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.

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.

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.

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.

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.

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

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film 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.

A Converging Slot-Hole Film-Cooling Geometry—Part 2: Transonic Nozzle Guide Vane Heat Transfer and Loss

2002 ◽  
Vol 124 (3) ◽  
pp. 461-471 ◽  
Author(s):  
J. E. Sargison ◽  
S. M. Guo ◽  
M. L. G. Oldfield ◽  
G. D. Lock ◽  
A. J. Rawlinson
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Wide Band ◽  
Nozzle Guide Vane ◽  
Cooling Hole ◽  
Hole Geometry ◽  
Temperature Levels ◽  
Nozzle Guide

This paper presents the first experimental measurements on an engine representative nozzle guide vane, of a new film-cooling hole geometry, a con¯vergings¯lot-hole¯ or console. The patented console geometry is designed to improve the heat transfer and aerodynamic performance of turbine vane and rotor blade cooling systems. These experiments follow the successful validation of the console design in low-speed flat-plate tests described in Part 1 of this paper. Stereolithography was used to manufacture a resin model of a transonic, engine representative nozzle guide vane in which seven rows of previously tested fan-shaped film-cooling holes were replaced by four rows of consoles. This vane was mounted in the annular vane ring of the Oxford cold heat transfer tunnel for testing at engine Reynolds numbers, Mach numbers and coolant to mainstream momentum flux ratios using a heavy gas to simulate the correct coolant to mainstream density ratio. Heat transfer data were measured using wide-band thermochromic liquid crystals and a modified analysis technique. Both surface heat transfer coefficient and the adiabatic cooling effectiveness were derived from computer-video records of hue changes during the transient tunnel run. The cooling performance, quantified by the heat flux at engine temperature levels, of the console vane compares favourably with that of the previously tested vane with fan-shaped holes. The new console film-cooling hole geometry offers advantages to the engine designer due to a superior aerodynamic efficiency over the fan-shaped hole geometry. These efficiency measurements are demonstrated by results from midspan traverses of a four-hole pyramid probe downstream of the nozzle guide vane.

scholarly journals The Effects of Combustor Cooling Features on Nozzle Guide Vane Film Cooling Experiments

2018 ◽  
Author(s):  
Nicholas E. Holgate ◽  
Peter T. Ireland ◽  
Eduardo Romero
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Stagnation Region ◽  
Airfoil Surface ◽  
Cooling Performance ◽  
Nozzle Guide Vane ◽  
Effectiveness Measurement ◽  
Extraneous Radiation ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Recent advances in experimental methods have allowed researchers to study nozzle guide vane film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region, due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface. Also explored are the implications of endwall film cooling for infrared vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.

Sign in / Sign up

Export Citation Format

Share Document