An experimental study was performed to measure the heat transfer distributions and frictional losses in rotating ribbed channels with an aspect ratio of 4:1. Angled, discrete angled, V-shaped, and discrete V-shaped ribs were investigated, as well as the newly proposed W-shaped and discrete W-shaped ribs. In all cases, the ribs are placed on both the leading and trailing surfaces of the channel, and they are oriented 45° to the mainstream flow. The rib height-to-hydraulic diameter ratio (e/D) is 0.078, and the rib pitch-to-height ratio (P/e) is 10. The channel orientation with respect to the direction of rotation is 135°. The range of flow parameters includes Reynolds number (Re = 10000–40000), rotation number (Ro = 0.0–0.15), and inlet coolant-to-wall density ratio (Δρ/ρ = 0.12). Both heat transfer and pressure measurements were taken, so the overall performance of each rib configuration could be evaluated. It was determined that the W-shaped and discrete W-shaped ribs had the superior heat transfer performance in both non-rotating and rotating channels. However, these two configurations also incurred the greatest frictional losses while the discrete V-shaped and discrete angled ribs resulted in the lowest pressure drop. Based on the heat transfer enhancement and the pressure drop penalty, the discrete V-shaped ribs and the discrete W-shaped ribs exhibit the best overall thermal performance in both rotating and non-rotating channels. These configurations are followed closely by the W-shaped ribs. The angled rib configuration resulted in the worst performance of the six configurations of the present study.Copyright © 2004 by ASME

Thermal Performance of Angled, V-Shaped, and W-Shaped Rib Turbulators in Rotating Rectangular Cooling Channels (AR=4:1)

Film cooling is widely used to protect modern gas turbine blades and vanes from the ever increasing inlet temperatures. Film cooling involves a very complex turbulent flow-field, the characterization of which is necessary for reliable and economical design. Several experimental studies have focused on gas turbine blade, vane and end-wall film cooling over the past few decades. Measurements of heat transfer coefficients, film cooling effectiveness values and heat flux ratios using several different experimental methods have been reported. The emphasis of this current review is on the Pressure Sensitive Paint (PSP) mass transfer analogy to determine the film cooling effectiveness. The theoretical basis of the method is presented in detail. Important results in the open literature obtained using the PSP method are presented, discussing parametric effects of blowing ratio, momentum ratio, density ratio, hole shape, surface geometry, free-stream turbulence on flat plates, turbine blades, vanes and end-walls. The PSP method provides very high resolution contours of film cooling effectiveness, without being subject to the conduction error in high thermal gradient regions near the hole.

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Turbine blade film cooling using psp technique

Gas turbines are used extensively for aircraft propulsion,
land-based power generation, and industrial applications.
Developments in turbine cooling technology play a critical
role in increasing the thermal efficiency and power output of
advanced gas turbines. Gas turbine blades are cooled internally
by passing the coolant through several rib-enhanced
serpentine passages to remove heat conducted from the outside
surface. External cooling of turbine blades by film cooling
is achieved by injecting relatively cooler air from the
internal coolant passages out of the blade surface in order
to form a protective layer between the blade surface and
hot gas-path flow. For internal cooling, this presentation focuses
on the effect of rotation on rotor blade coolant passage
heat transfer with rib turbulators and impinging jets. The
computational flow and heat transfer results are also presented
and compared to experimental data using the RANS
method with various turbulence models such as k-e, and
second-moment closure models. This presentation includes
unsteady high free-stream turbulence effects on film cooling
performance with a discussion of detailed heat transfer coef-
ficient and film-cooling effectiveness distributions for standard
and shaped film-hole geometry using the newly developed
transient liquid crystal image method.

/pdf/recent-studies-in-turbine-blade-cooling-tcu5ygwol0.pdf

Recent Studies in Turbine Blade Cooling

Experimental investigations are performed to measure the detailed heat transfer coefficient and static pressure distributions on the squealer tip of a gas turbine blade in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a modem first stage gas turbine rotor blade with a blade tip profile of a GE-E(sup 3) aircraft gas turbine engine rotor blade. A squealer (recessed) tip with a 3.77% recess is considered here. The data on the squealer tip are also compared with a flat tip case. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Two different turbulence intensities of 6.1% and 9.7% at the cascade inlet are also considered for heat transfer measurements. Static pressure measurements are made in the mid-span and near-tip regions, as well as on the shroud surface opposite to the blade tip surface. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and an exit Reynolds number based on the axial chord of 1.1 x 10(exp 6). A transient liquid crystal technique is used to measure the heat transfer coefficients. Results show that the heat transfer coefficient on the cavity surface and rim increases with an increase in tip clearance. 'Me heat transfer coefficient on the rim is higher than the cavity surface. The cavity surface has a higher heat transfer coefficient near the leading edge region than the trailing edge region. The heat transfer coefficient on the pressure side rim and trailing edge region is higher at a higher turbulence intensity level of 9.7% over 6.1 % case. However, no significant difference in local heat transfer coefficient is observed inside the cavity and the suction side rim for the two turbulence intensities. The squealer tip blade provides a lower overall heat transfer coefficient when compared to the flat tip blade.

/pdf/heat-transfer-and-flow-on-the-squealer-tip-of-a-gas-turbine-16g4gmvqar.pdf

Heat Transfer and Flow on the Squealer Tip of a Gas Turbine Blade

The tip leakage flow characteristics for flat and squealer turbine tip geometries are studied in the von Karman Institute Isentropic Light Piston Compression Tube facility, CT-2, at different Reynolds and Mach number conditions for a fixed value of the tip gap in a non-rotating, linear cascade arrangement. To the best knowledge of the authors, these are among the very few high-speed tip flow data for the flat tip and squealer tip geometries. Oil flow visualizations and static pressure measurements on the blade tip, blade surface, and on the corresponding endwall provide insight to the structure of the two different tip flows. Aerodynamic losses are measured for the different tip arrangements, also. The squealer tip provides a significant decrease in velocity through the tip gap with respect to the flat tip blade. For the flat tip, an increase in Reynolds number causes an increase in tip velocity levels, but the squealer tip is relatively insensitive to change in Reynolds number.Copyright © 2004 by ASME

Comparison of Turbine Tip Leakage Flow for Flat Tip and Squealer Tip Geometries at High-Speed Conditions

Heat transfer coefficient and static pressure distributions are experimentally investigated on a gas turbine blade tip in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a first stage gas turbine rotor blade with a blade tip profile of a GE-E(sup 3) aircraft gas turbine engine rotor blade. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1 x 10(exp 6). The middle 3-blade has a variable tip gap clearance. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Heat transfer measurements are also made at two different turbulence intensity levels of 6.1 % and 9.7% at the cascade inlet. Static pressure measurements are made in the mid-span and the near-tip regions as well as on the shroud surface, opposite the blade tip surface. Detailed heat transfer coefficient distributions on the plane tip surface are measured using a transient liquid crystal technique. Results show various regions of high and low heat transfer coefficient on the tip surface. Tip clearance has a significant influence on local tip beat transfer coefficient distribution. Heat transfer coefficient also increases about 15-20% along the leakage flow path at higher turbulence intensity level of 9.7% over 6.1 %.

/pdf/heat-transfer-and-pressure-distributions-on-a-gas-turbine-3o1c41f0m3.pdf

Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip

The detailed heat transfer coefficient and film cooling effectiveness distributions as well as tile detailed coolant jet temperature profiles on the suction side of a gas turbine blade A,ere measured using a transient liquid crystal image method and a traversing cold wire and a traversing thermocouple probe, respectively. The blade has only one row of film holes near the gill hole portion on the suction side of the blade. The hole geometries studied include standard cylindrical holes and holes with diffuser shaped exit portion (i.e. fanshaped holes and laidback fanshaped holes). Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity was 5.3 x 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. Coolant blowing ratio was varied from 0.4 to 1.2. Results show that both expanded holes have significantly improved thermal protection over the surface downstream of the ejection location, particularly at high blowing ratios. However, the expanded hole injections induce earlier boundary layer transition to turbulence and enhance heat transfer coefficients at the latter part of the blade suction surface. In general, the unsteady wake tends to reduce film cooling effectiveness.

/pdf/effect-of-film-hole-shape-on-turbine-blade-film-cooling-16xxevw7jm.pdf

Effect of Film-Hole Shape on Turbine-Blade Film-Cooling Performance

Measurements of detailed heat transfer coefficient distributions on a turbine blade tip were performed in a large-scale, low-speed wind tunnel facility, Tests were made on a five-blade linear cascade. The low-speed wind tunnel is designed to accommodate the 107.49 deg turn of the blade cascade. The mainstream Reynolds number based on cascade exit velocity was 5.3 × 10 5 . Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. The central blade had a variable tip gap clearance, Measurements were made at three different tip gap clearances of about 1.1 percent, 2.1 percent, and 3 percent of the blade span. Static pressure distributions were measured in the blade mid-span and on the shroud surface. Detailed heat transfer coefficient distributions were measured on the blade tip surface using a transient liquid crystal technique

Detailed heat transfer coefficient distributions on a large-scale gas turbine blade tip

The film effectiveness and coolant jet temperature profile on the suction side of a gas turbine blade were measured using a transient liquid crystal and a cold-wire technique, respectively. The blade has only one row of film holes near the gill hole portion on the suction side of the blade. Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity was 5.3×10 5 . Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. Coolant blowing ratio was varied from 0.6 to 1.2. Wake Strouhal number was kept at 0 and 0.1. Results show that unsteady wake reduces film cooling effectiveness. Results also show that film injection enhances local heat transfer coefficient while the unsteady wake promotes earlier boundary-layer transition. The development of coolant jet temperature profiles could be used to explain the film cooling performance.

/pdf/unsteady-wake-effect-on-film-temperature-and-effectiveness-2hsl9a7dqs.pdf

Unsteady wake effect on film temperature and effectiveness distributions for a gas turbine blade

Detailed heat-transfer coefe cient distributions on the suction side of a gas turbine blade were measured using a transient liquid crystal image method. The blade has only one row of e lm holes near the gill-hole portion on the suction side of the blade. Studies on three different kinds of e lm-cooling hole shapes were presented. The hole geometriesstudied includestandard cylindricalholesand holeswith adiffuser-shaped exitportion (i.e., fan-shaped holes and laidback fan-shaped holes ). Tests were performed on a e ve-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on the cascade exit velocity was 5 :3 ££ 10 5 . Upstream unsteady wakes were simulated using a spoke-wheel-type wake generator. The wake Strouhal number was kept at 0 and 0.1. The coolant-to-mainstream blowing ratio was varied from 0.4 to 1.2. The results show that unsteady wake generally tends to induce earlier boundary-layer transition and enhance the surface heat-transfer coefe cients. When compared to the cylindrical hole case, both the expanded hole injections have much lower heat-transfer coefe cients over the surface downstream of the injection location, particularly at high blowing ratios. However, the expanded hole injections induce earlier boundary-layer transition to turbulence and enhance heat-transfer coefe cients at the latter part of the blade suction surface.

Shuye Teng

Papers

Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip

Effect of Film-Hole Shape on Turbine-Blade Film-Cooling Performance

Detailed heat transfer coefficient distributions on a large-scale gas turbine blade tip

Unsteady wake effect on film temperature and effectiveness distributions for a gas turbine blade

Effect of Film-Hole Shape on Turbine-Blade Heat-Transfer Coefficient Distribution