Effect of Sweep on Performance of an Axial Compressor With Casing Grooves

2016 ◽  
Author(s):  
Shraman Narayan Goswami ◽  
M. Govardhan
Keyword(s):  
Gas Turbine ◽  
Interaction Effect ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Public Domain ◽  
Operating Conditions ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Stall Margin ◽  
Flow Variables

The need of increased stall margin is very high for aero gas turbine engines, as they operate under varied operating conditions. A number of different options are being used to increase the stall margin of gas turbine engines. Circumferential casing groove, in the compressor section of a gas turbine engine, is one of such methods. Incorporation of the grooves on the shroud increases the stall margin of the compressor, but this generally gives rise to loss of performance, such as efficiency and pressure ratio. By employing 3D blading techniques for rotor blades as well as stator vanes, performance of a compressor can be increased. 3D blading helps in reducing secondary flow losses and hence increased performance. Sweep and lean are examples of 3D blading, which is very common in any modern gas turbine compressor. A number of literatures are available in public domain, giving detailed understanding of effect of circumferential casing grooves and 3D blade features, but the interaction effect of sweep and casing grooves are not well published in public domain literature. In this work, an effort is made to understand, numerically, the interaction effect of sweep with circumferential grooves, using Computational Fluid Dynamics (CFD). Any numerical tool needs thorough validation before the results of numerical analyses can be used for analyzing the underlying physics. NASA Rotor37 is used to validate current CFD methodology. Mesh sensitivity is carried out to get mesh independence solution. Different turbulence models are used to get the best turbulence model for the problem in hand. 1D averaged performance data as well as hub to shroud variation of various flow parameters are compared to have full confidence on the CFD methodology. A baseline axial compressor rotor, without sweep and lean is generated, as the first step of this study. This rotor is created by using hub and tip profiles of NASA Rotor37. The profiles are stacked along a radial line through their center of gravities, which has resulted in rotor geometry without any sweep and lean. Modifications are done to the tip profile of the baseline rotor, in terms of stagger angle, to get comparable performance w.r.t. NASA Rotor37. Casing of the NASA Roto37 is used as the redesigned compressor casing. Circumferential casing grooves, with five grooves between leading edge to trailing edge, are created as per industry standards. Meshing and modeling are done according to the best practices developed while validating CFD methodology. It is to be noted that the casing grooves and the main flow domain are meshed with one to one mesh connectivity, in order to avoid any numerical losses due to interface interpolations. This is considered very critical in this work, as the vortices from the tip is expected to have a strong interaction with grooves. This interaction is expected to create high gradients of flow variables in this region. Valuable flow information might be lost, if flow variables are interpolated in this region. Baseline rotor is analyzed with and without casing grooves from choke to stall at 100% corrected speed. As expected, introduction of casing grooves has resulted in increased stall margin. A number of rotor geometries are created with different amount of sweeps. In the current study, blades are swept in the direction of chord, in order to avoid introduction of any sweep induced lean. The span location, where sweep starts, is also changed to understand the localized and global effect of this blade design features. Results obtained from numerical simulations of these geometries are presented in this paper. The performance and flow features are compared with respect to baseline rotor, with and without circumferential grooves, in an attempt to understand the underlying flow physics.

The Effects of Ambient Conditions on Gas Turbine Emissions—Generalized Correction Factors

1978 ◽  
Vol 100 (4) ◽  
pp. 640-646 ◽  
Author(s):  
P. Donovan ◽  
T. Cackette
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Turbine Engines ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Correction Factors ◽  
Oxides Of Nitrogen ◽  
Ambient Conditions ◽  
Nitrogen Emission ◽  
Wide Range

A set of factors which reduces the variability due to ambient conditions of the hydrocarbon, carbon monoxide, and oxides of nitrogen emission indices has been developed. These factors can be used to correct an emission index to reference day ambient conditions. The correction factors, which vary with engine rated pressure ratio for NOx and idle pressure ratio for HC and CO, can be applied to a wide range of current technology gas turbine engines. The factors are a function of only the combustor inlet temperature and ambient humidity.

Thermo-Economic Optimization in the Design of Small-Scale and Residential Cogeneration Systems

2009 ◽  
Author(s):  
C. P. Lea˜o ◽  
S. F. C. F. Teixeira ◽  
A. M. Silva ◽  
M. L. Nunes ◽  
L. A. S. B. Martins
Keyword(s):  
Gas Turbine ◽  
Urban Areas ◽  
Pressure Ratio ◽  
Optimization Method ◽  
Turbine Engines ◽  
Small Scale ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Economic Optimization ◽  
Micro Turbine

In recent years, gas-turbine engines have undergone major improvements both in efficiency and cost reductions. Several inexpensive models are available in the range of 30 to 250 kWe, with electrical efficiencies already approaching 30%, due to the use of a basic air-compressor associated to an internal air pre-heater. Gas-turbine engines offer significant advantages over Diesel or IC engines, particularly when Natural Gas (NG) is used as fuel. With the current market trends toward Distributed Generation (DG) and the increased substitution of boilers by NG-fuelled cogeneration installations for CO2 emissions reduction, small-scale gas turbine units can be the ideal solution for energy systems located in urban areas. A numerical optimization method was applied to a small-scale unit delivering 100 kW of power and 0.86 kg/s of water, heated from 318 to 353K. In this academic study, the unit is based on a micro gas-turbine and includes an internal pre-heater, typical of these low pressure-ratio turbines, and an external heat recovery system. The problem was formulated as a non-linear optimisation model with the minimisation of costs subject to the physical and thermodynamic constraints. Despite difficulties in obtaining data for some of the components cost-equations, the preliminary results indicate that the optimal compressor pressure ratio is about half of the usual values found in large installations, but higher than those of the currently available micro-turbine models, while the turbine inlet temperature remains virtually unchanged.

Improved criteria for stall-free preliminary design of axial compressor of aero gas turbine engines

2018 ◽  
Vol 233 (9) ◽  
pp. 3286-3297 ◽  
Author(s):  
MR Aligoodarz ◽  
A Mehrpanahi ◽  
M Moshtaghzadeh ◽  
A Hashiehbaf
Keyword(s):  
Gas Turbine ◽  
Large Scale ◽  
Axial Compressor ◽  
Axial Flow ◽  
Design Criterion ◽  
Flow Coefficient ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Parameter Ranges ◽  
Axial Flow Compressors

A worldwide effort has been devoted to developing highly efficient and reliable gas turbine engines. There exist many prominent factors in the development of these engines. One of the most important features of the optimal design of axial flow compressors is satisfying the allowable range for various parameters such as flow coefficient, stage loading, the degree of reaction, De-Haller number, etc. But, there are some applicable cases that the mentioned criteria are exceeded. One of the most famous parameters is De-Haller number, which according to literature data should not be kept less than 0.72 in any stage of the axial compressor. A deep insight into the current small- or large-scale axial flow compressors shows that a discrepancy will occur among design criterion for De-Haller number and experimental measurements in which the De-Haller number is less than the design limit but no stall or surge is observed. In this paper, an improved formulation is derived based on one-dimensional modeling for predicting the stall-free design parameter ranges especially stage loading, flow coefficient, etc. for various combinations. It was found that the current criterion is much more accurate than the De-Haller criterion for design purposes.

Design Study of an Advanced Concept Simple Gas Turbine for Possible Use in Low Emission Automobiles

1972 ◽  
Author(s):  
H. C. Eatock ◽  
M. D. Stoten
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Simple Cycle ◽  
Turbine Engines ◽  
Nox Emissions ◽  
Preliminary Design ◽  
Gas Turbine Engines ◽  
High Pressure Ratio

United Aircraft Corporation studied the potential costs of various possible gas turbine engines which might be used to reduce automobile exhaust emissions. As part of that study, United Aircraft of Canada undertook the preliminary design and performance analysis of high-pressure-ratio nonregenerated (simple cycle) gas turbine engines. For the first time, high levels of single-stage component efficiency are available extending from a pressure ratio less than 4 up to 10 or 12 to 1. As a result, the study showed that the simple-cycle engine may provide satisfactory running costs with significantly lower manufacturing costs and NOx emissions than a regenerated engine. In this paper some features of the preliminary design of both single-shaft and a free power turbine version of this engine are examined. The major component technology assumptions, in particular the high pressure ratio centrifugal compressor, employed for performance extrapolation are explained and compared with current technology. The potential low NOx emissions of the simple-cycle gas turbine compared to regenerative or recuperative gas turbines is discussed. Finally, some of the problems which might be encountered in using this totally different power plant for the conventional automobile are identified.

Advanced Multi-Cup Fuel Injector Technology for Environmentally Responsible Aviation Gas Turbine Engines

2015 ◽  
Author(s):  
Erlendur Steinthorsson ◽  
Adel Mansour ◽  
Brian Hollon ◽  
Michael Teter ◽  
Clarence Chang
Keyword(s):  
Gas Turbine ◽  
Direct Injection ◽  
Pressure Ratio ◽  
Test Facility ◽  
Turbine Engines ◽  
Practical Implementation ◽  
Gas Turbine Engines ◽  
Fuel Injector ◽  
Ambient Conditions ◽  
Environmentally Responsible

Participating in NASA’s Environmentally Responsible Aviation (ERA) Project, Parker Hannifin built and tested multipoint Lean Direct Injection (LDI) fuel injectors designed for NASA’s N+2 55:1 Overall Pressure-Ratio (OPR) gas turbine engine cycles. The injectors are based on Parker’s earlier three-zone injector (3ZI) which was conceived to enable practical implementation of multipoint LDI schemes in conventional aviation gas turbine engines. The new injectors offer significant aerodynamic design flexibility, excellent thermal performance, and scalability to various engine sizes. The injectors built for this project contain 15 injection points and incorporate staging to enable operation at low power conditions. Ignition and flame stability were demonstrated at ambient conditions with ignition air pressure drop as low as 0.3% and fuel-to-air ratio (FAR) as low as 0.011. Lean Blowout (LBO) occurred at FAR as low as 0.005 with air at 460 K and atmospheric pressure. A high pressure combustion testing campaign was conducted in the CE-5 test facility at NASA Glenn Research Center at pressures up to 250 psi and combustor exit temperatures up to 2,033 K (3,200 °F). The tests demonstrated estimated LTO cycle emissions that are about 30% of CAEP/6 for a reference 60,000 lbf thrust, 54.8-OPR engine. This paper presents some details of the injector design along with results from ignition, LBO and emissions testing.

scholarly journals EFFECT OF HYSTERESIS IN AXIAL COMPRESSORS OF GAS-TURBINE ENGINES

Aviation ◽  
2012 ◽  
Vol 16 (4) ◽  
pp. 97-102 ◽  
Author(s):  
Mykola Kulyk ◽  
Ivan Lastivka ◽  
Yuri Tereshchenko
Keyword(s):  
Gas Turbine ◽  
Axial Compressor ◽  
Separated Flow ◽  
Aerodynamic Performance ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Bladed Disks ◽  
Axial Compressors

The phenomenon of separated flow hysteresis in the process of the streamlining the axial compressor of gas-turbine engines is considered. Generalised results of research on the occurrence of hysteresis in the aerodynamic performance of compressor grids and its influence on the performance of the bladed disks of compressors that operate in real conditions of periodic circular non-uniformity are demonstrated.

Fuel Flexibility in LM2500 and LM6000 Dry Low Emission Engines

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
John Blouch ◽  
Hejie Li ◽  
Mark Mueller ◽  
Richard Hook
Keyword(s):  
Gas Turbine ◽  
Operating Conditions ◽  
Inert Gases ◽  
Effect Of Temperature ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Kinetics Analysis ◽  
Fuel Flexibility ◽  
Low Emission ◽  
Commercial Operation

The LM2500 and LM6000 dry-low-emissions aeroderivative gas turbine engines have been in commercial service for 15 years and have accumulated nearly 10 × 106 hours of commercial operation. The majority of these engines utilize pipeline quality natural gas predominantly comprised of methane. There is; however, increasing interest in nonstandard fuels that contain varying levels of higher hydrocarbon species and/or inert gases. This paper reports on the demonstrated operability of LM2500 and LM6000 DLE engines with nonstandard fuels. In particular, rig tests at engine conditions were performed to demonstrate the robustness of the dual-annular counter-rotating swirlers premixer design, relative to flameholding with fuels containing high ethane, propane, and N2 concentrations. These experiments, which test the ability of the hardware to shed a flame introduced into the premixing region, have been used to expand the quoting limits for LM2500 and LM6000 gas turbine engines to elevated C2+ levels. In addition, chemical kinetics analysis was performed to understand the effect of temperature, pressure, and fuel compositions on flameholding. Test data for different fuels and operating conditions were successfully correlated with Damkohler number.

Fuel Flexibility in LM2500 and LM6000 Dry Low Emission Engines

2011 ◽  
Author(s):  
John Blouch ◽  
Hejie Li ◽  
Mark Mueller ◽  
Richard Hook
Keyword(s):  
Gas Turbine ◽  
Operating Conditions ◽  
Inert Gases ◽  
Effect Of Temperature ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Kinetics Analysis ◽  
Fuel Flexibility ◽  
Low Emission ◽  
Commercial Operation

The LM2500 and LM6000 dry-low-emissions (DLE) aeroderivative gas turbine engines have been in commercial service for 15 years and have accumulated nearly 10 million hours of commercial operation. The majority of these engines utilize pipeline quality natural gas predominantly comprised of methane. There is, however, increasing interest in nonstandard fuels that contain varying levels of higher hydrocarbon species and/or inert gases. This paper reports on the demonstrated operability of LM2500 and LM6000 DLE engines with nonstandard fuels. In particular, rig tests at engine conditions were performed to demonstrate the robustness of the dual-annular counter-rotating swirlers (DACRS) premixer design, relative to flameholding with fuels containing high ethane, propane, and N2 concentrations. These experiments, which test the ability of the hardware to shed a flame introduced into the premixing region, have been used to expand the quoting limits for LM2500 and LM6000 gas turbine engines to elevated C2+ levels. In addition, chemical kinetics analysis was performed to understand the effect of temperature, pressure, and fuel compositions on flameholding. Test data for different fuels and operating conditions were successfully correlated with Damkohler number.

scholarly journals ВПЛИВ ГУСТОТИ РЕШІТКИ НА АЕРОДИНАМІЧНУ НАВАНТАЖЕНІСТЬ РОБОЧОГО КОЛЕСА ОСЬОВОГО ВЕНТИЛЯТОРА

2020 ◽  
pp. 38-43
Author(s):  
Екатерина Викторовна Дорошенко ◽  
Михаил Владимирович Хижняк ◽  
Юрий Матвеевич Терещенко
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Turbine Engines ◽  
Average Radius ◽  
Gas Turbine Engines ◽  
Axial Fan ◽  
Aerodynamic Loading ◽  
Wide Range ◽  
Axial Fans ◽  
Bypass Ratio

The main requirements that apply to axial fans and axial compressors of aircraft gas turbine engines include minimum dimensions and weight; high aerodynamic load; high coefficient of performance; wide range of steady work; high reliability. For gas turbine engines, the requirements of minimum weight and dimensions are especially important, since the engines must provide flights at high velocities and altitudes. This study aims to assess the effect of the solidity of the impeller fan on the average radius on the aerodynamic loading of the impeller of an axial fan for an engine with a high bypass ratio. The object of the study is the impeller of the fan. The solidity of the impeller fan on the average radius varied in the range from 1.8 to 0.82, the number of blades of the impeller fan varied from 33 to 15, respectively. The studies in this work were carried out by the method of numerical experiment. The flow in the axial fans was simulated by solving the system of Navier-Stokes equations, which were closed by the SST turbulent viscosity model. Based on the analysis of the results of the study, an assessment is made of the influence of the solidity of the impeller fan at an average radius on the aerodynamic loading of the impeller of an axial fan for an engine with a high bypass ratio. The research results showed that with a decrease in the solidity of the impeller fan at an average radius of 1.8 to 0.82 in operating modes with an axial inlet velocity of 80 to 120 m / s, the impeller fan pressure ratio decreases by 0.11 ... 3.2 %. The maximum decrease in the fan pressure ratio increase for the fan impeller with the parameters studied is 3.2 %, with a decrease in the number of fan blades from 33 to 15, while the total weight of the blades decreases by 54.55 %. The decrease in the solidity on the average radius of the impeller of the studied fan leads to a decrease in the relative sizes of the low-velocity zones at the sleeve and on the periphery and to a decrease in the level of flow unevenness. A further reduction in the level of flow non-uniformity behind the fan is possible when using the boundary layer control in the fan - this is the task of subsequent studies.

scholarly journals МЕТОД РАСЧЕТА ТЕРМОГАЗОДИНАМИЧЕСКИХ ПАРАМЕТРОВ ТУРБОВАЛЬНОГО ГТД НА ОСНОВЕ ПОВЕНЦОВОГО ОПИСАНИЯ ЛОПАТОЧНЫХ МАШИН. ЧАСТЬ 1. ОСНОВНЫЕ УРАВНЕНИЯ

2018 ◽  
pp. 48-58
Author(s):  
Людмила Георгиевна Бойко ◽  
Олег Владимирович Кислов ◽  
Наталия Владимировна Пижанкова
Keyword(s):  
Gas Turbine ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Turbine Engines ◽  
Geometrical Parameters ◽  
Gas Turbine Engines ◽  
Balance Equations ◽  
Wide Range ◽  
In Series ◽  
Multistage Compressor

Gas turbine engines processes mathematic simulations are widely used in different steps of its living cycle. All engine simulations may be divided into different difficulty levels: higher simulation level allows doing a more pre­cise description of physical processes in main units of gas turbine engines and their elements. It gives the oppor­tunity for getting better arrangement of calculation results and experimental data, reduce the quality of factors, which are traditionally used in determine engine operational characteristics with 1-level models.The purpose of the article is to describe the thermogasdynamic parameters and maintenance perfomances cal­culation method, which based on second level mathematic simulation. Its main feature is blade-to-blade turbomachines description (multistage compressor and multistage cooling gas turbine), which allows to take into account blade and flow path geometrical parameters. Their changing during the gas turbine engine design and de­velopment processes influence its performances: thrust, fuel consumption, efficiency as functions of values of flow rate, rotational speed, engine entrance conditions and so on. All these dependences could be defined by using proposed calculation method.In distinction from methods which are noted, this method allows to concede compressor or turbine incidence angles, drag values, pressure ratio, surge margin in design and off-design  engine regimes. The opportunity to take into account by-passing and air bleeding from compressor blade channels and their engine parameters influence is very important also.The article includes calculation method main points, block-scheme, equations system, which gives the opportunity of alignment the engine units and their elements in wide range of state working regimes. Set of equations consists of flow rate balance equations through the stages of multistage compressor and turbine, combustion chamber and connected channels. Also system includes power balance equations, by-passing, air bleeding from compressor stages channels, its admission into the cooling turbine stages and ac­counts their thermodynamic parameters. Compressors and turbines maps parameters are calculated with main turbomachinery theory lows and semi-empirical dependences.This article is the first in series of articles, which considers this problem

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