Gas Turbine Blade Cooling Structure with Dimples and Turbulators
Find Innovative SolutionsGenerate Solutions
Solution Overview
Problem
Conventional gas turbine blade cooling structures face a trade-off between increasing the heat transfer coefficient and reducing pressure loss, as features like pins or dimples enhance heat transfer but increase pressure loss, potentially declining cooling performance.
Innovation Solution
The proposed blade cooling structure incorporates dimples in the downstream region of secondary flows caused by turbulators, creating vortices to enhance heat transfer without increasing pressure loss by optimizing the arrangement of turbulators and dimples.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If pins or dimples are provided in the cooling passage to increase the heat transfer coefficient, then heat transfer is enhanced, but the pressure loss of the cooling medium increases
Solution Approach 1:
The patent applies local quality by providing dimples only in specific regions of the cooling passage wall rather than uniformly across the entire surface. The dimples are strategically positioned to enhance heat transfer in areas where it is most needed, while leaving other regions smooth to minimize pressure loss. This localized modification allows the system to improve heat transfer coefficient without incurring the penalty of widespread surface roughness.
Solution Approach 2:
The patent implements partial action by introducing dimples only in certain portions of the cooling passage rather than covering the entire surface. This partial modification is sufficient to achieve the desired heat transfer enhancement while avoiding the excessive pressure loss that would result from complete surface roughening. The dimples are provided in a controlled manner to balance heat transfer improvement with pressure loss mitigation.
2Power
If the temperature of the combustion gas is raised to achieve high output and high efficiency, then output and efficiency increase, but the strength of the stationary blades and moving blades decreases
Solution Approach 1:
The patent applies segmentation by dividing the blade structure into functional zones: an internal cooling passage system for thermal management and an external aerodynamic structure for mechanical strength. The cooling passage is further segmented into regions with and without dimples, allowing different thermal management strategies in different areas. This segmentation enables the blade to withstand higher combustion gas temperatures while maintaining structural integrity.
Solution Approach 2:
The patent uses cooling medium (air or steam) as an intermediary to transfer heat away from the blade surfaces. The cooling passage acts as an intermediary channel between the hot combustion gas environment and the cooler blade interior. By introducing this thermal intermediary system, the blade can operate in high-temperature environments for increased power output while the cooling medium maintains the blade metal temperature within safe limits for structural strength.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This design reduces pressure loss while maintaining the heat transfer coefficient, ensuring efficient cooling of gas turbine blades without compromising performance.
Implementation Method 1
causing turbulence to the cooling medium flowing into the cooling passage, and also forming secondary flows running along the turbulators
Implementation Method 2
the amount of heat exchange with the wall surface of the cooling passage is increased to increase the heat transfer coefficient
Implementation Method 3
creating vortices to enhance heat transfer without increasing pressure loss by optimizing the arrangement of turbulators and dimples
Implementation Method 4
provide a cooling passage, intended for flowing a cooling medium such as air or steam, in the interior of each of the stationary blades and the moving blades
Data Source
Figure 1
Figure 2
Figure 3
AI summary
A blade cooling structure of a gas turbine, which can reduce the pressure loss of a cooling medium without decreasing the heat transfer coefficient, is provided. The blade cooling structure comprises a cooling passage (15) for flowing cooling air (A) from a proximal end portion (12) toward a blade portion (14) of a moving blade (11), a plurality of turbulators (21) arranged, on both wall surfaces of the cooling passage (15) opposing each other, in such a manner as to be inclined with respect to the flowing direction of the cooling air (A), and a plurality of dimples (22) formed in a downstream region (N) downstream of a center position (O) in the flowing direction of the cooling air (A) on the wall surface of the cooling passage (15) between the adjacent turbulators (21).