Evaluation of Superposition Predictions for Showerhead Film Cooling on a Vane

2015 ◽  
Vol 137 (4) ◽  
Author(s):  
Joshua B. Anderson ◽  
James R. Winka ◽  
David G. Bogard ◽  
Michael E. Crawford
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Edge Region ◽  
Current Investigation ◽  
Engine Design ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

The leading edge of a turbine vane is subject to some of the highest temperature loading within an engine, and an accurate understanding of leading edge film coolant behavior is essential for modern engine design. Although there have been many investigations of the adiabatic effectiveness for showerhead film cooling of a vane leading edge region, there have been no previous studies in which individual rows of the showerhead were tested with the explicit intent of validating superposition models. For the current investigation, a series of adiabatic effectiveness experiments were performed with a five-row and three-row showerhead. The experiments were repeated separately with each individual row of holes active. This allowed evaluation of superposition methods on both the suction side of the vane, which was moderately convex, and the pressure side of the vane, which was mildly concave. Superposition was found to accurately predict performance on the suction side of the vane at lower momentum flux ratios, but not at higher momentum flux ratios. On the pressure side of the vane, the superposition predictions were consistently lower than measured values, with significant errors occurring at the higher momentum flux ratios. Reasons for the underprediction by superposition analysis are presented.

Evaluation of Superposition Predictions for Showerhead Film Cooling on a Vane

2014 ◽  
Author(s):  
Joshua B. Anderson ◽  
James R. Winka ◽  
David G. Bogard ◽  
Michael E. Crawford
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Edge Region ◽  
Current Investigation ◽  
Engine Design ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

The leading edge of a turbine vane is subject to some of the highest temperature loading within an engine, and an accurate understanding of leading edge film coolant behavior is essential for modern engine design. Although there have been many investigations of the adiabatic effectiveness for showerhead film cooling of a vane leading edge region, there have been no previous studies in which individual rows of the showerhead were tested with the explicit intent of validating superposition models. For the current investigation, a series of adiabatic effectiveness experiments were performed with a five-row and three-row showerhead. The experiments were repeated separately with each individual row of holes active. This allowed evaluation of superposition methods on both the suction side of the vane, which was moderately convex, and the pressure side of the vane, which was mildly concave. Superposition was found to accurately predict performance on the suction side of the vane at lower momentum flux ratios, but not at higher momentum flux ratios. On the pressure side of the vane the superposition predictions were consistently lower than measured values, with significant errors occurring at the higher momentum flux ratios. Reasons for the under-prediction by superposition analysis are presented.

Numerical Investigation of Coolant-to-Mainstream Scaling Parameters With Film Cooling on Pressure and Suction Side of a Gas Turbine Blade

2017 ◽  
Author(s):  
Lingyu Zeng ◽  
Xueying Li ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Momentum Flux ◽  
Velocity Ratio ◽  
Density Ratio ◽  
Flux Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Adiabatic Effectiveness

Most experiments of blade film cooling are conducted with density ratio lower than that of turbine conditions. In order to accurately model the performance of film cooling under a high density ratio, choosing an appropriate coolant to mainstream scaling parameter is necessary. The effect of density ratio on film cooling effectiveness on the surface of a gas turbine twisted blade is investigated from a numerical point of view. One row of film holes are arranged in the pressure side and two rows in the suction side. All the film holes are cylindrical holes with a pitch to diameter ratio P/d = 8.4. The inclined angle is 30°on the pressure side and 34° on the suction side. The steady solutions are obtained by solving Reynolds-Averaged-Navier-Stokes equations with a finite volume method. The SST turbulence model coupled with γ-θ transition model is applied for the present simulations. A film cooling experiment of a turbine vane was done to validate the turbulence model. Four different density ratios (DR) from 0.97 to 2.5 are studied. To independently vary the blowing ratio (M), momentum flux ratio (I) and velocity ratio (VR) of the coolant to the mainstream, seven conditions (M varying from 0.25 to 1.6 on the pressure side and from 0.25 to 1.4 on the suction side) are simulated for each density ratio. The results indicate that the adiabatic effectiveness increases with the increase of density ratio for a certain blowing ratio or a certain momentum flux ratio. Both on the pressure side and suction side, none of the three parameters listed above can serve as a scaling parameter independent of density ratio in the full range. The velocity ratio provides a relative better collapse of the adiabatic effectiveness than M and I for larger VRs. A new parameter describing the performance of film cooling is introduced. The new parameter is found to be scaled with VR for nearly the whole range.

Film Cooling Jet Injection Effect in Heat Transfer Coefficient Augmentation for the Pressure Side Cooling of Turbine Vane

2014 ◽  
Author(s):  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Mats Kinell ◽  
Esa Utriainen
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Time Resolved ◽  
Turbine Vane

Improving film cooling performance of turbine vanes and blades is often achieved through application of multiple arrays of cooling holes on the suction side, the showerhead region and the pressure side. This study investigates the pressure side cooling under the influence of single and multiple rows of cooling in the presence of a showerhead from a heat transfer coefficient augmentation perspective. Experiments are conducted on a prototype turbine vane working at engine representative conditions. Transient IR thermography is used to measure time-resolved surface temperature and the semi-infinite method is utilized to calculate the heat transfer coefficient on a low conductive material. Investigations are performed for cylindrical and fan-shaped holes covering blowing ratio 0.6 and 1.8 at density ratio of about unity. The freestream turbulence is approximately 5% close to the leading edge. The resulting heat transfer coefficient enhancement, the ratio of HTC with to that without film cooling, from different case scenarios have been compared to showerhead cooling only. Findings of the study highlight the importance of showerhead cooling to be used with additional row of cooling on the pressure side in order to reduce heat transfer coefficient enhancement. In addition, it is shown that extra rows of cooling will not significantly influence heat transfer augmentation, regardless of the cooling hole shape.

Predicting Adiabatic Film Effectiveness of a Turbine Vane by Two-Equation Turbulence Models

2015 ◽  
Author(s):  
Prasert Prapamonthon ◽  
Huazhao Xu ◽  
Jianhua Wang ◽  
Ge Li
Keyword(s):  
Turbulence Intensity ◽  
Film Cooling ◽  
Free Stream ◽  
Turbulence Models ◽  
Leading Edge ◽  
Suction Side ◽  
Operating Conditions ◽  
Pressure Side ◽  
Free Stream Turbulence ◽  
Turbine Vane

The thermal efficiency of gas turbine engines increases with turbine inlet temperature (TIT) directly. However, the TIT is limited by the allowable temperature of current blade materials. Film cooling technique is an effective method to maintain turbine vane working smoothly under high TIT conditions. The adiabatic film effectiveness has been widely employed to understand film cooling mechanism. Therefore, the prediction of the adiabatic effectiveness of gas turbine engines under real operating conditions is essential. The showerhead film cooled turbine vane reported by L. P. Timko (NASA CR-168289) is adopted in the present study. There are two rows of film holes on the leading edge, three rows on the pressure side, and two rows on the suction side. All holes are cylindrical, which are placed at an angle of 45 degrees to the vane surface in the span-wise direction. This numerical investigation discusses the influences of free stream turbulence intensity on the adiabatic film effectiveness in the vane leading edge region and its vicinity. Five two-equation turbulence models based on Reynolds Averaged Navier-Stokes (RANS) are employed to predict the adiabatic film effectiveness under real operating conditions at a blowing ratio (BR) of 1.41 and three free stream turbulence intensities (Tu=3.3, 10, and 20%). The adiabatic film effectiveness on the vane surface at 8, 52.5, and 89% span in an x/C range between −0.4 and 0.4 is presented. Obviously, the numerical results predicted by all five models show that on the suction side, the increasing free stream turbulence intensity can reduce film effectiveness except at 8% span. On the pressure side, the RNG k-ε, Realizable k-ε and SST k-ω models predict the same trend of the adiabatic film effectiveness, especially the RNG k-ε and SST k-ω models. Those three models predict that the locally adiabatic film effectiveness (especially near film holes) can be improved when turbulence intensity increases. However, at a span of 89% within the x/C range between −0.4 and −0.2, all k-ε models and SST k-ω model predict that the increase of turbulence intensity can reduce the adiabatic film effectiveness. In addition, the film effectiveness contours show a significant variation of film effectiveness predicted by the five turbulence models on the leading edge when turbulence intensity increases. For the near-pressure side, all models except the Standard k-ω model predict that the high turbulence intensity can reduce the film spreading from film holes dramatically.

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

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 ◽  
Film Cooling Effectiveness ◽  
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.

Bump and Trench Modifications to Film-Cooling Holes at the Vane-Endwall Junction

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first-stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases toward it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely, trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

Adiabatic Effectiveness Measurements of Endwall Film-Cooling for a First Stage Vane

2004 ◽  
Author(s):  
D. G. Knost ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Extreme Environment ◽  
Wall Region ◽  
Pressure Side ◽  
Stagnation Region ◽  
Pressure Surface ◽  
Adiabatic Effectiveness ◽  
Endwall Film Cooling

In gas turbine development, the direction has been towards higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first stage nozzle guide vane. This paper presents measurements of two endwall film-cooling hole patterns combined with cooling from a flush slot that simulates leakage flow between the combustor and turbine sections. Adiabatic effectiveness measurements showed the slot flow adequately cooled portions of the endwall. Measurements also showed two very difficult regions to cool including the leading edge and pressure side-endwall junction. As the momentum flux ratios were increased for the film-cooling jets in the stagnation region, the coolant was shown to impact the vane and wash down onto the endwall surface. Along the pressure side of the vane in the upstream portion of the passage, the jets were shown to separate from the surface rather than penetrate to the pressure surface. In the downstream portion of the passage, the jets along the pressure side of the vane were shown to impact the vane thereby eliminating any uncooled regions at the junction. The measurements were also combined with computations to show the importance of considering the trajectory of the flow in the near-wall region, which can be highly influenced by slot leakage flows.

Bump and Trench Modifications to Film-Cooling Holes at the Vane Endwall Junction

2007 ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases towards it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

The Effects of Varying the Combustor-Turbine Gap

2006 ◽  
Author(s):  
N. D. Cardwell ◽  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
High Pressure ◽  
Thermal Expansion ◽  
Mass Flow ◽  
Momentum Flux ◽  
Opposite Effect ◽  
Flux Ratio ◽  
Pressure Side ◽  
Coverage Area ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

To protect hot turbine components, cooler air is bled from the high pressure section of the compressor and routed around the combustor where it is then injected through the turbine surfaces. Some of this high pressure air also leaks through the mating gaps formed between assembled turbine components where these components experience expansions and contractions as the turbine goes through operational cycles. This study presents endwall adiabatic effectiveness levels measured using a scaled up, two-passage turbine vane cascade. The focus of this study is evaluating the effects of thermal expansion and contraction for the combustor-turbine interface. Increasing the mass flow rate for the slot leakage between the combustor and turbine showed increased local adiabatic effectiveness levels while increasing the momentum flux ratio for the slot leakage dictated the coverage area for the cooling. With the mass flow held constant, decreasing the combustor-turbine interface width caused an increase in uniformity of coolant exiting the slot, particularly across the pressure side endwall surface. Increasing the width of the interface had the opposite effect thereby reducing coolant coverage on the endwall surface.

Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
M. D. Barringer ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Secondary Flows ◽  
Pressure Side ◽  
Surface Deposition ◽  
Hot Gas ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes ◽  
Effect Of Surface

Film cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film cooling along the pressure side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the pressure side of the airfoil. Deposits having a semielliptical shape were placed along the pressure side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets toward the vane endwall junction on the pressure side. Results also indicated that there was a steady decrease in adiabatic effectiveness levels with a sequential increase in the number of rows with the deposits.

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