Abstract: A turbine nozzle includes outer and inner bands between which extend a plurality of vanes for channeling combustion gases. Each of the vanes includes leading and trailing edges, and pressure and suction sides extending therebetween, and also a bow along the trailing edge to increase pressure in the gases adjacent the inner band. The vanes also include a thermal barrier coating (TBC) selectively disposed solely along the suction side between the leading and trailing edges.

Abstract: A gas turbine engine rotor blade 18 includes a dovetail 22 and integral airfoil 24. The airfoil includes a pair of sidewalls 28,30 extending between leading and trailing edges 32,34, and longitudinally between a root 36 and tip 38. The sidewalls are spaced laterally apart to define a flow channel 40 for channeling cooling air through the airfoil. The tip includes a floor 48 atop the flow channel, and a pair of ribs 50,52 laterally offset from respective sidewalls. The ribs are longitudinally tapered for increasing cooling conduction thereof.

Abstract: A turbine nozzle includes outer and inner bands between which extend a plurality of vanes for channeling combustion gases. Each of the vanes includes leading and trailing edges, and pressure and suction sides extending therebetween, and also a bow along the trailing edge to increase pressure in the gases adjacent the inner band. The vanes also include a thermal barrier coating (TBC) selectively disposed solely along the suction side between the leading and trailing edges.

Abstract: A turbine blade including an airfoil section having a double-wall construction for side-wall impingement cooling on the pressure side and a multi-pass serpentine along the suction side of the blade, is described. More particularly, and in one embodiment, the airfoil section includes a pressure side wall and a suction side wall which are joined together at a leading edge and a trailing edge. The blade also includes a leading edge, or tip, and a trailing edge, or tail. The airfoil section also includes a leading edge cavity having a plurality of radial film air holes, and an inner cavity which is a three pass serpentine. As cooling air flows along the passageways, it convectively cools the portions of the turbine blade adjacent these passageways. The airfoil section further includes a trailing edge cavity to cool the trailing edge flow region of the airfoil section.

Abstract: A turbine seal includes a first arcuate segment defining a flowpath boundary between combustion gases and air, and includes a radially outwardly extending rail at one end thereof. A second arcuate segment is disposed coaxially with the first segment for defining a continuation of the flowpath boundary, and has a radially extending face adjoining the rail. A leaf seal bridges the rail and the face for sealing leakage therebetween, and a plurality of pins extend through the leaf seal for providing mounting to the rail. Each pin includes a head and an opposite tip, with the pins being fixedly joined to the rail solely about the pin tips for freely supporting the pin heads in a cantilever without obstruction therearound for eliminating supporting tabs. A leaf spring is mounted on the pins between the heads and the leaf seal to pre-load the leaf seal against the rail and face to effect the sealing therebetween.

Abstract: A wall adapted for use in a gas turbine engine between a first fluid and a second hotter fluid includes a first side over which is flowable the first fluid, and an opposite second side over which is flowable the second fluid. An elongate slot extends partly inwardly and perpendicularly from the second side toward the first side and is provided with the first fluid through a plurality of longitudinally spaced apart holes. The holes are aligned coplanar with the slot and longitudinally inclined to effect longitudinal overlapping of the first fluid inside the slot prior to discharge therefrom as a substantially continuous film.

Abstract: A joint connection for abutting, circumferentially extending segments in a gas turbine engine, such as turbine nozzle segments, includes a sealing member which is insertable between longitudinally extending slots formed in the abutting side edges of two adjacent segments. The segments are each formed with a number of grooves which are longitudinally spaced along the length of the slots therein, and which extend beneath the sealing member carried within the slots. A cooling air flow path is thus formed extending from one side of the sealing member, into the longitudinal slot of each segment and then through the grooves to the opposite side of the sealing member so that the entire joint connection or seal region between the abutting side edges of the segments is cooled.

Abstract: The cooling insert for a movable vane in a jet engine is divided into plural overlapping forward and aft segregated members, each member being coolant fed through separate vane trunnion areas and each insert is of successively decreasing cross-sectional area. Problems of locally inadequate vane cooling, coolant match point movement, and unequal coolant supply pressures are addressed by the disclosed apparatus. Reuse of the trunnion supplied impingement cooling air for additional cooling functions is also disclosed.

Abstract: A multiple-impingement cooled structure, such as for use as a turbine shroud assembly. The structure includes a plurality of baffles which define with an element to be cooled, such as a shroud, a plurality of cavities. Impingement cooling air is directed through holes in one of the baffles to impinge upon only the portion of the shroud in a first cavity. That cooling air is then directed to impinge again upon the portion of the shroud in a second cavity.

Abstract: A multiple-impingement cooled structure, such as for use as a turbine shroud assembly. The structure includes a plurality of baffles which define with an element to be cooled, such as a shroud, a plurality of cavities. Impingement cooling air is directed through holes in one of the baffles to impinge upon only the portion of the shroud in a first cavity. That cooling air is then directed to impinge again upon the portion of the shroud in a second cavity.

Abstract: A gas turbine engine comprising a number of annular wall sectors which form a complete circular wall defining a hot gas passage is provided with a cooling system incorporating a plurality of hollow impingement vessels disposed in a circular array within an annular chamber behind the wall. Cooling fluid is routed to the self-contained impingement pressure vessels which are provided with perforations to disperse the cooling fluid into impingement upon the wall. Unlike prior systems, the impingement vessels are not physically attached to the band but, rather, are supported by a structural frame which partially defines the annular chamber and, in the preferred embodiment, also supports the wall sectors. Since the impingement vessel, and not the backside of the wall, is the pressure vessel for the pressurized cooling fluid, leakage between adjacent wall sectors is greatly reduced with no loss in cooling effectiveness when compared to conventional cooling systems. Thus, turbine cycle performance is improved.