Performance Assessment of an Annular S-Shaped Duct

An experimental investigation has been carried out to determine the aerodynamic performance of an annular S-shaped duct representative of that used to connect the compressor spools of aircraft gas turbine engines. For inlet conditions in which boundary layers are developed along an upstream entry length the static pressure, shear stress and velocity distributions are presented. The data shows that as a result of flow curvature significant streamwise pressure gradients exist within the duct, with this curvature also affecting the generation and suppression of turbulence. The stagnation pressure loss within the duct is also assessed and is consistent with the measured distributions of shear stress. More engine representative conditions are provided by locating a single stage compressor at inlet to the duct. Relative to the naturally developed inlet conditions the flow within the duct is less likely to separate, but mixing out of the compressor blade wakes increases the measured duct loss. With both types of inlet conditions the effect of a radial strut, such as that used for carrying loads and engine services, is also described both in terms of the static pressure distribution along the strut and its contribution to overall loss.

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Performance Assessment of an Annular S-Shaped Duct

Journal of Turbomachinery ◽

10.1115/1.2841003 ◽

1997 ◽

Vol 119 (1) ◽

pp. 149-156 ◽

Cited By ~ 24

Author(s):

D. W. Bailey ◽

K. M. Britchford ◽

J. F. Carrote ◽

S. J. Stevens

Keyword(s):

Shear Stress ◽

Static Pressure ◽

Compressor Blade ◽

Turbine Engines ◽

Pressure Gradients ◽

Gas Turbine Engines ◽

Entry Length ◽

Static Pressure Distribution ◽

Flow Curvature ◽

Inlet Conditions

An experimental investigation has been carried out to determine the aerodynamic performance of an annular S-Shaped duct representative of that used to connect the compressor spools of aircraft gas turbine engines. For inlet conditions in which boundary layers are developed along an upstream entry length, the static pressure, shear stress and velocity distributions are presented. The data show that as a result of flow curvature, significant streamwise pressure gradients exist within the duct, with this curvature also affecting the generation and suppression of turbulence. The stagnation pressure loss within the duct is also assessed and is consistent with the measured distributions of shear stress. More engine representative conditions are provided by locating a single-stage compressor at inlet to the duct. Relative to the naturally developed inlet conditions, the flow within the duct is less likely to separate, but mixing out of the compressor blade wakes increases the measured duct loss. With both types of inlet conditions, the effect of a radial strut, such as that used for carrying loads and engine services, is also described both in terms of the static pressure distribution along the strut and its contribution to overall loss.

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Experience With the TF40B Engine in the LCAC Fleet

Volume 2: Aircraft Engine; Marine; Microturbines and Small Turbomachinery ◽

10.1115/92-gt-270 ◽

1992 ◽

Author(s):

Edward M. House

Keyword(s):

Gas Turbine ◽

Hot Corrosion ◽

Compressor Blade ◽

Turbine Engines ◽

Gas Turbine Engines ◽

Improvement Program ◽

Air Cushion ◽

Engine Components ◽

Carbon Erosion ◽

The U.S

Four Textron Lycoming TF40B marine gas turbine engines are used to power the U.S. Navy’s Landing Craft Air Cushion (LCAC) vehicle. This is the first hovercraft of this configuration to be put in service for the Navy as a landing craft. The TF40B has experienced compressor blade pitting, carbon erosion of the first turbine blade and hot corrosion of the hot section. Many of these problems were reduced by changing the maintenance and operation of the LCAC. A Component Improvement Program (CIP) is currently investigating compressor and hot section coatings better suited for operation in a harsh marine environment. This program will also improve the performance of some engine components such as the bleed manifold and bearing seals.

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Experimental Investigation of Flow Resistance and Wall Shear Stress in the Interior Subchannel of a Triangular Array of Parallel Rods

Journal of Fluids Engineering ◽

10.1115/1.3449003 ◽

1979 ◽

Vol 101 (4) ◽

pp. 429-434 ◽

Cited By ~ 12

Author(s):

M. Fakory ◽

N. Todreas

Keyword(s):

Shear Stress ◽

Wall Shear Stress ◽

Friction Factor ◽

Static Pressure ◽

Reynolds Numbers ◽

Triangular Array ◽

Flow Area ◽

Static Pressure Distribution ◽

The Common ◽

Wall Shear

A simulated model of a triangular array of rods with pitch to diameter ratio of 1.1 with air flow was used to study the hydraulic parameters of the liquid metal fast breeder reactor (LMFBR) fuel geometry. The wall shear stress distribution, static pressure distribution, turbulence intensity, and friction factor were measured in the central subchannel from Reynolds numbers of 4 × 103 to 36 × 103. Our results show that the maximum wall shear stress occurs at the largest flow area, the static pressure is not uniform around the rod periphery, there is no detectable presence of secondary flow from the wall shear stress measurements, and the friction factor derived from the measured wall shear stress is less than the common friction factor derived from pressure drop measurement.

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The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 1—A Unique Experimental Facility

Journal of Turbomachinery ◽

10.1115/1.3262142 ◽

1987 ◽

Vol 109 (4) ◽

pp. 520-526 ◽

Cited By ~ 4

Author(s):

S. Deutsch ◽

W. C. Zierke

Keyword(s):

Boundary Layer ◽

Flow Field ◽

Pressure Distribution ◽

Static Pressure ◽

Compressor Blade ◽

Boundary Layer Transition ◽

Two Dimensional ◽

Suction Surface ◽

Pressure Surface ◽

Static Pressure Distribution

A unique cascade facility is described which permits the use of laser-Doppler velocimetry (LDV) to measure blade boundary layer profiles. Because of the need for a laser access window, the facility cannot reply on continuous blade pack suction to achieve two-dimensional, periodic flow. Instead, a strong suction upstream of the blade pack is used in combination with tailboards to control the flow field. The distribution of the upstream suction is controlled through a complex baffling system. A periodic, two–dimensional flow field is achieved at a chord Reynolds number of 500,000 and an incidence angle of 5 deg on a highly loaded, double circular arc, compressor blade. Inlet and outlet flow profiles, taken using five-hole probes, and the blade static-pressure distribution are used to document the flow field for use with the LDV measurements (see Parts 2 and 3). Inlet turbulence intensity is measured, using a hot wire, to be 0.18 percent. The static-pressure distribution suggests both separated flow near the trailing edge of the suction surface and an initially laminar boundary layer profile near the leading edge of the pressure surface. Probe measurements are supplemented by sublimation surface visualization studies. The sublimation studies place boundary layer transition at 64.2 ± 3.9 percent chord on the pressure surface, and indicate separation on the suction surface at 65.6 percent ± 3.5 percent chord.

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The Influence of Inlet Swirl on the Flow Within an Annular S-Shaped Duct

Volume 1: Turbomachinery ◽

10.1115/96-gt-060 ◽

1996 ◽

Cited By ~ 3

Author(s):

D. W. Bailey ◽

J. F. Carrotte

Keyword(s):

Experimental Investigation ◽

Streamwise Velocity ◽

Stagnation Pressure ◽

Turbine Engines ◽

Pressure Gradients ◽

Gas Turbine Engines ◽

Mean Flow ◽

Inlet Swirl ◽

Tangential Momentum ◽

The Mean

An experimental investigation has been carried out to determine the effects of inlet swirl on the flow field that develops within an annular S-shaped duct. The duct is representative of that used to connect the compressor spools on multi-spool gas turbine engines. By removing the outlet guide vanes from an upstream single stage compressor swirl angles in excess of 30° were generated. Results show that within the S-shaped duct tangential momentum (Wr) is conserved, leading to increasing swirl velocities through the duct as the radius decreases. Furthermore, this component influences the streamwise velocity as pressure gradients are established to ensure the mean flow follows the duct curvature. Consequently in the critical region adjacent to the inner casing, where separation is likely to occur, higher streamwise velocities are observed. Within the duct substantial changes also occur to the turbulence field which results in an increased stagnation pressure loss between duct inlet and exit. Data is also presented showing the increasing swirl angles through the duct which has consequences both for the design of the downstream compressor spool and of any radial struts which may be located within the duct.

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Numerical Insight Into Flow and Thermal Patterns Within an Inlet Profile Generator Comparing to Experimental Results

Volume 3: Heat Transfer, Parts A and B ◽

10.1115/gt2006-90276 ◽

2006 ◽

Cited By ~ 4

Author(s):

V. R. Kunze ◽

M. Wolff ◽

M. D. Barringer ◽

K. A. Thole ◽

M. D. Polanka

Keyword(s):

Computational Study ◽

Three Dimensional ◽

Turbine Engines ◽

Gas Turbine Engines ◽

Benchmark Tests ◽

Annular Combustor ◽

Inlet Conditions ◽

Air Force Base ◽

Turbine Design ◽

Insight Into

Historically the design of gas turbine engines have not considered the interaction between the combustor and turbine stages. High pressure turbine vane stages have been designed assuming inlet conditions consistent with a standard turbulent boundary layer profile. However, combustor exit flow entering the vane is known to be highly non-uniform in both the primary and secondary flow regimes. In order to develop higher performance, more efficient, longer life stages, turbine design must take into account combustor exit non-uniformities. The Turbine Research Facility (TRF) at Wright-Patterson Air Force Base has installed a non-reactive full scale annular combustor simulator or more accurately a turbine inlet profile generator to study combustor-vane interaction. Several benchmark tests have been performed on the profile generator consisting of a Taguchi type matrix wherein nine independent variables were adjusted. Supplementing the experimental research at the TRF, a steady state, unstructured, fully three-dimensional CFD analysis was performed. This paper will make comparisons between the CFD and experimental profiles generated by the simulator. Furthermore, the computational study will help to give an understanding of the aerodynamic and aerothermal environment within the generator that experimental instrumentation alone cannot.

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Enhanced Static-Dynamic Pressure Transducer for the Detection of Acoustic Level Flow Instabilities in Gas Turbine Engines

Volume 3: Controls, Diagnostics and Instrumentation; Education; Electric Power; Microturbines and Small Turbomachinery; Solar Brayton and Rankine Cycle ◽

10.1115/gt2011-45750 ◽

2011 ◽

Author(s):

Adam M. Hurst ◽

Scott Goodman ◽

Boaz Kochman ◽

Alex Ned

Keyword(s):

High Temperature ◽

Gas Turbine ◽

Pressure Transducer ◽

Static Pressure ◽

Dynamic Pressure ◽

Flow Instabilities ◽

Sensor Data ◽

Turbine Engines ◽

Gas Turbine Engines ◽

Low Level

The push to advance the performance and longevity of gas turbine engines requires better characterization of flow instabilities within the compressor and most importantly the combustor. Detecting the earliest onset of these flow instabilities can help engineers either manipulate the flow to restabilize it or make informed design changes to the engine. The pressures within gas turbine engines are typically composed of an undesired, low-level oscillatory pressure of less than 1kPa to several kPa superimposed on top of a large, relatively constant pressure of several thousand kPa [1–7]. The high-pressure transducers used to measure the pressures within these environments are often unable to resolve these low-level oscillatory pressures that characterize the flow instabilities because the signal output for such pressures is often the same level as the noise within the sensor-data acquisition system. This paper presents an engine test ready, high temperature, combined static and dynamic pressure transducer that uses static pressure compensation in order to measure these low-level dynamic pressures with an excellent signal to noise ratio and, at the same time, captures the overall static pressure within a gas turbine [8–10]. Test bench experiments demonstrate the static-dynamic transducer’s unique ability to capture both large static or quasi-static pressures of 1,380kPa or greater and simultaneously measure the acoustic-level dynamic pressures superimposed on top of these pressures. The static-dynamic transducer achieves this advanced sensitivity through the use of a low-pass acoustic filter that passes the large static pressure to the reference port of a high sensitivity dynamic pressure sensor within the transducer such that the overall static pressures cancel out and the sensor measures all acoustic-level dynamic pressures. These bench tests additionally demonstrate the transducer’s ability to operate reliably when exposed to the harsh, high temperature environment (up to 500°C) within a gas turbine [8–10].

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Structural synthesis of complex multi-coherent engineering process of compressor blade processing of gas turbine engines with function-oriented coatings

Science intensive technologies in mechanical engineering ◽

10.30987/2223-4608-2020-2020-1-40-48 ◽

2020 ◽

Vol 2020 (1) ◽

pp. 40-48

Author(s):

Alexander Mikhailov ◽

D. Mikhaylov ◽

E. Sheiko ◽

V. Mikhailov

Keyword(s):

Gas Turbine ◽

Technological Process ◽

Compressor Blade ◽

Turbine Engines ◽

Gas Turbine Engines ◽

Erosion Wear ◽

Engineering Process ◽

Structural Synthesis ◽

Turbine Engine ◽

Abrasive Erosion

It is shown that the compressor of a gas turbine engine consists of a group of blades which experience uneven abrasive-erosion wear characterized with unevenness of three types. To increase compressor blade life it is offered to use function-oriented coatings and to ensure their properties on the basis of the principle of equal resource. In the work there are solved the problems of a structural synthesis of a complex multi-coherent technological process of function-oriented blade group sputtering.

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