Gas turbine engines are known to include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine for expanding the hot combustion gases to produce mechanical shaft power. Combustors operate at temperatures that may exceed 2,500 degrees Fahrenheit, thereby exposing the turbine blade and vane assemblies to these high temperatures. As a result, the turbine airfoils must be made of materials capable of withstanding such high temperatures. In addition, the airfoils often contain cooling systems for prolonging the life of the airfoils and reducing the likelihood of failure as a result of excessive temperatures.
Gas turbine airfoils have an outer skin defining the desired airfoil shape including a leading edge and a trailing edge and extending along a chord length. An outer skin of metal may by coated with a ceramic thermal barrier coating material for additional protection, especially in the first few rows of airfoils within the turbine, which are exposed to the highest temperatures and greatest fluid velocities. Inner structures of the airfoils typically define cooling channels for directing cooling fluid against the backside of the outer skin. The cooling fluid may be air extracted from the compressor/combustor flow path or it may be steam in some combined cycle plant applications. The cooling channels often include multiple flow paths designed to maintain all regions of the airfoil below a design temperature value, including impingement plates and holes for directing cooling fluid against the back side of the outer skin and film cooling holes through the outer skin for directing a layer of cooling air across the outer surface of the airfoil. See, for example, U.S. Pat. No. 5,511,937 issued on Apr. 30, 1996, and U.S. Pat. No. 4,153,386 issued on May 8, 1979. Centrifugal forces and flow boundary layers sometimes prevent certain areas of the airfoils from being adequately cooled, resulting in the formation of localized hot spots. Furthermore, contaminants in the cooling fluid can clog impingement orifices and film cooling orifices, resulting in additional localized hot spots. Also, debonding and/or spallation of the thermal barrier coating can result in such hot spots, as the thermal insulation material chips off, leaving the airfoil unprotected. Such hot spots can result in a premature failure of the airfoil and thereby necessitate replacement of the part. When an airfoil fails, portions of the airfoil may break off and strike downstream components of the turbine engine, thereby causing collateral damage that may be extremely costly.
A variety of systems have been used to monitor the performance of an airfoil during operation of a gas turbine engine. U.S. Pat. No. 4,595,298 issued on Jun. 17, 1986, describes a temperature detection system used on the exterior of a film cooled turbine airfoil. U.S. Pat. No. 4,983,034 issued on Jan. 8, 1991, describes a sensing fiber used to monitor strain levels at one or more locations of a composite member. U.S. Pat. No. 5,442,285 issued on Aug. 15, 1995, describes a stationary eddy current sensor used to examine a passing turbine blade. U.S. Pat. No. 6,838,157 issued on Jan. 4, 2005, describes the embedding of sensors within a ceramic thermal barrier coating of a gas turbine component. All of the patents mentioned in this Background section are incorporated by reference herein.