Variable Heat Transfer Collector Baffle for Gas Turbine Airfoil Cooling
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Solution Overview
Problem
Ceramic matrix composite (CMC) airfoils in gas turbine engines face challenges with thermal gradients due to their lower stress capability compared to metallic airfoils, and existing cooling schemes often result in over-cooling of regions away from the stagnation point, leading to inefficient heat transfer and temperature distribution.
Innovation Solution
A cooling scheme featuring a radially extending cavity with a baffle that includes a forward chamber with impingement holes for cooling fluid, axial ribs to direct fluid aft, and an aft chamber for radial flow, along with thermal heat transfer features to enhance cooling, tailored to mimic external thermal heat transfer distribution and minimize thermal gradients.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If cooling air is flowed radially through a leading edge cavity to create high heat transfer at the stagnation point, then heat transfer at the stagnation point is improved, but flow distributes to other areas of the cavity causing over-cooling of regions away from the stagnation point
Solution Approach 1:
The baffle structure creates different flow paths and heat transfer characteristics in different regions of the cavity. By positioning the baffle to restrict flow distribution and concentrate cooling air through impingement holes at the stagnation point, the system achieves localized high heat transfer where needed while preventing unnecessary cooling in other regions.
Solution Approach 2:
The cavity is divided into distinct flow regions by the baffle structure. The baffle segments the cooling air flow to ensure that cooling air is directed specifically to the stagnation point through controlled impingement holes, rather than distributing uniformly across the entire cavity surface.
2Temperature
If cooling air is distributed to balance heat transfer across the cavity, then heat transfer distribution is improved, but excessive cooling air is required leading to inefficient heat transfer
Solution Approach 1:
The baffle structure enables localized cooling air delivery to regions with highest heat load. By concentrating cooling air through impingement holes at the stagnation point rather than distributing it uniformly, the system achieves effective heat transfer balance with reduced total cooling air quantity.
Solution Approach 2:
The baffle modifies the flow distribution parameters within the cavity, changing the velocity distribution and flow direction to concentrate cooling air where most needed. This parameter change allows achieving balanced heat transfer with lower overall cooling air consumption.
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 cooling scheme effectively manages thermal gradients and cooling fluid temperature, providing targeted cooling to high-temperature regions while reducing the need for excessive cooling air, thus enhancing the performance and reliability of CMC airfoils in gas turbine engines.
Implementation Method 1
flowing cooling air radially through a leading edge cavity of the airfoil to create high heat transfer at the stagnation point
Implementation Method 2
thermal heat transfer features to enhance cooling, tailored to mimic external thermal heat transfer distribution
Implementation Method 3
Fluid exiting the baffle is directed axially aft using axial ribs
Data Source
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AI summary
A gas turbine engine component includes an airfoil (65) defined by a leading edge (66), a trailing edge (67), a pressure sidewall (71), and a suction sidewall (72). An internal cavity (69) extends radially through the airfoil (65) and is partially defined by an inner surface (73) of the pressure sidewall (74) and an inner surface (74) of the suction sidewall (72). A baffle (75; 175; 275) is disposed within the internal cavity (69), and includes a baffle wall (76; 176) conformal with adjacent surfaces of the internal cavity (69), a divider (77) separating a forward chamber (78) and an aft chamber (79), and a plurality of orifices (80) extending through the baffle wall (76; 176) at the forward chamber (78). At least one axial rib (84) is disposed between the baffle wall (76; 176) and the internal cavity (69), and is positioned at a discrete spanwise location of the airfoil (65).