Abstract: Sounds are generated by an aerial vehicle during operation. For example, the motors and propellers of an aerial vehicle generate sounds during operation. Disclosed are systems, methods, and apparatus for actively adjusting the position of one or more propeller blade treatments of a propeller blade of an aerial vehicle during operation of the aerial vehicle. For example, the propeller blade may have one or more propeller blade treatments that may be adjusted between two or more positions. Based on the position of the propeller blade treatments, the airflow over the propeller is altered, thereby altering the sound generated by the propeller when rotating. By altering the propeller blade treatments on multiple propeller blades of the aerial vehicle, the different sounds generated by the different propeller blades may effectively cancel, reduce, and/or otherwise alter the total sound generated by the aerial vehicle.

Abstract: An injection mold system for injecting slurry into an internal portion of an injection mold component core is disclosed. The system may include a fill clip for injecting slurry into an internal portion of an injection mold component core including a fill member for injecting slurry into the internal portion of the injection mold component core, and a latch member for removably coupling the fill member with the injection mold component core. The system may also include a fill reservoir for collecting excess slurry from the internal portion of the injection mold component core. The fill reservoir may also redistribute the excess slurry back into the internal portion of the injection mold component core.

Abstract: The present invention provides a stator component of a vacuum pump, which is suitable for reducing the fracture energy (energy of fracture that occurs when a rotor of the pump is damaged during its rotation) and the size of the pump, and also provides a vacuum pump having this stator component. In the vacuum pump, a spacer or of a thread groove pump stator, which is a stator component forms a gap satisfying the following <<condition>> between an outer circumferential surface of each of housed in a pump case of the vacuum pump, and an inner circumferential surface of the pump case, with the stator component being housed in the pump case. <<Condition>> 2d/D??max, where D is the outer diameter of the stator component (spacer or thread groove pump stator), d is the width of the gap, and ?max is the breaking elongation of the stator component.

Abstract: An airfoil for a gas turbine engine includes a substrate and a sheath providing an edge. A cured adhesive secures the sheath to the substrate. The cured adhesive has a fillet that extends beyond the edge that includes a mechanically worked finished surface. A method of manufacturing the airfoil includes the steps of securing a sheath to a substrate with adhesive, curing the adhesive, and mechanically removing a portion of the adhesive extending beyond the sheath.

Abstract: A turbomachine rotor blade includes an outer part at its distal end. The outer part includes a platform defining an outside surface of a passage for gas passing through a turbomachine and presenting first and second opposite side edges; and upstream and downstream sealing wipers extending outwards from the platform, each wiper extending between two lateral faces situated respectively at the first and second side edges. The two lateral faces of the upstream or the downstream wiper are covered at least in part in an anti-wear material.

Abstract: A partially coated blade for a gas turbine engine, including a fillet surface surrounding the airfoil section and connecting it to the platform section. A radially outermost portion of the pressure side and leading edge is covered by a thermal barrier coating. This portion extends radially from a first limit to the blade tip. The first limit is located at a radial distance from the platform of at most 21% of the maximum span. The fillet surface is free or substantially free of the thermal barrier coating. In another embodiment, a second portion of the pressure side and of the leading edge is free or substantially free of the thermal barrier coating, extending radially from the platform section to a second limit located a radial distance from the platform section corresponding to at least 5% of the maximum span. A method of applying a thermal barrier coating is also discussed.

Abstract: A fan blade for an aircraft engine, including a leading edge, a trailing edge, a suction side, a pressure side and a blade tip, is provided. The fan blade has a large-area elastomer layer which takes up at least 20% of the surface of the suction side of the fan blade.

Abstract: A method for patching a damaged polymeric erosion coating (250) on a gas turbine engine component. The method comprises removing a portion of the coating around a damage site (200; 202; 204) and applying a pre-formed coating patch (220).

Abstract: A turbine blade having a root and a tip and an airfoil with a core section made of composite material and a protective sheath or layer joined to the composite core at a location or locations exposed to erosion is described with the composite core of the airfoil being continuous from a location in vicinity of the root to a location in vicinity of the tip of the blade and including a section which widens with distance from the rotational axis along the length of the airfoil and which ends at the location in vicinity of the tip of the blade. The protective sheath or layer is secured by interference fit with widening section of the core.

Abstract: A gas turbine component is provided. The gas turbine component includes an airfoil having a leading edge, a trailing edge, a suction side extending from the leading edge to the trailing edge, and a pressure side extending from the leading edge to the trailing edge opposite the suction side. The gas turbine component also includes a thermal barrier coating applied to the airfoil pressure side such that an uncoated margin is defined on the pressure side at the trailing edge.

Abstract: A composite turbine blade for high-temperature applications such as gas turbines or the like includes a root for mounting the blade in a corresponding circumferential assembly groove of a rotor and an airfoil connected to said root. An inner carrying structure is provided extending at least over a portion of the root as well as at least a portion of said airfoil. The inner carrying structure is made of a high strength eutectic ceramic and the airfoil is made of a ceramic matrix composite (CMC) material.

Abstract: A method of making a fan annulus filler for use in a fan assembly of a gas turbine engine may include an annulus filler body formed of a polymer. The annulus filler body may include a surface coated with a nanocrystalline coating to form a flow surface.

Abstract: A scroll conveyor tile protection or any other material subject to wear with a weldable support and a wear resistant insert, not necessary weldable. The support is fit with a flat surface and shoulder allowing for precise positioning of tile on a scroll flight to be protected. At least one opening, preferably at its center, allows for insertion and assembly of a complementary shaped insert. The unit setting up on the scroll flight and the contact between the flight faces and the unit faces allows for creating a dove tail locking system. Welding of the support permanently insures locking of the system and insert being maintained without being otherwise fixed beyond this mechanical clamping.

Abstract: A fan assembly for gas assembly engine is disclosed. The fan assembly includes a rotor that is coupled to at least one fan blade. The fan assembly also includes a spinner that extends towards the rotor and the fan blades so that an edge of the spinner engages the fan blades. The edge of the spinner thereby serves as a grounding element for the fan blade. If the spinner is fabricated from a composite or a plastic material that is not sufficiently conductive for grounding purposes, the edge may be coated with a conductive material. Alternatively, the spinner may be fabricated from a metallic material, which would eliminate the need for a conductive coating on the edge that engages the fan blades.

Abstract: An integrally bladed rotor includes a rim integral with a web that extends radially inward to a bore. The rim provides an end wall from which integral blades extend radially outward to a tip. The blades have an airfoil that extends in a chord-wise direction from a leading edge to a trailing edge. A fillet circumscribes the airfoil and joins the airfoil to the end wall. The fillet is at one of the leading edge and the trailing edge is truncated at least 50% in an axial direction to provide a face of the rim.

Abstract: A metallic leading edge protective strip adapted to provide impact protection for a leading edge of an airfoil of a turbomachine. The protective strip is formed of a stainless steel or a nickel-based alloy and is denser and provides increased strength and elasticity characteristics as compared to an identical protective strip formed of a titanium-based alloy. The protective strip is particularly suitable for use with composite blades and allows for thinner airfoils, thereby improving engine efficiency.

Abstract: A ceramic matrix composite member used as a turbine blade includes a principal part forming a blade part and a dovetail part, and a subordinate part forming a platform part. A principal fiber in a ceramic fiber fabric forming the principal part is a continuous fiber. An extension direction of the principal fiber is in parallel with a direction in which stress is applied. The ceramic fiber fabrics respectively forming the principal part and the subordinate part are joined together and formed into an integrated three-pronged fiber fabric. The ceramic fiber fabric forming the principal part and the ceramic fiber fabric forming the subordinate part are integrated together by being set into a mold with the ceramic fiber fabric forming the subordinate part folded at a desired angle to the ceramic fiber fabric forming the principal part. Then, a ceramic matrix is formed in the obtained molded body.

Abstract: Fans having fan blades with lightweight blade tips are disclosed herein. In one embodiment, for example, each fan blade includes a blade tip attached to an end portion of a main airfoil. The blade tip includes a shaped body that is molded over a support structure to form the shape and exterior surface of the blade tip.

Abstract: A fan rotor blade includes a blade body made from a composite material of a resin and fibers, and having a pressure surface facing one side in a rotation direction and a suction surface facing the other side in the rotation direction. The blade body is provided with a sheath which is more rigid than the blade body and is fixed to the blade body in a state where the sheath covers a leading edge portion of the blade body, which is located on an upstream side in the direction of intake of external air, and the vicinity thereof. A tip end portion of the sheath extends from the leading edge portion side of the blade body to a trailing edge portion side of the blade body. An end portion of the tip end portion of the sheath is flush with the trailing edge portion of the blade body.

Abstract: A component of a gas turbine engine is provided. The component has a leading edge with two air-washed surfaces extending rearwardly therefrom to define opposing sides of the component. The component has a leading edge protector formed as a plurality of protector segments which succeed in a line along the leading edge. Each protector segment includes a sheath which covers the leading edge and the air-washed surfaces at regions adjacent the leading edge. Each protector segment overlaps with the neighboring sheath of the succeeding segment.

Abstract: A root section of a rotor blade for interacting with working fluid upon rotating the rotor blade is provided. The root section includes a curved cooling passage for guiding a cooling fluid within the root section. A cooling fluid entry plenum has an entry aperture for introducing the cooling fluid into the cooling passage. A platform is located at a radially outer end of the root section. The curved cooling passage penetrates through the platform, and the following condition is satisfied at least in a portion of a radial extent of the cooling passage: 0.25*dr<rc<1.5*dr, where dr is a radial distance in the radial direction between the platform of the root section and the aperture of the entry plenum and rc is the radius of curvature of the curved cooling passage.

Abstract: A composite airfoil includes a core that has a three-dimensional network of fibers. The core defines an airfoil section and a root section. A composite skin covers a portion of the core excluding the root section.

Abstract: An airfoil includes an airfoil body that defines a longitudinal axis. The airfoil body includes a leading edge and a trailing edge and a first side wall and a second side wall that is spaced apart from the first side wall. The first side wall and the second side wall join the leading edge and the trailing edge and at least partially define a cavity in the airfoil body. A lattice network connects the first side and the second side. The lattice network includes at least one enlarged node spaced apart from the first side wall and the second side wall and ribs that extend from the at least one enlarged node. Each of the ribs connects to one of the first side wall and the second side wall.

Abstract: A rotor blade assembly for a wind turbine has a rotor blade. A strip-member blade accessory is mounted to one of the rotor blade surfaces that stretches under normal operating conditions of the wind turbine. The blade accessory has a base portion. An attachment layer connects the base portion to the rotor blade surface and has a lower shear modulus than the base portion to allow for shear slippage between the base portion and the underlying rotor blade surface.

Abstract: This invention relates to a stator aerofoil for a gas turbine engine, comprising: a body portion; at least one panel which forms at least part of one of either the pressure or section surface of the aerofoil; at least one internal chamber which is bounded by the body portion and panel; and, at least one elastomeric component within the at least one internal chamber, wherein the elastomeric component at least partially defines a cell within the internal chamber, wherein the cell is in fluid communication with the exterior of the aerofoil.

Abstract: An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a total chord length and a span position and corresponds to a curve that has an increasing total chord length from the 0% span position to a first peak. The first peak occurs in the range of 45-65% span position, and the curve either remains generally constant or has a decreasing total chord length from the first peak to the 100% span position. The total chord length is at the 0% span position in the range of 8.2-10.5 inches (20.8-26.7 cm).

Abstract: A blade insert for coupling a first blade segment to a second blade segment is disclosed. The blade insert may generally include an aerodynamic body extending between a forward end configured to be coupled to the first blade segment and an aft end configured to be coupled to the second blade segment. The aerodynamic body may include a top side extending between a forward edge and an aft edge. The top side may define a top scarfed section at its forward edge. The aerodynamic body may further include a bottom side extending between a forward edge and an aft edge. The bottom side may define a bottom scarfed section at its forward edge. Additionally, at least a portion of the forward edge of the top side may be configured to be offset relative to the forward edge of the bottom side.

Abstract: In a method of making a leading edge component for an airfoil device, a first elongate sheet is provided and has a shape substantially conforming to a surface on one of a suction side or a pressure side of an airfoil. A second elongate sheet is provided and has a shape substantially conforming to a surface on the other of the suction side or the pressure side of the airfoil. The elongate sheets can be welded together at least along respective elongate edges thereof, in such a way as to form a leading edge base structure comprising a weld extending the length of the elongate sheets. Material deposits can be provided onto an outer surface of the leading edge base structure in such a way as to substantially cover the weld, whereby the leading edge component is formed.

Abstract: A turbine airfoil includes an airfoil body with a leading edge, trailing edge, and an exterior surface including a suction side and a pressure side located opposite to the suction side. The exterior surface shows away from the interior of the airfoil body. A thermal barrier coating system is present on the exterior surface of the airfoil body in a coated surface region. The airfoil also includes an uncoated surface region where a thermal barrier coating system is not present. The uncoated region extends on the suction side of the exterior surface from the trailing edge towards the leading edge to a boundary line between the coated surface region and the uncoated surface region, wherein the boundary line is located on the suction side between the leading edge and the trailing edge. The airfoil body also includes a step in the exterior surface extending along the boundary line.

Abstract: A sectional blade for a wind turbine, the blade extending lengthwise between a tip and a root where the blade is attachable to a hub of a wind turbine, crosswise between a leading edge and a trailing edge, and thickness wise between a windward and a leeward outer surface. The sectional blade comprises a first blade portion and a second blade portion. The blade portions are formed as separate components and arranged to extend lengthwise in opposite directions from a junction where the blade portions join. The blade forms, in a cross section perpendicular to the lengthwise direction and through the junction, a maximum thickness location between the leading and trailing edges, the maximum thickness location being where the distance between the windward and leeward outer surfaces is larger than at any other location in this cross section.

Abstract: An article may include an array of features formed in a substrate and may be coated with a coating layer. The array of features may mitigate stress experienced by the coated article. In particular, the array of features may reduce or limit crack propagation at the interface between the substrate and the coating layer. In some examples, the article is an airfoil that includes a tip that defines an edge. An array of features is formed on the surface of the tip, where the array of features is proximate to the edge, and the array of features does not intersect the edge. The airfoil includes a coating layer formed on the surface of the tip and the array of features.

Abstract: A component, designed so that the longitudinal grain which experiences the greatest stress has a larger grain width than the other grains, is provided. This longitudinal grain is usually on the edge of the component. The component includes longitudinal grains that are directionally solidified in a rod-crystalline form along the longitudinal axis of the component.

Abstract: Superalloy components are joined by mating a recess formed in one component with a corresponding projection formed in another component along a contact surface. The components are compressed along the contact surface and resistance heat brazed to each other. Current is passed between the components at a selected flow rate and application time until brazing alloy melting occurs along the contact surface, and they are mutually affixed to each other. When repairing a damaged surface portion of a superalloy material component, the damaged portion is removed to form an excavated recess. A repair splice is formed, preferably of a same material with similar mechanical structural properties, having a mating projection with profile conforming to the corresponding recess profile. The splice and substrate are resistance heat brazed under compression pressure until brazing alloy melting occurs along the contact surface, so that they are mutually affixed.

Abstract: Structural repair of cracks and other defects in superalloy components, such as steam or gas turbine blades in stationary or aero gas turbines, are performed by heating the blade substrate to an isothermal hold temperature below the substrate's incipient melting point and filling the crack with molten superalloy filler material. The molten filler solidifies into a casting and bonds with the component substrate at the isothermal hold temperature. Heat treatment processes are completed, so that the former crack is filled with cast superalloy material having identical or similar structural properties as the adjoining substrate superalloy material. The casting repair method may be utilized universally for all types of superalloy component defects, including those previously repaired by cosmetic, lower strength welding or brazing methods.

Abstract: A fan blade is disclosed comprising a lightweight metallic airfoil portion and a high-strength sheath portion. The airfoil portion has a forward airfoil edge, a first airfoil surface, and a second airfoil surface. The sheath portion has a sheath head section, a first sheath flank, and a second sheath flank, both flanks extending chordwise from the forward sheath section. The sheath portion is bonded to the airfoil portion such that the sheath head section covers the forward airfoil edge, defining a blade leading edge. The first sheath flank covers a portion of the first airfoil surface proximate the airfoil forward edge, jointly defining a blade suction surface. The second sheath flank covers a portion of the second airfoil surface proximate the airfoil forward edge, jointly defining a blade pressure surface.

Abstract: A method of forming a leading edge protection component includes depositing particles using a cold spray process on a mandrel to form a leading edge protection component; and removing the leading edge protection structure from the mandrel. The leading edge protection can be formed in one or more pieces and involve using one or more mandrels.

Abstract: A method of manufacturing a longitudinally extending composite structure including a shell part comprising a fiber reinforced polymer material including a polymer matrix and fiber material embedded in the polymer material is described. The shell part is manufactured in a closed mold comprising at least a first outer mold part having a first forming surface and a second outer mold part having a second forming surface. The method comprises the steps of: arranging a first fiber material in the first forming surface of the first outer mold part, arranging a pre-fabricated longitudinally extending reinforcement element, such as a beam or a web, on top of the first fiber material, arranging a second fiber material in the second forming surface of the second outer mold part, and sealing a polymer foil above the second fiber material so as to retain the second fiber material against the second forming surface.

Abstract: An interface element to be mounted between a blade root and a blade root housing provided in a turbine disc of a gas turbine engine to limit wear between the root and the housing. The interface element includes: a base, configured to be aligned with a lower part of the root; and first and second upper side walls connected to the base and configured to surround the blade root up to an upper portion thereof, the first upper side wall including at least one ventilation opening to allow a flow of cooling air flowing through the rotor disc housing to flow over the upper portion of the root via the ventilation opening.

Abstract: A propeller blade assembly including a propeller blade having a leading edge, and a trailing edge extending between a tip and a hub, and an electrically conductive band extending longitudinally on either or both of the leading edge and trailing edge. The electrically conductive band secured to the leading edge of the propeller serves as an entry point on the aircraft and conductive channel into the airframe thereby avoiding the potentially damaging effect of the lightening on the carbon fiber composite propeller.

Abstract: An erosion shield for a wind turbine blade is described, the erosion shield having a plurality of layers of erosion resistant material. The layers of erosion resistant material have an adhesive bond strength between adjacent layers less than the cohesive tensile strength of the layers, such that the outer layers of erosion resistant material are arranged to peel away or delaminate from the erosion shield under the action of the wind once the particular layer is ruptured or eroded. This dynamic removal of the outer layers of the erosion shield provides for increased shield lifetime, and a reduction in the maintenance operations required for a wind turbine blade having such an erosion shield.

Abstract: A method of manufacturing a filled aerofoil component (100) for a gas turbine engine (10) comprises using a capping panel (200) to cover a pocket (310) in a pocketed aerofoil body (300). During manufacture, a pre-formed filler insert (400) is provided to support the capping panel (200) in the correct position. This ensures that the outer surface of the capping panel (200) is located as accurately as possible. This means that the capping panel (200) can be made to be as thin as possible, which in turn reduces weight and material wastage. The pre-formed filler insert (400) provides a lightweight core for the finished aerofoil component in use.

Abstract: A protective cap for a trailing edge of a rotor blade of a wind turbine for use during transportation, handling, or maintenance of the rotor blade is disclosed. The protective cap includes a body having a first leg, a second leg, and a cap member. The cap member connects the first and second legs. Further, the cap member may be configured to cover at least a portion of a trailing edge of the rotor blade and may be configured to provide a gap between an inner surface of the protective cap and the trailing edge. In addition, the first and second legs may resiliently engage the rotor blade.

Abstract: A fan blade of a gas turbine engine includes an airfoil, a root received in a slot of a hub, and a wear pad covering at least a portion of the root and made of an integrally bonded woven composite laminate.

Abstract: A fan rotor blade includes a blade body made from a composite material of a resin and fibers, and having a pressure surface facing one side in a rotation direction and a suction surface facing the other side in the rotation direction. The blade body is provided with a sheath which is more rigid than the blade body and is fixed to the blade body in a state where the sheath covers a leading edge portion of the blade body, which is located on an upstream side in the direction of intake of external air, and the vicinity thereof. A tip end portion of the sheath extends from the leading edge portion side of the blade body to a trailing edge portion side of the blade body. An end portion of the tip end portion of the sheath is flush with the trailing edge portion of the blade body.

Abstract: A fan rotor blade includes a blade body made from a composite material of a resin and fibers, and having a pressure surface and a suction surface. A dovetail capable of being fitted into a fitting groove in a fan disk is provided in a base end of the blade body. The blade body is provided with a sheath. The sheath is more rigid than the blade body and is fixed to the blade body in a state where the sheath covers a leading edge portion of the blade body and the vicinity thereof. The sheath extends in the longitudinal direction of the blade body from a blade root side to a tip end side of the blade body, and a base end portion of the sheath located on the blade root side of the blade body extends to the dovetail.

Abstract: A wind turbine blade includes a multilayer composite structure including a first reflective layer, and a second layer including a plurality of resistive circuit analog (CA) elements. The CA elements are tuned so as to interact with the first layer to provide absorption of electromagnetic (EM) energy over a desired frequency range. The parameters of the CA elements can be varied to provide for frequency tuning and to maintain absorption at a specific frequency range despite varying layer separation, while at the same time ensuring that the mechanical properties of the CA layer are compatible with integration into the turbine blade.

Abstract: A method of making a metal reinforcing member that is to be mounted on a leading edge or a trailing edge of a composite blade of a turbine engine, the method including: shaping two metal sheets, positioning them on either side of a core including at least one recess that is to form a mold for a spacer for positioning the reinforcing member, assembling them together under a vacuum, conforming them against the core by hot isostatic compression, and cutting them to separate the reinforcing member and release the core.

Abstract: A method of fabricating a near-net shape erosion shield, a method of forming a shielded article, and a near-net shape erosion shield are provided. The method of fabricating a near-net shape erosion shield includes providing a base, positioning an energy source relative to the base, and depositing at least one wear resistant material over the base with an energy beam from the energy source. The at least one wear resistant material deposited on the base forms the near-net shape erosion shield configured to be positioned on a turbine component. The method of forming a shielded article includes removing the base from the near-net shape erosion shield, and securing the near-net shape erosion shield to a turbine component. The near-net shape erosion shield includes a near-net shape erosion-resistant portion configured to be positioned on a turbine component.

Abstract: One embodiment of the present application includes a hot section component of a gas turbine engine having a covering. The covering includes a protrusion and is attached to the hot section component though a flexible retainer. In one form the covering is made from ceramic matrix composite. The flexible retainer has a closed position and an open position. The retainer secures the protrusion to the hot section component when it engages part of the protrusion when in the closed position.

Abstract: A protected turbine blade includes a first turbine blade shell and a second turbine blade shell, being fiber-reinforced, are connected by a cured resin and provided with a protective cover at a leading edge. The protective cover is a composite comprising a layer of UV-resistant thermoplastic material and a layer of cured epoxy resin, the UV-resistant thermoplastic material having the following properties a surface free energy of less than 35 mJ/m2, and a erosion resistance of the thermoplastic material larger than the erosion resistance of the cured resin connecting the turbine blade shells. The layer of cured epoxy-resin covers the cured resin over at least part of the length of the distal half of the turbine blade, and is bonded to said turbine blade. A method of manufacturing such a turbine blade, and to a turbine having such turbine blades is also provided.