Abstract: A turbofan engine including an axially extending inlet wall surrounding an inlet flow path. A radial distance between the inlet wall and the inner wall adjacent the fan defines a downstream height of the inlet flow path. A plurality of vanes are circumferentially spaced around the inlet, each of the vanes extending radially inwardly from the inlet wall, a maximum radial distance between a tip of each of the vanes and the inlet wall defining a maximum height of the vane. The maximum height of the vane is at most 50% of the downstream height of the flow path. In another embodiment, the maximum height of the vane is at most 50% of the maximum fan blade span. A method of reducing a relative Mach number at fan blade tips is also discussed.

Abstract: A rotating circular aerofoil system that includes a disc body that has an aerofoil shape in cross section. The aerofoil shape has an angle of attack that is symmetrical. The disc body's includes a circular inner edge which forms a circular center opening and the disc body's leading edge. The disc body also includes an outer edge that forms the trailing edge. Intersecting ribs extend across the center opening and include a shaft opening that attaches to a rotating shaft. Formed on the top and bottom surfaces of the disc body are evenly spaced apart, extending T-shaped or L-shaped fins. In one embodiment, an air blower or impeller is attached to a hollow shaft linked to an air pump. When the air pump is activated, air is delivered to the air blower and impeller that distributes the air over the top surface of the disc body to improve lift.

Abstract: An exemplary section of a gas turbine engine according to this disclosure includes, among other things, a first array of airfoils including a first number of airfoils, and a second array of airfoils downstream of the first array of airfoils. The second array includes a second number of airfoils. The second number of airfoils is at least 1.19 times the first number of airfoils thereby providing a predetermined modal.

Abstract: A compressor airfoil of a gas turbine engine includes a pressure side and a suction side of the airfoil extending downstream from a stagnation point, the suction side including a suction side surface portion within a leading edge region, and a main suction side airfoil surface downstream from the suction side surface portion and extending contiguously therewith. The suction side surface portion having a compound curvature profile which includes at least a leading edge having a first curvature profile and a chamfered surface having a second curvature profile different from the first curvature profile. The chamfered surface being contiguous with and extending immediately downstream from the leading edge. The first curvature profile being curved. The second curvature profile of the chamfered surface being substantially flat and defining a substantially straight-line profile in a cross-section transverse to the span-wise axis of the airfoil.

Abstract: Embodiments of an impeller shroud support for disposition around an impeller are provided, as are embodiments of gas turbine engine including impeller shroud supports. In one embodiment, the impeller shroud support includes a shroud body, a support arm joined to and extending around the shroud body, and a plurality of Mid-Impeller Bleed (MIB) flow passages. Each MIB flow passage includes, in turn, an inlet formed in the shroud body and configured to receive bleed air extracted from the impeller, a throat portion, an outlet formed in the support arm and through which the bleed air is discharged, and a curved intermediate section between the inlet and the outlet. During usage of the impeller shroud support, the curved intermediate section turns the bleed air flowing through the MIB passage in a radially outward direction prior to discharge from the outlet of the MIB flow passage.

Abstract: A turbine stage assembly includes a disc carrying a cascade of blades and an annular seal-plate that is secured to the disc by a first connection. One or more of the blades include a root portion configured to be received in a complementarily shaped radially extending slot in the disc such that a face of the root portion faces the plate. A terminal portion of the root portion is cut away adjacent the face to present an open space between a radially inner wall of the slot and a wall of the cutaway root portion. A first part of a second connector extends radially inwardly from the wall of the cutaway root portion. The first part of the second connector is configured to engage with a complementing second part of the second connector provided on the seal-plate.

Abstract: A rotary wing aircraft is provided including a dual counter-rotating, coaxial rotor system having an upper rotor system and a lower rotor system rotatable about a common axis. A plurality of blade assemblies is mounted to a portion of either the upper rotor system or the lower rotor system. A plurality of individually controllable actuators is coupled to each of the plurality of blade assemblies. Each of the plurality of actuators is configured to control movement of the coupled blade assembly about a pitch axis. The rotary-wing aircraft additionally includes a sensor system within an airframe. A higher harmonic control (HHC) controller is arranged in communication with the sensor system and the plurality of actuators to individually control the upper rotor system and the lower rotor system to reduce vibration.

Abstract: A method of determining and controlling the attack angle of a fixed-speed wind turbine blade involves determination of an appropriate angle of attack for optimal extraction of wind energy at low to medium wind speed, which includes three steps: 1—defining basic parameters of the turbine including blade length and width, fixed rotational speed, rated wind speed, start-up speed and the lowest speed at which the turbine is forced to stop, and a relationship with wasted power; 2—defining wind speed based on a set of attack angles calculated to form an overall optimal attack angle of the wind turbine blade; and 3—calculating necessary physical parameters to come up with the most effective method of controlling the turbine blade so that the turbine can be directly connected to the grid, making wind power costs as low as those of other conventional energy sources.

Abstract: A multi-stage, self-priming centrifugal pump assembly includes at least two pump stages (4) which are consecutive in a main flow direction (32), and a backflow channel (13) which lies parallel to at least one a pump stage (4). The backflow channel (13) runs out downstream of the first or a further pump stage (4), in the main flow direction (32).

Abstract: A casting core for a Blade Outer Air Seal includes a heat exchange cavity core section in communication with a first plenum section and a second plenum section, the first plenum and the second plenum section are of a thickness greater than the heat exchange cavity section.

Abstract: A vortex generator for a wind turbine blade to be mounted to a wind turbine blade includes: a platform portion to be mounted to a surface of the wind turbine blade; and at least one fin disposed upright on the platform portion. The platform portion has a cross section having a curved convex shape, at least along a blade spanwise direction of the wind turbine blade.

Abstract: A fan assembly includes a fan having an outlet side, a fan duct positioned at the fan outlet side, and a hollow cylindrical chamber positioned within the fan duct, wherein the chamber has two closed ends, a cylindrical side wall, and an inlet port. The dimensions of the chamber are determined to suppress one or more noise frequencies emitted by the fan. In one example, one or more dimensions of a chamber disposed in an outlet duct of an axial fan is automatically adjusted to suppress a noise frequency that is a function of a rotational speed of the axial fan.

Abstract: The Invention provides a spiral flow constant pressure pump, comprising a pump body, a pump casing and an impeller; the pressure drop between the end face of pump rear cover and the rear end face of impeller auxiliary blade is used for providing wash water to sealing device; a hole is opened in the high pressure area of the end face of pump rear cover; the position of the hole for entering the seal cavity is close to the sealing device, leading expectant high pressure water into the seal cavity; a hole is opened in the low pressure area of the end face of pump rear cover; the position of the hole for exiting the seal cavity is far away from the sealing device; the expectant high pressure water that enters the seal cavity will flow back to low pressure area through the drainage hole along flow direction after washing the sealing device; the pump recycles a part of energy, improves the hydraulic efficiency, decreases the turbulent loss and takes way the gas might exist in the seal cavity; the extended orifice s

Abstract: A heating pump includes a pump housing with an inlet, an outlet, a heater, an impeller, and a motor for driving the impeller. The impeller is received in the pump housing. The heater includes connecting heads and a main body for generating heat. The main body is bent into an arc. The pump housing includes a cylindrical sidewall extending in an axial direction of the motor and a top plate sealed to an axial opening of the sidewall. The top plate is made of metal. The heater extends through and is fixed to the top plate. The connecting heads extend to an exterior of the pump housing. The main body is located in an interior of the pump housing and in contact with the top plate.

Abstract: A gas turbine engine includes a turbine section. The turbine section includes a disk that is rotatable about an axis. A plurality of turbine blades are mounted around a periphery of the disk, and a plurality of seals are arranged between the turbine blades and the periphery of the disk. Each of the seals includes, with respect to the axis, a radially outer surface and a radially inner surface. The radially inner surface includes a plurality of protrusions.

Abstract: According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is no more than 0.45.

Abstract: A turbine exhaust case has an outer housing to be secured within a gas turbine engine and a central hub. Struts extend between the outer housing and the central hub. The struts are formed at least in part of a first material. The central hub is formed at least in part of a second material.

Abstract: A blade or vane arrangement for a gas turbine engine has an array of aerofoils mounted to respective platforms about an axis and defining a passage through which a working gas flow passes. The arrangement has a datum and the aerofoil has a radial span. Each aerofoil has pressure side, a suction side, a leading edge region and a leading edge foot extending from the leading edge region, the leading edge foot has a ridge line. The platform defines a channel and a platform leading edge, the channel has a minimum radial height line, and the platform leading edge partly defines an outlet through which a secondary flow passes. The ridge line is aligned generally in the direction of the working gas flow and the minimum radial height line is aligned generally in the direction of the secondary flow.

Abstract: A rotor blade for a gas turbine configured for use within a row of samely configured rotor blades. The rotor blade may further include: an airfoil defined between pressure and suction faces; and a midspan shroud comprising a pressure wing and a suction wing extending from the airfoil. The pressure wing and the suction wing of the midspan shroud may be configured so to cooperatively form an interface between installed neighboring ones of the rotor blades within the row of samely configured rotor blades. One of the pressure wing and the suction wing may be designated a first wing, and the first wing may include a chamber hollowed through a first surface of the first wing. The first surface of the first wing may include one of: a circumferential face formed at a distal end of the first wing; and a contact face of the first wing.

Abstract: A clearance control ring having at least two segments is disclosed. Each of the segments interlock with adjacent segments to form a full hoop clearance control ring. Separate carriers and seals or one-piece carriers and seals may be mounted on the clearance control ring. The segmented structure of the clearance control ring allows for simpler assembly with segmented cases for gas turbine engines than prior art one-piece clearance control rings. The segmented structure also may be used with one-piece pre-assembled and multi-stage rotors.