Investigation of Heat Transfer to a Turbine Blade Cascade Using a Shock Tube.
Abstract
In order to increase the turbine inlet temperatures at which gas turbine engines operate it is important to understand the heat transfer mechanisms which govern turbine blade heating. This study used a shock tube to generate high temperature gas flows which were allowed to pass through a turbine blade cascade. A Germanium surface thermocouple was used to provide temperature historics at five locations on a turbine blade for a range of flow temperature to blade temperature ratios. Heat transfer rates were determined from these temperature histories using a finite differencing scheme to approximate the heat equation. It was found that the rate of heat transfer along the pressure side of the blade decreased with chordwise position from a maximum value at the leading edge. On the blade suction side, heat transfer rates were found to be considerably greater at the 1//4 and 1/2 chord positions than at the leading edge.
Document Details
- Document Type
- Technical Report
- Publication Date
- Dec 01, 1984
- Accession Number
- ADA153090
Entities
People
- J. E. Gochenaur
Organizations
- Air Force Institute of Technology