Sweeping Jet Film Cooling at High Blowing Ratio on a Turbine Vane

2019 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Boundary Layer ◽  
Film Cooling ◽  
Thermal Field ◽  
Guide Vane ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Blowing Ratio ◽  
Shaped Hole ◽  
D 12 ◽  
Nozzle Guide

Abstract Experimental and numerical investigations were conducted to study the effects of high blowing ratios and high freestream turbulence on sweeping jet film cooling. Experiments were conducted on a nozzle guide vane suction surface in a low-speed linear cascade. Experiments were performed at blowing ratios of 0.5–3.5 and freestream turbulence of 0.6% and 14.3%. Infrared thermography was used to estimate the adiabatic cooling effectiveness. Thermal field and boundary layer measurement were conducted at a cross-plane (x/D = 12) downstream of the hole exit. Results were compared with a baseline 777-shaped hole and showed that sweeping jet hole has a better cooling performance at high blowing ratios. The Thermal field data revealed that the coolant separates from the surface at high blowing ratios for the 777-shaped hole while the coolant remains attached for the sweeping jet hole. Boundary layer measurement further confirmed that due to the sweeping action of the jet, the jet momentum of the sweeping jet hole is much lower than that of a 777-shaped hole. Thus the coolant remains closer to the wall even at high blowing ratios. Large Eddy Simulations (LES) were performed for both sweeping jet and the 777-shaped hole to evaluate the interaction between the coolant and the freestream at the near hole regions. Results showed that 777-shaped hole has a strong jetting action at high blowing ratio that originates inside the hole breakout edges thus causing the jet to blow off from the surface. In contrast, the sweeping jet hole does not show this behavior due to its internal geometry and the sweeping action of the jet.

Sweeping Jet Film Cooling at High Blowing Ratio on a Turbine Vane

2020 ◽  
Vol 142 (12) ◽  
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Boundary Layer ◽  
Film Cooling ◽  
Thermal Field ◽  
Guide Vane ◽  
Freestream Turbulence ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Blowing Ratio ◽  
Shaped Hole ◽  
D 12

Abstract Experimental and numerical investigations were performed to study the effects of high blowing ratios and high freestream turbulence on sweeping jet film cooling. Experiments were conducted on a nozzle guide vane suction surface in a low-speed linear cascade at a range of blowing ratios of 0.5–3.5 and freestream turbulence of 0.6% and 14.3%. Infrared thermography was used to estimate the adiabatic cooling effectiveness. Thermal field and boundary layer measurements were conducted at a cross-plane at x/D = 12 downstream of the hole exit. Results were compared with a baseline 777-shaped hole and showed that the sweeping jet hole has improved cooling effectiveness at high blowing ratios (M > 1). The thermal field data revealed that the coolant separates from the surface at high blowing ratios for the 777-shaped hole while the coolant remains attached for the sweeping jet hole. Boundary layer measurements further confirmed that due to the sweeping motion of the jet, the effective jet momentum of the sweeping jet hole remains much lower than that of a 777-shaped hole. Thus, the coolant remains closer to the wall even at high blowing ratios. Large Eddy simulations (LES) were performed for both the sweeping jet and 777-shaped hole to evaluate the interaction between the coolant jet and the freestream in the near hole regions. Results showed that 777-shaped hole has a strong jetting at high blowing ratio that originates inside the hole breakout edges thus causing the jet to blow-off from the surface. In contrast, the sweeping jet does not show this behavior due to its internal geometry and flapping motion of the jet.

The Influence of the Boundary Layer State and Reynolds Number on Film Cooling and Heat Transfer on a Cooled Nozzle Guide Vane

2000 ◽  
Author(s):  
Hans Reiss ◽  
Albin Bölcs
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Flow Conditions ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Film cooling and heat transfer measurements were carried out on a cooled nozzle guide vane in a linear cascade, using a transient liquid crystal technique. Three flow conditions were realized: the nominal operating condition of the vane with an exit Reynolds number of 1.47e6, as well as two lower flow conditions: Re2L = 1.0e6 and 7.5e5. The vane model was equipped with a single row of inclined round film cooling holes with compound angle orientation on the suction side. Blowing ratios ranging form 0.3 to 1.5 were covered, all using foreign gas injection (CO2) yielding an engine-representative density ratio of 1.6. Two distinct states of the incoming boundary layer onto the injection station were compared, an undisturbed laminar boundary layer as it forms naturally on the suction side, and a fully turbulent boundary layer which was triggered with a trip wire upstream of injection. The aerodynamic flow field is characterized in terms of profile Mach number distribution, and the associated heat transfer coefficients around the uncooled airfoil are presented. Both detailed and spanwise averaged results of film cooling effectiveness and heat transfer coefficients are shown on the suction side, which indicate considerable influence of the state of the incoming boundary layer on the performance of a film cooling row. The influence of the mainstream flow condition on the film cooling behavior at constant blowing ratio is discussed for three chosen injection regimes.

Experimental Investigation of Innovative Cooling Schemes on an Additively Manufactured Engine Scale Turbine Nozzle Guide Vane

2020 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Jet Impingement ◽  
Development Effort ◽  
Suction Surface ◽  
Trailing Edge ◽  
Pin Fin ◽  
Nozzle Guide ◽  
Film Cooling Holes

Abstract This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

Adiabatic Effectiveness Measurements for a Baseline Shaped Film Cooling Hole

2014 ◽  
Author(s):  
Robert P. Schroeder ◽  
Karen A. Thole
Keyword(s):  
Film Cooling ◽  
Freestream Turbulence ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Cylindrical Holes ◽  
Shaped Hole ◽  
Adiabatic Effectiveness ◽  
Improved Performance ◽  
Density Ratios ◽  
Film Cooling Holes

Film cooling on airfoils is a crucial cooling method as the gas turbine industry seeks higher turbine inlet temperatures. Shaped film cooling holes are widely used in many designs given the improved performance over that of cylindrical holes. Although there have been numerous studies of shaped holes, there is no established baseline shaped hole to which new cooling hole designs can be compared. The goal of this study is to offer the community a shaped hole design, representative of proprietary and open literature holes that serves as a baseline for comparison purposes. The baseline shaped cooling hole design includes the following features: hole inclination angle of 30° with a 7° expansion in the forward and lateral directions; hole length of 6 diameters; hole exit-to-inlet area ratio of 2.5; and lateral hole spacing of 6 diameters. Adiabatic effectiveness was measured with this new shaped hole and was found to peak near a blowing ratio of 1.5 at density ratios of 1.2 and 1.5 as well as at both low and moderate freestream turbulence of 5%. Reductions in area-averaged effectiveness due to freestream turbulence at low blowing ratios were as high as 10%.

Experimental Investigation of Innovative Cooling Schemes on an Additively Manufactured Engine Scale Turbine Nozzle Guide Vane

2021 ◽  
pp. 1-39
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Jet Impingement ◽  
Development Effort ◽  
Suction Surface ◽  
Trailing Edge ◽  
Pin Fin ◽  
Nozzle Guide ◽  
Film Cooling Holes

Abstract This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375 K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

The Aerodynamic Development of a Highly Loaded Nozzle Guide Vane

1986 ◽  
Vol 108 (2) ◽  
pp. 261-268 ◽  
Author(s):  
N. C. Baines ◽  
M. L. G. Oldfield ◽  
J. P. Simons ◽  
J. M. Wright
Keyword(s):  
Boundary Layer ◽  
High Speed ◽  
Guide Vane ◽  
Layer Growth ◽  
Suction Surface ◽  
Pressure Surface ◽  
Unsteady Separation ◽  
Surface Contour ◽  
Boundary Layer Development ◽  
Nozzle Guide

A series of high-pressure turbine nozzle guide vanes has been designed for progressively increasing blade loading and reduction in blade solidity without additional loss penalty. Early members of the series achieved this by changes to the suction surface contour, but for the latest design the pressure surface contour was extensively modified to reduce the velocities on this surface substantially. Cascade testing revealed that this vane had a higher loss than its predecessor, and this appears to be largely due to a long region of boundary layer growth on the suction surface and possibly also an unsteady separation. These tests demonstrated the value of a flattened pitot tube held against the blade surface in determining the boundary layer state. By using a pitot probe of only modest frequency response (of order 100 Hz) it was possible to observe significant qualitative differences in the raw signals from laminar, transitional and turbulent boundary layers, which have previously been observed only with much higher frequency instruments. The test results include a comparison of boundary layer measurements on the same cascade test section in two different high-speed wind tunnels. This comparison suggests that freestream turbulence can have a large effect on boundary layer development and growth.

Suction and Pressure Side Film Cooling Influence on Vane Aero Performance in a Transonic Annular Cascade

2013 ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Pressure Side ◽  
Aerodynamic Loss ◽  
Cooling Flow ◽  
Blowing Ratio ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Cooling Air

An experimental study on a film cooled nozzle guide vane has been conducted in a transonic annular sector to observe the influence of suction and pressure side film cooling on aerodynamic performance. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at an exit reference Mach number of 0.89. The aerodynamic results using a five hole miniature probe are quantified and compared with the baseline case which is uncooled. Results lead to a conclusion that the aerodynamic loss is influenced substantially with the change of the cooling flow rate regardless the positions of the cooling rows. The aerodynamic loss is very sensitive to the blowing ratio and a value of blowing ratio higher than one leads to a considerable higher loss penalty. The suction side film cooling has larger influence on the aerodynamic loss compared to the pressure side film cooling. Pitch-averaged exit flow angles around midspan remain unaffected at moderate blowing ratio. The secondary loss decreases (greater decrease in the tip region compared to the hub region) with inserting cooling air for all cases compared to the uncooled case.

Effect of In-Hole Roughness on Film Cooling From a Shaped Hole

2016 ◽  
Author(s):  
Robert P. Schroeder ◽  
Karen A. Thole
Keyword(s):  
Film Cooling ◽  
Thermal Field ◽  
Core Flow ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Flow Through ◽  
Shaped Hole ◽  
Adiabatic Effectiveness ◽  
Cool Gas ◽  
Interior Walls

While much is known about how macro-geometry of shaped holes affects their ability to successfully cool gas turbine components, little is known about the influence of surface roughness on cooling hole interior walls. For this study a baseline shaped hole was tested with various configurations of in-hole roughness. Adiabatic effectiveness measurements at blowing ratios up to three showed that in-hole roughness caused decreased adiabatic effectiveness relative to smooth holes. Decreases in area-averaged effectiveness grew more severe with larger roughness size and with higher blowing ratios for a given roughness. Decreases of more than 60% were measured at a blowing ratio of three for the largest roughness values. Thermal field and flowfield measurements showed that in-hole roughness caused increased velocity of core flow through the hole, which increased the jet penetration height and turbulence intensity resulting in increased mixing between coolant and the mainstream. Effectiveness reductions due to roughness were also observed when roughness was isolated to only the diffused outlet of holes, and when the mainstream was highly turbulent.

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.

Experimental and CFD Studies of NGV Endwall Cooling

2013 ◽  
Author(s):  
Mitra Thomas ◽  
Benjamin Kirollos ◽  
Dougal Jackson ◽  
Thomas Povey
Keyword(s):  
Boundary Layer ◽  
Film Cooling ◽  
Large Scale ◽  
Guide Vane ◽  
Complex Nature ◽  
Outlet Temperature ◽  
Double Row ◽  
The Difference ◽  
Layer Flow ◽  
Nozzle Guide

For engines operating at high turbine entry temperatures it is increasingly important to cool the high pressure nozzle guide vane (HP NGV) endwalls. This is particularly so for low NOx combustors operating with flatter outlet temperature distributions. Double-row arrangements of film/ballistic cooling holes upstream of the NGV passage have been employed in production engines. Optimisation of such systems is non-trivial, however, due to the complex nature of the flow in the endwall region. Previous studies have reported that strong cross passage pressure gradients lead to migration of coolant flow and boundary layer flow within the passage. In addition the vane potential field effects lead to non-uniform blowing ratios for holes upstream of the vanes. It has also been reported that inlet total pressure and turbulence profiles have a significant effect on the development of the film cooling layer. In this study, endwall film cooling flows are studied experimentally in a large-scale low-speed cascade tunnel with engine-realistic combustor geometry and turbulence profiles. At very low blowing ratios mild cross-passage migration effects are observed. At higher blowing ratios more realistic of the engine situation no cross-passage migration is observed. This finding is somewhat contrary to the classical view of endwall secondary flow, which is presented as significant at the scale of the vane passage by several authors. The difference arises in part because of the thinning of the boundary layer due to strong acceleration in the vane inlet contraction. The findings are further supported by CFD simulations. Methods of improving conventional double-row systems to offer improved cooling of the endwall are also discussed.

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