Endwall Film Cooling Using the Staggered Combustor-Turbine Gap Leakage Flow

2014 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Local Pressure ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blind Area ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the Nozzle Guide Vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the pressure side gill region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with upstream staggered slot, simulating the combustor-turbine leakage gap flow. The test cascades model was manufactured according to the GE-E3 nozzle guide vane scaled model, with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. The staggered slots were positioned upstream of the cascades to simulate the combustor-turbine gap leakage. The Pressure Sensitive Painting (PSP) technique was used to detect the film cooling effectiveness distribution on the endwall surface. Through the investigation, the following results could be achieved: 1) the film-cooling effectiveness on the endwall surface downstream the slot and along the pitchwise direction increased, with the highest parameter at Z/Pitch = 0.6; 2) a larger cooled region developed towards the suction side as the blowing ratio increased; 3) the advantage of the staggered slot was apparent on the endwall surface near the inlet area, while the coolant film was obviously weakened along the axial chord at a low blowing ratio. The influence of the staggered slots could only be detected in the downstream area of the endwall surface at the higher blowing ratio.

Turbine Endwall Film Cooling With Pressure Side Radial Holes

2013 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Local Pressure ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Nozzle Guide Vane ◽  
Cylindrical Holes ◽  
Nozzle Guide

A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the nozzle guide vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the side gill pressure region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with radial cylindrical holes on the pressure side. The test cascades model was manufactured according to the GE-E3 nozzle guide vane scaled model, with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. Four rows of staggered radial film-cooling holes were placed at the pressure side gill region. The diameter of the cylindrical holes was 1 mm and the length was 5 d, with a hole space of 6 d. The spanwise angle of the cooling holes was 35 ° and the radial angle was 90 °. Three blowing ratios were chosen as the test conditions in the experiment, M = 0.7, M = 1.0 and M = 1.3. The film-cooling effectiveness was probed using PSP (pressure sensitive painting) technology and the post processing was performed by means of a mass and heat transfer analogy. Through the investigation, the following results could be achieved: 1) the film-cooling effectiveness on the endwall surface near the pressure side gill region increased, with the highest parameter at X/Cax = 0.3; 2) a double-peak cooled region developed towards the suction side as the blowing ratio increased; 3) the advantage of the pressure side radial cooling holes was apparent on the endwall surface near the gill region, while the coolant film was obviously weakened along the axial chord at a low blowing ratio. The influence of the pressure film cooling could only be detected in the downstream area of the endwall at the higher blowing ratio.

Application of Unsteady CFD Methods to Trailing Edge Cutback Film Cooling

2014 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Navier Stokes ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film 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.

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

2017 ◽  
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 ◽  
Film 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.

Comparison of Film Cooling Performance for Different Purge Slot Configurations in a Cylindrical and State-of-the-Art Nozzle Guide Vane

2021 ◽  
Author(s):  
Christian Landfester ◽  
Gunther Müller ◽  
Robert Krewinkel ◽  
Clemens Domnick ◽  
Martin Böhle
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
State Of The Art ◽  
Guide Vane ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract This comparative study is concerned with the advances in nozzle guide vane (NGV) design developments and their influence on the film cooling performance by injecting coolant through the purge slot. An experimental study compares the film cooling effectiveness as well as the aerodynamic effects for different purge slot configurations on both a flat and an axisymmetrically contoured endwall of a NGV. While the flat endwall cascade was equipped with four cylindrical vanes, the contoured endwall cascade consisted of four modern NGVs which represent state-of-the-art high-pressure turbine design standards. Geometric variations, e.g. the purge slot width and injection angle, as well as different blowing ratios (BR) at an engine-like density ratio (DR = 1.6) were realized to investigate the real-life effect of thermal expansion, design modifications and the interaction between secondary flow and coolant. The mainstream flow parameters were set to meet real engine conditions with regard to Reynolds and Mach numbers. The Pressure Sensitive Paint (PSP) technique was used to determine the adiabatic film cooling effectiveness. Five-hole probe measurements (DR = 1.0) were performed to measure the flow field with its characteristic vortex structures as well as the loss distribution in the vane wake region. For a more profound insight into the origin and development of the secondary flows, oil dye visualizations were carried out on both endwalls. The measurement results will be discussed based on a side-by-side comparison of the distribution of film cooling effectiveness on the endwall, its area-averaged values as well as the two-dimensional distribution of total pressure losses and the secondary flow field. The results of this study show that the advances in NGV design development have had a significantly positive influence on the distribution of the coolant. This has to be attributed to lesser disturbance of the coolant propagation by secondary flow for the optimized NGV design, since the design features are intended to suppress the formation of secondary flow. In contrast to the results of the cylindrical profile, sufficient cooling can be already provided with a perpendicular injection in the case of the modern NGV. It is therefore advisable to take these effects into account when designing the film cooling system of a modern high-pressure turbine.

Nozzle Guide Vane Film Cooling Effectiveness for Radial Showerheads With Restricted Cooling Hole Surface Angles

2017 ◽  
Author(s):  
Nicholas E. Holgate ◽  
Peter T. Ireland ◽  
Kevin P. Self
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Stagnation Region ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Cooling Hole ◽  
Leading Edges ◽  
Nozzle Guide ◽  
Surface Angle

Adiabatic film cooling effectiveness measurements are made on nozzle guide vane leading edges in an engine-realistic flow environment. The tested leading edges feature radial showerheads with different spanwise distributions of hole surface angle. The showerheads blow towards the midspan, except for one model with showerhead holes orthogonal to the vane surface. The results show that low surface angle radial showerhead holes generate high effectiveness within their rows and further downstream, but neglect the stagnation region lying between the two most upstream cooling hole rows. This downstream effectiveness gain is due to both the continued surface attachment of this coolant as it progresses downstream, and its beneficial interactions with downstream cooling jets. Moderate radial showerhead surface angles cause moderate coolant jet penetration into the mainstream, which promotes near-surface mixing of the coolant with the mainstream, increasing stagnation region effectiveness. The mixing effect is enhanced by the intense turbulence generated by combustor dilution jets. High surface angles may cause the stagnation region coolant to penetrate too far for either of these gains to be realised. Considering also the presence of endwall film cooling, these effects, taken together, suggest the superiority of radial showerheads which blow towards the midspan, as against those which blow towards each endwall. Surface temperature data is acquired by a novel infrared thermography technique which permits measurement of both heat transfer coefficient and film effectiveness from a single heated test.

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

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 ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

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.

Experimental Study of the Endwall Heat Transfer of a Transonic Nozzle Guide Vane With Upstream Jet Purge Cooling: Part 2 — Effect of Combustor-NGV Misalignment

2020 ◽  
Author(s):  
Ridge A. Sibold ◽  
Shuo Mao ◽  
Wing F. Ng ◽  
Hongzhou Xu ◽  
Michael Fox
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Backward Facing Step ◽  
Cooling Effectiveness ◽  
Blowing Ratio ◽  
Nozzle Guide Vane ◽  
Forward Facing Step ◽  
Lift Off ◽  
Nozzle Guide

Abstract A misalignment between the combustor exit and the nozzle guide vane (NGV) platform commonly exists due to manufacturing tolerances and thermal transience. This study experimentally investigated the effect of the combustor-turbine misalignment on the heat transfer for an axisymmetric converging endwall with a jet purge cooling scheme at transonic conditions. Tests were conducted at engine-representative freestream exit Mach number of 0.85, inlet turbulence intensity of 16%, film cooling blowing ratio of 2.5 (design condition) and 3.5, and density ratio of 1.95. Three different step misalignments, combustor exit 4.9% span higher than turbine inlet (backward-facing), no step (baseline), and combustor exit 4.9% span lower than turbine inlet (forward-facing), were tested to demonstrate the misalignment effect on endwall Nusselt number, adiabatic film cooling effectiveness, and net heat flux reduction. A supportive numerical simulation was conducted to provide insight into the flow field. Experimental and numerical results indicated a blowing ratio of 2.5 amid a backward-facing step leads to more turbulent mixing between the coolant and mainstream due to the horseshoe vortex and step-induced cavity vortex, reducing coolant coverage and adversely affecting coolant performance. For the same blowing ratio, a forward-facing step induced more lift-off and shifted the coverage downstream and toward the pressure side, ultimately slightly enhancing the cooling performance. Nevertheless, the forward-facing step leads to a large cooling effectiveness gradient in the pitch-wise direction. At a blowing ratio of 3.5, the baseline no-step case provided superior coolant performance, however, the effect of the step was much less significant as compared to the lower blowing case. The forward-facing step amid a blowing ratio of 3.5 results in severe jet lift-off, diminishing the returns traditionally associated with increasing coolant mass flow. Based on this study, a backward-facing step should be avoided to prevent endwall burnout in the nozzle guide. Additionally, a forward-facing step should be avoided, especially amid high blowing ratios.

COMPARISON OF FILM COOLING PERFORMANCE FOR DIFFERENT PURGE SLOT CONFIGURATIONS IN A CYLINDRICAL AND STATE-OF-THE-ART NOZZLE GUIDE VANE

2021 ◽  
pp. 1-13
Author(s):  
Christian Landfester ◽  
Gunther Mueller ◽  
Robert Krewinkel ◽  
Clemens Domnick ◽  
Martin Böhle
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
State Of The Art ◽  
Guide Vane ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract This comparative study is concerned with the advances in nozzle guide vane (NGV) design developments and their influence on endwall film cooling performance by injecting coolant through the purge slot. This experimental study compares the film cooling effectiveness and the aerodynamic effects for different purge slot configurations on both a flat and an axisymmetrically contoured endwall of a NGV. While the flat endwall cascade was equipped with cylindrical vanes, the contoured endwall cascade consisted of modern NGVs which represent state-of-the-art high-pressure turbine design standards. Geometric variations, e.g. the slot width and injection angle, as well as different blowing ratios were realized. The mainstream flow parameters were set to meet real engine conditions with regard to Reynolds and Mach numbers. Pressure Sensitive Paint was used to determine the adiabatic film cooling effectiveness. Five-hole probe measurements were performed to measure the flow field in the vane wake region. For a more profound insight into the origin of the secondary flows, oil dye visualizations were carried out. The results show that the advances in NGV design have a significantly positive influence on the distribution of the coolant. This has to be attributed to lesser disturbance of the coolant propagation by secondary flow for the optimized NGV design, since the design features are intended to suppress the formation of secondary flow. It is therefore advisable to take these effects into account when designing the film cooling system of a modern high-pressure turbine.

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

2019 ◽  
Author(s):  
Luke Luehr ◽  
Ridge Sibold ◽  
Shuo Mao ◽  
Wing F. Ng ◽  
Zhigang Li ◽  
...  
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

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.

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