Luzeng Zhang

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.

Heat Transfer in Rotating Multipass Rectangular Ribbed Channel With and Without a Turning Vane

This paper experimentally investigates the effect of a turning vane in hub region on heat transfer in a multipass rectangular channel with rib-roughed wall at high rotation numbers. The experimental data were taken in the second and the third passages (aspect ratio = 2:1) connected by an 180 deg U-bend. The flow was radial inward in the second passage and was radial outward after the 180 deg U-bend in the third passage. The square-edged ribs with P/e = 8, e/Dh = 0.1, and α = 45 deg were applied on the leading and trailing surfaces of the second and the third passages. Results showed that rotation increases heat transfer on the leading surface but decreases it on the trailing surface in the second passage. In the third passage, rotation decreases heat transfer on the leading surface but increases it on the trailing surface. Without a turning vane, rotation reduces heat transfer on the trailing surface and increases it on the leading surface in the hub 180 deg turn region. After adding a half-circle-shaped turning vane, heat transfer coefficients do not change in the second passage before-turn while they are different in the turn region and after-turn region in the third passage. Regional heat transfer coefficients are correlated with rotation numbers for multipass rectangular ribbed channel with and without a turning vane.

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

Turbine Platform Cooling and Blade Suction Surface Phantom Cooling From Simulated Swirl Purge Flow

This paper presents the swirl purge flow on a platform and a modeled land-based turbine rotor blade suction surface. Pressure-sensitive paint (PSP) mass transfer technique provides detailed film-cooling effectiveness distribution on the platform and phantom cooling effectiveness on the blade suction surface. Experiments were conducted in a low-speed wind tunnel facility with a five-blade linear cascade. The inlet Reynolds number based on the chord length is 250,000. Swirl purge flow is simulated by coolant injection through 50 inclined cylindrical holes ahead of the blade leading edge row. Coolant injections from cylindrical holes pass through nozzle endwall and a dolphin nose axisymmetric contour before reaching the platform and blade suction surface. Different “coolant injection angles” and “coolant injection velocity to cascade inlet velocity” result in various swirl ratios to simulate real engine conditions. Simulated swirl purge flow uses coolant injection angles of 30 deg, 45 deg, and 60 deg to produce swirl ratios of 0.4, 0.6, and 0.8, respectively. Traditional purge flow has a coolant injection angle of 90 deg to generate swirl ratio of 1. Coolant to mainstream mass flow rate (MFR) ratio is 0.5%, 1.0%, and 1.5% for all the swirl ratios. Coolant to mainstream density ratio maintains at 1.5 to match engine conditions. Most of the swirl purge and purge coolant approach the platform; however, a small amount of the coolant migrates to the blade suction surface. Swirl ratio of 0.4 has the highest relative motion between rotor and coolant and severely decreases film cooling and phantom cooling effectiveness. Higher MFR of 1% and 1.5% cases suffers from apparent decrement of the effectiveness while increasing relative motion.

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 )

Turbine Blade Surface Phantom Cooling from Upstream Nozzle Trailing-Edge Ejection

This paper presents upstream nozzle trailing edge coolant ejection on downstream uncooled blades. Pressure sensitive paint (PSP) mass transfer technique provides detailed phantom cooling effectiveness distribution on a modeled land-based turbine rotor blade surfaces. Cavity purge and tip leakage flows are excluded, and a uniform blade inlet temperature is adopted in the current study. Experiments have completed in a low speed wind tunnel facility with a five blade linear cascade. The inlet Reynolds numbers based on chord length are 100,000 and 200,000. Nozzle trailing edge coolant ejection on rotor blade is simulated by a spoked wheel-type rotating facility with 32 hollow rods equipped with coolant ejection from 128 holes per rod. Coolant to mainstream density ratio maintains at 1.5 to match engine conditions. Nozzle coolant discharge velocity to nozzle mainstream velocity ratio varies from 0.4 to 1.4. Velocity ratios from 0.4 to 0.6 are closest to typical engine conditions. Coolant to mainstream mass flow rate ratio (MFR) is from 0.67% to 2.94%. Higher phantom cooling effectiveness occurs on suction and pressure surfaces at the velocity ratio from 0.4 to 0.6 and over 1.0, respectively. Velocity ratio effect impacts on phantom cooling effectiveness distribution more than MFR effect. Most of trailing edge coolant migrates toward blade inner and outer spans than blade mid-span. Further investigation of the trailing edge coolant ejection including cavity purge and tip leakage flows is essential for understanding real applications.Copyright

Experimental Investigation of Disc Cavity Leakage Flow and Hub Endwall Contouring in a Linear Rotor Cascade

University of Minnesota M.S. thesis. April 2010. Major: Mechanical Engineering. Advisor: Terrence W. Simon. 1 computer file (PDF); xv, 163 pages, appendix A. Ill. (some col.)