Abstract: A turbomachinery component with a surface that includes a bounded wear coat, the component includes: a body; a contact surface defined by the body; a recess extending into the body and communicating with the contact surface; and a wear coat positioned in the recess.

Abstract: An autonomous unmanned aerial vehicle for land, sea and air use. The autonomous unmanned aerial vehicle is more specifically related to an unmanned aerial vehicle, wherein the autonomous unmanned aerial vehicle is configured to vertically take off and vertically land, fly with fixed wings and stay in the air silently for a long time by means of a balloon inflated behind it.

Abstract: Blades (20a to 20c) of a propeller fan (10) have different circumferential pitches ?1, ?2, and ?3. The blades (20a to 20c) have different masses so that the center of gravity of the propeller fan (10) is positioned on a rotational center axis (11) of the propeller fan (10). Blade body portions (42c) of the blades (20a to 20c) have different thicknesses. In contrast, camber lines of the blades (20a to 20c) in blade cross section have the same shape, projections of the blades (20a to 20c) on a plane perpendicular to the rotational center axis (11) of the propeller fan (10) have the same shape, and leading edge portions (41a to 41c) of the blades (20a to 20c) have the same shape. As a result, a propeller fan (10) having reduced noise and vibrations can be achieved.

Abstract: Disclosed is a gas turbine including a housing, a rotor rotatably provided in the housing to transfer a rotary force to a compressor, the compressor receiving the rotary force from the rotor and compressing air, a combustor mixing a fuel with the compressed air supplied from the compressor and igniting the mixture of the fuel and the air to generate combustion gas, and a turbine receiving the rotary force caused by the combustion gas generated by the combustor and rotating the rotor by using the received rotary force.

Abstract: Method for constructing impingement/effusion cooling features in a component of a combustion turbine engine is provided. A pocket 102 may be arranged between an outer wall 104 and an inner wall 106 of the component. A lasing device 108 allows drilling through the component to form an effusion hole 110. The lasing device further allows welding closed an opening 117 formed at outer wall 104 of the component during the drilling with the lasing device through the component. Lasing device 108 further allows drilling through outer wall 104 of the component to form an impingement hole 118 for the impingement/effusion cooling feature. The proposed methodology in a multi-panel arrangement, for example, eliminates a need of having to pre-drill such holes in individual panels prior to the bonding and forming of the component, which overcomes various drawbacks commonly associated with such pre-drilling.

Abstract: A gas turbine engine includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, a first compressor coupled with the first shaft, and a second turbine coupled with the second shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

Abstract: A turbulence generating structure for liner cooling enhancement is a liner cooling structure applied to a double-structured side portion formed by a liner and a flow sleeve and includes the liner and a first turbulence generator protruding from a surface space of the liner and including a plurality of ribs arranged in an axial direction, each of the ribs comprising blocks arranged at regular distances in the axial direction and interspaced by cooling holes. The ribs include a first set of adjacent ribs separated by a first passage distance, the first passage distance being repeated in a circumferential direction, and a second set of adjacent ribs separated by a second passage distance, the second passage distance being repeated in the circumferential direction, wherein the first and second sets of adjacent ribs have exactly one rib in common, and the first passage distance is greater than the second passage distance.

Abstract: In an embodiment, an apparatus includes: a gearbox including: a housing having an opening; an input pinion disposed in the housing; a moving feature mechanically coupled to the input pinion, the moving feature and the input pinion being operable to convert between power and torque; and first sensors disposed on the moving feature, the first sensors having batteries, where the moving feature is operable to rotate about a common axis, each of the first sensors being exposed by the opening as the moving feature rotates about the common axis.

Abstract: Provided is a machining device for machining a rotor arranged in a turbine housing which includes two parts, thereby forming a joint, in situ while the upper housing half is removed. The machining device includes a main part that releasably secures the main part to the lower housing half in the region of the joint, at least one tool holder for receiving at least one machining tool which is secured on the main part directly or indirectly so as to be alignable relative to the main part, multiple telescopic stabilizing elements which can be locked into different longitudinal settings preferably in a continuous manner.

Abstract: An improved combustion section for a gas turbine engine is disclosed. A fuel nozzle includes new features which provide improved injection patterns of oil fuel and cooling water, resulting in better control of combustion gas temperature and NOx emissions, and eliminated impingement of cooling water on walls of the combustor. A new combustor includes a plate-fin design which provides improved cooling, while the combustor also makes more efficient use of available cooling air and has an improved component life. A new transition component has a smoother shape which reduces stagnation of combustion gas flow and impingement of combustion gas on transition component walls, improved materials and localized thickness increases for better durability, and improved cooling features for more efficient usage of cooling air.

Abstract: A turbine unit for a hydraulic installation, and more in particular it deals with the hub of the turbine unit. The present invention proposes to provide means for adjusting a gap extent formed between the hub and the inner edge of the blade, this way dramatically increasing the performance of the turbine.

Abstract: A process of coating a substrate containing silicon with an environmental barrier coating, comprising altering a surface of the substrate and applying an environmental barrier layer to the surface of the substrate.

Abstract: An example high-performance system may include an example high-performance component. The high-performance component may include a substrate defining a channel. The channel defines a leading ramp and a trailing ramp. The example high-performance component includes an abradable track between the leading and the trailing ramps. The abradable track includes a porous abradable composition. The example high-performance system may include a rotating component configured to contact and abrade the abradable track. An example technique for forming the abradable track includes thermal spraying a precursor composition at the channel to form the abradable track.

Abstract: A method of producing an abrasive tip for a turbine blade includes producing or obtaining a metal powder that is mixed with an abrasive ceramic powder and producing or obtaining a metallic mold that is in the shape of an airfoil. The metallic mold includes a hollow interior portion. The method further includes sealing the metal and ceramic powder mixture within the hollow interior portion of the metallic mold under vacuum and subjecting the sealed mold to a hot isostatic pressing process. The hot isostatic pressing process compacts and binds the metal and ceramic powder mixture together into a solid article in the shape of the airfoil. Still further, the method includes slicing the solid article into a plurality of airfoil-shaped slices and bonding one slice of the plurality of airfoil-shaped slices to a tip portion of a turbine blade.

Abstract: A rotor (100) for a turbomachine is provided, having at least two bladed, detachably interconnected rotor stages (1, 3, 6), a flange (17) of a first rotor stage (1) being attached to a rotor disk (25) of a second, adjacent rotor stage (3) via at least one fastening element (7). The rotor disk (25) has a through bore (11) for connecting the two rotor stages (1, 3) by the fastening element (7); the through bore (11) having a cross section that is larger in the circumferential direction (u) of the rotor disk (25) than in the radial direction (r) thereof.

Abstract: A gas engine turbine has a transition duct (100) that has improved cooling features and a method for forming the cooling features. A continuous exit section cooling channel (130) is formed through the transition duct panel (112), the exit frame (114) and the connection (116). The continuous exit section cooling channel (130) reduces the need for effusion channels in the transition duct. The continuous exit section cooling channel (130) further reduces costs and improves the emissions associated with the transition duct (100) of the gas engine turbine.

Abstract: A method of shaping the profile of a fan casing having an inside surface, the method including placing the casing around a surface of revolution of a drum of shaping tooling; interposing at least one bladder that is inflatable under the action of a fluid under pressure between a portion of the inside surface of the casing and the drum, the bladder extending over all or part of the surface of revolution of the drum; stoving the assembly including the casing, the tooling, and the at least one bladder at a predetermined temperature; and during the stoving, applying isostatic pressure via the at least one bladder so as to impart a cylindrical profile to the portion of the inside surface of the casing facing the at least one bladder.

Abstract: A means for attaching a metallic component to a non-metallic component using a compliant material having thermal properties intermediate those of the metallic component to a non-metallic component is provided. The method can accommodate CTE mismatches and wear-type problems common to many assemblies of dissimilar materials. In particular, the method provides a sufficient wear surface to accommodate relative motion and provide a durable wear surface that does not excessively wear/gall/mico-weld itself together and provides the necessary damping and motion for proper operation in aeronautical applications.

Abstract: A method of manufacturing a superalloy part including a first portion including a majority by weight of a first superalloy and a second portion including a majority by weight of a second superalloy, the second portion extending from the first portion, the method including depositing the second portion on the first portion by a direct metal deposition method, deposition of the second portion including depositing a first layer, then depositing a second layer on the first layer, the first layer including the first and second superalloys, the first layer presenting a content by weight of the first superalloy that is strictly greater than that content by weight of the second layer and strictly less than that content by weight of the first portion, the second layer presenting a content by weight of the second superalloy that is strictly greater than that content by weight of the first layer.

Abstract: An airfoil includes a first airfoil piece and a second airfoil piece bonded to the first airfoil piece at a joint. The first airfoil piece and the second airfoil piece are formed of aluminum alloys. At least one of the aluminum alloys is an aluminum alloy composition that has greater than 0.8% by weight of zinc.

Abstract: The present disclosure is directed to a nozzle for a turbomachine. The nozzle includes an inner side wall, an outer side wall radially spaced apart from the inner side wall, and an airfoil extending radially from the inner side wall to the outer side wall. The airfoil defines a cavity that extends radially through the nozzle. The cavity is at least partially defined by a cavity wall. The cavity wall at least partially defines a pocket in fluid communication with the cavity. A cooling passage is defined by one of the inner side wall or the outer side wall. The cooling passage is in fluid communication with the cavity via the pocket.

Abstract: A system and methods are provided for pre-stressing blade elements for a gas turbine engine. In one embodiment, a method includes rotating a blade element relative to an axis. The method may also include controlling rotational speed of the blade element to generate residual stress in the blade element. The method may also include rotating multiple blade elements and fan blade units to generate residual stress. Blade elements may be rotated to exceed a maximum operating speed of the blade element to 120% of the maximum operating speed of the blade element.

Abstract: A turbine engine with a casing including opposing surfaces and a manifold opening passing through the surfaces. A surface cooler is provided adjacent one of the surfaces. A cooler manifold passes through the manifold opening. A seal is provided between the surface cooler and the casing. At least one bolt is used to secure the surface cooler to the casing with the seal in between.

Abstract: Aspects of the disclosure are directed to a sync ring assembly associated with an engine, comprising: a sync ring, and a clevis coupled to the sync ring, where the clevis includes a rib that resists a deflection of the sync ring assembly when the engine is operated, and where the clevis is made of titanium.

Abstract: An airfoil structure for a gas turbine engine includes an airfoil that includes a suction side cooling circuit with at least two segments that are connected by at least one impingement passage.

Abstract: A rotor disc for a gas turbine engine includes an annular disc body configured to support a circumferential array of blades and having a plurality of passages defined therethrough. The passages form coils within the disc body and/or have a packing density of at least 0.1 in cross-sectional plane containing the central axis, the packing density being defined by a ratio between an open area of the passages and a solid area of the disc in the cross-sectional plane. A method of manufacturing a rotor disc is also discussed.

Abstract: A gas turbine drive shaft arrangement, comprising: a hollow compressor drive shaft, a first hollow turbine drive shaft for driving a first turbine stage and a second hollow turbine drive shaft for driving a second turbine stage, wherein the compressor drive shaft has a spline on an internal surface that is meshed with a spline on an external surface of the first hollow turbine drive shaft, the first turbine drive shaft further comprising a spline on an internal surface thereof that is meshed with a spline on an external surface of the second hollow turbine drive shaft.

Abstract: A gas turbine engine includes a fluid intake. A turbine section includes a turbine case. A firewall is located upstream of the turbine section. A conduit is configured to direct a fluid from the fluid intake to the turbine case. A valve is located on a first side of the firewall opposite from the turbine section and is configured to regulate a flow of the fluid through the conduit.

Abstract: A mid-turbine frame for a gas turbine engine includes an inner frame case that includes a bolt opening and at least one spoke for connecting an outer frame case to the inner frame case that includes an inlet passage that extends in a radial direction. A central bolt extends through the bolt opening for securing at least one spoke to the inner frame case.

Abstract: A method of assembling a rotor stack comprising a plurality of component parts, the method comprising: determining swash measurements for interfacing surfaces of the plurality of parts; calculating a runout estimate for a plurality of relative orientations of the parts; applying an optimization algorithm to identify an optimal orientation from the plurality of relative orientations based on the runout estimates; and assembling the parts in the optimal orientation.

Abstract: A method of manufacturing a gas turbine engine, the method including providing a fan having a plurality of fan blades circumscribed by a fan track liner and attaching a cutter to the fan track liner, the cutter is arranged towards the blades of the fan. The fan blades are rotated so that one or more of the fan blades are trimmed by the cutter. Once one or more of the fan blades have been trimmed by a desired amount, the cutter is removed from the fan track liner.

Abstract: A cylindrical case main body, and a connecting ring body fitted and fixed to a first end of the case main body and having an annular groove open in the centrifugal direction, are provided. The connecting ring body has a notch open to a side opposite to a second end of the case main body. Into the notch of the connecting ring body, a reinforcement body is fitted, and the reinforcement body has an arc-shaped groove continuing to the annular groove of the connecting ring body in a state where the reinforcement body is fitted into the notch of the connecting ring body. In addition to attaining high rigidity, reduction in weight and reduction in manufacturing cost can also be achieved.

Abstract: A cylindrical case body made of a composite material; a linking ring made of aluminum, fitted and fixed to a rear end portion of the case body and having an annular groove which receives a reverse thrust load from a reverse thrust transmission body; and ring constituent bodies made of titanium alloy, each of which is placed in the rear end portion of the case body and has an arcuate groove which undertakes a reverse thrust load larger than that received by the linking ring, are provided in a fan case. The ring constituent bodies are fitted in notches of the linking ring and are brought into contact with outward receiving walls of the linking ring. The ring constituent bodies are positioned relative to the case body in the axis direction and in the radial direction, and the annular groove and the arcuate grooves continue to each other.

Abstract: A combustor of a gas turbine engine including a combustor shell having an interior surface defining a combustion chamber, a first panel mounted to the interior surface at a first position, the first panel having a first surface and a first rail extending from the first surface toward the combustor shell, the first rail configured at a first angle relative to the first surface, and a second panel mounted to the interior surface at a second position axially adjacent to the first panel, the second panel having a second surface and a second rail extending from the second surface toward the combustor shell, the second rail configured at a second angle relative to the second surface. The first and second rails are proximal to each other and define a circumferential gap there between and at least one of the first or second angles is an acute angle.

Abstract: The structural case has an annular body having a central axis and including a plurality of boss sections circumferentially interspaced from one another around the axis by a plurality of arcuate panel sections, each panel section having: two parallel arcuate structural flange members being axially interspaced from one another; a sheet metal wall extending between and interconnecting the two flange members; and at least one rib having an edge welded to the sheet-metal wall.

Abstract: A nacelle assembly for a high-bypass gas turbine engine includes a fan variable area nozzle movable relative a fan nacelle to vary a fan nozzle exit area to reduce a fan instability.

Abstract: A synchronizing ring assembly for a variable guide vane system is disclosed. The synchronizing ring assembly is disposed about an engine axis and includes a synchronizing ring having a radially outer surface with a clevis bracket mounted to the radially outer surface of the synchronizing ring and defining radially extending openings through the clevis bracket. A plurality of fasteners are configured and adapted to be received within the radially extending openings in the clevis bracket for securing the clevis bracket to the radially outer surface of the synchronizing ring.

Abstract: A method of, and apparatus for, urging blades of a gas turbine engine radially outwardly is disclosed. The method may be used to grind blade tips of blades of a rotor stage of a gas turbine engine. The method comprises locating a fluid-tight bag which is in a radial gap formed between a radially inner surface of a respective blade root and a slot in a disc which the blade root cooperates. The method comprises inflating the fluid-tight bag and rotating the rotor stage relative to a grinding surface so as to grind any blade tips that contact the grinding surface during rotation. This results in more accurate positioning of the blades during the grinding process and/or during operation.

Abstract: A screw pump for transporting a fluid, with a drive spindle having a drive spindle profile and with at least one running spindle having a running spindle profile, the running spindle engages with its running spindle profile at least partially in the drive spindle profile of the drive spindle, the drive spindle profile and/or the running spindle profile is formed as a rolled profile, and a method for producing a spindle for such a screw pump, and a method for producing a groove in such a spindle are disclosed.

Abstract: The invention relates to a rotor shaft adapted to rotate about a rotor axis thereof. The rotor shaft includes a rotor cavity configured concentrically or quasi-concentrically to the rotor axis inside the rotor shaft, and a plurality of cooling bores extending radially or quasi-radially outward from the inside to an outside of the rotor shaft. Each cooling bore having a bore inlet location and a distal bore outlet portion, the respective bore inlet location being adapted to abut on the rotor cavity. At least one side or part-side of the cooling bore inlet location is provided with an asymmetric edge fillet in order to maximize the wall thickness between two adjacent cooling bores.

Abstract: A compressor wheel for a turbocharger has an improved geometry that provides improved performance as compared with a comparably sized conventional compressor wheel by having the radially inward portion of the primary blades sweep upwardly into the upper region of the hub portion. The through bore and integral nut of conventional compressor wheels is eliminated and replaced with a blade extension that has an angled upper edge which extends downwardly and radially outwardly from the hub portion, effectively increasing the size of the blade, the amount of air movement and the amount of power generated, without increasing the overall cost, size or mass of the compressor wheel.

Abstract: The present technology provides methods and systems for reducing ice buildup in a closed engine breather system. The systems comprise an air inlet duct, a compressor wheel and a compressor cover. The compressor cover comprises a breather gas port. The breather gas port is placed in the compressor cover at a location within close proximity to the leading edge of the compressor wheel so that the compressor wheel is capable of removing ice that accumulates at the breather gas port. The present technology also provides methods that include providing an air inlet duct, a compressor wheel, and a compressor cover with a breather gas port in the compressor cover. In this manner the compressor wheel can grind or shave off ice that accumulates at the breather gas port.

Abstract: An wheel fastening inspection method includes: directly or indirectly applying vibration to a turbine shaft to be inserted to a bearing housing, the turbine shaft being provided with wheels at both ends protruding from the bearing housing, at least one of the wheels being fastened by a fastening member, and a rotary member being fastened by a fastening force generated by the fastening member to the turbine shaft to be integrally rotated with the wheel; measuring the vibration of the turbine shaft; and determining whether or not a vibration frequency at which a peak of the measured vibration of the turbine shaft is given is included in a setting range previously set.

Abstract: The present invention relates to a stator of an axial compressor stage of a turbomachine featuring a radially outer blade ring forming an outer ring surface, a radially inner blade ring forming an inner ring surface, and several stator blades connected to the blade rings. It is provided that the outer ring surface and/or the inner ring surface has at least in a partial area a changing radius relative to a central axis of the stator both in the axial direction and in the circumferential direction.

Abstract: A bulkhead (22) of a wind turbine (10) to be arranged on a rotor blade connection of a rotor blade (14), especially on a rotor hub (9). The bulkhead (22) has a core body (30). A layer (31, 32) of fiberglass-reinforced plastic (31, 32) is arranged on the core body (30) on both sides respectively and a metal layer body (33) is arranged on one side of the layer of fiberglass-reinforced plastic (31). A method for producing a bulkhead (22) of a wind turbine (9), which is arranged on a rotor blade connection of a rotor blade (14) and a use of a bulkhead (22) of a wind turbine (10).

Abstract: A component for a turbine engine includes a substrate that includes a first surface, and an insert coupled to the substrate proximate the substrate first surface. The component also includes a channel. The channel is defined by a first channel wall formed in the substrate and a second channel wall formed by at least one coating disposed on the substrate first surface. The component further includes an inlet opening defined in flow communication with the channel. The inlet opening is defined by a first inlet wall formed in the substrate and a second inlet wall defined by the insert.

Abstract: A system for inspecting surfaces of rotor blades for a surface characteristic. The system may include an assembly having a movable arm and, mounted on the movable arm, a scanner. A row of rotor blades may be positioned near the assembly for inspection. The row of rotor blades may include a plurality of the rotor blades circumferentially spaced about a center axis. The row of rotor blades and the assembly may be moved relative to the other so as to index the row of rotor blades relative to the assembly.

Abstract: An aero engine rotor assembling method and device based on concentricity and verticality measurement belongs to mechanical assembly technology. The present invention effectively solves the problem of poor coaxality after the aero engine rotor is assembled and has the characteristics of high coaxality after the rotor is assembled, reduced vibration, mounting easiness, high flexibility and improved engine performance. The measurement and device is: determining rotary reference; determining the angular positioning of a rotary table; extracting the radial error of the radial mounting plane and the inclination error of the axial mounting plane of the rotor based on the four-probe measuring device to obtain the influencing weight of this rotor to the assembled rotor on coaxality; measuring respectively all the rotors to obtain the influencing weight of each rotor to the assembled rotor on coaxality; vector optimizing the weight of each rotor to obtain the assembling angle of each rotor.

Abstract: A treated component and methods for forming a treated component are disclosed. The methods include providing an untreated component having an untreated creep strength. The untreated component is formed by a three-dimensional printing process, and is treated to yield the treated component having a treated creep strength. The treated component comprises an arrangement formed by the three-dimensional printing process, wherein the arrangement has been subjected to treating to increase creep strength.

Abstract: A method for fabricating turbine engine blade or vane made of composite material includes: performing three-dimensional weaving to make a single-piece fiber blank; shaping the fiber blank to obtain a single-piece fiber preform having a first portion forming a preform for at least a blade/vane airfoil, at least one second portion forming a preform for an inner part of a blade/vane inner platform or for an outer part of a blade/vane outer platform, and at least one third portion forming a preform for an outer part of a blade/vane inner platform or for an inner part of a blade/vane outer platform; and densifying the fiber preform with a matrix to obtain a composite material blade, and forming a single piece with an inner and/or outer platform(s) incorporated therein.