Aerodynamic Implications of Reduced Vane Count

2015 ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Mats Annerfeldt
Keyword(s):  
Gas Turbines ◽  
Fossil Fuels ◽  
Positive Impact ◽  
Guide Vane ◽  
Suction Side ◽  
Combined Cycle ◽  
Significant Deficit ◽  
Nozzle Guide ◽  
The Cost ◽  
Cooling Air

Given the shortage of fossil fuels and the growing greenhouse effect, one strive in modern gas turbines is to make maximum usage of the burnt fuel. By reducing the number of vanes or blades and thereby increasing the loading per vane (or blade) it is possible to spend less cooling air, which will have a positive impact on the combined cycle efficiency. It also reduces the number of components and usage of metal and thereby also the cost of the engine. These savings should be achieved without any efficiency deficit in aerodynamic efficiency. Based on the fact, aerodynamic investigations were performed to see the aerodynamic implications of reduced vane number in a transonic annular sector cascade. The number of new nozzle guide vane was reduced with 24% compared to a previous design with higher vane count. The investigated vanes were two typical high pressure gas turbine vanes. Results regarding the loading indicated an expected increase with the reduced vane case. The minimum static pressure at the suction side is lower and at an earlier location for the reduced vane case and therefore, an extension of the trailing edge deceleration zone is observed for the reduced vane case. Results regarding losses indicate that even though the losses produced per vane significantly increases for the reduced vane case, a comparison of mass averaged losses between the reduced vane case and previous vane case show similar spanwise loss distributions. Assessing results leads to a conclusion that the reduction of the number of vanes in the first stage seems to be a useful method to save cooling flow as well as material costs without any significant deficit in overall efficiency.

Suction and Pressure Side Film Cooling Influence on Vane Aero Performance in a Transonic Annular Cascade

2013 ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Pressure Side ◽  
Aerodynamic Loss ◽  
Cooling Flow ◽  
Blowing Ratio ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Cooling Air

An experimental study on a film cooled nozzle guide vane has been conducted in a transonic annular sector to observe the influence of suction and pressure side film cooling on aerodynamic performance. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at an exit reference Mach number of 0.89. The aerodynamic results using a five hole miniature probe are quantified and compared with the baseline case which is uncooled. Results lead to a conclusion that the aerodynamic loss is influenced substantially with the change of the cooling flow rate regardless the positions of the cooling rows. The aerodynamic loss is very sensitive to the blowing ratio and a value of blowing ratio higher than one leads to a considerable higher loss penalty. The suction side film cooling has larger influence on the aerodynamic loss compared to the pressure side film cooling. Pitch-averaged exit flow angles around midspan remain unaffected at moderate blowing ratio. The secondary loss decreases (greater decrease in the tip region compared to the hub region) with inserting cooling air for all cases compared to the uncooled case.

Characterisation of Static Strip Seal Flow

2007 ◽  
Author(s):  
Arash Farahani ◽  
Peter Childs
Keyword(s):  
Gas Flow ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Main Stream ◽  
Cfd Modelling ◽  
Height Range ◽  
And Performance ◽  
Nozzle Guide ◽  
The Cost ◽  
Cooling Air

Strip seals are used in gas turbine engines between two static elements or between components which do not move relative to each other, such as Nozzle Guide Vanes (NGVs). The key role of a strip seal between NGV segments is sealing between the flow through the main stream annulus and the internal air system, a further purpose is to limit the inter-segmental movements. In general the shape of the strip seal is a rectangular strip that fits into two slots in adjacent components. The minimum clearance required for static strip seals must be found by accounting for thermal expansion, misalignment, and application, to allow correct fitment of the strip seals. Any increase in leakage raises the cost due to an increase in the cooling air use, which is linked to specific fuel consumption, and it can also alter gas flow paths and performance. The narrow path within the seal assembly, especially the height has the most significant affect on leakage. The height range of the narrow path studied in this paper is 0.01–0.06 mm. The behaviour of the flow passing through the narrow path has been studied using CFD modelling and measurements in a bespoke rig. The CFD and experimental results show that normalized leakage flow increases with pressure ratio before reaching a maximum. The main aim of this paper is to provide new experimental data to verify the CFD modelling for static strip seals. The typical flow characteristics validated by CFD modelling and experiments can be used to predict the flow behaviour for future static strip seal designs.

scholarly journals Flow Inhomogeneities in a Realistic Aeronautical Gas-Turbine Combustor: Formation, Evolution and Indirect Noise

2018 ◽  
Author(s):  
Andrea Giusti ◽  
Luca Magri ◽  
Marco Zedda
Keyword(s):  
Gas Turbines ◽  
Transfer Functions ◽  
Guide Vane ◽  
Flame Structure ◽  
Combustion Noise ◽  
Moment Closure ◽  
Combustion Model ◽  
Conditional Moment ◽  
Eddy Simulation ◽  
Nozzle Guide

Indirect noise generated by the acceleration of combustion inhomogeneities is an important aspect in the design of aeroengines because of its impact on the overall noise emitted by an aircraft and the possible contribution to combustion instabilities. In this study, a realistic rich-quench-lean combustor is numerically investigated, with the objective of quantitatively analyzing the formation and evolution of flow inhomogeneities and determine the level of indirect combustion noise in the nozzle guide vane (NGV). Both entropy and compositional noise are calculated in this work. A high-fidelity numerical simulation of the combustion chamber, based on the Large-Eddy Simulation (LES) approach with the Conditional Moment Closure (CMC) combustion model, is performed. The contributions of the different air streams to the formation of flow inhomogeneities are pinned down and separated with seven dedicated passive scalars. LES-CMC results are then used to determine the acoustic sources to feed an NGV aeroacoustic model, which outputs the noise generated by entropy and compositional inhomogeneities. Results show that non-negligible fluctuations of temperature and composition reach the combustor’s exit. Combustion inhomogeneities originate both from finite-rate chemistry effects and incomplete mixing. In particular, the role of mixing with dilution and liner air flows on the level of combustion inhomogeneities at the combustor’s exit is highlighted. The species that most contribute to indirect noise are identified and the transfer functions of a realistic NGV are computed. The noise level indicates that indirect noise generated by temperature fluctuations is larger that the indirect noise generated by compositional inhomogeneities, although the latter is not negligible and is expected to become louder in supersonic nozzles. It is also shown that relatively small fluctuations of the local flame structure can lead to significant variations of the nozzle transfer function, whose gain increases with the Mach number. This highlights the necessity of an on-line solution of the local flame structure, which is performed in this paper by CMC, for an accurate prediction of the level of compositional noise. This study opens new possibilities for the identification, separation and calculation of the sources of indirect combustion noise in realistic aeronautical gas turbines.

A Conjugate 3-D Flow and Heat Transfer Analysis of a Thermal Barrier Cooled Turbine Guide Vane

1998 ◽  
Author(s):  
Dieter E. Bohn ◽  
Volker J. Becker
Keyword(s):  
Heat Transfer ◽  
Gas Turbines ◽  
Guide Vane ◽  
Suction Side ◽  
Heat Transfer Analysis ◽  
Thermal Barrier ◽  
Trailing Edge ◽  
Flow And Heat Transfer ◽  
Turbine Guide Vane ◽  
Basic Configuration

This paper presents the numerical investigations of the flow and heat transfer of two configurations of a transonic turbine guide vane. The basic configuration is a vane with convection cooling. The second configuration is additionally coated with a thermal barrier consisting of ZrO2. The results are obtained with a conjugate heat transfer and flow computer code that has been developed at the Institute of Steam and Gas Turbines. Measurement data is available for the basic configuration and the computational results are compared to the experimental results. The results show very good agreement between calculated and measured vane surface temperatures. The trailing edge turns out to be subjected to high thermal loads as it is too thin to be cooled effectively. Secondary flow phenomena like the passage vortex and the corner vortex and their impact on the temperature distribution are discussed. The ZrO2 coating is calculated for a thickness of 300μm. The substrate material temperatures are lowered by about 20 K–29 K in the stagnation point area and by about 27 K–43 K in the shock area on the suction side. At the trailing edge, the coating on the suction side and on the pressure side hardly influences the metal temperature.

Solar Upgrading of Fuels for Generation of Electricity

2001 ◽  
Vol 123 (2) ◽  
pp. 160-163 ◽  
Author(s):  
Rainer Tamme ◽  
Reiner Buck ◽  
Michael Epstein ◽  
Uriyel Fisher ◽  
Chemi Sugarmen
Keyword(s):  
Gas Turbines ◽  
Large Scale ◽  
Fossil Fuels ◽  
High Efficiency ◽  
Combined Cycle ◽  
Test Facility ◽  
Small Scale ◽  
Start Up ◽  
Efficiency Conversion ◽  
Solar Field

This paper presents a novel process comprising solar upgrading of hydrocarbons by steam reforming in solar specific receiver-reactors and utilizing the upgraded, hydrogen-rich fuel in high efficiency conversion systems, such as gas turbines or fuel cells. In comparison to conventionally heated processes about 30% of fuel can be saved with respect to the same specific output. Such processes can be used in small scale as a stand-alone system for off-grid markets as well as in large scale to be operated in connection with conventional combined-cycle plants. The complete reforming process will be demonstrated in the SOLASYS project, supported by the European Commission in the JOULE/THERMIE framework. The project has been started in June 1998. The SOLASYS plant is designed for 300 kWel output, it consists of the solar field, the solar reformer and a gas turbine, adjusted to operate with the reformed gas. The SOLASYS plant will be operated at the experimental solar test facility of the Weizmann Institute of Science in Israel. Start-up of the pilot plant is scheduled in April 2001. The midterm goal is to replace fossil fuels by renewable or non-conventional feedstock in order to increase the share of renewable energy and to establish processes with only minor or no CO2 emission. Examples might be upgrading of bio-gas from municipal solid waste as well as upgrading of weak gas resources.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

Advanced Gas Turbine and High Performance Gas-Steam Combined Cycle Plant With Blade Cooling Air Controlling

2006 ◽  
Author(s):  
Tadashi Tsuji
Keyword(s):  
Power Generation ◽  
Heat Exchanger ◽  
Gas Turbine ◽  
Thermal Efficiency ◽  
Gas Turbines ◽  
Combined Cycle ◽  
Hot Water ◽  
Blade Cooling ◽  
Combined Cycle Plant ◽  
Cooling Air

Air cooling blades are usually applied to gas turbines as a basic specification. This blade cooling air is almost 20% of compressor suction air and it means that a great deal of compression load is not converted effectively to turbine power generation. This paper proposes the CCM (Cascade Cooling Module) system of turbine blade air line and the consequent improvement of power generation, which is achieved by the reduction of cooling air consumption with effective use of recovered heat. With this technology, current gas turbines (TIT: turbine inlet temperature: 1350°C) can be up-rated to have a relative high efficiency increase. The increase ratio has a potential to be equivalent to that of 1500°C Class GT/CC against 1350°C Class. The CCM system is designed to enable the reduction of blade cooling air consumption by the low air temperature of 15°C instead of the usual 200–400°C. It causes the turbine operating air to increase at the constant suction air condition, which results in the enhancement of power and thermal efficiency. The CCM is installed in the cooling air line and is composed of three stage coolers: steam generator/fuel preheater stage, heat exchanger stage for hot water supplying and cooler stage with chilled water. The coolant (chilled water) for downstream cooler is produced by an absorption refrigerator operated by the hot water of the upstream heat exchanger. The proposed CCM system requires the modification of cooling air flow network in the gas turbine but produces the direct effect on performance enhancement. When the CCM system is applied to a 700MW Class CC (Combined Cycle) plant (GT TIT: 135°C Class), it is expected that there will be a 40–80MW increase in power and +2–5% relative increase in thermal efficiency.

An Advanced Single-Stage Turbine Facility for Investigating Non-Axisymmetric Contoured Endwalls in the Presence of Purge Flow

2019 ◽  
Author(s):  
Robin R. Jones ◽  
Oliver J. Pountney ◽  
Bjorn L. Cleton ◽  
Liam E. Wood ◽  
B. Deneys J. Schreiner ◽  
...  
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Optical Measurement ◽  
Guide Vane ◽  
Single Stage ◽  
Test Facility ◽  
Pressure Distributions ◽  
Mainstream Flow ◽  
Purge Flow ◽  
Nozzle Guide

Abstract In modern gas turbines, endwall contouring (EWC) is employed to modify the static pressure field downstream of the vanes and minimise the growth of secondary flow structures developed in the blade passage. Purge flow (or egress) from the upstream rim-seal interferes with the mainstream flow, adding to the loss generated in the rotor. Despite this, EWC is typically designed without consideration of mainstream-egress interactions. The performance gains offered by EWC can be reduced, or in the limit eliminated, when purge air is considered. In addition, EWC can result in a reduction in sealing effectiveness across the rim seal. Consequently, industry is pursuing a combined design approach that encompasses the rim-seal, seal-clearance profile and EWC on the rotor endwall. This paper presents the design of, and preliminary results from a new single-stage axial turbine facility developed to investigate the fundamental fluid dynamics of egress-mainstream flow interactions. To the authors’ knowledge this is the only test facility in the world capable of investigating the interaction effects between cavity flows, rim seals and EWC. The design of optical measurement capabilities for future studies, employing volumetric velocimetry and planar laser induced fluorescence are also presented. The fluid-dynamically scaled rig operates at benign pressures and temperatures suited to these techniques and is modular. The facility enables expedient interchange of EWC (integrated into the rotor bling), blade-fillet and rim-seals geometries. The measurements presented in this paper include: gas concentration effectiveness and swirl measurements on the stator wall and in the wheel-space core; pressure distributions around the nozzle guide vanes at three different spanwise locations; pitchwise static pressure distributions downstream of the nozzle guide vane at four axial locations on the stator platform.

Nozzle Guide Vane Static Strip Seals

2006 ◽  
Author(s):  
Arash Farahani ◽  
Peter Childs
Keyword(s):  
Gas Turbines ◽  
Guide Vane ◽  
Main Stream ◽  
Guide Vanes ◽  
Advantages And Disadvantages ◽  
Air System ◽  
Improved Performance ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Assembly Constraints

Sealing of components where there is no relative motion between the elements concerned remains a significant challenge in many gas turbine engine applications. Loss of sealing and cooling air from the internal air system through seals impacts on specific fuel consumption and can lead to undesirable flow interactions with resultant cost implications. For gas turbines, various strip seal types have been developed for use between Nozzle Guide Vanes in order to limit the flow of gas between the main stream annulus and the internal air system. Many different types of design have been proposed for overcoming strip seal problems such as misalignment of the grooves due to manufacturing and assembly constraints. In this paper various methods, with a particular focus on patents, for minimising the amount of leakage caused by such problems for strip seals between nozzle guide vanes are reviewed. By considering the advantages and disadvantages of each technique it is concluded that although apparently new strip seal designs for NGVs have improved performance, none of them can be considered to be ideal. This paper reviews the techniques and makes recommendations for future designs.

Gas Turbine Part Load Performance Testing: Comparison of Test Methodologies

2008 ◽  
Author(s):  
Thomas P. Schmitt ◽  
Herve Clement
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Guide Vane ◽  
Performance Testing ◽  
Load Level ◽  
Combined Cycle ◽  
Inlet Guide Vane ◽  
Performance Guarantee ◽  
Part Load ◽  
Advantages And Disadvantages

Current trends in usage patterns of gas turbines in combined cycle applications indicate a substantial proportion of part load operation. Commensurate with the change in operating profile, there has been an increase in the propensity for part load performance guarantees. When a project is structured such that gas turbines are procured as equipment-only from the manufacturer, there is occasionally a gas turbine part load performance guarantee that coincides with the net plant combined cycle part load performance guarantee. There are several methods by which to accomplish part load gas turbine performance testing. One of the more common methods is to operate the gas turbine at the specified load value and construct correction curves at constant load. Another common method is to operate the gas turbine at a specified load percentage and construct correction curves at constant percent load. A third method is to operate the gas turbine at a selected load level that corresponds to a predetermined compressor inlet guide vane (IGV) angle. The IGV angle for this third method is the IGV angle that is needed to achieve the guaranteed load at the guaranteed boundary conditions. The third method requires correction curves constructed at constant IGV, just like base load correction curves. Each method of test and correction embodies a particular set of advantages and disadvantages. The results of an exploration into the advantages and disadvantages of the various performance testing and correction methods for part load performance testing of gas turbines are presented. Particular attention is given to estimates of the relative uncertainty for each method.

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