Advanced Cooled Turbine Airfoil Aerodynamic Investigation
Abstract
Several connectively cooled trailing edge designs were investigated with the objective of eliminating the need for film cooling on the airfoil suction side. A 43.4% reduced solidity first stage turbine vane having potential application for an advanced Air Force fighter engine was selected for the evaluation. The final design eliminates film cooling on the suction side and uses the wavy criss-cross slot as the cooling scheme for the trailing edge section. The cooling design was incorporated into a cascade test airfoil using the radial wafer fabrication technique. The airfoil was constructed by photoetching the cooling design into the individual wafers, bonding the wafers together and machining the bonded block into the airfoil shape. The airfoil was subsequently evaluated in an airfoil cascade test to determine the aerodynamic and cooling performance. The aerodynamic profile loss of the reduced solidity radial wafer airfoil was reduced 56% relative to a baseline 43.4% reduced solidity configuration with film cooled suction surface and was 30% under the program goal. The wavy criss-cross slot design used in the trailing edge section proved to be an efficient cooling technique, and eliminated the need for suction side film cooling.
Document Details
- Document Type
- Technical Report
- Publication Date
- Feb 01, 1977
- Accession Number
- ADA041137
Entities
People
- W. G. Hess
Organizations
- Pratt & Whitney