Effects of Purge Jet Momentum on Sealing Effectiveness

2016 ◽  
Vol 139 (3) ◽  
Author(s):  
Kenneth Clark ◽  
Michael Barringer ◽  
Karen Thole ◽  
Carey Clum ◽  
Paul Hiester ◽  
...  
Keyword(s):  
Flow Rate ◽  
Momentum Flux ◽  
Reynolds Numbers ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Swirl Velocity ◽  
Purge Flow ◽  
Secondary Air ◽  
Flow Tracer ◽  
Mach And Reynolds Numbers

Driven by the need for higher cycle efficiencies, overall pressure ratios for gas turbine engines continue to be pushed higher thereby resulting in increasing gas temperatures. Secondary air, bled from the compressor, is used to cool turbine components and seal the cavities between stages from the hot main gas path. This paper compares a range of purge flows and two different purge hole configurations for introducing the purge flow into the rim cavities. In addition, the mate face gap leakage between vanes is investigated. For this particular study, stationary vanes at engine-relevant Mach and Reynolds numbers were used with a static rim seal and rim cavity to remove rotational effects and isolate gas path effects. Sealing effectiveness measurements, deduced from the use of CO2 as a flow tracer, indicate that the effectiveness levels on the stator and rotor side of the cavity depend on the mass and momentum flux ratios of the purge jets relative to the swirl velocity. For a given purge flow rate, fewer purge holes resulted in better sealing than the case with a larger number of holes.

Effects of Purge Jet Momentum on Sealing Effectiveness

2016 ◽  
Author(s):  
Kenneth Clark ◽  
Michael Barringer ◽  
Karen Thole ◽  
Carey Clum ◽  
Paul Hiester ◽  
...  
Keyword(s):  
Flow Rate ◽  
Momentum Flux ◽  
Reynolds Numbers ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Swirl Velocity ◽  
Purge Flow ◽  
Secondary Air ◽  
Flow Tracer ◽  
Mach And Reynolds Numbers

Driven by the need for higher cycle efficiencies, overall pressure ratios for gas turbine engines continue to be pushed higher thereby resulting in increasing gas temperatures. Secondary air, bled from the compressor, is used to cool turbine components and seal the cavities between stages from the hot main gas path. This paper compares a range of purge flows and two different purge hole configurations for introducing the purge flow into the rim cavities. In addition, the mate face gap leakage between vanes is investigated. For this particular study, stationary vanes at engine relevant Mach and Reynolds numbers were used with a static rim seal and rim cavity to remove rotational effects and isolate gas path effects. Sealing effectiveness measurements, deduced from the use of CO2 as a flow tracer, indicate that the effectiveness levels on the stator and rotor side of the cavity depend on the mass and momentum flux ratios of the purge jets relative to the swirl velocity. For a given purge flow rate, fewer purge holes resulted in better sealing than the case with a larger number of holes.

Prediction of Compressible Flow Pressure Losses in 30–150 Deg Sharp-Cornered Bends

1995 ◽  
Vol 117 (4) ◽  
pp. 589-592 ◽  
Author(s):  
Nia Haidar
Keyword(s):  
Compressible Flow ◽  
Turbine Engines ◽  
Air Cooling ◽  
Gas Turbine Engines ◽  
Pressure Losses ◽  
Flow Pressure ◽  
Secondary Air ◽  
Limiting Condition ◽  
Experimental Findings ◽  
Wide Speed Range

This paper considers the measurement and prediction of the additional total pressure losses of subsonic steady air flow in sharp-cornered bends, similar to those present in the secondary air cooling systems of gas turbine engines. The bends examined ranged between 30 to 150 in 30 deg increments and were circular in cross section. Experimental results covering a wide speed range up to choking are presented for five different bend geometries. An analytical flow model provided results in fairly good agreement with the measurements obtained and equally compared favourably with the experimental findings of other researchers at low Mach numbers. The highest attainable upstream Mach number (MU) of the average upstream flow was 0.57 for the 30 deg bend. The maximum possible values of MU represent a limiting condition dictated by downstream choking of the flow. The compressible flow coefficients, caused by the presence of the bends, can be expected to be between 10 to 20 percent higher than those for incompressible flow.

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

Using a Tracer Gas to Quantify Sealing Effectiveness for Engine Realistic Rim Seals

2016 ◽  
Author(s):  
Kenneth Clark ◽  
Michael Barringer ◽  
Karen Thole ◽  
Carey Clum ◽  
Paul Hiester ◽  
...  
Keyword(s):  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Tracer Gas ◽  
Engine Efficiency ◽  
Hot Gas ◽  
Area Of Interest ◽  
Sensitivity Studies ◽  
Secondary Air ◽  
Effective Use ◽  
Validation Experiments

As overall pressure ratios increase in gas turbine engines, both the main gas path and cooling temperatures increase leading to component durability concerns. At the same time effective use of the secondary air for both cooling and sealing becomes increasingly important in terms of engine efficiency. To fully optimize these competing requirements, experiments at engine-relevant conditions are required to validate new designs and computational tools. A test turbine has been commissioned in the Steady Thermal Aero Research Turbine (START) lab. The test turbine was designed to be a 1.5 stage turbine operating under continuous flow simulating engine-relevant conditions including Reynolds and Mach numbers with hardware true to engine scale. The first phase of research conducted using the test turbine, which was configured for a half-stage (vane only), was to study hot gas ingestion through turbine rim seals. This paper presents a series of facility benchmarks as well as validation experiments conducted to study ingestion using a tracer gas to quantify the performance of rim seals and purge flows. Sensitivity studies included concentration levels and sampling flow rates in flow regimes that ranged from stagnant to compressible depending upon the area of interest. The sensitivity studies included a range of purge and leakage flow conditions for several locations in the rim seal and cavity areas. Results indicate reasonable sampling methods were used to achieve isokinetic sampling conditions.

scholarly journals Experimental Deposition of NaCl Particles From Turbulent Flows at Gas Turbine Temperatures

2019 ◽  
Vol 141 (2) ◽  
Author(s):  
Peter R. Forsyth ◽  
David R. H. Gillespie ◽  
Matthew McGilvray
Keyword(s):  
Gas Turbine ◽  
Turbulent Flows ◽  
Size Range ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Horizontal Pipe ◽  
Experimental Campaign ◽  
System Temperature ◽  
Secondary Air

The ingestion and deposition of solid particulates within gas turbine engines has become a very significant concern for both designers and operators in recent times. Frequently aircraft are operated in environments where sand, ash, dust, and salt are present, which can drive damage mechanisms from long term component degradation to in-flight flame-out. Experiments are presented to assess deposition characteristics of sodium chloride (NaCl) at gas turbine secondary air system temperature conditions in horizontal pipe flow. Monodisperse NaCl particles were generated in the size range 2.0–6.5 µm, with gas temperatures 390–480 °C, and metal temperatures 355–730 °C. Two engine-representative surface roughnesses were assessed. An experimental technique for the measurement of deposited NaCl based on solution conductivity was developed and validated. Experiments were carried out under isothermal and nonisothermal/thermophoretic conditions. An initial experimental campaign was conducted under ambient and isothermal conditions; high temperature isothermal results showed good similarity. Under thermophoretic conditions, deposition rates varied by up to several orders of magnitude compared to isothermal rates.

Heat Transfer and Pressure Loss Measurements in Additively Manufactured Wavy Microchannels

2016 ◽  
Vol 139 (1) ◽  
Author(s):  
Kathryn L. Kirsch ◽  
Karen A. Thole
Keyword(s):  
Heat Transfer ◽  
Additive Manufacturing ◽  
Pressure Loss ◽  
Reynolds Numbers ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Pressure Losses ◽  
Wavy Channels ◽  
The Waves

The role of additive manufacturing for the hot section components of gas turbine engines grows ever larger as progress in the industry continues. The opportunity for the heat transfer community is to exploit the use of additive manufacturing in developing nontraditional cooling schemes to be built directly into components. This study investigates the heat transfer and pressure loss performance of additively manufactured wavy channels. Three coupons, each containing channels of a specified wavelength (length of one wave period), were manufactured via direct metal laser sintering (DMLS) and tested at a range of Reynolds numbers. Results show that short wavelength channels yield high pressure losses, without corresponding increases in heat transfer, due to the flow structure promoted by the waves. Longer wavelength channels offer less of a penalty in pressure drop with good heat transfer performance.

Performance Measurements of a Unique Louver Particle Separator for Gas Turbine Engines

2012 ◽  
Author(s):  
Grant O. Musgrove ◽  
Karen A. Thole ◽  
Eric Grover ◽  
Joseph Barker
Keyword(s):  
Secondary Flow ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Collection Efficiency ◽  
Reynolds Numbers ◽  
Solid Particles ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Performance Measurements ◽  
Particle Separator

Solid particles, such as sand, ingested into gas turbine engines reduce the coolant flow in the turbine by blocking cooling channels in the secondary flow path. One method to remove solid particles from the secondary flow path is to use an inertial particle separator because of its ability to incur minimal pressure losses in high flow rate applications. In this paper, an inertial separator is presented that is made up of an array of louvers followed by a static collector. The performance of two inertial separator configurations was measured in a unique test facility. Performance measurements included pressure loss and collection efficiency for a range of Reynolds numbers and sand sizes. To complement the measurements, both two-dimensional and three-dimensional computational results are presented for comparison. Computational predictions of pressure loss agreed with measurements at high Reynolds numbers, whereas predictions of sand collection efficiency for a sand size range 0–200μm agreed within 10% of experimental measurements over the range of Reynolds numbers. Collection efficiency values were measured to be as high as 35%, and pressure loss measurements were equivalent to less than 1% pressure loss in an engine application.

An Advanced Multi-Configuration Stator Well Cooling Test Facility

2010 ◽  
Author(s):  
D. D. Coren ◽  
N. R. Atkins ◽  
J. R. Turner ◽  
D. E. Eastwood ◽  
S. Davies ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Flow Rate ◽  
Coolant Flow ◽  
Coolant Flow Rate ◽  
Test Facility ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Displacement Sensors ◽  
Cooling Air ◽  
Coolant Delivery

Optimisation of cooling systems within gas turbine engines is of great interest to engine manufacturers seeking gains in performance, efficiency and component life. The effectiveness of coolant delivery is governed by complex flows within the stator wells and the interaction of main annulus and cooling air in the vicinity of the rim seals. This paper reports the development of a test facility which allows the interaction of cooling air and main gas paths to be measured at conditions representative of those found in modern gas turbine engines. The test facility features a two stage turbine with an overall pressure ratio of approximately 2.6:1. Hot air is supplied to the main annulus using a Rolls-Royce Dart compressor driven by an aero-derivative engine plant. Cooling air can be delivered to the stator wells at multiple locations and at a range of flow rates which cover bulk ingestion through to bulk egress. The facility has been designed with adaptable geometry to enable rapid changes of cooling air path configuration. The coolant delivery system allows swift and accurate changes to the flow settings such that thermal transients may be performed. Particular attention has been focused on obtaining high accuracy data, using a radio telemetry system, as well as thorough through-calibration practices. Temperature measurements can now be made on both rotating and stationary discs with a long term uncertainty in the region of 0.3 K. A gas concentration measurement system has also been developed to obtain direct measurement of re-ingestion and rim seal exchange flows. High resolution displacement sensors have been installed in order to measure hot running geometry. This paper documents the commissioning of a test facility which is unique in terms of rapid configuration changes, non-dimensional engine matching and the instrumentation density and resolution. Example data for each of the measurement systems is presented. This includes the effect of coolant flow rate on the metal temperatures within the upstream cavity of the turbine stator well, the axial displacement of the rotor assembly during a commissioning test, and the effect of coolant flow rate on mixing in the downstream cavity of the stator well.

Thermal Details in a Rotor–Stator Cavity at Engine Conditions With a Mainstream

1992 ◽  
Vol 114 (2) ◽  
pp. 446-453 ◽  
Author(s):  
S. H. Ko ◽  
D. L. Rhode
Keyword(s):  
Stokes Equations ◽  
Numerical Study ◽  
Navier Stokes ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
Flow Conditions ◽  
Navier Stokes Equations ◽  
Flow Parameters ◽  
Purge Flow ◽  
Compressible Turbulent Flow

This investigation involves a numerical study of enclosed Rotor–Stator cavities of gas turbine engines. The complete elliptic form of the 2-D, axisymmetric Navier–Stokes equations for compressible turbulent flow were solved. Included are the complete fluid and thermal effects of the hot mainstream gas interacting with the cooling cavity purge flow at actual engine flow conditions for generalized geometries. Additional flow conditions above and below those for engine nominal conditions are also considered. The relationships among the important flow parameters are investigated by examining the entire set of computations. The predictions reveal that a small recirculation zone in the stator shroud axial gap region is the primary mechanism for the considerable thermal transport from the mainstream to the turbine blade root/retainer region of the rotor.

scholarly journals Aerodynamic improvement of an aircraft gas-turbine engine fan

2021 ◽  
Vol 2021 (3) ◽  
pp. 23-29
Author(s):  
Yu.A. Kvasha ◽  
◽  
N.A. Zinevych ◽  
Keyword(s):  
Gas Turbine ◽  
Flow Rate ◽  
Turbine Engines ◽  
Gas Flows ◽  
Gas Turbine Engines ◽  
Air Flow Rate ◽  
Profile Angle ◽  
Turbine Engine ◽  
Blade Height ◽  
Second Stage

This work is concerned with the development of approaches to the aerodynamic improvement of axial-flow compressors for gas-turbine engines. The aim of this work is the aerodynamic improvement of an aircraft gas-turbine engine two-stage fan by numerical simulation of 3D turbulent gas flows. The approach used in this study features: varying the spatial shape of the fan blades for the first- and the second-stage impeller by varying the profile angle along the blade height; formulating quality criteria as the mean integral values of the power characteristics of each impeller of the fan over the operating range of the air flow rate through the impeller; and searching for advisable values of the impeller blade parameters by scanning the independent variable range at points that form a uniformly distributed sequence of small length. The basic tool is a numerical method developed at the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine, which simulates 3D turbulent gas flows using the complete averaged Navier¬–Stokes equations and a two-parameter turbulence model. It is shown that varying the profile angle along the blade height for the fan second-stage impeller allows one to increase the air compression ratio in the fan by about 2 percent throughout the operating range of the fan air flow rate without affecting the adiabatic efficiency of the fan. On the whole, by the example of the fan under study, the paper considers the assumption that the aerodynamic improvement of compressors at the initial stage can be made on an impeller by impeller basis. It is shown that in further analysis providing the gas-dynamic stability of the compressor should be accounted for. The results obtained are intended to be used in the aerodynamic improvement of multistage compressors for aircraft gas-turbine engines and various power plant.

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