Impact of the Combustor-Turbine Interface Slot Orientation on the Durability of a Nozzle Guide Vane Endwall

2013 ◽  
Vol 135 (4) ◽  
Author(s):  
Alan Thrift ◽  
Karen Thole ◽  
Satoshi Hada
Keyword(s):  
Guide Vane ◽  
Leading Edge ◽  
Time Resolved ◽  
Turbine Engine ◽  
External Cooling ◽  
Image Velocimetry ◽  
Nozzle Guide Vane ◽  
Edge Vortex ◽  
Nozzle Guide ◽  
Stagnation Plane

The combustor-turbine interface is an essential component in a gas turbine engine as it allows for thermal expansion between the first stage turbine vanes and combustor section. Although not considered as part of the external cooling scheme, leakage flow from the combustor-turbine interface can be utilized as coolant. This paper reports on the effects of orientation of a two-dimensional leakage slot, simulating the combustor-turbine interface, on the net heat flux reduction to a nozzle guide vane endwall. In addition to adiabatic effectiveness and heat transfer measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed in the vane stagnation plane. Four interface slot orientations of 90 deg, 65 deg, 45 deg, and 30 deg located at 17% axial chord upstream of a first vane in a linear cascade were studied. Results indicate that reducing the slot angle to 45 deg can provide as much as a 137% reduction to the average heat load experienced by the endwall. Velocity measurements indicate the formation of a large leading edge vortex for coolant injected at 90 deg and 65 deg while coolant injected at 45 deg and 30 deg flows along the endwall and washes up the vane surface at the endwall junction.

Impact of the Combustor-Turbine Interface Slot Orientation on the Durability of a Nozzle Guide Vane Endwall

2012 ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Guide Vane ◽  
Leading Edge ◽  
Time Resolved ◽  
Turbine Engine ◽  
External Cooling ◽  
Image Velocimetry ◽  
Nozzle Guide Vane ◽  
Edge Vortex ◽  
Nozzle Guide ◽  
Stagnation Plane

The combustor-turbine interface is an essential component in a gas turbine engine as it allows for thermal expansion between the first stage turbine vanes and combustor section. Although not considered as part of the external cooling scheme, leakage flow from the combustor-turbine interface can be utilized as coolant. This paper reports on the effects of orientation of a two-dimensional leakage slot, simulating the combustor-turbine interface, on the net heat flux reduction to a nozzle guide vane endwall. In addition to adiabatic effectiveness and heat transfer measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed in the vane stagnation plane. Four interface slot orientations of 90°, 65°, 45°, and 30° located at 17% axial chord upstream of a first vane in a linear cascade were studied. Results indicate that reducing the slot angle to 45° can provide as much as a 137% reduction to the average heat load experienced by the endwall. Velocity measurements indicate the formation of a large leading edge vortex for coolant injected at 90° and 65° while coolant injected at 45° and 30° flows along the endwall and washes up the vane surface at the endwall junction.

Effects of Orientation and Position of the Combustor-Turbine Interface on the Cooling of a Vane Endwall

2011 ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Vortex Formation ◽  
Horseshoe Vortex ◽  
Time Resolved ◽  
Image Velocimetry ◽  
Thermal Measurements ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Stagnation Plane

First stage, nozzle guide vanes and accompanying endwalls are extensively cooled by the use of film cooling through discrete holes and leakage flow from the combustor-turbine interface gap. While there are cooling benefits from the interface gap, it is generally not considered as part of the cooling scheme. This paper reports on the effects of the position and orientation of a two-dimensional slot on the cooling performance of a nozzle guide vane endwall. In addition to surface thermal measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed at the vane stagnation plane. Two slot orientations, 90° and 45°, and three streamwise positions were studied. Effectiveness results indicate a significant increase in area averaged effectiveness for the 45° slot relative to the 90° orientation. Flowfield measurements show dramatic differences in the horseshoe vortex formation.

Effects of Orientation and Position of the Combustor-Turbine Interface on the Cooling of a Vane Endwall

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Vortex Formation ◽  
Horseshoe Vortex ◽  
Time Resolved ◽  
Image Velocimetry ◽  
Thermal Measurements ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Stagnation Plane

First stage, nozzle guide vanes and accompanying endwalls are extensively cooled by the use of film cooling through discrete holes and leakage flow from the combustor-turbine interface gap. While there are cooling benefits from the interface gap, it is generally not considered as part of the cooling scheme. This paper reports on the effects of the position and orientation of a two-dimensional slot on the cooling performance of a nozzle guide vane endwall. In addition to surface thermal measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed at the vane stagnation plane. Two slot orientations, 90 deg and 45 deg, and three streamwise positions were studied. Effectiveness results indicate a significant increase in area averaged effectiveness for the 45 deg slot relative to the 90 deg orientation. Flowfield measurements show dramatic differences in the horseshoe vortex formation.

Influence of Pre-History and Leading Edge Contouring on Aero-Performance of a 3D Nozzle Guide Vane

2013 ◽  
Author(s):  
Ranjan Saha ◽  
Boris I. Mamaev ◽  
Jens Fridh ◽  
Björn Laumert ◽  
Torsten H. Fransson
Keyword(s):  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Stagnation Line ◽  
Wide Range ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Loss Coefficients ◽  
Exit Mach Number ◽  
Nozzle Guide

Experiments are conducted to investigate the effect of the pre-history in the aerodynamic performance of a three-dimensional nozzle guide vane with a hub leading edge contouring. The performance is determined with two pneumatic probes (5 hole and 3 hole) concentrating mainly on the endwall. The investigated vane is a geometrically similar gas turbine vane for the first stage with a reference exit Mach number of 0.9. Results are compared for the baseline and filleted cases for a wide range of operating exit Mach numbers from 0.5 to 0.9. The presented data includes loading distributions, loss distributions, fields of exit flow angles, velocity vector and vorticity contour, as well as, mass-averaged loss coefficients. The results show an insignificant influence of the leading edge fillet on the performance of the vane. However, the pre-history (inlet condition) affects significantly in the secondary loss. Additionally, an oil visualization technique yields information about the streamlines on the solid vane surface which allows identifying the locations of secondary flow vortices, stagnation line and saddle point.

Aerothermal Performance of Shielded Vane Design

2017 ◽  
Vol 139 (11) ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Total Pressure ◽  
High Speed ◽  
Guide Vane ◽  
Leading Edge ◽  
Experimental Facility ◽  
Numerical Studies ◽  
Nozzle Guide Vane ◽  
Inlet Conditions ◽  
Nozzle Guide

This paper presents an experimental investigation of the concept of using the combustor transition duct wall to shield the nozzle guide vane leading edge. The new vane is tested in a high-speed experimental facility, demonstrating the improved aerodynamic and thermal performance of the shielded vane. The new design is shown to have a lower average total pressure loss than the original vane, and the heat transfer on the vane surface is overall reduced. The peak heat transfer on the vane leading edge–endwall junction is moved further upstream, to a region that can be effectively cooled as shown in previously published numerical studies. Experimental results under engine-representative inlet conditions showed that the better performance of the shielded vane is maintained under a variety of inlet conditions.

Heat Transfer Studies on the Interior Surfaces of Cooled Nozzle Guide Vane in a Linear Cascade

2015 ◽  
Author(s):  
Arun Kumar Pujari ◽  
Bhamidi Prasad ◽  
Nekkanti Sitaram
Keyword(s):  
Heat Transfer ◽  
Surface Temperature ◽  
Guide Vane ◽  
Internal Heat ◽  
Leading Edge ◽  
Internal Surface ◽  
Transfer Coefficients ◽  
Computational Data ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Experimental and computational heat transfer investigations are reported in the interior side of a nozzle guide vane (NGV) subjected to combined impingement and film cooling. The domain of study is a two dimensional five-vane cascade having four passages. Each vane has a chord length of 228 mm and the pitch distance between the vanes is 200 mm. The vane internal surface is cooled by dry air supplied through the two impingement inserts: the front and the aft. The mass flow through the impingement chamber is varied, for a fixed spacing (H) to jet diameter (d) ratio of 1.2. The surface temperature distributions, at certain locations of the vane interior, are measured by pasting strips of liquid crystal sheets. The vane interior surface temperature distribution is also obtained by computations carried out by using Shear stress transport (SST) k-ω turbulence model in the ANSY FLUENT-14 flow solver. The computational data are in good agreement with the measured values of temperature. The internal heat transfer coefficients are thence determined along the leading edge and the mid span region from the computational data.

Failure Analysis of Nozzle Guide Vane of a Low Pressure Turbine in an Aero Gas Turbine Engine

2014 ◽  
Vol 14 (5) ◽  
pp. 578-587 ◽  
Author(s):  
R. K. Mishra ◽  
Johney Thomas ◽  
K. Srinivasan ◽  
Vaishakhi Nandi ◽  
Raghavendra Bhat
Keyword(s):  
Failure Analysis ◽  
Gas Turbine ◽  
Guide Vane ◽  
Low Pressure ◽  
Gas Turbine Engine ◽  
Turbine Engine ◽  
Low Pressure Turbine ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

scholarly journals Gas Turbine Testing on Naval Distillate Fuel

1971 ◽  
Author(s):  
J. S. Siemietkowski
Keyword(s):  
Gas Turbine ◽  
Diesel Fuel ◽  
Guide Vane ◽  
Engine Operation ◽  
Accelerated Corrosion ◽  
Turbine Engine ◽  
Base Materials ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Distillate Fuel

A Pratt & Whitney FT4A Marine Gas Turbine Engine rated at 22,600 hp, 3600 rpm was run at the Naval Ship Engineering Center, Philadelphia Division for 1000 hr. Fuel used was naval distillate having a vanadium level of 0.5 ppm. Basically there was no problem with engine operation on naval distillate when compared to diesel fuel. The smoke level was barely visible at high powers. Coalescent fuel filters are a problem due to their relatively short (100–130 hr) life. The corrosion rate was accelerated when compared to navy diesel fuel. The fuel parameter suspect is vanadium, however other parameters may be at fault. Additional efforts are required into definitely determining the cause of accelerated corrosion and also into optimizing nozzle guide vane and turbine blade base materials and coatings.

Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Convective Heat Transfer Measurements

2010 ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Heat Transfer ◽  
Guide Vane ◽  
Leading Edge ◽  
Critical Factor ◽  
Suction Side ◽  
Leakage Flow ◽  
Cross Sectional ◽  
Nozzle Guide Vane ◽  
Linear Slope ◽  
Nozzle Guide

Heat transfer is a critical factor in the durability of gas turbine components, particularly in the first vane. An axisymmetric contour is sometimes used to contract the cross sectional area from the combustor to the first stage vane in the turbine. Such contouring can lead to significant changes in the endwall flows thereby altering the heat transfer. This paper investigates the effect of axisymmetric contouring on endwall heat transfer of a nozzle guide vane. Heat transfer measurements are performed on the endwalls of a planar and contoured passage whereby one endwall is modified with a linear slope in the case of the contoured passage. Included in this study is upstream leakage flow issuing from a slot normal to the inlet axis. Each of the endwalls within the contoured passage presents a unique flowfield. For the contoured passage, the flat endwall is subject to an increased acceleration through the area contraction while the contoured endwall includes both increased acceleration and a turning of streamlines due to the slope. Results indicate heat transfer is reduced on both endwalls of the contoured passage relative to the planar passage. In the case of all endwalls, increasing leakage mass flow rate leads to an increase in heat transfer near the suction side of the vane leading edge.

Nozzle Passage Endwall Effectiveness Values With Various Combustor Coolant Flowrates—Part 2: Endwall and Vicinity Surface Effectiveness Measurements

2021 ◽  
Vol 143 (4) ◽  
Author(s):  
Kedar P. Nawathe ◽  
Rui Zhu ◽  
Enci Lin ◽  
Yong W. Kim ◽  
Terrence W. Simon
Keyword(s):  
Gas Turbines ◽  
Guide Vane ◽  
Suction Surface ◽  
Turbine Engine ◽  
Pressure Surface ◽  
Coolant Injection ◽  
Measurement Results ◽  
Nozzle Guide Vane ◽  
Effectiveness Measurement ◽  
Nozzle Guide

Abstract Effective coolant schemes are required for providing cooling to the first-stage stator vanes of gas turbines. To correctly predict coolant performance on the endwall and vane surfaces, these coolant schemes should also consider the effects of coolant streams introduced upstream in the combustor section of a gas turbine engine. This two-part paper presents measurements taken on a first-stage nozzle guide vane cascade that includes combustor coolant injection. The first part of this paper explains how coolant transport and coolant-mainstream interaction in the vane passage is affected by changing the combustor coolant and endwall film coolant flowrates. This paper explains how those flows affect the coolant effectiveness on the endwall and vane surfaces. Part one showed that a significant amount of coolant injected upstream of the endwall is present along the pressure surface of the vanes as well as over the endwall. Part two shows effectiveness measurement results taken in this study on the endwall and pressure and suction surfaces of the vanes. Sustained endwall coolant effectiveness is observed along the whole passage for all cases. It is uniform in the pitch-wise direction. Combustor coolant flow significantly affects cooling performance even near the trailing edge. The modified flowfield results in the pressure surface being cooled more effectively than the suction surface. While the effectiveness distribution on the pressure surface varies with combustor and film coolant flowrates, the distribution along the suction surface remains largely unchanged.

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