H. K. Moon

Effects of Large Scale High Freestream Turbulence and Exit Reynolds Number on Turbine Vane Heat Transfer in a Transonic Cascade

This paper experimentally and numerically investigates the effects of large scale high freestream turbulence intensity and exit Reynolds number on the surface heat transfer distribution of a turbine vane in a 2D linear cascade at realistic engine Mach numbers. A passive turbulence grid was used to generate a freestream turbulence level of 16% and integral length scale normalized by the vane pitch of 0.23 at the cascade inlet. The base line turbulence level and integral length scale normalized by the vane pitch at the cascade inlet were measured to be 2% and 0.05, respectively. Surface heat transfer measurements were made at the midspan of the vane using thin film gauges. Experiments were performed at exit Mach numbers of 0.55, 0.75, and 1.01, which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 910 5 , 1.0510 6 , and 1.510 6 based on a vane chord. The experimental results showed that the large scale high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the vane as compared to the low freestream turbulence case and promoted a slightly earlier boundary layer transition on the suction surface for exit Mach 0.55 and 0.75. At nominal conditions, exit Mach 0.75, average heat transfer augmentations of 52% and 25% were observed on the pressure and suction sides of the vane, respectively. An increased Reynolds number was found to induce an earlier boundary layer transition on the vane suction surface and to increase heat transfer levels on the suction and pressure surfaces. On the suction side, the boundary layer transition length was also found to be affected by increase changes in Reynolds number. The experimental results also compared well with analytical correlations and computational fluid dynamics predictions. DOI: 10.1115/1.2952381

The Effects of Freestream Turbulence, Turbulence Length Scale, and Exit Reynolds Number on Turbine Blade Heat Transfer in a Transonic Cascade

This paper experimentally investigates the effect of high freestream turbulence intensity, turbulence length scale, and exit Reynolds number on the surface heat transfer distribution of a turbine blade at realistic engine Mach numbers. Passive turbulence grids were used to generate freestream turbulence levels of 2%, 12%, and 14% at the cascade inlet. The turbulence grids produced length scales normalized by the blade pitches of 0.02, 0.26, and 0.41, respectively. Surface heat transfer measurements were made at the midspan of the blade using thin film gauges. Experiments were performed at the exit Mach numbers of 0.55, 0.78, and 1.03, which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 6 × 10 5 , 8 × 10 5 , and 11 × 10 5 , based on true chord. The experimental results showed that the high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the blade as compared with the low freestream turbulence case. At nominal conditions, exit Mach 0.78, average heat transfer augmentations of 23% and 35% were observed on the pressure side and suction side of the blade, respectively.

An Experimental Investigation of Showerhead Film Cooling Performance in a Transonic Vane Cascade at Low and High Freestream Turbulence

This experiment investigates the effects of blowing ratio on the film cooling performance of a showerhead-cooled first-stage turbine vane at low freestream turbulence (Tu = 2%) and an integral length scale normalized by vane pitch (Λx /P) of 0.05. The exit Reynolds number based on vane true chord is 1.1 × 106 . The effect of freestream turbulence at high Mach number (Mex = 0.76) and blowing ratio (BR = 0, 1.5, 2.0) is also explored by comparing results with high freestream turbulence measurements (Tu = 16%) previously performed in the same cascade. To characterize film cooling performance, platinum thin-film gauges were used to measure Nusselt number and film cooling effectiveness distributions at the midspan of the vane. Net heat flux reduction is also addressed. The primary effects of coolant injection were augmentation of Nusselt number and reduction of adiabatic wall temperature on the vane surface. In general, increasing blowing ratio showed increases in Nusselt number augmentation over the vane surface and an increase in film cooling effectiveness as well. Both Nusselt number and film cooling effectiveness trends were influenced by a strong favorable pressure gradient and resulting flow acceleration on the suction surface. Comparing low freestream turbulence results with high freestream turbulence measurements showed that large-scale, high freestream turbulence can decrease heat transfer coefficient and film cooling effectiveness downstream of injection.© 2009 ASME

Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet.Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.Copyright

NUMERICAL INVESTIGATION OF AEROTHERMAL CHARACTERISTICS OF THE BLADE TIP AND NEAR-TIP REGIONS OF A TRANSONIC TURBINE BLADE

In modern gas turbine engines, the blade tips and near-tip regions are exposed to high thermal loads caused by the tip leakage flow. The rotor blades are therefore carefully designed to achieve optimum work extraction at engine design conditions without failure. However, very often gas turbine engines operate outside these design conditions which might result in sudden rotor blade failure. Therefore, it is critical that the effect of such off-design turbine blade operation be understood to minimize the risk of failure and optimize rotor blade tip performance. In this study, the effect of varying the exit Mach number on the tip and near-tip heat transfer characteristics was numerically studied by solving the steady Reynolds Averaged Navier Stokes (RANS) equation. The study was carried out on a highly loaded flat tip rotor blade with 1% tip gap and at exit Mach numbers of Mexit = 0.85 (Reexit = 9.75 x 10 5 ) and Mexit = 1.0 (Reexit = 1.15 x 10 6 ) with high freestream turbulence (Tu = 12%). The exit Reynolds number was based on the rotor axial chord. The numerical results provided detailed insight into the flow structure and heat transfer distribution on the tip and near-tip surfaces. On the tip surface, the heat transfer was found to generally increase with exit Mach number due to high turbulence generation in the tip gap and flow reattachment. While increase in exit Mach number generally raises he heat transfer over the whole blade surface, the increase is significantly higher on the neartip surfaces affected by leakage vortex. Increase in exit Mach number was found to also induce strong flow relaminarisation on the pressure side near-tip. On the other hand, the size of the suction surface near-tip region affected by leakage vortex was insensitive to changes in exit Mach number but significant increase in local heat transfer was noted in this region. NOMENCLATURE

Heat Transfer Performance of a Showerhead and Shaped Hole Film Cooled Vane at Transonic Conditions

An experimental study was performed to measure surface Nusselt number and film cooling effectiveness on a film cooled first stage nozzle guide vane (NGV) at high freestream turbulence, using a transient thin film gauge (TFG) technique. The information presented attempts to further characterize the performance of shaped hole film cooling by taking measurements on a row of shaped holes downstream of leading edge showerhead injection on both the pressure and suction surfaces (hereafter PS and SS) of a first stage NGV. Tests were performed at engine representative Mach and Reynolds numbers and high inlet turbulence intensity and large length scale at the Virginia Tech 2D Linear Transonic Cascade facility. Three exit Mach/Reynolds number conditions were tested: 1.0/ 1,400,000, 0.85/1,150,000, and 0.60/850,000 where Reynolds number is based on exit conditions and vane chord. At Mach/Reynolds numbers of 1.0/1,450,000 and 0.85/ 1,150,000, three blowing ratio conditions were tested: BR ¼1.0, 1.5, and 2.0. At a Mach/ Reynolds number of 0.60/850,000, two blowing ratio conditions were tested: BR ¼1.5 and 2.0. All tests were performed at inlet turbulence intensity of 12% and length scale normalized by the cascade pitch of 0.28. Film cooling effectiveness and heat transfer results compared well with previously published data, showing a marked effectiveness improvement (up to 2.5� ) over the showerhead-only NGV and also agreement with published showerhead-shaped hole data. Net heat flux reduction (NHFR) was shown to increase substantially (average 2.6 � ) with the addition of shaped holes with an increase (average 1.6� ) in required coolant mass flow. Based on the heat flux data, the boundary layer transition location was shown to be within a consistent region on the suction side regardless of blowing ratio and exit Mach number. [DOI: 10.1115/1.4006666]

PERFORMANCE OF A SHOWERHEAD AND SHAPED HOLE FILM COOLED VANE AT HIGH FREESTREAM TURBULENCE AND TRANSONIC CONDITIONS

An experimental study was performed to measure surface Nusselt number and film cooling effectiveness on a film cooled first stage nozzle guide vane using a transient thin film gauge (TFG) technique. The information presented attempts to further characterize the performance of shaped hole film cooling by taking measurements on a row of shaped holes downstream of leading edge showerhead injection on both the pressure and suction surfaces (hereafter PS and SS) of a 1st stage NGV. Tests were performed at engine representative Mach and Reynolds numbers and high inlet turbulence intensity and large length scale at the Virginia Tech Transonic Cascade facility. Three exit Mach/Reynolds number conditions were tested: 1.0/1,400,000; 0.85/1,150,000; and 0.60/850,000 where Reynolds number is based on exit conditions and vane chord. At Mach/Reynolds numbers of 1.0/1,450,000 and 0.85/1,150,000 three blowing ratio conditions were tested: BR = 1.0, 1.5, and 2.0. At a Mach/Reynolds number of 0.60/850,000, two blowing ratio conditions were tested: BR = 1.5 and 2.0. All tests were performed at inlet turbulence intensity of 12% and length scale normalized by the cascade pitch of 0.28. Film cooling effectiveness and heat transfer results compared well with previously published data, showing a marked effectiveness improvement (up to 2.5x) over the showerhead only NGV and agreement with published showerhead-shaped hole data. Net heat flux reduction was shown to increase substantially (average 2.6x) with the addition of shaped holes, with an increase (average 1.6x) in required coolant mass flow. Boundary layer transition location was shown to be within a consistent region on the suction side regardless of blowing ratio and exit Mach number.Copyright

An Experimental and Numerical Study on the Aerothermal Characteristics of a Ribbed Transonic Squealer-Tip Turbine Blade With Purge Flow

Detailed heat transfer coefficient (HTC) and film cooling effectiveness (Eta) distribution on a squealer tipped first stage rotor blade were measured using an infrared (IR) technique. The blade tip design, obtained from a Solar Turbines Inc. gas turbine, consisted of double purge hole exits and four ribs within the squealer cavity, with a bleeder exit port on the pressure side close to the trailing edge. The tests were carried out in a transient linear transonic wind tunnel facility under land-based engine representative Mach/Reynolds number. Measurements were taken at an inlet turbulent intensity of Tu =12%, with exit Mach numbers of 0.85 (Reexit=9.75x10 5 ) and 1.0 (Reexit = 1.15x10 6 ) with the Reynolds number based on the blade axial chord and the cascade exit velocity. The tip clearance was fixed at 1% (based on engine blade span) with a purge flow blowing ratio BR = 1.0. At each test condition, an accompanying numerical study was performed using Reynolds Averaged Navier Stokes (RANS) equations solver ANSYS Fluent to further understand the tip flow characteristics. The results showed that the tip purge flow has a blocking effect on the leakage flow path. Furthermore, the ribs significantly altered the flow (and consequently heat transfer) characteristics within the squealer tip cavity resulting in a significant reduction in film cooling effectiveness. This was attributed to increased coolant-leakage flow mixing due to increased recirculation within the squealer cavity. Overall, the peak heat transfer coefficient on the cavity floor increased with exit Mach/Reynolds number. NOMENCLATURE BR Averaged blowing Ratio (BR = ρcUc /ρ∞ U∞,avg )

A Transient Infrared Technique for Measuring Surface and Endwall Heat Transfer in a Transonic Turbine Cascade

This paper describes a method for obtaining surface and endwall heat transfer in an uncooled transonic cascade facility using infrared thermography measurements. Midspan heat transfer coefficient results are first presented for an engine representative first stage nozzle guide vane at exit Mach number of 0.77, Reynolds number of 1.05×106 and freestream turbulence intensity of 16%. The results obtained from infrared thermography are compared with previously published results using thin film gauges in the same facility on the same geometry. There is generally good agreement between the two measurement techniques in both trend and overall level of heat transfer coefficient over the vane surface. Stanton number contours are then presented for a blade endwall at exit Mach number of 0.88, Reynolds number of 1.70×106 and freestream turbulence intensity of 8%. Infrared thermography results are qualitatively compared with results from a published work obtained with liquid crystals at similar flow conditions. Results are qualitatively in agreement.Copyright

The Effects of Freestream Turbulence, Turbulence Length Scale and Exit Reynolds Number on Turbine Blade Heat Transfer in a Transonic Cascade

This paper experimentally investigates the effect of high freestream turbulence intensity, turbulence length scale, and exit Reynolds number on the surface heat transfer distribution of a turbine blade at realistic engine Mach numbers. Passive turbulence grids were used to generate freestream turbulence levels of 2%, 12%, and 14% at the cascade inlet. The turbulence grids produced length scales normalized by the blade pitch of 0.02, 0.26, and 0.41, respectively. Surface heat transfer measurements were made at the midspan of the blade using thin film gauges. Experiments were performed at exit Mach numbers of 0.55, 0.78 and 1.03 which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 6 × 105 , 8 × 105 , and 11 × 105 , based on true chord. The experimental results showed that the high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the blade as compared to the low freestream turbulence case. At nominal conditions, exit Mach 0.78, average heat transfer augmentations of 23% and 35% were observed on the pressure side and suction side of the blade, respectively.© 2007 ASME