Abstract: A heated guide vane for turbomachinery includes a guide vane having two major surfaces joined about their periphery by edges and an electric heater element, wherein the electric heater element is secured to at least one major surface of the guide vane.

Abstract: A method of operating a heated guide vane assembly for turbomachinery, the heated guide vane assembly including a plurality of guide vanes each having two major surfaces joined about their periphery by edges and an associated electric heater element secured to at least one major surface of the guide vanes. The method includes the steps of energizing heater elements on at least one of the guide vanes, de-energizing heater elements on at least one of the guide vanes, and energizing heater elements on at least one of the guide vanes which was not energized in the first energizing step.

Abstract: Methods and apparatus for a balance weight access assembly are provided. The assembly includes an access cover, and an access tube including a first opening, a second opening, and a substantially hollow body extending therebetween. The first opening is positioned proximate to a balance weight retainer, the second opening is positioned proximate to the access cover, and the body is positioned in substantial alignment with an installation axis of the balance weight retainer.

Abstract: A method for producing apertures in hot section components of gas turbine engines made from ceramic matrix composites that have at least one oxidizable component. The method involves forming the apertures using a laser beam controlled by parameters that ablate the ceramic matrix composite in the path of the beam, while simultaneously heating the matrix material, SiC or SiN, to a sufficient temperature to oxidize it to form a silica. Sufficient heat is supplied by the beam to melt the silica to cause it to flow. The melted silica is quickly solidified as recast silica along the walls of the newly created aperture before it has an opportunity to flow and form undesirable geometries. The wall of the aperture is formed of recast silica that is a smooth surface and that forms an oxidation barrier to inhibit any further oxidation of the underlying composite as it is exposed to the high temperatures and oxidative, corrosive atmosphere of an operating gas turbine.

Abstract: A method for producing apertures in hot section components of gas turbine engines made from ceramic matrix composites that have at least one oxidizable component. The method involves forming the apertures using a laser beam controlled by parameters that ablate the ceramic matrix composite in the path of the beam, while simultaneously heating the matrix material, SiC or SiN, to a sufficient temperature to oxidize it to form a silica. Sufficient heat is supplied by the beam to melt the silica to cause it to flow. The melted silica is quickly solidified as recast silica along the walls of the newly created aperture before it has an opportunity to flow and form undesirable geometries. The wall of the aperture is formed of recast silica that is a smooth surface and that forms an oxidation barrier to inhibit any further oxidation of the underlying composite as it is exposed to the high temperatures and oxidative, corrosive atmosphere of an operating gas turbine.

Abstract: A method for producing apertures in hot section components of gas turbine engines made from ceramic matrix composites that have at least one oxidizable component. The method involves forming the apertures using a laser beam controlled by parameters that ablate the ceramic matrix composite in the path of the beam, while simultaneously heating the matrix material, SiC or SiN, to a sufficient temperature to oxidize it to form a silica. Sufficient heat is supplied by the beam to melt the silica to cause it to flow. The melted silica is quickly solidified as recast silica along the walls of the newly created aperture before it has an opportunity to flow and form undesirable geometries. The wall of the aperture is formed of recast silica that is a smooth surface and that forms an oxidation barrier to inhibit any further oxidation of the underlying composite as it is exposed to the high temperatures and oxidative, corrosive atmosphere of an operating gas turbine.

Abstract: A combustor having liners made from ceramic matrix composite materials (CMC's) that are capable of withstanding higher temperatures than metallic liners. The ceramic matrix composite liners are used in conjunction with mating components that are manufactured from superalloy materials. To permit the use of a combustor having liners made from CMC materials in conjunction with metallic materials used for the mating forward cowls, and aft seals with attached seal retainer over the broad range of temperatures of a combustor, the combustor is designed to allow for the differential thermal expansion of the differing materials at their interfaces in a manner that does not introduce stresses into the liner as a result of thermal expansion and also balances the flow of cooling air as a result of the thermal expansion.