Abstract: A method of maintaining rotor tip clearance and compressor operational line during a transient operation of a gas turbine engine is disclosed. In various embodiments, the method includes applying high spool auxiliary power to a high speed spool for a first time period, applying low spool auxiliary power to a low speed spool for a second time period, sensing one or more operational parameters of the gas turbine engine during the transient operation, and ceasing application of power to the high speed spool, based on the one or more operational parameters.

Abstract: Described are an airfoil array segment (100, 200, 300) having at least two airfoils (20, 30) and a platform (10) that features an axis asymmetrical platform surface (12). This platform surface features an elevation (110, 210, 310) that extends from the pressure side (21) of the first to the suction side (32) of the second airfoil (30). A highest point (111, 211, 311) of the elevation is more proximate to the suction side 32 of the second airfoil (30) than to the pressure side (21) of the first airfoil (20). Also described are an airfoil, a platform, an airfoil passage and a turbomachine.

Abstract: The present invention relates to a lightweight composite propeller for an outboard motor, wherein the propeller has a separate hub and blades which are easily repaired when damaged, improves fuel efficiency because a lightweight composite material is used therefor, and is easily manufactured in large quantities.

Abstract: High performance wedge diffusers utilized within compression systems, such as centrifugal and mixed-flow compression systems employed within gas turbine engines, are provided. In embodiments, the wedge diffuser includes a diffuser flowbody and tapered diffuser vanes, which are contained in the diffuser flowbody and which partition or separate diffuser flow passages or channels extending through the flowbody. The diffuser flow channels include, in turn, flow channel inlets formed in an inner peripheral portion of the diffuser flowbody, flow channel outlets formed in an outer peripheral portion of the diffuser flowbody, and flow channel throats fluidly coupled between the flow channel inlets and the flow channel outlets. The diffuser vanes include a first plurality of vane sidewalls, which transition from linear sidewall geometries to non-linear sidewall geometries at locations between the flow channel inlets and the flow channel outlets.

Abstract: A gas turbine and nozzle system is provided that includes a radial inflow turbine rotor and a volute providing a flow path to deliver a pressurized gas to a circumference of the radial turbine rotor. The volute incorporates a shape which substantially conforms to a radial turbine shroud contour. The volute includes at least first and second parts. A mating surface between the first and second parts is substantially aligned with a direction of pressurized gas flow in the volute.

Abstract: An aviation turbine blade extending in the radial direction and presenting a pressure side and a suction side, including a plurality of pressure side cavities extending radially at the pressure side of the blade, a plurality of suction side cavities extending radially at the suction side of the blade, and at least one central cavity located in the central portion of the blade and surrounded by pressure side cavities and by suction side cavities, the blade further including a plurality of cooling circuits, in which at least a first cooling circuit comprises: a first cavity and a second cavity, the first and second cavities communicating with each other at a radially inner end and at a radially outer end of the blade.

Abstract: A radial compressor, with a rotor-side impeller, a stator-side housing, an inflow passage via which medium to be compressed can be fed to the impeller in the axial direction, a diffuser via which compressed medium can be conducted in the radial direction away from the impeller in the direction of a spiral housing section of the housing The diffuser radially outside in a transition region to the spiral housing section has an enlarging flow cross section.

Abstract: A blade outer air seal (BOAS) for a gas turbine engine includes a seal ring body having a radially inner face and a radially outer face that axially extend between a leading edge portion and a trailing edge portion and a segmented spline that extends from the radially outer face of the seal ring body, the seal secured to the radially inner face of the seal ring body.

Abstract: An airfoil extends in a radial direction a span in a range 3.97-4.27 inch (100.9-108.6 mm). A chord length extends in a chordwise direction from a leading edge to a trailing edge at 50% span in a range 1.28-1.58 inch (32.4-40.0 mm). The airfoil includes at least two of a first mode with a frequency of 243±10% Hz, a second mode with a frequency of 374±10% Hz, a third mode with a frequency of 741±10% Hz, a fourth mode with a frequency of 1198±10% Hz, a fifth mode with a frequency of 1663±10% Hz, a sixth mode with a frequency of 2411±10% Hz, a seventh mode with a frequency of 3734±10% Hz, an eighth mode with a frequency of 6738±10% Hz and a ninth mode with a frequency of 8977±10% Hz.

Abstract: A gas turbine engine has an engine static structure and at least one component rotatable relative to the engine static structure about an engine axis of rotation. A fan is coupled to at least one component for rotation about the engine axis of rotation. An actuator is mounted to the engine static structure, wherein the actuator is activated to prevent the fan from rotation and is inactivated to allow the fan to rotate. A method for preventing rotation of a fan in a gas turbine engine is also disclosed.

Abstract: A dual stage grinder pump (10) includes a housing (12) with a liquid inlet (20) and an outlet (26). An impeller (36) includes first impeller vanes (52) on a first axial side, and second impeller vanes (60) on a second axial side. A grinder (40) operates to reduce the size of suspended solids in the liquid that enters the pump inlet. Liquid passes from the inlet to a first fluid passage (30) and is acted upon by first vanes of the impeller in a first stage, and then passes through a second fluid passage (72). Liquid is acted upon by the second vanes of the impeller in a second stage and is passed through a third fluid passage (88) to the outlet.

Abstract: A propeller pitch control system for a turboprop engine of an aircraft includes a negative torque sensor, a full feather control valve assembly, and a feather valve assembly. The feather valve assembly is responsive to fluid pressures from the negative torque sensor and the control valve assembly to move between a plurality of positions. An engine control unit or a manual user interface can also be used to control the position of the feather.

Abstract: The invention relates to a pitch bearing, a blade, an impeller of a wind turbine and to a connecting method for the wind turbine. There is provided a pitch bearing of a wind turbine comprising: an inner bearing ring, an outer bearing ring and an extension portion, wherein one of the inner bearing ring and the outer bearing ring is connected to the extension portion. With the pitch bearing, blade, impeller of the wind turbine and the connecting method for the wind turbine according to the disclosure, it is possible to connect the blade to the pitch bearing without bolts, allow a lower weight for the wind turbine, increase strength and fatigue life of the wind turbine, and allow a lower manufacturing cost and difficulty for the wind turbine.

Abstract: A fan housing includes an inner housing portion centered on a central axis, an outer housing portion centered on a central axis, and a disk portion connecting the first end of the outer housing portion to the first end of the inner housing portion. The inner housing portion has a first end and a second end positioned axially away from the first end. The outer housing portion has a first end and a second end positioned axially away from the first end. The outer housing portion includes a first thickened wall section with a first end and a second end, a stator interface section with a first end connected to the second end of the first thickened wall section and a second end, and a second thickened wall section with a first end connected to the second end of the stator interface section and a second end.

Abstract: A blade outer air seal includes a base portion that extends between a leading edge and a trailing edge. A forward wall and an aft wall extend radially outward from the base portion to a radially outer portion. The radially outer portion is spaced from the base portion.

Abstract: A sealing assembly for a gas turbine is provided, especially an aircraft gas turbine, including: a first component having a first flange portion and a second component having a second flange portion; in the area of the first flange portion and the second flange portion, the first component and the second component being detachably intercoupled by a bolt connection; between the first component, especially a rounded portion, and the second component, especially the second flange portion, a variable interspace being provided that is at least partially sealed by at least one sealing element. The sealing element has a first sealing portion that rests against the first component, especially against the rounded portion, and a second sealing portion that rests against the second component, especially against the second flange portion; the first sealing portion and the second sealing portion being interconnected and extending obliquely relative to each other.

Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y, Zadim given in Table 1, in which the coordinate Zadim is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of said profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Abstract: A structure adapted to traverse a fluid environment exerting an ambient fluid pressure is provided. The structure includes an elongate body extending from a root to a wingtip and encapsulating at least one interior volume containing an interior fluid exerting an interior fluid pressure that is different from the ambient fluid pressure. A method of retrofitting a structure adapted to traverse a fluid environment exerting an ambient fluid pressure, the structure comprising an elongate body extending from a root to a wingtip and having at least one interior volume is also provided. The method includes sealing the elongate body to encapsulate the at least one interior volume containing an interior fluid; associating at least one valve with the at least one interior volume; and modifying interior fluid content via the at least one valve to produce an interior fluid pressure that is different from the ambient fluid pressure.

Abstract: A system for controlling the clearance distance between an impeller blade tip of a centrifugal compressor and a radially inner surface of an impeller shroud in a turbine engine. The system comprises a thermal driver coupled between the impeller shroud and engine casing by hinged linkages. The thermal driver is coupled to a closed form refrigeration system having an evaporator, a compressor, a condenser, an expansion valve, and a refrigerant contained therein. The thermal driver is cooled by the refrigeration system, and the rate of cooling causes expansion and contraction of the thermal driver that is translated by linkages into axially forward and aft motion of the shroud.

Abstract: Embodiments of systems and methods for operating a gas turbine engine defining a bowed rotor condition are generally provided. The systems and methods include rotating a rotor assembly defining a bowed rotor condition from approximately zero revolutions per minute (RPM) to within a bowed rotor mitigation speed range, in which the bowed rotor mitigation speed range is defined by a lower speed limit greater than zero RPM and an upper speed limit less than or equal to an idle speed condition of the gas turbine engine; applying a load at the rotor assembly via an energy storage device; adjusting the load to limit rotational speed or acceleration of the rotor assembly to within the bowed rotor mitigation speed range for a period of time; and removing the load to enable rotation of the rotor assembly to the idle speed condition following the period of time.