Abstract: A propeller. The blades of the propeller are characterized by a blade center axis which corresponds to a generator line of a cone to which the blade, or blade thrust surface, conforms. The propeller hub is fixed to the blade at the root, in line with the blade center axis, such that the hub axis and blade center axis lie in the same plane, and the leading edge of the blade is positioned forward (referring to the upstream direction of movement caused by the propeller) of the trailing edge of the blade.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). The hub has a negative hade angle (?) with respect to the rotational axis at an axial position coincident with the leading edge of the blades.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). A plurality of outlet guide vanes are positioned aft of the fan. An afterbody is located aft of the plurality of outlet guide vanes and which tapers to an apex having an apex angle with respect to the rotational axis of between 35 and 45 degrees.

Abstract: A duct for a ducted-rotor aircraft includes a hub, a duct ring, and a plurality of stators that extend outward from the hub. The duct ring defines a trailing edge. The duct includes one or more fittings, such as a system of fittings. Each fitting has a body that defines a first attachment interface that is configured to couple to structure of the duct ring and a second attachment interface that is configured to couple to one of the plurality of stators. The first and second attachment interfaces are spaced from each other such that when the fittings are coupled to the duct ring and the plurality of stators, all or substantially all of each of the plurality of stators are located aft of the trailing edge of the duct ring.

Abstract: An airfoil includes an airfoil wall that defines a leading end, a trailing end, and first and second sides that join the leading end and the trailing end. A rib connects the first and second sides of the airfoil wall. The rib defines a tube portion that circumscribes a rib passage, and first and second connector arms that solely join the tube portion to, respectively, the first and second sides of the airfoil wall.

Abstract: A drive gearbox for turbomachine equipment, the drive gearbox having a main shaft intended to be rotationally connected to a power transmission shaft of a turbomachine. The main shaft has at least one bevel gear rotationally driving at least one piece of equipment. The gearbox has gear set with cylindrical gearwheels that have an input connected to the rotational drive to the main shaft and a plurality of outputs each intended to drive a piece of equipment. The gear set extends over an angular sector and the outputs are angularly spaced apart along the angular sector.

Abstract: A ring-segment surface-side member is included in a ring segment disposed at a position opposite to a turbine blade on a stationary side of a gas turbine, serves as a part of a combustion gas flow path through which combustion gas flows, and is formed of a ceramic matrix composite. The ring-segment surface-side member includes: a surface portion forming the combustion gas flow path; a turned-back portion including a first part extending outward in a radial direction of the gas turbine from the surface portion and a second part extending toward a central line of the surface portion from an end part of the first part; and a protrusion portion extending outward from the turned-back portion in the radial direction of the gas turbine.

Abstract: An engine assembly with an electric motor that includes a cooling system of which a heat sink portion is arranged in an exhaust cone downstream of the fan driven by the electric motor, that is designed to effect thermal exchanges with a wall of the exhaust cone. Such an electric rear fan arrangement with boundary layer ingestion enables a cold air flow coming from the fan when the engine assembly is in operation to be used to cool the different components of the engine assembly.

Abstract: A blower wheel having a plurality of blower wheel blades arranged in a blade ring, which are connected to a disc covering the blower wheel blades, at least in sections, on at least one axial side, wherein a connection between the blower wheel blades and the disc determines a transition geometry, which has a rounded curve of a quadratic function when viewed in the cross-section, at least on one side of the blower wheel blades, particularly a side facing radially inward with respect to an axis of rotation of the blower wheel.

Abstract: A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct.

Abstract: Airfoil assemblies for gas turbine engines are provided. For example, an airfoil assembly comprises an airfoil, an inner band defining an inner opening shaped complementary to an inner end of the airfoil, and an outer band defining an outer opening shaped complementary to an outer end of the airfoil. The airfoil inner end is received with the inner opening, and the airfoil outer end is received within the outer opening. A strut extends radially through an airfoil cavity. A first pad is defined at a first radial location within the cavity. A second pad is defined within the cavity at a second, different radial location. In some embodiments, the airfoil assembly inner band includes a first inner flange, through which the inner band is secured to a support structure, and the outer band includes a first outer flange, through which the outer band is secured to a support structure.

Abstract: The invention concerns a cooling device (21) extending circumferentially around a turbo machine housing, such as a turbine housing, having at least one cooling module (30a, 30b, 30c) comprising a tube holder (31, 32) having at least one plate, preferably a radially inner plate (31) and a radially outer plate (32) fixed to each other, the tube holder (31, 32) defining at least two axially spaced housings (34a), each housing (34a) receiving a tube (23) and a sleeve (38) mounted around each tube (23), each tube (23) extending circumferentially around the housing, each cooling module (30a, 30b, 30c) having first fastening means capable of firmly holding its tube holder (31, 32) with respect to the housing, the fixing means comprising at least one spacer (41) having a radially inner end intended to bear against the housing and a radially outer end bearing on the single or on one of the sheets of the tube support (31, 32).

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle (501) defining a duct, and a fan located therewithin. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached thereto. Each blade has a span (r) from a root at the hub defining a 0 percent span position (r=0) to a tip defining a 100 percent span position (r=1) and a plurality of span positions therebetween (r ? [0, 1]), and leading and trailing edges defining, for each span position, a chord therebetween to having a chord length (c). For each of said plurality of blades, the ratio of chord length at the 0 percent span position (chub) to chord length at the 100 percent span position (ctip) is 1 or greater.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). A hub-tip ratio of the fan, defined as the ratio of the diameter of the hub to the diameter of the fan measured at the leading edge of the blades, is from 0.45 to 0.55.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle (501) defining a duct (502), and a fan (503) located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub. Each blade has a span (r) from a root at the hub defining a 0 percent span position (r=0) to a tip defining a 100 percent span position (r=1) and a plurality of span positions therebetween (r ? [0, 1]), and a stagger angle at the 0 percent span position (?hub) relative to the rotational axis of 40 degrees or greater.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft includes a nacelle (501) defining a duct (502), and a fan (503) located therewithin. The fan comprises a hub which rotates around a rotational axis (A-A) and a plurality of blades attached thereto. Each blade has a span (r) from a root at the hub defining a 0 percent span position (r=0) to a tip defining a 100 percent span position (r=1) and a plurality of span positions therebetween (r?[0, 1]), a leading edge and a trailing edge defining, for each span position, a chord therebetween having a chord length (c), and a blade thickness (t) defined for each span position thereof. For each blade, a ratio of thickness at the 0 percent span position (thub) to chord length is 0.1 or greater.

Abstract: A gas turbine engine is provided. The gas turbine engine includes an engine casing structure and a part retained relative to the engine casing structure by a channel-cooled hook. The channel-cooled hook includes at least a portion of a hook cooling channel. A vane assembly for the gas turbine engine is also provided.

Abstract: The blades of the propeller are characterized by a blade center axis which corresponds to a generator line of a cone to which the blade, or blade thrust surface, conforms. The propeller hub is fixed to the blade at the root, in line with the blade center axis, such that the hub axis and blade center axis lie in the same plane, and the leading edge of the blade is positioned forward (referring to the upstream direction of movement caused by the propeller) of the trailing edge of the blade.

Abstract: A wind turbine includes at least one wind speed sensor, a number of pitch-adjustable rotor blades, and a control system for changing the pitch of the rotor blades and/or a generator torque. The the control system determines at time intervals an error parameter as the difference between an estimated wind speed and a measured wind speed as measured by the wind speed sensor. Then, based on a number of pre-defined wind speed intervals, a group of error parameters is obtained over time for each wind speed interval. For each wind speed interval and for each group of error parameters a wind speed offset is determined based on the average of the error parameters within the group which wind speed offsets are used in adjusting the measured wind speed.

Abstract: A thrust producing unit for producing thrust in a predetermined direction, comprising at least two rotor assemblies and a shrouding that accommodates at most one of the at least two rotor assemblies, wherein the shrouding defines a cylindrical air duct that is axially delimited by an air inlet region and an air outlet region, and wherein the air inlet region exhibits in circumferential direction of the cylindrical air duct an undulated geometry.