Abstract: A gas turbine engine combustor is described which includes outer and inner annular combustor liners formed of sheet metal. The exit duct circumscribes an annular combustor exit and defines a combustion gas path. The exit duct includes a large exit duct having an annular forged metal section which is butt welded at an upstream end to the outer combustor liner to form a first annular joint. The annular forged metal section is fixed at a downstream end to an annular sheet metal wall to form a second annular joint. Methods of forming and of repairing a gas turbine engine combustor are also disclosed.

Abstract: A combustor floating collar assembly having a floating collar disposed about an opening in an outer surface of a combustor liner. The floating collar is displaceable relative to the outer surface. A retention plate is attached to the outer surface of the combustor liner and traps the floating collar between the outer surface and the retention plate. At least a portion of the retention plate engages the floating collar and biases it toward the outer surface.

Abstract: A gas turbine engine combustor is described which includes outer and inner annular combustor liners formed of sheet metal. The exit duct circumscribes an annular combustor exit and defines a combustion gas path. The exit duct includes a large exit duct having an annular forged metal section which is butt welded at an upstream end to the outer combustor liner to form a first annular joint. The annular forged metal section is fixed at a downstream end to an annular sheet metal wall to form a second annular joint. Methods of forming and of repairing a gas turbine engine combustor are also disclosed.

Abstract: A vane structure of a gas turbine engine includes a plurality of vanes extending radially between outer and inner shrouds. At least the outer shroud is formed substantially from a single-piece annular skin of sheet metal. A plurality of reinforcing plates are placed against and are affixed to an outer surface of the skin of the outer shroud in respective joining locations where a radial outer end of each of the respective vanes joins the skin.

Abstract: A gas turbine engine combustor includes an exit duct having annular first and second exit duct walls radially spaced apart to define therebetween the combustor exit opening. The first and/or second exit duct walls has a double-skin wall section which includes an inner hot wall facing the combustor exit opening and an outer cold wall fastened to the inner hot wall and radially spaced away therefrom to define a radial gap therebetween. The outer cold wall has a coefficient of thermal expansion greater than that of the inner hot wall. This helps reduce thermal growth mismatch between the outer cold wall and the inner hot wall during operation of the combustor, and reduces thermal stress at the joint between the hot and cold walls.

Abstract: A turbine vane ring has radially outer and inner annular shrouds defining therebetween an annular gaspath. Circumferentially spaced-apart airfoil vanes extend radially across the gaspath between the outer and the inner shrouds. The radially outer shroud has a circumferentially continuous cylindrical wall extending axially from a leading edge to a trailing edge. A set of circumferentially distributed stress relieving slots is defined in the leading edge of the cylindrical wall at locations adjacent to the leading edge of at least some of said airfoil vanes. The stress relieving slots extend radially through the cylindrical wall from the radially inner surface to the opposed radially outer surface thereof.

Abstract: A turbine vane ring has a radially outer and inner annular shrouds defining therebetween an annular gaspath. Circumferentially spaced-apart airfoil vanes extend radially across the gaspath between the outer and the inner shrouds. The radially outer shroud has a circumferentially continuous cylindrical wall extending axially from a leading edge to a trailing edge. A set of circumferentially distributed stress relieving slots is defined in the leading edge of the cylindrical wall at locations adjacent to the leading edge of at least some of said airfoil vanes. The stress relieving slots extend radially through the cylindrical wall from the radially inner surface to the opposed radially outer surface thereof.

Abstract: A vane structure of a gas turbine engine includes a plurality of vanes extending radially between outer and inner shrouds. At least the outer shroud is formed substantially from a single-piece annular skin of sheet metal. A plurality of reinforcing plates are placed against and are affixed to an outer surface of the skin of the outer shroud in respective joining locations where a radial outer end of each of the respective vanes joins the skin.