investigation of two automotive centrifugal compressors with a radial vaneless diffusers has been conducted. Static pressure measurements together with 1-D flow assumptions have been used to determine the fluid properties throughout the diffuser and this has provided quantitative estimates of the location and value of the losses occurring in this region.
A correlation of diffuser loss for the two specific geometry’s has been developed. This correlation has been used to extend a 1-D prediction technique to include the vaneless diffuser, enabling the operating range and performance of the diffuser to be predicted from the geometry.
Analysis of Diffusers With Different Number of Vanes in a Centrifugal Compressor Stage
Authors: He N. , Tourlidakis A. 
Published: July 2014

In this paper, a computational analysis of a high-speed centrifugal compressor stage for turbocharger applications is presented. Emphasis is focused on the effect of different number of diffuser vanes, and for this reason four different designs of the vaned diffuser are analysed. The first three of the diffusers consist of 11, 22 and 33 vanes, respectively, with their leading edge at a radius of 1.075 times the radius of the impeller tip. The fourth one consists of 22 vanes with its leading edge at 1.15 times the radius of the impeller tip. All the above vane designs are of double circular arc shape.
A steady CFD analysis is carried out using the Reynolds-Averaged Navier-Stokes solver TASCflow at design and off-design operating conditions. An averaging approach is used at the interface between the impeller and the diffuser. A detailed comparison between the predicted and the available experimental data is performed in terms of pressure rise and efficiency characteristics, and very good agreement is accomplished. In addition, detailed flow distributions are compared and critically analysed.
One of the most important conclusions is that while maintaining the overall throat area and the location of the leading and trailing edges of the diffuser, as the number of diffuser vanes increases, the pressure recovery coefficient at the semi-vaneless space at surge condition was found to reduce, the wake pattern becomes more pronounced and the velocity distribution at vaneless and semi-vaneless space becomes more distorted when passing the same mass flow rate and therefore the diffuser has a narrower flow range. On the other hand, it was found that the diffuser outlet to throat area ratio is not the dominant factor to influence the flow range when the number of vanes changes.
Improved Compressor Exit Diffuser for an Industrial Gas Turbine
Authors: Orth U. , Ebbing H. , Krain H. , Weber A. , Hoffmann B. 
Published: July 2014

Cycle studies carried out for the medium pressure ratio gas turbine THM 1304 of 10 MW power output manufactured by MAN Turbomaschinen AG GHH BORSIG predicted that the overall efficiency of the multi stage compressor, composed of a 10 stage axial and a single stage centrifugal compressor, would improve by 0.8% if the efficiency of the back stage centrifugal unit could be raised by 4%. It was expected that this would result in a noticeable improvement of the thermal gas turbine efficiency.
The paper describes the aerodynamical design process used for the stage optimization applying today’s advanced design tools for blade generation and three dimensional aerodynamic calculation methods. Additionally it describes the manufacturing procedure for the resulting three-dimensional blades and the experimental verification of the design approach.
Redesign of a 12-Stage Axial-Flow Compressor Using Multistage CFD
Authors: Wellborn Steven R. , Delaney Robert A. 
Published: July 2014

A multistage turbomachinery computational fluid dynamics (CFD) system was developed and used to redesign an industrial core compressor for improved performance. Overall performance improvement goals of a 1% increase in the design point adiabatic efficiency and a 5% increase in high-speed surge margin were established for this first attempt at using this new system. Outlined in this paper are the details of how advanced multistage tools were incorporated into the existing design system. Data from the baseline and redesigned configurations are presented that illustrate the capabilities of the system. The redesign activities are reviewed, with emphasis placed on the ability of the tools to adequately predict axial stage matching.
A simple and economical means to increase compressor performance was found using the new CFD system. Test data confirmed that the redesign of the target compressor was successful. Engine results demonstrated a 2% increase in design speed compressor efficiency and a 4% drop in the high speed operating line with the redesigned blading installed, thereby exceeding the original efficiency improvement goal. This increase in compressor efficiency produced a 4% increase in engine shaft horsepower with a 16% decrease in the heat rate near the design turbine inlet temperature.
Efficiency Upgrading of an Ambient Pressure MCFC Plant Through the Introduction of an Indirect Heated Gas Turbine
Authors: Lunghi Piero , Ubertini Stefano 
Published: July 2014

The efficient end environmentally friendly production of electricity is undoubtedly one of the 21st century priorities. Since renewable sources will be able to guarantee only a share of the future demand, the present research activity must focus on innovative energy devices and improved conversion systems and cycles. Great expectations are reserved to fuel cell systems. The direct conversion from chemical to electrical energy eliminates environmental problems connected with combustion and bypass the stringent efficiency limit due to Carnot’s principle.
Still in infancy high temperature Fuel cells present the further advantage of feasible cycle integration with steam or gas turbines.
In this paper, a Molten Carbonate Fuel Cell plant is simulated in a cycle for power generation. The introduction of an external combustion Gas Turbine is evaluated with the aim of efficiency and net power output increase. The results show that the proposed cycle can be conveniently used as a source of power generation. As compared to internal combustion Gas Turbine hybrid cycles found in literature the plant is characterized by fuel cell greater simplicity, due to the absence of pressurization, and gas turbine increased complexity, due to the presence of the heat exchange system.
Degradation Effects on Combined Cycle Power Plant Performance: Part 1 — Gas Turbine Cycle Component Degradation Effects
Authors: Zwebek A. , Pilidis P. 
Published: July 2014

This paper describes the effects of degradation of the main gas path components of the gas turbine topping cycle on the Combined Cycle Gas Turbine (CCGT) plant performance. Firstly the component degradation effects on the gas turbine performance as an independent unit are examined. It is then shown how this degradation is reflected on a steam turbine plant of the CCGT and on the complete Combined Cycle plant.
TURBOMATCH, the gas turbine performance code of Cranfield University was used to predict the effects of degraded gas path components of the gas turbine have on its performance as a whole plant. To simulate the steam (Bottoming) cycle, another Fortran code was developed. Both codes were used together to form a complete software system that can predict the CCGT plant design point, off-design, and deteriorated (due to component degradation) performances.
The results show that the overall output is very sensitive to many types of degradation, specially in the turbine of the gas turbine. Also shown is the effect on gas turbine exhaust conditions and how this affects the steam cycle.
Performance Simulation of Turboprop Engine for Basic Trainer
Authors: Kong Changduk , Ki Jayoung 
Published: July 2014

A performance simulation program for the turboprop engine (PT6A-62), which is the power plant of the first Korean indigenous basic trainer KT-1, was developed for more precise performance prediction, development of an EHMS (Engine Health Monitoring System) and the flight simulator.
Characteristics of components including compressors, turbines, the combustor and the constant speed propeller were required for the steady state and transient performance analysis with on and off design point analysis. In most cases, these were substituted for what scaled from similar engine components’ characteristics with the scaling law.
The developed program was evaluated with the performance data provided by the engine manufacturer and with analysis results of GASTURB program, which is well known for the performance simulation of gas turbines. Performance parameters such as mass flow rate, compressor pressure ratio, fuel flow rate, specific fuel consumption ratio and turbine inlet temperature were discussed to evaluate validity of the developed program at various cases. The first case was the sea level static standard condition and other cases were considered with various altitudes, flight velocities and part loads with the range between idle and 105% rotational speed of the gas generator.
In the transient analysis, the Continuity of Mass Flow Method was utilized under the condition that mass stored between components is ignored and the flow compatibility is satisfied, and the Modified Euler Method was used for integration of the surplus torque.
The transient performance analysis for various fuel schedules was performed. When the fuel step increase was considered, the overshoot of the compressor turbine inlet temperature occurred. However, in case of the fuel ramp increase longer than the fuel step increase, the overshoot of the compressor turbine inlet temperature was effectively reduced.
Modelling the Air-Cooled Gas Turbine: Part 2 — Coolant Flows and Losses
Authors: Young J. B. , Wilcock R. C. 
Published: July 2014

This paper is Part II of a study concerned with developing a formal framework for modelling air-cooled gas turbine cycles. It deals with the detailed specification of coolant flowrates and losses. For accurate performance assessment it is necessary to divide the turbine expansion into individual stages with stator and rotor rows being treated separately. Particular care is needed when deriving the equations for the rotor and it is shown how all required flow variables can be estimated from minimal data if design values are unavailable.
Specification of the cooling flowrates is based on a modified Holland and Thake procedure which can be formalised in terms of averaged parameters. Thermal barrier coatings can be included if present. The importance of allowing for fluctuations in combustor outlet temperature is stressed and procedures for dealing with end-wall and disk cooling are suggested.
There is confusion in the literature concerning cooling losses and it is shown how these may be defined and sub-divided in a consistent way. The importance of representing losses in terms of irreversible entropy creation rather than total pressure loss is stressed. A set of models for the components of the cooling loss are presented and sample calculations are used to illustrate the division and magnitude of the loss.
Pressure Surface Separations in Low Pressure Turbines: Part 2 of 2 — Interactions With the Secondary Flow
Authors: Brear Michael J. , Hodson Howard P. , Gonzalez Paloma , Harvey Neil W. 
Published: July 2014

This paper describes a study of the interaction between the pressure surface separation and the secondary flow on low pressure turbine blades. It is found that this interaction can significantly affect the strength of the secondary flow and the loss that it creates. Experimental and numerical techniques are used to study the secondary flow in a family of four low pressure turbine blades in linear cascade. These blades are typical of current designs, share the same suction surface and pitch, but have differing pressure surfaces.
A mechanism for the interaction between the pressure surface separation and the secondary flow is proposed and is used to explain the variations in the secondary flows of the four blades. This mechanism is based on simple dynamical secondary flow concepts and is similar to the aft-loading argument commonly used in modern turbine design.
TERTS: A Generic Real-Time Gas Turbine Simulation Environment
Authors: Visser Wilfried P. J. , Broomhead Michael J. , van der Vorst Jasper 
Published: July 2014

Real-time simulation of gas turbine engine performance is used in a variety of aerospace applications. For simulation of propulsion system performance in flight-simulators, fidelity requirements become increasingly stringent. Significant improvements in simulation fidelity can be obtained when using thermodynamic models instead of the customary (piece-wise) linear real-time models. However, real-time thermodynamic models require sophisticated methods to efficiently solve the model equations on a real-time basis with sufficient speed.
NLR has developed the ‘Turbine Engine Real-Time Simulator’ (TERTS) generic real-time engine simulation environment for full thermodynamic simulation of various gas turbine engine configurations. At NLR’s National Simulator Facility (NSF), research is performed on pilot-in-the-loop simulation of complex aircraft and helicopter configurations such as thrust-vectoring and Integrated Flight Propulsion Control (IFPC) concepts. For this application, high-fidelity real-time gas turbine models are required. TERTS has an efficient method for solving the engine model equations real-time. The system is implemented in Matlab-Simulink®, which offers advantages in terms of control system modeling flexibility. With TERTS, detailed thermodynamic real-time engine models can easily be implemented in NSF providing an excellent means to analyze a variety of engine effects on pilot-in-the-loop aircraft performance.
In this paper the TERTS modeling environment will be described including the numerical solutions used to comply with the real-time requirements. A TERTS model of a military afterburning turbofan will be presented including simulation results.
Oxidation Behavior of Prospective Silicon Nitride Materials for Advanced Microturbine Applications
Authors: Schenk Bjoern , Strangman Tom , Opila Elizabeth J. , Robinson R. Craig , Fox Dennis S. , Klemm Hagen , Taut Christine , More Karren , Torterelli Peter 
Published: July 2014

Various laboratory tests have shown that high-pressure water vapor environments combined with elevated temperatures and intermediate gas velocities (current facilities limited to about 50 m/s) can cause grain boundary degradation and material recession in silica formers. Recent tests include burner rig testing conducted by NASA [1], Honeywell Engines & Systems [2], Siemens Power Generation [3], CRIEPI in Japan [4, 5], “Keiser rig” testing at Oak Ridge National Laboratory (ORNL) [6], and engine testing in the Allison 501K industrial gas turbine [7]. This paper presents a summary of oxidation test data of candidate silicon nitride materials for advanced microturbine applications. These data are of interest to microturbine component designers in order to determine the limits of safe unprotected component operation with respect to the given turbine environment, as well as to understand the behavior of ceramic microturbine components once local spallation of the protective environmental barrier coating has occurred.
This paper intends to give materials and engine development engineers some guidance with respect to the different test facility capabilities and the prevailing oxidation/recession mechanisms to better understand/interprete the oxidation test results when developing new ceramic material compositions and environmental barrier coating systems.
Numerical and Experimental Investigations on Endwall Contouring in a Four-Stage Turbine
Authors: Bonn Dieter E. , Funke Harald , Sürken Norbert , Kreitmeier Franz 
Published: July 2014

Secondary flows and leakage flows create complex vortex structures in the 3-D flow field of a turbine stage. Aerodynamic losses are the consequence. Reducing the aerodynamic losses by endwall contouring is subject of an actual investigation of the flow field in a 4-stage test turbine with repeating stages.
Numerical 4-stage simulations are performed for a reference case of a turbine without endwall modifications and two different geometric configurations with endwall contouring. The numerical results for the reference case are compared to corresponding experimental investigations. Both, the experiment and the CFD focus on the stage exit flow field of the second, the third and the fourth stage of the actual four stage turbine. The 3-D flow field is calculated by application of a steady 3-D Navier-Stokes code.
The numerical results of an arc-like endwall contouring at the casing are presented a) with a maximum deviation from the reference contour in the axial gap within the stages (“arc contour”) and b) with a maximum deviation in the axial gap between the stages (“off-set arc contour”). The results show a significant influence of the bumps on the blade’s profile pressure distribution near the radial gap, the leakage flow and the radial pressure field. A detailed secondary flow analysis shows the influence of the different endwall contours on the leakage vortex development.
Finally, the aerodynamic efficiencies of the geometric configurations are compared. It is predicted that the off-set arc contour has a remarkable positive influence on the machine’s performance.
Midspan Flow-Field Measurements for Two Transonic Linear Turbine Cascades at Off-Design Conditions
Authors: Jouini D. B. M. , Sjolander S. A. , Moustapha S. H. 
Published: July 2014

The paper presents detailed mid-span experimental results from two transonic linear turbine cascades. The blades for the two cascades were designed for the same service and differ mainly in their leading-edge geometries. One of the goals of the study was investigate the influence of the leading-edge metal angle on the sensitivity of the blade to positive off-design incidence. Measurements were made for incidence values of −10.0°, 0.0°, +4.5°, +10.0°, and +14.5° relative to design incidence. The exit Mach numbers varied roughly from 0.5 to 1.2 and the Reynolds numbers from about 4×105 to 106. The measurements include the midspan losses, blade loadings and base pressures. In addition, the axial-velocity-density ratio (AVDR) was extracted for each operating point The AVDR was found to vary from about 0.98 at −10.0° of incidence to about 1.27 at +14.5°. Thus, the data set also provides some evidence of the influence AVDR on axial turbine blade performance.
Detailed experimental results for turbine blade performance at off-design incidence are very scarce in the open literature, particularly for transonic conditions. Among other things, the present results are intended to expand the database available in the open literature. To this end, the key aerodynamic results are presented in tabular form, along with the detailed geometry of the cascades. The results could be used in the development of new or improved correlations for use in the early stages of design. They could also be used to evaluate the ability of current CFD codes to capture reliably the variation in losses and other aerodynamic quantities with variations in blade incidence.
Development of H Series Gas Turbine
Authors: Maekawa A. , Uematsu K. , Ito E. , Ohya T. , Hirokawa K. , Fukuizumi Y. 
Published: July 2014

Mitsubishi Heavy Industries, Ltd. developed the M501H gas turbine to improve the efficiency and power of G series using closed steam cooling of stage 1 and 2 blades and vanes following the firm’s experience in the construction of the D, F and G series. Development included elementary tests for pre-verification and trial operation using a combined plant verification plant to reach 220MW in May 1999 and to verify the stability of the steam cooling rotor, cooling properties of blades and vanes, and the efficiency of major components. The few remaining items can be easily taken care of.
The next operation is planned to complete verification.