Curvature Effects on the Heat Transfer Performance of Three-Dimensional Film Cooling of Gas Turbine Blades.
Abstract
Film cooling is used extensively for the blades of high-performance, high-temperature gas turbines, especially for aircraft turbines. In this method, a film of coolant is injected into the boundary layer covering the skin of the blades and creating a cool layer which separates the blade surface from the hot gas stream and, in this way, reduces the blade temperature. For best performance the coolant should be injected through a slot or a strip of porous material. This, however, is not possible for turbine blades because of strength considerations, and the coolant is injected through one or several rows of holes. For aircraft gas turbines, air is used as a coolant. The present investigation, therefore, is concerned with the cooling performance of film cooling when cooling air is injected into the boundary through one or two rows of holes. A standard configuration of the coolant holes is used because it has been used in previous investigations and because configurations in actual turbine blades are close to it. The cooling holes are arranged at a distance apart equal to three times the hole diameter. For injection through two rows of holes, the two rows are staggered and the centers of the holes are on the corners of equilateral triangles. The channels which end at the blade skin in the cooling holes are inclined by an angle of 35 deg against the skin surface in the downstream direction.
Document Details
- Document Type
- Technical Report
- Publication Date
- Aug 01, 1982
- Accession Number
- ADA119729
Entities
People
- E. R. G. Eckert
- R. J. Goldstein
Organizations
- University of Minnesota