Abstract: A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (?3??1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

Abstract: A method, device and system for determining angle-to-wind deviation and correcting angle-to-wind; the method for determining angle-to-wind deviation comprises: obtaining historical operation data of a wind turbine group during a specific time period (S101); determining an angle-to-wind deviation value for each wind speed segment on the basis of the acquired historical operation data (S102); for any wind speed segment, determining the angle-to-wind deviation value on the basis of the actual angle-to-wind measurement value and the output power value of an environmental wind speed value at a time point within the specific time period.

Abstract: A wind turbine blade and associated method of manufacture is described. The blade comprises an outer shell formed of first and second half shells joined together. A shear web is arranged inside the outer shell. The shear web has a web panel disposed between first and second longitudinally-extending mounting flanges. The shear web is bonded to inner surfaces of the respective half shells via a first adhesive bond line between the first mounting flange and the inner surface of the first half shell and a second adhesive bond line between the second mounting flange and the inner surface of the second half shell. One or more bond spacers are provided in the second bond line, and optionally in the first bond line. The bond spacers are compressed between a shear web mounting flange and an inner surface of a half shell and are plastically deformed. The method of making the shear web involves compressing and the one or bond spacers in the bond line(s) such that they undergo plastic deformation.

Abstract: Provided is a method of manufacturing wind turbine rotor blades, wherein each rotor blade includes an inboard section and an outboard section, and wherein an inboard blade section including a root end and a transition region is manufactured using a first casting process; and an outboard blade section including an airfoil region is manufactured using a second casting process, which second casting process is different from the first casting process. The invention further describes a wind turbine rotor blade manufactured using such a method.

Abstract: A lubrication system for a gas turbine engine, the system comprising: a gearbox, the gearbox comprising a sump; an oil tank; a primary gearbox lubrication system configured to pump oil from the oil tank to lubricate the gearbox with a gearbox primary feed; a secondary gearbox lubrication system, configured to lubricate the gearbox with oil from the sump when the oil level in the sump reaches a predetermined level; wherein the system is configured to increase the oil level in the sump to at least the predetermined level in response to a failure of the primary gearbox lubrication system.

Abstract: The invention relates to a method of separating at least one portion of a metallic part (3) glued to a composite material part (2) of the carbon-epoxy type, comprising a step to degrade the glued interface between the metallic part (3) and the composite material part (2). The metallic part (3) and the composite material part (2) are electrically connected to a dc electrical voltage generator so that an electrical potential difference can be applied to them to generate partial discharges in the glued interface to degrade the interface.

Abstract: A centrifugal pump includes at least one pump stage, with a rotatable impeller (5) with a suction port (6) which is sealed with respect to a stationary pump part (1) by way of a sealing arrangement. The sealing arrangement includes a sealing ring (9) between the impeller (5) and the stationary pump part (1). The sealing arrangement is configured such that at least on delivery operation of the pump, the sealing arrangement has sealing sections which are distanced to the counter sealing surface and sealing sections which bear on the counter sealing surface, in an alternatingly successive manner considered in the peripheral direction of the sealing ring (9).

Abstract: A method of controlling a delivery pressure of an aircraft engine fuel system includes operating a rotodynamic fuel pump hydraulically connected to the fuel system to pump a fluid from an inlet of the rotodynamic pump and out of an outlet of the rotodynamic pump for delivery to the fuel system. The method further includes varying an angular relationship between the inlet and the outlet of the pump to control the delivery pressure. An aircraft engine fuel system and a rotodynamic pump are also described.

Abstract: An aerial vehicle includes an aerial vehicle body, and first and second rotor assemblies. The first rotor assembly includes a first hub coupled to first blades, and a first drive shaft coupled to the first hub via fastening features. The first drive shaft is configured to rotate the first hub in a first direction such that the first blades rotate in a first rotation plane. The second rotor assembly includes a second hub coupled to second blades, and a second drive shaft coupled to the second hub via fastening features. The second drive shaft is configured to rotate the second hub in a second direction opposite to the first direction, such that the second blades rotate in a second rotation plane. The first and second rotation planes are at opposite sides of the aerial vehicle body.

Abstract: The present disclosure relates to a fan blade for a gas turbine engine, the fan blade comprising an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span. A maximum thickness of the cross-sections through the aerofoil portion from a suction surface of the aerofoil portion to a pressure surface of the aerofoil portion perpendicular to the camber line decreases along the blade span from the root to the tip. There is a discontinuity in the rate of decrease of maximum thickness between a radius at 30% of the blade span from the aerofoil root and 70% of the blade span from the aerofoil root. The rate of decrease of the maximum thickness before the discontinuity is less than the rate of decrease of the maximum thickness after the discontinuity.

Abstract: A balancing system and method for an engine of an aircraft includes a housing retaining a magnetorheological fluid including magnetic particles within a carrier fluid. Electromagnets are coupled to the housing. The electromagnets are associated with fan blades of the engine. A balancing control unit is in communication with the electromagnets and sensors of the fan blades.

Abstract: A turbofan engine includes a first spool including a first turbine and a second spool including a second turbine. A superposition gear system includes a plurality of intermediate gears engaged to a sun gear and supported in a carrier and a ring gear circumscribing the intermediate gears. A second tower shaft is engaged to drive the sun gear. A starter is selectively coupled to the sun gear through a starter clutch. A first electric motor and a second electric motor are coupled to the superposition gear system and are operable to input power into a corresponding one of the first and second spools.

Abstract: A gas turbine rotor includes: a first rotor in a shaft shape extending by a predetermined length in an axial direction, and provided with an insertion hole at an end of the first rotor; a second rotor in a shaft shape extending by a predetermined length in the axial direction, and provided with an insertion part extending by a predetermined length at an end of the second rotor and corresponding to the insertion hole such that the insertion part is inserted into the insertion hole; and a connecting member mounted between the insertion hole and the insertion part to connect the first rotor and the second rotor together, and made of a material having a thermal expansion coefficient different from a thermal expansion coefficient of both the first rotor and the second rotor.

Abstract: A method of improving a blade and also an improved blade and a advancement propeller including the improved blade. The radius of the initial leading edge circle of each airfoil of the blade is increased, and its leading edge is moved away from a pressure side half-airfoil towards a suction side half-airfoil, thereby modifying the airfoil of each cross-section of the blade and modifying the camber of each airfoil. Consequently, the absolute value of the negative stall angle of attack of the blade is increased, thus making it possible to increase the aerodynamic performance of the blade under a negative angle of attack compared with a blade that is not modified, and without significantly degrading its aerodynamic performance under a positive angle of attack.

Abstract: There is provided a wind turbine comprising a tower which has a wall, a cooling device having an air inlet unit and an air outlet unit which are provided in the lower region of the tower. The air inlet unit and the air outlet unit each have an outer portion, a central portion and a shutter unit. The outer portions are provided outside the wall of the tower, the central portions are provided in the region of the wall of the tower and the shutter units are provided within the tower. The air inlet unit and the air outlet unit have fixing units, by means of which the central portions can be fixed to the wall of the tower from the interior of the tower.

Abstract: A gas turbine engine includes a device for providing sealing between one rotor section and one stator section. The device includes a coating made of an abradable material attached to the stator section. The device further includes a lip on a portion of the rotor section. The lip is configured to form a seal with the abradable material. The gas turbine engine further includes passages for a gaseous fluid and means for blowing such gaseous fluid. The passages open into the rotor section provided with the lip, so that blown gaseous fluid can be present in a zone radially located between the coating and the lip.

Abstract: An aircraft gas turbine engine has an engine core having a turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine and having a turbine length being the distance between the roots of the most upstream turbine blade at its leading edge and of the most downstream turbine blade trailing edge, and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine; and a gearbox receiving an input from the core shaft and outputting drive to the fan. The engine core has three bearings to support the core shaft, the three bearings having a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings, and wherein further a minor span to turbine length ratio of: minor ? ? span tu ? ? rbine ? ? length is equal to or less than 1.05.

Abstract: Controlling a wind turbine comprising a wind direction sensor, a yawing system, and a control system for turning the wind turbine rotor relative to the wind. Over one or more time intervals a data set comprising a direction of the wind relative to the wind turbine as measured by the wind direction sensor and a wind power parameter determined as one of a power, a torque, or a blade load of the wind turbine are obtained. The data sets are sorted into a number of bins of different intervals of wind power parameter. or each power bin, a statistical representation of the wind power parameter as a function of the relative wind direction is determined and then used in estimating a wind direction offset corresponding to the relative wind direction where the wind power parameter attains a peak value. The relative wind direction is then adjusted as a function of the set of wind direction offsets to yield more accurate wind direction data which can be used in controlling the turbine.

Abstract: A method of health monitoring of a gas turbine engine includes mounting a detection system configured to detect an aeromechanical damping characteristic of a row of airfoils of a gas turbine. An actual aeromechanical damping characteristic of the row of airfoils is measured with the detection system. An output signal is generated indicative of the actual aeromechanical damping characteristic of the row of airfoils. A current flutter characteristic is determined based on the output signal indicative of the actual aeromechanical damping characteristic of the row of airfoils. An airfoil health monitoring system for gas turbine engine and a gas turbine engine are also disclosed.

Abstract: A fan blade for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that has a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, and extends radially between a root portion and a tip portion. A tip pocket is defined by collapsible external walls of the tip portion, and has a pocket opening at a radially outer face of the tip portion.