Abstract: Method for monitoring a turbomachine comprising a stator and a rotor, device, system, aircraft and computer program product. The method comprises a step of acquiring an input signal which represents a deformation of the stator or rotor of the turbomachine. The input signal has been acquired by a deformation gauge which is attached to the stator or rotor. The input signal comprises a first component which represents deformations of the stator or rotor which are caused by rotation of the rotor vanes in relation to the stator, and a second component which represents deformations which are caused by different elements of the rotor vanes. The method comprises a step of re-sampling the input signal in order to obtain a re-sampled input signal comprising a predefined whole number of samples per revolution of the turbomachine rotor.

Abstract: A gas turbine engine article includes a silicon-containing ceramic wall that has an external combustion gaspath side and an internal side that borders a cooling air cavity. The external combustion gaspath side has an associated combustion gas flow direction there along. An array of cooling holes extends through the silicon-containing ceramic wall and connects the internal side with the external combustion gaspath side. The cooling holes are oriented to discharge cooling air to the external gaspath side in a direction counter to the combustion gas flow direction.

Abstract: A passive flow modulation device for a machine defining an axial direction and a radial direction, the passive flow modulation device including: a first ring with a first coefficient of thermal expansion; a second ring disposed coaxially with the first ring and positioned at least partially inward of the first ring along the radial direction, spaced from the first ring along the axial direction, or both, the first ring, the second ring, or both defining at least in part one or more passages, the second ring with a second coefficient of thermal expansion that is less than the first coefficient of thermal expansion to passively modulate a size of the one or more passages during operation.

Abstract: A gas turbine engine comprises at least a power output turbine unit (POT), which is rotatably arranged inside an outer housing unit, and a compressor-turbine unit (CTU), which is rotatably arranged inside the POT, and the CTU, POT and outer housing unit are arranged about a common axis of rotation (CL). The POT and the CTU are arranged in such close proximity (d) that a dynamic friction coupling is generated between the POT and the CTU.

Abstract: According to an embodiment, a turbine casing comprises a cooling air supply unit configured to supply a cooling air to an interior space of a casing of a gas turbine, and the cooling air supply unit includes: a first supply unit disposed in an upper half of the casing so as to face a first region and configured to supply the cooling air to the first region, where the first region is a region on a radially outer side of a plurality of combustors arranged annularly around a rotor; and a second supply unit disposed so as to face a second region and configured to supply the cooling air to the second region, where the second region is a region on a radially inner side of the plurality of combustors.

Abstract: A high temperature flange joint in a gas turbine engine includes a first flange formed on a first component abutting a second flange formed on a second component. The flange joint includes multiple adjacently spaced bolt connections. Each bolt connection includes a first spacer plate bearing against the first flange and a second spacer plate bearing against the second flange. First and second lock washers are provided that bear against the first and second spacer plates respectively. A bolt is inserted through the first and second flanges, the first and second spacer plates and the first and second lock washers. The bolt is preloaded to clamp the first flange to the second flange. Each spacer plate has a respective thickness and is sized to enhance a bearing surface in contact with the respective flange. Thereby, a bolt preload is maintained during operation of the gas turbine engine.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a case structure and an air conduit. The case structure extends circumferentially about and axially along an axis. The air conduit has a conduit centerline, a conduit first end and a conduit second end. The air conduit extends longitudinally along the conduit centerline between the conduit first end and the conduit second end. The conduit first end is connected to the case structure at a first location. The conduit second end is connected to the case structure at a second location. The air conduit is displaced from the case structure longitudinally between the conduit first end and the conduit second end. At least a majority of the conduit centerline follows a non-straight trajectory.

Abstract: A cooling system for cooling hot components of a radial or axial gas turbine engine, which includes a recuperator heat exchanger, provides engine cooling without loss of thermal efficiency. Air flow leaving a compressor is split between a recuperator flow path and a bleed flow path. Air in the bleed flow path flows through the hot parts of the engine, thereby cooling the engine and heating the air. The air in the bleed flow path is combined with the output flow from a combustor and directed into a turbine inlet. A reduction of air flow in the recuperator flow path increases the thermal effectiveness of the recuperator heat exchanger by increasing a ratio of hot and cold flows inside the heat exchanger. The increase in thermal effectiveness of the heat exchanger compensates for energy losses incurred by diverting a portion of the compressor air flow for cooling.

Abstract: The film cooling structure includes a wall part and a cooling hole inclined such that an outlet is positioned rearward of an inlet. The cooling hole includes a straight-tube part and a diffuser part. The diffuser part includes a flat surface, a curved surface curved rearward and forming, together with the flat surface, a semicircular or semi-elliptical channel cross section larger than that of the straight-tube part, a first section and a second section extending from the first section toward the outlet. In the first section, an area of the channel cross section increases as it approaches the outlet. In the second section, the area of the channel cross section increases as it approaches the outlet at an increase rate smaller than that of the first section or is constant. The diffuser part has a width equal to or twice greater than the depth of the diffuser part.

Abstract: A dual-wall airfoil configured for coolant transfer includes a spar having a pressure side wall and a suction side wall each including raised features on an outer surface thereof. An interior of the spar includes coolant cavities. An inner surface of a pressure side coversheet is in contact with the raised features on the outer surface of the pressure side wall so as to define pressure side flow pathways between the pressure side wall and the pressure side coversheet, and an inner surface of a suction side coversheet is in contact with the raised features on the outer surface of the suction side wall so as to define suction side flow pathways between the suction side wall and the suction side coversheet. The pressure side flow pathways and/or the suction side flow pathways include cooling circuit(s) configured to transfer coolant between the coolant cavities.

Abstract: The invention concerns an air inlet duct (10) for a compressor (12) of a gas or fuel oil turbine, including: two transition sections (S3, S4) in fluid communication with one another for the circulation of a flow of air through said sections (S3, S4), each of said sections (S3, S4) being self-supporting, a structure (20) for injecting a mist of liquid particles, configured to be disposed between said sections (S3, S4) and in contact with said sections (S3, S4), the structure (20) being removable independently of demounting said sections (S3, S4). The invention also concerns a gas or fuel oil turbine assembly comprising an inlet duct (10) of this type and a method of maintaining an inlet duct (10) of this type.

Abstract: A cleaning system for cleaning gas paths in an engine core of a gas turbine engine is provided. The cleaning system includes a source of an engine cleaning liquid; an engine cleaning mist forming unit that vapourises the engine cleaning liquid to form an engine cleaning mist and delivers the engine cleaning mist into the engine core of the gas turbine engine; at least one delivery device configured to deliver the engine cleaning liquid to the engine cleaning mist forming unit; a pump configured to draw the engine cleaning mist through the engine core to clean the gas paths within the engine core; and a mist collecting arrangement including a condensing chamber. The mist collecting arrangement is configured to collect the engine cleaning mist that has passed through the engine core and condense the collected engine cleaning mist in the condensing chamber.

Abstract: A turbine engine with at least a compressor section, combustor section, turbine section and a set of airfoils. The airfoils include geometric characteristics to create a high contraction ratio (CR), a low blade turning (BT) at a radially inward location the airfoil, a low solidity, or a low aspect ratio (AR).

Abstract: A blade for a gas turbine includes an external structure including a plurality of seating grooves which are separately disposed in a chord direction toward a trailing edge from a leading edge, an internal structure received in the external structure and including a plurality of protrusions protruding toward an internal side of the external structure, a plurality of porous strips combined to the seating groove in an attachable/detachable way, and a coolant channel formed for a coolant to flow among the porous strip, the neighboring protrusions, and an external side of the internal structure.

Abstract: A turbine housing is provided with an inner pipe forming a spiral-shaped exhaust gas flow path, an exhaust-air-inlet-side flange serving as an exhaust gas inlet to the inner pipe, and an outer pipe that, together with the exhaust-air-inlet-side flange, covers and seals the inner pipe. The outer pipe has an opening through which the washing fluid can pass through and a closing member that closes the opening, and a space through which the washing fluid can pass through is formed between the inner pipe and the exhaust-air-inlet-side flange.

Abstract: A turbine assembly adapted for use with a gas turbine engine includes an outer case, a blade track segment, and a carrier. The outer case extends circumferentially at least partway around an axis of the engine. The blade track segment is configured to define a portion of a gas path of the turbine assembly. The carrier is coupled with the outer case and the blade track segment to support the blade track segment in position radially relative to the axis. The carrier is coupled with the outer case for movement with the outer case in response to thermal expansion and contraction of the outer case during use of the turbine assembly.

Abstract: A Francis turbine includes a waterway, an impeller disposed in the waterway, a distributor apparatus disposed in the waterway and an additional element for adjusting the flow of water through the turbine. The additional element is disposed in the waterway upstream of the impeller in the water flow direction, and the additional element includes a perforated metal sheet that may be brought into the waterway. A method for operating such a turbine is also provided.

Abstract: An engine component includes a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit. The component further includes a plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body. The plurality of cooling holes includes a first cooling hole with a metering portion with a constant cross-sectional area and a cross-sectional shape having a maximum height that is offset relative to a longitudinal centerline of the metering portion; and a diffuser portion extending from the metering portion to the external surface of the body.

Abstract: A turbo engine rotor disc that has a rotation axis and includes, at the periphery thereof, a plurality of slots regularly distributed around the rotation axis, wherein at least one slot of the plurality of slots has a base that has a plurality of plates arranged in staggered rows and protruding from the base, each plate extending mainly along a direction perpendicular to a radial direction.

Abstract: An impingement plate for providing cooling to a gas turbine engine component includes a first surface with a second surface opposite the first surface. The first surface and the second surface are surrounded by a perimeter defining an edge. A first plurality of holes extend through the impingement plate. A plurality of protrusions extend outward from one of the first surface or the second surface and have a distal surface transverse to one of the first surface or the second surface. At least one second impingement hole extends through the distal surface.

Abstract: An assembly in a turbine engine having a dynamically moveable flowpath boundary member for encasing a rotatable bladed disc and maintaining a clearance gap between the flowpath boundary member and the blade tips of the bladed disc. The assembly comprises a static casing, a flowpath boundary member carried by the casing, and a refrigeration system configured to remove heat from the flowpath boundary member to thereby thermally expand and contract the flowpath boundary member to control blade tip clearance.

Abstract: A clearance control system includes one or more clearance control devices that include features for controlling the clearances between rotating and stationary components of an engine. In one exemplary aspect, a clearance control device utilizes a fluid to pressurize an actuation chamber defined by a compliant member of the clearance control device. The pressurization of the actuation chamber causes the actuation chamber to expand in a direction that changes the clearance between the stationary and rotating components of the engine.

Abstract: A disclosed micro gas turbine system includes a micro gas turbine apparatus and an extracting cycle apparatus. The micro gas turbine apparatus includes a first compressor, a burner, and a first turbine. The first turbine expands a combustion gas generated by the burner. The extracting cycle apparatus includes a second compressor and a second turbine. The second compressor receives a flow of extracted air that is generated by extracting a part of a working fluid discharged from the first compressor. The second turbine expands the working fluid discharged from the second compressor. The working fluid discharged from the second turbine cools down the first turbine.

Abstract: A steam turbine is provided. The steam turbine includes a housing, a first steam inlet configured to discharge a first steam flow within the housing, and a second steam inlet configured to provide a second steam flow. A rotor and stator are coupled to the housing and configured to form a first flow path therebetween and in flow communication with the first steam flow. The rotor includes a plurality of blades coupled to the rotor, at least one root of the plurality of blades has a first side, a second side and a passageway coupled in flow communication to the first side and the second side. The passageway is configured to receive the second steam flow within the at least one root. The at least one root includes an angel wing configured to seal the second steam flow from the first flow path.

Abstract: Retention assemblies and retention assembly cooling methods are provided. An exemplary retention assembly comprises an annular baffle, an annular attachment bracket comprising a gas turbine engine component, and first and second cooling passages defined by the baffle and attachment bracket. The first cooling passage is configured to receive a first airflow and the second cooling passage is configured to receive a second airflow. The first airflow has a lower pressure than the second airflow. An exemplary method comprises flowing a first airflow to a first cooling passage defined by a baffle and an attachment bracket of a retention assembly, and flowing a second airflow to a second cooling passage defined by the baffle and the attachment bracket. The second cooling passage is separate from the first cooling passage. The first airflow cools radially outward structures of the retention assembly, and the attachment bracket comprises a CMC component.

Abstract: An airfoil includes an airfoil body that has a peripheral wall that defines an exterior side and an interior side that bounds an internal cavity in the airfoil body. The peripheral wall has first and second wall sections joined by a transition section. The first wall section is thicker than the second wall section. The transition section provides a change in thickness between the first wall section and the second wall section. The second wall section includes a cooling hole that has a first end that opens to the internal cavity at the interior side and a second end that opens to the exterior side.

Abstract: A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed.

Abstract: An airfoil cooling system may be provided. An airfoil may have a pressure side and a suction side that are separated by a leading edge. The leading edge may pass through a stagnation point of the airfoil. A spar may have an outer surface comprising standoffs. The standoffs may define a cooling channel that extends across the leading edge on the outer surface of the spar, from the pressure side to the suction side. The cooling channel may have a first portion and a second portion defined by the standoffs. The first portion may be located closer to a base or a tip of the airfoil than the second portion. The spar may further comprise an inlet on the pressure side or the suction side. The inlet may be configured to convey a cooling fluid from a passageway located inside of the spar to the cooling channel.

Abstract: A method of measuring the clearance between a rotating component and a static component in a rotating machine. The method comprises installing a rub pin on an interior surface of the static component, operating the rotating machine such that the rotating component impinges on the rub pin, securing the rotating machine, and inserting an optical device to visually ascertain the length of rub pin remaining. The method reduces the time and cost of clearance determinations by avoiding disassembly or the degree of disassembly of the rotating machine required to conduct a clearance determination.

Abstract: A turbine ring assembly includes ring sectors made as a single piece of ceramic matrix composite material and being for mounting on a metal ring support structure, the ring support structure having two tabs extending radially towards a gas flow passage, each ring sector having a first portion forming an annular base with a radially inside face defining the inside face of the turbine ring and an outside face from which there extend two tab-forming portions, the ring sectors having a section that is substantially ? shaped. The tabs of each ring sector fit axially over the tabs of the ring support structure with contact between the tabs of the ring support structure and the tabs of each ring sector when cold. The tabs of each ring sector are held to the ring support structure by pegs passing through holes formed in the tabs of each ring sector.

Abstract: A device to fragment solidified bulk material is disclosed. The device comprises a hydraulic motor, a stationary assembly and rotating assemblies, wherein the rotating assemblies includes, a rotational upper whip mount assembly adapted to rotate in a direction, a rotational middle assembly perimeter adapted to rotate in the same direction as the rotational upper whip mount assembly, and a rotational lower whip mount assembly adapted to rotate in a direction opposite the rotational direction of the upper whip mount and middle perimeter assemblies, and a plurality of flails configured to fracture hardened, solidified bulk material while balancing the torque forces to more accurately keep the dual whip-head head in a desired location when operationally engaged with the bulk material.

Abstract: A fan inlet diffuser housing includes a housing body of composite material and includes a heat exchanger interface portion positioned between an ejector housing portion and a bypass housing portion. A first transition region is formed between the heat exchanger interface portion and the ejector housing portion including an air cycle machine end reinforcement patch proximate to a heat exchanger interface. The air cycle machine end reinforcement patch includes a first patch thickness and a second patch thickness, and a ratio of the first patch thickness to the second patch thickness is between 2.02 and 3.11. An ejector is formed having an ejector gap width between a nozzle portion and a diffuser portion within the ejector housing portion of the housing body. The diffuser portion has a downstream ejector gap width, and a ratio of the downstream ejector gap width to the ejector gap width is between 4.62 and 5.01.

Abstract: Methods and systems are provided for injecting water stored in a water reservoir at a plurality of locations in an engine system, including directly into engine cylinders, upstream of a turbocharger turbine, and at an exhaust manifold, and controlling a water level of the water reservoir. In one example, a method may include supplying the water stored in the water reservoir to one or more of a water injection system, a windshield wiper system, an engine coolant system, and a drinking water system based on water supply conditions and responsive to engine operating conditions. When the water level is low (e.g., below a threshold), supply to the water injection system may be prioritized.

Abstract: A method and system to provide a fan inlet diffuser housing including a housing body formed from a composite material, wherein the composite material includes a base fiberglass material having a first conductivity, a conductive material having a second conductivity, wherein the second conductivity is greater than the first conductivity, and a binder that bonds the base fiberglass material and the conductive material.

Abstract: A stator heat shield for a gas turbine having a hot gas flow path, is disclosed. The stator heat shield includes a first surface configured to face the hot gas flow path of the gas turbine; a second surface opposite to the first surface; cooling channels for directing cooling fluid from the second surface towards the first surface; and cavities arranged at the first surface for receiving the cooling fluid from at least a part of the cooling channels; wherein at least a part of the cavities each have at least two corresponding cooling channels open thereto, the at least two corresponding cooling channels being inclined towards each other. In use, a vortex is created in the cavity.

Abstract: A wall arrangement for the main gas path of a gas turbine engine, including: a wall segment which defines the main gas path, the wall segment having a gas path side, an outboard side and a support wall extending from the outboard side towards a supporting structure, and a channel member abutting the support wall and having one or more channels defined by the abutment of the support wall and channel member, the one or more channel having radially separated inlet and outlet.

Abstract: A blade for a water current turbine rotor, which extends along a radial direction and comprises an outer surface, an inner surface, a leading edge, and a trailing edge. The leading edge is the edge of the blade that extends along the radial direction and is disposed upstream in the direction in which the water flows along the blade, while the trailing edge is the edge of the blade opposite the leading edge and is disposed downstream in the direction of flow. At least one portion of the blade has a profile comprising a thick portion and a thin portion, the cross section being perpendicular to the radial direction. The thick portion and the thin portion each have a maximum thickness along a direction perpendicular to the outer surface, the maximum thickness of the thick portion being at least four times greater than the maximum thickness of the thin portion.

Abstract: According to one aspect, an air cycle machine includes a fan rotor, a diffuser cone axially aligned with the fan rotor, and a strut plate assembly. The strut plate assembly includes a first strut plate and a second strut plate with an ejector formed between the first strut plate and the second strut plate. The first strut plate is axially positioned between the fan rotor and the ejector, and the second strut plate is axially positioned between the fan rotor and the diffuser cone. The first strut plate includes a plurality of outer struts and is absent inner struts between the fan rotor and the ejector. The second strut plate includes a plurality of inner struts and outer struts, where the inner struts are axially positioned between the ejector and the diffuser cone.

Abstract: A vane seal assembly for a gas turbine engine comprises of a case including a first connector. A notch in the case adjoins the groove. A vane having a second connector mates with the first connector. A seal assembly is provided between the vane and the case to provide a sealed cavity adjoining the notch.

Abstract: An aircraft gas turbine is constituted by accommodating a compressor (14), a combustor (15), and a turbine (16) in a cylindrical main unit casing (12). A thick wall part (52) is provided on an outer periphery side of rotor blades (34) in the main unit casing. A cooling passage (53), for cooling the thick wall part (52) by circulating compressed air compressed by the compressor, is provided in the thick wall part. Also provided is a discharge passage (55) for discharging compressed air having circulated in the cooling passage to a combustion gas passage A. The structural strength of the thick wall part is ensured by appropriately cooling the thick wall part of the casing, while simplifying the structure and preventing a decrease in efficiency, thereby ensuring effective containment performance and an appropriate clearance between the casing and the rotor blades.

Abstract: A shroud segment assembly for use within a turbine shroud of a gas turbine engine may generally include a forward shroud portion having a forward outer arm and a forward inner arm extending from a forward wall and an aft shroud portion having an aft outer arm and an aft inner arm extending from an aft wall. Additionally, the shroud segment assembly may include a first expansion joint positioned between the forward and aft shroud portions such that the first expansion joint extends circumferentially between the forward outer arm of the forward shroud portion and the aft outer arm of the aft shroud portion and a second expansion joint positioned between the forward and aft shroud portions such that the second expansion joint extends circumferentially between the forward inner arm of the forward shroud portion and the aft outer arm of the aft shroud portion.

Abstract: A heat transfer apparatus for a gas turbine engine includes: a component having a wall structure defining a flow bounding surface; a chamber formed in the component, the chamber including a wicking structure, a vapor channel, and a working fluid.

Abstract: The present invention discloses a novel and improved device and method for securing a combustion liner within a gas turbine combustor. The improved configuration comprises a plurality of equally spaced support tab assemblies secured about the combustion liner and positioned radially outward of the combustion zone. The support tab assemblies comprise parallel liner tabs which provide increased flexibility, greater structural support, and lower operating stresses.

Abstract: A sealing element is described for positioning in a turbomachine, for the self-securing of a sealing element in a mounting position, its support structure having a safeguard which has a springy catch section, whose mounting surface is inclined at a smaller angle in the mounting direction than a securing surface in the opposite direction, a sealing unit, and a turbomachine.

Abstract: A shroud apparatus for a gas turbine engine includes: an annular metallic hanger; a shroud segment disposed inboard of the hanger, comprising low-ductility material and having a cross-sectional shape defined by opposed forward and aft walls, and opposed inner and outer walls, the walls extending between opposed first and second end faces, wherein the inner wall defines an arcuate inner flowpath surface; and a retainer mechanically coupled to the hanger which engages the shroud segment to retain the shroud segment to the hanger while permitting movement of the shroud segment in a radial direction.

Abstract: Provided is a film cooling structure capable of suppressing a cooling medium film from being separated from a wall surface, to increase a film efficiency on the wall surface and thereby-cool the wall surface effectively. One or more pairs of injection holes are formed on a wall surface facing a passage of high-temperature gas to inject a cooling medium to the passage. A single supply passage is formed inside the wall to supply the cooling medium to the injection holes. A separating section is provided between the injection holes in a location forward relative to rear ends of the injection holes to separate the cooling medium into components flowing to the injection holes. An injection direction of the cooling medium is inclined relative to a gas flow direction so that the cooling medium forms swirl flows that push the cooling medium against the wall surface.

Abstract: A seal for sealing a combustor heat shield against an interior surface of a combustor shell, the seal comprising: a first sealing surface on the interior surface of the combustor shell; and a second sealing surface on a rail on an edge of a heat shield, wherein each of the first and second sealing surfaces include first and second projections defining a non-linear leakage path between the projections.

Abstract: The present application provides a stage of a gas turbine engine. The stage may include a bucket extending radially about a longitudinal axis of the gas turbine engine, a shroud facing the bucket, the shroud including a fore end portion including a radially inner surface spaced a first distance from the longitudinal axis, and a forward step honeycomb seal positioned on the shroud downstream of the fore end portion and facing the bucket. The forward step honeycomb seal may include a first linear portion including a radially inner surface spaced a second distance from the longitudinal axis, and a forward step portion positioned adjacent to and downstream of the first linear portion, the forward step portion including a radially inner surface spaced a third distance from the longitudinal axis, wherein the second distance is greater than the third distance, and wherein the third distance is greater than the first distance.

Abstract: An ambient air intake duct of a static gas turbine is provided, having at least one filter, which is arranged in the intake duct, for cleaning the ambient air (A) that can flow through the intake duct. A method for operating a static gas turbine which is equipped with a filter for cleaning the ambient air (A) is also provided. To rapidly provide a higher level of gas turbine power to a generator, it is provided that, by means of a bypass or by means of flaps arranged downstream of the filter, partially to completely unfiltered ambient air (A) can temporarily flow into the compressor inlet.

Abstract: A Blade Outer Air Seal (BOAS) includes a body manufactured of a metal alloy, the body includes a face opposite a forward interface and an aft interface, the face includes a cavity. A non-metallic insert within the cavity such that the insert is flush with the face.