Effect of grooved cooling passage near the trailing edge region for HP stage gas turbine blade -A numerical investigation

This paper numerically investigated the impact of the holes and their location on the flow and tip internal heat transfer in a U-bend channel (aspect ratio = 1:2), which is applicable to the cooling passage with dirt purge holes in the mid-chord region of a typical gas turbine blade. Six different tip ejection configurations are calculated at Reynolds numbers from 25,000 to 200,000. The detailed three-dimensional flow and heat transfer over the tip wall are presented, and the overall thermal performances are evaluated. The topological methodology, which is first applied to the flow analysis in an internal cooling passage of the blade, is used to explore the mechanisms of heat transfer enhancement on the tip wall. This study concludes that the production of the counter-rotating vortex pair in the bend region provides a strong shear force and then increases the local heat transfer. The side-mounted single hole and center-mounted double holes can further enhance tip heat transfer, which is attributed to the enhanced shear effect and disturbed low-energy fluid. The overall thermal performance of the optimum hole location is a factor of 1.13 higher than that of the smooth tip. However, if double holes are placed on the upstream of a tip wall, the tip surface cannot be well protected. The results of this study are useful for understanding the mechanism of heat transfer enhancement in a realistic gas turbine blade and for efficient designing of blade tips for engine service.

Download Full-text

Design Modification of a Heavy Duty Industrial Gas Turbine Blade for Life Improvement

Volume 7: Structures and Dynamics, Parts A and B ◽

10.1115/gt2012-69963 ◽

2012 ◽

Author(s):

Chris Hutchison ◽

Anthony Chan ◽

Dan Stankiewicz

Keyword(s):

Crack Initiation ◽

Gas Turbine ◽

Turbine Blade ◽

Low Cycle Fatigue ◽

Inlet Temperature ◽

Gas Turbine Blade ◽

Heavy Duty ◽

Trailing Edge ◽

Design Modification ◽

Industrial Gas Turbine

Cracking at the trailing edge of a heavy duty industrial gas turbine blade has been observed on a number of serviced parts. The cracking usually occurs within 1.0″ of the platform. The trailing edge (TE) cracks have been found to propagate through the airfoil, leading to airfoil separation and severe engine damage. Liburdi Turbine Services has undertaken an independent metallurgical and stress analysis of the blade to determine the cause of the cracking. This paper covers the stress and low cycle fatigue (LCF) analysis of a platform undercut modification designed to mitigate crack initiation and thus increase part life. A finite element model of the blade was developed. Thermal loading was applied from a conjugate heat and mass transfer analysis between the blade, gas path flow and internal cooling flow. Base load conditions were used at turbine inlet temperature 2482°F. Results showed that the peak stress was present in the TE cooling slot corner, and was large enough to cause local yielding and LCF. The geometry of the modification was shown to strongly influence stress in the TE airfoil region and in the undercut region. Thus a balance was found to provide sufficiently low stresses in both regions and still be practical for machining. The modification was found to decrease stress in the TE cooling slot by a factor of 0.71 relative to that of the current OEM design, and increase life by 1.79 times. A viable modification has been demonstrated to extend blade life by reducing local stress and thus mitigating crack initiation at the airfoil TE.

Download Full-text

Effect of Rotation on Heat Transfer in AR = 2:1 and AR = 4:1 Channels Connected by a Series of Crossover Jets

Journal of Turbomachinery ◽

10.1115/1.4053237 ◽

2021 ◽

pp. 1-19

Author(s):

Srivatsan Madhavan ◽

Prashant Singh ◽

Srinath V. Ekkad

Keyword(s):

Heat Transfer ◽

Gas Turbine ◽

Turbine Blade ◽

Flow Direction ◽

Gas Turbine Blade ◽

Trailing Edge ◽

Cooling Flow ◽

Liquid Crystal Thermography ◽

Pin Fins ◽

Rib Turbulators

Abstract Detailed heat transfer measurements using transient liquid crystal thermography were performed on a novel cooling design covering the mid-chord and trailing edge region of a typical gas turbine blade under rotation. The test section comprised of two channels with aspect ratio (AR) of 2:1 and 4:1, where the coolant was fed into the AR = 2:1 channel. Rib turbulators with a pitch-to-rib height ratio (p/e) of 10 and rib height-to-channel hydraulic diameter ratio (e/Dh) of 0.075 were placed in the AR = 2:1 channel at 60° relative to flow direction. The coolant after entering this section was routed to the AR = 4:1 section through a set of crossover jets. The 4:1 section had a realistic trapezoidal shape that mimics the trailing edge of an actual gas turbine blade. The pin fins were arranged in a staggered array with a center-to-center spacing of 2.5 times pin diameter. The trailing edge section consisted of radial and cutback exit holes for flow exit. Experiments were performed for Reynolds number of 20,000 at Rotation numbers (Ro) of 0, 0.1 and 0.14. The channel averaged heat transfer coefficient on trailing side was ~28% (AR = 2:1) and ~7.6% (AR = 4:1) higher than the leading side for Ro = 0.1. It is shown that the combination of crossover jets and pin-fins can be an effective method for cooling wedge shaped trailing edge channels over axial cooling flow designs.

Download Full-text