Abstract: An airfoil includes a blade having a leading edge, a trailing edge, a pressure side, a suction side, a tip, and a scoop. The scoop extends along the tip of the blade. The scoop comprises a difference in a radial height of the blade from a pressure side to a suction side of the blade. The radial height of the blade at the pressure side is less than the radial height of the blade at the suction side.

Abstract: An airfoil includes a blade having a pocket recess therein and one or more features disposed within the pocket recess. The one or more features are configured to disrupt pressure oscillations within the pocket recess. In another embodiment, a blade is disclosed having a first wall and a second wall. The first wall is disposed on a suction side of the blade and the second wall is disposed on a pressure side of the blade. The second wall is connected to the first wall at a leading edge of the blade. Together the first wall and the second wall form a portion of a pocket recess and the pocket recess is disposed asymmetrically with respect to a camber line of the blade.

Abstract: A rotor for a gas turbine engine includes a plurality of turbine blades extending radially outwardly of a rotor body. A plurality of purge blades are positioned to rotate with the rotor body, and to drive air radially outwardly toward the turbine blades.

Abstract: A turbine blade comprises an airfoil having a pressure side and a suction side, and extending from a leading edge to a trailing edge. The airfoil has a tip remote from a mounting root, and a pocket extending inwardly of the tip. The pocket has spaced walls with one wall associated with the pressure side of the airfoil, and an opposed wall associated with the suction side. A pin extends across the pocket and connects the opposed walls. A slot is formed in the pin at a location intermediate ends of the pin which connect to the opposed walls. A method for identifying a location for the pin along a distance between a leading edge and a trailing edge of the pocket utilizes a modal analysis, and seeks to find a location where both a reaction force and a moment are lower than they might be at other locations.

Abstract: A turbine wheel is disposed about an axis and has a back face including a plurality of lobes disposed about a periphery of the back face. The lobes define scalloped areas therebetween. The scalloped areas are further defined by a radius BR2 that blends into a first lobe and into a radius BR1 that also blends into a flat area.

Abstract: A turbine wheel is disposed about an axis and has a back having a separator disposed thereon, an inner undercut disposed between the separator and the axis, and an outer undercut disposed between the separator and an outer periphery of the back face. The inner undercut is defined by a first radius blending toward the axis and into a first flat section, and further defined by a second radius blending into the first flat section and into the separator.

Abstract: A turbine apparatus comprises a rotor with a hub section defined about a rotational axis and a plurality of blades attached to the hub section. The plurality of blades comprises a first group having a first angular spacing in a first circumferential sector of the rotor, and a second group having a second angular spacing in a second circumferential sector of the rotor. The first angular spacing is different from the second angular spacing, and the rotor blades are asymmetric about the rotational axis.

Abstract: An example method of designing blade lobes of a turbomachine blade and corresponding disk lobes includes determining contact areas between the blade lobes on a blade model and the disk loads on a disk model when the turbomachine blade is in a loaded position. The method adjusts the blade lobes, the disk lobes, or both, so that gaps are established between the blade lobes and the disk lobes at the contact areas when the turbomachine blade is in an unloaded position. The size of the gaps varies.

Abstract: A gas turbine engine includes a radial compressor with first and second blades. The first and second blades have tuned leading edges that prevent natural frequencies from exciting at speeds within an expected operating speed range.

Abstract: A gas turbine engine includes a radial compressor having first and second blades. The first blade has a tuned leading edge that prevents either blade from exciting at a natural frequency at speeds within an expected operating speed range.

Abstract: A method for controlling a flow of fuel to a starting gas turbine engine includes controlling the flow of fuel to the engine by controlling fuel pressure to the engine, and modulating the flow of fuel to the engine if engine exhaust gas temperature approaches a given exhaust gas temperature to lower the engine exhaust gas temperature below the given exhaust gas temperature.

Abstract: An impeller shroud is configured to receive an impeller. The impeller shroud establishes a plurality of air-bleed holes configured to communicate air with an impeller. The air-bleed holes are circumferentially distributed about the impeller shroud. The circumferential spacing between some adjacent air-bleed holes within the plurality of air-bleed holes is different than the circumferential spacing between other adjacent air-bleed holes within the plurality of air-bleed holes. A diffuser vane is configured to direct pressurized air to the combustor, the diffuser vanes are circumferentially distributed about the blade exducer. The circumferential spacing between some adjacent diffuser vanes within the plurality of diffuser vanes is different than the circumferential spacing between other adjacent diffuser vanes within the plurality of diffuser vanes.

Abstract: A variable vane system includes a plurality of vanes each being pivotal about an axis. A mechanical linkage drives the plurality of vanes to rotate about the axis. The mechanical linkage includes a ring gear to rotate, and in turn drive the plurality of vanes. There is at least one rod to drive the ring gear to rotate. The rod is driven by a hydraulic servo motor. A spring bias force is provided in the mechanical linkage to resist either translational or rotational oscillation.

Abstract: An example gas turbine engine fuel injector nozzle assembly includes a nozzle tip secured relative to a combustion area within a gas turbine engine. The nozzle establishes a plurality of first apertures that are configured to communicate a fuel to the combustion area. The nozzle establishes at least one second aperture that is configured to communicate a fluid to the combustion area. The fluid is different than the fuel. An example method of providing fuel to a combustion area within a gas turbine engine includes communicating a fuel through a first aperture in a nozzle tip to a combustion area in a gas turbine engine. The nozzle tip establishes an axis. The method also includes influencing fuel moving from the nozzle tip using a fluid directed through a second aperture in the nozzle tip. The fluid is different than the fuel. A portion of the second aperture is radially closer to the axis than the first aperture.

Abstract: A rotor bearing system for a rotating machine includes a housing having a bore that provides an inner surface. A bearing assembly is disposed within the bore and includes an outer surface. An annular cavity is provided radially between the outer surface and the inner surface. At least one protrusion extends radially outwardly from at least one of the inner and outer surfaces to an apex and into the annular cavity. A radial gap is arranged between the apex and the opposite surface from which the protrusion extends. In the disclosed example, the annular cavity is filled with an oil to provide a squeeze film damper between the housing and the bearing assembly. The protrusions exert a hydrodynamic preload on the bearing assembly, which reduces vibration during operation of the rotating machine.

Abstract: A gas turbine engine includes a radial compressor with first and second blades. The first and second blades have tuned leading edges that prevent natural frequencies from exciting at speeds within an expected operating speed range.

Abstract: A gas turbine engine includes a radial compressor having first and second blades. The first blade has a tuned leading edge that prevents either blade from exciting at a natural frequency at speeds within an expected operating speed range.

Abstract: A fuel system for a gas turbine engine comprises a set of starting fuel injectors that operate in a fuel pressurised atomisation mode to atomise fuel that flows therethrough during an ignition and a post-ignition phase and a set of main fuel injectors that operate in an air blast mode to atomise fuel that flows therethrough during the post-ignition phase and a high altitude ready to load phase and an air assist mode to control the pattern of fuel atomisation during a loaded phase.

Abstract: A ceramic matrix composite part having elongated cooling channels within the wall thereof is manufactured by inserting decomposable inserts within a woven ceramic fiber preform. The inserts are tows of continuous carbon fibers surrounded by a carbonaceous filler, and are inserted where the channels are desired. The preform, with the inserts in place, is disposed within a mold. A ceramic matrix material is added and the fiber preform is consolidated with the ceramic matrix material. The consolidated part is then heated to thermally decompose the inserts to create the elongated channels within the part. The inserts may be flexible and woven into the preform using an automated weaving loom, or they may have limited flexibility and be inserted by machine or by hand.

Abstract: A ceramic matrix composite part having elongated cooling channels within the wall thereof is manufactured by inserting decomposable inserts within a woven ceramic fiber preform. The inserts are tows of continuous carbon fibers surrounded by a carbonaceous filler, and are inserted where the channels are desired. The preform, with the inserts in place, is disposed within a mold. A ceramic matrix material is added and the fiber preform is consolidated with the ceramic matrix material. The consolidated part is then heated to thermally decompose the inserts to create the elongated channels within the part. The inserts may be flexible and woven into the preform using an automated weaving loom, or they may have limited flexibility and be inserted by machine or by hand.