Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Adiabatic Effectiveness Measurements

Adiabatic Effectiveness ◽

Improved Performance ◽

Nozzle Guide ◽

Gas turbine designs seek improved performance by modifying the endwalls of nozzle guide vanes in the engine hot section. Within the nozzle guide vanes these modifications can be in the form of an axisymmetric contour as the area contracts from the combustor to the turbine. This paper investigates the effect of axisymmetric endwall contouring on the cooling performance of a film cooled endwall. Adiabatic effectiveness measurements were performed in a planar passage for comparison to a contoured passage whereby the exit Reynolds numbers was matched. For the contoured passage, measurements were performed on both the flat endwall and on the contoured endwall. Fully expanded film cooling holes were distributed on the endwall surface preceded by a two-dimensional slot normal to the inlet axis. Results indicated that the coolant coverage from the upstream leakage slot was spread over a larger area of the contoured endwall in comparison to the flat endwall of the planar passage. Film cooling effectiveness on the flat endwall of the contoured passage showed minimal differences relative to the planar passage results. The contracting endwall of the contoured passage, however, showed a significant reduction with average film cooling effectiveness levels approximately 40% lower than the planar passage at low film cooling flow rates. In the case of all endwalls, increasing leakage and film cooling mass flow rates led to an increase in cooling effectiveness and coolant coverage.

Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Adiabatic Effectiveness Measurements

10.1115/1.4002965 ◽

2011 ◽

Vol 133 (4) ◽

Cited By ~ 7

Author(s):

A. A. Thrift ◽

K. A. Thole ◽

S. Hada

Keyword(s):

Film Cooling ◽

Guide Vane ◽

Cooling Effectiveness ◽

Flow Rates ◽

Guide Vanes ◽

Adiabatic Effectiveness ◽

Improved Performance ◽

Nozzle Guide ◽

Gas turbine designs seek improved performance by modifying the endwalls of nozzle guide vanes in the engine hot section. Within the nozzle guide vanes, these modifications can be in the form of an axisymmetric contour as the area contracts from the combustor to the turbine. This paper investigates the effect of axisymmetric endwall contouring on the cooling performance of a film cooled endwall. Adiabatic effectiveness measurements were performed in a planar passage for comparison to a contoured passage, whereby the exit Reynolds numbers were matched. For the contoured passage, measurements were performed both on the flat endwall and on the contoured endwall. Fully expanded film cooling holes were distributed on the endwall surface preceded by a two-dimensional slot normal to the inlet axis. Results indicated that the coolant coverage from the upstream leakage slot was spread over a larger area of the contoured endwall in comparison to the flat endwall of the planar passage. Film cooling effectiveness on the flat endwall of the contoured passage showed minimal differences relative to the planar passage results. The contracting endwall of the contoured passage, however, showed a significant reduction with average film cooling effectiveness levels approximately 40% lower than the planar passage at low film cooling flow rates. In the case of all endwalls, increasing leakage and film cooling mass flow rates led to an increase in cooling effectiveness and coolant coverage.

The Effect of Step Misalignment on Purge Flow Cooling of Nozzle Guide Vane at Transonic Conditions

10.1115/1.4047634 ◽

2020 ◽

Vol 142 (10) ◽

Author(s):

Luke Luehr ◽

Ridge Sibold ◽

Shuo Mao ◽

Wing F. Ng ◽

Zhigang Li ◽

...

Keyword(s):

Nusselt Number ◽

Film Cooling ◽

Guide Vane ◽

Cooling Effectiveness ◽

Blowing Ratio ◽

Purge Flow ◽

Exit Mach Number ◽

Nozzle Guide ◽

Abstract This study describes a detailed investigation on the effects that upstream step misalignment and upstream purge film cooling have on the endwall heat transfer for first stage nozzle guide vanes (NGVs) in a gas turbine at transonic conditions. Endwall Nusselt number and adiabatic film-cooling effectiveness distributions were experimentally measured and compared with flow visualization. Tests were conducted in a transonic linear cascade blowdown facility at an inlet freestream turbulence intensity of 16%, an exit Mach number of 0.85, and an exit Re = 1.5 × 106 based on axial chord. Varied upstream purge blowing ratios (BRs) and a no-blowing case were tested for three different upstream step geometries, the baseline (no misalignment), a span-wise upstream step of +4.86% span, and a step of −4.86% span. Experimentation shows that compared with no-blowing case, the addition of upstream purge film cooling increases the Nusselt number at injection upward of 50% but lowers it in the passage throat by approximately 20%. The backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow aiding to help keep the film attached to the endwall. Increasing the blowing ratio increases film-cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff.

The Effect of Step Misalignment on Purge Flow Cooling of Nozzle Guide Vane at Transonic Conditions

Volume 5B: Heat Transfer ◽

10.1115/gt2019-91810 ◽

2019 ◽

Author(s):

Luke Luehr ◽

Ridge Sibold ◽

Shuo Mao ◽

Wing F. Ng ◽

Zhigang Li ◽

...

Keyword(s):

Nusselt Number ◽

Film Cooling ◽

Guide Vane ◽

Cooling Effectiveness ◽

Blowing Ratio ◽

Purge Flow ◽

Exit Mach Number ◽

Nozzle Guide ◽

Abstract This study describes a detailed investigation on the effects that step misalignment and upstream purge film cooling have on the endwall heat transfer for 1st stage nozzle guide vanes in a land-based power generating gas turbine at transonic conditions. Endwall Nusselt number and adiabatic film cooling effectiveness distributions were experimentally measured and compared with flow visualization. Tests were conducted in a transonic linear cascade blowdown facility where data were gathered at an exit Mach number of 0.85 with a freestream turbulence intensity of 16% at an exit Re = 1.5 × 106 based on axial chord. Varied upstream purge blowing ratios and a no blowing case were tested for 3 different upstream step geometries, the baseline (no misalignment), a span-wise upstream step of +4.86% span, and a step of −4.86% span. Experimentation shows that the addition of upstream purge film cooling increases the Nusselt number at injection upwards of 50% but lowers it in the throat of the passage by approximately 20% compared to no blowing case. The addition of a backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow aiding to help keep the film attached to the endwall at higher blowing ratios. Increasing the blowing ratio increases film cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff.

Comparison of RANS and DES Modeling Against Measurements of Leading Edge Film Cooling on a First-Stage Vane

Volume 5C: Heat Transfer ◽

10.1115/gt2016-57567 ◽

2016 ◽

Author(s):

S. Ravelli ◽

G. Barigozzi

Keyword(s):

Film Cooling ◽

Guide Vane ◽

Leading Edge ◽

Spanwise Direction ◽

Intensity Level ◽

Edge Region ◽

Cooling Effectiveness ◽

Stagnation Line ◽

Adiabatic Effectiveness

The performance of a showerhead arrangement of film cooling in the leading edge region of a first stage nozzle guide vane was experimentally and numerically evaluated. A six-vane linear cascade was tested at an isentropic exit Mach number of Ma2s = 0.42, with a high inlet turbulence intensity level of 9%. The showerhead cooling scheme consists of four staggered rows of cylindrical holes evenly distributed around the stagnation line, angled at 45° towards the tip. The blowing ratios tested are BR = 2.0, 3.0 and 4.0. Adiabatic film cooling effectiveness distributions on the vane surface around the leading edge region were measured by means of Thermochromic Liquid Crystals technique. Since the experimental contours of adiabatic effectiveness showed that there is no periodicity across the span, the CFD calculations were conducted by simulating the whole vane. Within the RANS framework, the very widely used Realizable k-ε (Rke) and the Shear Stress Transport k-ω (SST) turbulence models were chosen for simulating the effect of the BR on the surface distribution of adiabatic effectiveness. The turbulence model which provided the most accurate steady prediction, i.e. Rke, was selected for running Detached Eddy Simulation at the intermediate value of BR = 3. Fluctuations of the local temperature were computed by DES, due to the vortex structures within the shear layers between the main flow and the coolant jets. Moreover, mixing was enhanced both in the wall-normal and spanwise direction, compared to RANS modeling. DES roughly halved the prediction error of laterally averaged film cooling effectiveness on the suction side of the leading edge. However, neither DES nor RANS provided the expected decay of effectiveness progressing downstream along the pressure side, with 15% overestimation of ηav at s/C =0.2.

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

10.1115/1.4004825 ◽

2012 ◽

Vol 134 (5) ◽

Cited By ~ 9

Author(s):

Giovanna Barigozzi ◽

Antonio Perdichizzi ◽

Silvia Ravelli

Keyword(s):

Film Cooling ◽

Guide Vane ◽

Pressure Side ◽

Trailing Edge ◽

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.

Interactions Between the Combustor Swirl and the High Pressure Stator of a Turbine

Volume 8: Turbomachinery, Parts A, B, and C ◽

10.1115/gt2012-69157 ◽

2012 ◽

Cited By ~ 9

Author(s):

Lucas Giller ◽

Heinz-Peter Schiffer

Keyword(s):

High Pressure ◽

Flow Field ◽

Film Cooling ◽

Field Measurements ◽

Cooling Effectiveness ◽

Nozzle Guide ◽

Nozzle Guide Vanes ◽

Film Cooling Holes ◽

The Impact

The interaction between the strongly swirling combustor outflow and the high pressure turbine nozzle guide vanes were investigated at the cascade test rig at Technische Universität Darmstadt. The test section of the rig consists of six swirl generators and five cascade vanes. The three middle vanes are equipped with film cooling holes at the leading edges. The swirler nozzles are aligned with the center of the cascade passages. The operating settings are defined by the swirl number, the distance between the swirler nozzles and the vanes, the blowing ratio and the radial angle of the film cooling holes. Flow field measurements using PIV downstream of the swirlers and five hole probe measurements at the inlet and outlet plane of the cascade were accomplished. Measurements using the ammonia diazo technique to determine the adiabatic film cooling effectiveness on the surface of the center cascade vane were also carried out. It is shown that a swirling inflow leads to a strong alteration of the flow field and the losses in the passages in comparison to an axial inflow. Furthermore, the impact of the swirl on the formation of the cooling film and it’s adiabatic film cooling effectiveness is presented.

Application of Unsteady CFD Methods to Trailing Edge Cutback Film Cooling

Volume 5B: Heat Transfer ◽

10.1115/gt2014-25435 ◽

2014 ◽

Cited By ~ 1

Author(s):

S. Ravelli ◽

G. Barigozzi

Keyword(s):

Film Cooling ◽

Guide Vane ◽

Navier Stokes ◽

Trailing Edge ◽

Cooling Effectiveness ◽

Nozzle Guide Vane ◽

Unsteady Rans ◽

Exit Mach Number ◽

Nozzle Guide

The main purpose of this numerical investigation is to overcome the limitations of the steady modeling in predicting the cooling efficiency over the cutback surface in a high pressure turbine nozzle guide vane. Since discrepancy between Reynolds-averaged Navier–Stokes (RANS) predictions and measured thermal coverage at the trailing edge was attributable to unsteadiness, Unsteady RANS (URANS) modeling was implemented to evaluate improvements in simulating the mixing between the mainstream and the coolant exiting the cutback slot. With the aim of reducing the computation effort, only a portion of the airfoil along the span was simulated at an exit Mach number of Ma2is = 0.2. Three values of the coolant-to-mainstream mass flow ratio were considered: MFR = 0.66%, 1.05%, and 1.44%. Nevertheless the inherent vortex shedding from the cutback lip was somehow captured by the URANS method, the computed mixing was not enough to reproduce the measured drop in adiabatic effectiveness η along the streamwise direction, over the cutback surface. So modeling was taken a step further by using the Scale Adaptive Simulation (SAS) method at MFR = 1.05%. Results from the SAS approach were found to have potential to mimic the experimental measurements. Vortices shedding from the cutback lip were well predicted in shape and magnitude, but with a lower frequency, as compared to PIV data and flow visualizations. Moreover, the simulated reduction in film cooling effectiveness toward the trailing edge was similar to that observed experimentally.

Nozzle Guide Vane Static Strip Seals

Volume 3: Heat Transfer, Parts A and B ◽

10.1115/gt2006-90185 ◽

2006 ◽

Cited By ~ 1

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.

Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 1—Experimental Technique and Results

10.1115/1.2928026 ◽

1992 ◽

Vol 114 (4) ◽

pp. 734-740 ◽

Cited By ~ 28

Author(s):

S. P. Harasgama ◽

C. D. Burton

Keyword(s):

Heat Transfer ◽

Film Cooling ◽

Guide Vane ◽

Density Ratio ◽

Suction Side ◽

Coolant Temperature ◽

Guide Vanes ◽

Nozzle Guide ◽

Annular Cascade ◽

Heat transfer and aerodynamic measurements have been made on the endwalls of an annular cascade of turbine nozzle guide vanes in the presence of film cooling. The results indicate that high levels of cooling effectiveness can be achieved on the endwalls of turbine nozzle guide vanes (NGV). The NGV were operated at the correct engine nondimensional conditions of Reynolds number, Mach number, gas-to-wall temperature ratio, and gas-to-coolant density ratio. The results show that the secondary flow and horseshoe vortex act on the coolant, which is convected toward the suction side of the NG V endwall passage. Consequently the coolant does not quite reach the pressure side/casing trailing edge, leading to diminished cooling in this region. Increasing the blowing rate from 0.52 to 1.1 results in significant reductions in heat transfer to the endwall. Similar trends are evident when the coolant temperature is reduced. Measured heat transfer rates indicate that over most of the endwall region the film cooling reduces the Nusselt number by 50 to 75 percent.

Overall Cooling Effectiveness Measurements on Pressure Side Surface of the Nozzle Guide Vane With Optimized Film Cooling Hole Arrangements

Volume 5A: Heat Transfer ◽

10.1115/gt2017-63421 ◽

2017 ◽

Cited By ~ 2

Author(s):

Dong-Ho Rhee ◽

Young Seok Kang ◽

Bong Jun Cha ◽

Sanga Lee

Keyword(s):

Film Cooling ◽

Guide Vane ◽

Side Surface ◽

Internal Cooling ◽

Pressure Side ◽

Cooling Effectiveness ◽

Nozzle Guide Vane ◽

Rib Turbulators ◽

Nozzle Guide

Most of the optimization researches on film cooling have dealt with adiabatic film cooling effectiveness on the surface. However, the information on the overall cooling effectiveness is required to estimate exact performance of the optimization configuration since hot components such as nozzle guide vane have not only film cooling but also internal cooling features such as rib turbulators, jet impingement and pin-fins on the inner surface. Our previous studies [1,2] conducted the hole arrangement optimization to improve adiabatic film cooling effectiveness values and uniformity on the pressure side surface of the nozzle guide vane. In this study, the overall cooling effectiveness values were obtained at various cooling mass flow rates experimentally for the baseline and the optimized hole arrangements proposed by the previous study [1] and compared with the adiabatic film cooling effectiveness results. The tests were conducted at mainstream exit Reynolds number based on the chord of 2.2 × 106 and the coolant mass flow rate from 5 to 10% of the mainstream. For the experimental measurements, a set of tests were conducted using an annular sector transonic turbine cascade test facility in Korea Aerospace Research Institute. To obtain the overall cooling effectiveness values on the pressure side surface, the additive manufactured nozzle guide vane made of polymer material and Inconel 718 were installed and the surface temperature was measured using a FLIR infrared camera system. Since the optimization was based on the adiabatic film cooling effectiveness, the regions with rib turbulators and film cooling holes show locally higher overall cooling effectiveness due to internal convection and conduction, which can cause non-uniform temperature distributions. Therefore, the optimization of film cooling configuration should consider the effect of the internal cooling to avoid undesirable non-uniform cooling.