Jeremy Clyde Bailey

Effect of Squealer Cavity Depth and Oxidation on Turbine Blade Tip Heat Transfer

An experimental study has been performed to investigate the effect of squealer cavity depth on the detailed distribution of convective heat transfer coefficients of a turbine blade tip surface. This paper presents full surface information on heat transfer coefficients within a blade cascade which develops an appropriate pressure distribution about an airfoil blade tip and shroud model. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.21, exit Mach number of 0.74, pressure ratio of 1.41, Reynolds number of 2.8•106, and total turning of about 100 degrees. The cascade inlet turbulence intensity level is 9%. Tip surface heat transfer coefficient distributions are first shown for a flat, square-edge tip with a clearance gap of 2.03 mm. Heat transfer distributions are then shown for full-perimeter squealer tip cavities having the same clearance gap above the squealer rim, and clearance-to-cavity depth ratios from 0.67 to 2. Regionally averaged heat transfer coefficients are analyzed to discern a relationship between tip heat transfer and cavity depth. Further tests demonstrate the effect of partial squealer rim oxidation, or material loss, on the surface heat transfer distributions.Copyright

Film Cooling Discharge Coefficient Measurements in a Turbulated Passage With Internal Crossflow

Gas turbine blades utilize internal geometry such as turbulator ribs for improved cooling. In some designs it may be desirable to benefit from the internal cooling enhancement of ribs as well as external film cooling. An experimental study has been performed to investigate the effect of turbulator rib placement on film hole discharge coefficient. In the study, a square passage having a hydraulic diameter of 1.27 cm is used to feed a single angled film jet. The film hole angle to the surface is 30 deg and the hole length-to-diameter ratio is 4. Turbulators were placed in one of three positions: upstream of film hole inlet, downstream of film hole inlet, and with the film hole inlet centered between turbulators. For each case 90 deg turbulators with a passage blockage of 15 percent and a pitch to height ratio of 10 were used. Tests were run varying film hole-to-crossflow orientation as 30, 90, and 180 deg, pressure ratio from 1.02 to 1.8, and channel crossflow velocity from Mach 0 to 0.3. Film hole flow is captured in a static plenum with no external crossflow. Experimental results of film discharge coefficients for the turbulated cases and for a baseline smooth passage are presented. Alignment of the film hole entry with respect to the turbulator is shown to have a substantial effect on the resulting discharge coefficients. Depending on the relative alignment and flow direction discharge coefficients can be increased or decreased 5-20 percent from the nonturbulated case, and in the worst instance experience a decrease of as much as 50 percent.

Heat Transfer and Friction in Channels With Very High Blockage 45° Staggered Turbulators

Heat transfer and friction coefficients have been measured within a rectangular passage of aspect ratio 0.4 containing 45-degree staggered turbulators of very high blockage. Using a constant pitch-to-height ratio of 10 for all geometries, turbulator height-to-channel hydraulic diameter ratios from 0.193 to 0.333 were investigated. This range of e/D creates actual channel blockage ratios e/H from 0.275 to 0.475, presenting significant flow area restrictions. A liquid crystal test technique is used to obtain both detailed heat transfer behavior on the surfaces between turbulators, as well as averaged fully developed heat transfer coefficients. Reynolds numbers from 20000 to 100000 were tested. Nusselt number enhancements of up to 3.6 were obtained over that of a smooth channel, with friction coefficient enhancements of as much as 65. In contrast to low-blockage turbulated channels, the 45-degree turbulated Nu is found to be lower than that at 90-degree orientation, given very similar e/D and e/H values.Copyright

Experimental and Numerical Study of Heat Transfer in a Gas Turbine Combustor Liner

Experiments and numerical simulations were conducted to understand the heat transfer characteristics of a stationary gas turbine combustor liner cooled by impingement jets and cross flow between the liner and sleeve. Heat transfer was also aided by trip-strip turbulators on the outside of the liner and in the flowsleeve downstream of the jets. The study was aimed at enhancing heat transfer and prolonging the life of the combustor liner components. The combustor liner and flow sleeve were simulated using a flat plate rig. The geometry has been scaled from actual combustion geometry except for the curvature. The jet Reynolds number and the mass-velocity ratios between the jet and cross flow in the rig were matched with the corresponding combustor conditions. A steady state liquid crystal technique was used to measure spatially resolved heat transfer coefficients for the geometric and flow conditions mentioned above. The heat transfer was measured both in the impingement region as well as over the turbulators. A numerical model of the combustor test rig was created that included the impingement holes and the turbulators. Using CFD, the flow distribution within the flow sleeve and the heat transfer coefficients on the liner were both predicted. Calculations were made by varying the turbulence models, numerical schemes, and the geometrical mesh. The results obtained were compared to the experimental data and recommendations have been made with regard to the best modeling approach for such liner-flow sleeve configurations.Copyright

Experimental Investigation of Aerodynamic Losses of Different Shapes of a Shrouded Blade Tip Section

An experimental investigation was conducted to study the effects on aerodynamic losses of different tip shroud shapes of a shrouded turbine blade. Pressures were measured on the airfoil surface near the tip and a plane downstream of the exit plane in a three-airfoil stationary cascade. The instrumented center airfoil and the two slave airfoils modeled the aerodynamic tip section of a blade and have the capability to vary tip clearance. The experiments were run at tip-clearances varying from 0.25% to 1.67% and at an exit Reynolds number of 1.25 × 106 and Mach Number of 0.95. The paper presents the influence of three tip-shroud shapes and five different tip-clearances on the aerodynamic losses.Copyright

Experimental and Computational Investigation of Heat Transfer Effectiveness and Pressure Distribution of a Shrouded Blade Tip Section

An experimental and computational investigation was conducted to study the detailed distribution of heat transfer effectiveness and pressure on an attached tip-shroud of a turbine blade. Temperatures and pressures were measured on the airfoil-side and gap-side surfaces of the shrouded tip in a three-airfoil stationary cascade. The instrumented center airfoil and the two slave airfoils modeled the aerodynamic tip section of a blade and have the capability to vary tip clearance. The experiments were run at gaps varying of 0.25% to 1.67% of blade span and at an airfoil exit Reynolds number of 1.26×106 and Mach number of 0.95. The effect of coolant flow through the radial-cooled airfoil was also studied. The experimental results are compared with a computational model using the commercially available code, CFX. This unique study presents the influence of gap and coolant flow on the pressure distribution and heat transfer effectiveness of an attached tip-shroud surface.Copyright

Turbine Blade Platform Impingement Cooling With Effects of Film Extraction and Roughness

With the movement of combustion systems to flatter exit temperature profiles for the purpose of reduced emissions, vane endwalls and blade platforms are experiencing higher gas temperatures than previously. While vane endwalls may be cooled by conventional methods extended from the airfoil techniques, blade platforms require simple but robust cooling methods that are amenable to the high rotational loads. This study presents a platform cooling solution that utilizes a single impingement jet directed to the underside of the platform from the forward shank face. Detailed local distributions of heat transfer coefficients are presented for the cooled platform region formed by the cavity between the pressure side and suction side of two adjacent blade castings. Fundamental understanding is provided for impingement into this confined cavity region, with two differing jet positions. The effects of jet impingement distance to the surface from Z/D of 1.5 to 8, and jet Reynolds number from 65,000 to 155,000 are investigated. Furthermore, to augment the cooled surface heat transfer coefficients, surface roughness is applied by a patented method appropriate to actual turbine hardware. The single cooling jet produces a platform cooling distribution that may be treated as two distinct regions. The primary impingement region is seen to behave in line with expectations from prior literature, with some cavity geometry effects. Impingement heat transfer coefficients are increased by as much as 50% through the use of an applied randomly roughened surface that models an actual braze alloy method. This roughness benefit is however lost if the impingement cooling jet is not relatively close to the platform surface. The non-impingement heat transfer regions of the platform are affected only by the total cooling flow rate supplied by the impingement jet.© 2005 ASME

In-Wall Network (Mesh) Cooling Augmentation of Gas Turbine Airfoils

The present study investigates a new class of in-wall mesh cooling techniques intended to produce significant wetted surface area heat transfer enhancement products, hA Awetted , with low to moderate pressure losses. Cooling networks (meshes) are presented using round pins and rounded diamond shaped pins with H/Dp ratios of 0.2 and S/Dp ratios of 1.5, as well as less dense rounded diamond pins of smaller diameter with H/Dp of 0.3 and S/Dp of 2.14. These geometries differ substantially from conventional pin fin arrays in which H/Dp ≥ 1 and S/Dp ratios are typically about 2.5. The objective of mesh cooling is to provide highly effective cooling inside component walls while maintaining very thin outer protective skin thickness. Special attention is paid to the combination of techniques including pin meshes, turbulators, and concavity arrays that can provide multiple design solutions for heat transfer and pressure loss goals. Average channel wetted-area heat transfer capabilities exceeding 3 times that of smooth channels have been demonstrated using actual surface area enhancements of no more than 20%. This cooling capability increase is also realized with a large decrease in channel material solidity, up to 30%, compared to conventional pin fin arrays. Friction factor enhancements in the range of 15 to 25 times that of a smooth channel turbulent flow accompany this performance. This investigation represents the first of its kind to demonstrate feasible in-wall cooling methods that may be realized in practice for investment cast turbine blades.Copyright

Film Cooling Discharge Coefficient Measurments in a Turbulated Passage With Internal Cross Flow

Gas turbine blades utilize internal geometry such as turbulator ribs for improved cooling. In some designs it may be desirable to benefit from the internal cooling enhancement of ribs as well as external film cooling. An experimental study has been performed to investigate the effect of turbulator rib placement on film hole discharge coefficient. In the study a square passage having a hydraulic diameter of 1.27 cm is used to feed a single angled film jet. The film hole angle to the surface is 30° and the hole length-to-diameter ratio is 4. Turbulators were placed in one of three positions: upstream of film hole inlet, downstream of film hole inlet, and with the film hole inlet centered between turbulators. For each case 90° turbulators with a passage blockage of 15% and a pitch to height ratio of 10 were used. Tests were run varying film hole-to-cross flow orientation as 30°, 90°, and 180°, pressure ratio from 1.02 to 1.8, and channel cross flow velocity from Mach 0 to 0.3. Film hole flow is captured in a static plenum with no external cross flow. Experimental results of film discharge coefficients for the turbulated cases and for a baseline smooth passage are presented. Alignment of the film hole entry with respect to the turbulator is shown to have a substantial effect on the resulting discharge coefficients. Depending on the relative alignment and flow direction, discharge coefficients can be increased or decreased 5 to 20% from the non-turbulated case, and in the worst instance experience a decrease of as much as 50%.© 2001 ASME