Abstract: A method of characterising mechanical and thermal growth of a rotor tip clearance arrangement and a method of controlling such an arrangement. The characterisation method comprising constructing a finite element model and calibration cases; run the model for the calibration cases to obtain displacements; and calculating component growth therefrom. The growth is characterised and scaled for the operating conditions. The clearance is calculated in a time step from the growths and input to a control arrangement for subsequent time steps.

Abstract: Systems and devices configured for active thermal control of turbine components are disclosed. In one embodiment, a thermal control system for a turbine includes: a thermal source shaped to connect to a turbine; a set of sensors disposed about the turbine and configured to obtain operational data from the turbine; and a computing device communicatively connected to the thermal source and the set of sensors, the computing device configured to regulate a thermal input of the thermal source to the turbine based on the operational data obtained by the set of sensors.

Abstract: A gas turbine engine includes an annular plenum defined with an outer skin and a perforated inner skin for receiving selective air flow to impinge a support case which supports shrouds of the rotor assemblies of the engine therein for active tip clearance control of the rotor assemblies. In one embodiment a bobbin-type transfer tube for supplying cooling air into the plenum, is provided between an outer case of the engine an the plenum such that the thermally induced relative movement of the outer case and the plenum is permitted.

Abstract: A triple hook ring segment including forward, midsection and aft mounting hooks for engagement with respective hangers formed on a ring segment carrier for supporting a ring segment panel, and defining a forward high pressure chamber and an aft low pressure chamber on opposing sides of the midsection mounting hook. An isolation plate is provided on the aft side of the midsection mounting hook to form an isolation chamber between the aft low pressure chamber and the ring segment panel. High pressure air is supplied to the forward chamber and flows to the isolation chamber through crossover passages in the midsection hook. The isolation chamber provides convection cooling air to an aft portion of the ring segment panel and enables a reduction of air pressure in the aft low pressure chamber to reduce leakage flow of cooling air from the ring segment.

Abstract: A feather seal is positioned in a mold and a shroud body is metal injection molded about a proximal end of the feather seal to provide a shroud segment with an integrated feather seal. A seal groove is provided in an opposite lateral side of the shroud body to receive the feather seal of a circumferentially adjacent shroud segment.

Abstract: The disclosure relates to a turbo-machine comprising a stator and a rotor arranged rotatable inside the stator as well as at least one electric heating device which is arranged on the surface of at least part of the stator for active clearance control. Besides the turbo-machine, a method for operating the active clearance control comprising electric heating devices is disclosed.

Abstract: A seal assembly includes a body having two circumferential sides, a leading end, and a trailing end. At least one of the circumferential sides includes a first channel sidewall, a second channel sidewall, and a channel bottom wall that together define a seal channel for receiving a seal. The seal channel includes a slot that cooperates with a tab on the seal to facilitate securing the seal within the seal channel.

Abstract: A turbine shroud sealing apparatus for a gas turbine engine includes: (a) an arcuate shroud segment comprising a 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 of the segment; and (b) a first seal assembly received in at least one slot formed in the first end face, the first seal assembly comprising one or more spline seals which protrude from the first end face and which are arranged to define a continuous sealing surface around the perimeter of the first end face.

Abstract: A casing component (2) of a turbomachine, the casing component (2) comprising: a plurality of casing elements (14) which define a diameter of the casing component; and an actuation means operable to change the diameter of the casing component (2), wherein the actuation means changes the diameter of the casing component (2) as a function of a rotational speed of a rotatable component (4) disposed within the casing component (2).

Abstract: A feather seal assembly includes a seal having a directional passage to direct an airflow generally non-perpendicular to the seal.

Abstract: An apparatus and method is disclosed herein in which a heat-transfer fluid can be directed through a gap defined between a fillet and a chamfer. The apparatus includes a first member having a body portion and a flange portion projecting from the body portion. The flange portion extends along an endless path encircling an axis to define a receiving aperture. A fillet is defined at a junction of the body portion and an inner surface of the flange portion. The apparatus also includes a second member having a plug portion receivable in the receiving aperture. The plug portion includes a first surface operable to abut the body portion and limit movement of the plug portion into the receiving aperture. The plug portion also includes a second surface slidably engageable with the inner surface of the flange portion to guide movement of the plug portion into the receiving aperture along the axis.

Abstract: A shroud for turbine engines. The shroud has an integrated anti-rotation device that prevents circumferential movement of the shroud during normal engine operation, and which allows for circumferential installation in split annular case designs. Since the anti-rotation device is an integral part of the shroud and/or annular split turbine case, no additional parts are necessary for assembly or disassembly. Moreover, existing annular split turbine cases can be reworked to accept the anti-rotation device and yet still be backwards compatible with original shroud designs.

Abstract: A seal structure for a gas turbine engine, the seal structure including first and second components located adjacent to each other and forming a barrier between high and low pressure zones. A seal cavity is defined in the first and second components, the seal cavity extending to either side of an elongated gap extending generally in a first direction between the first and second components. A seal member is positioned within the seal cavity and spans across the elongated gap. The seal member includes first and second side edges extending into each of the components in a second direction transverse to the first direction, and opposing longitudinal edges extending between the side edges generally parallel to the first direction. The side edges include a groove formed therein for effecting a reduction of gas flow around the seal member at the side edges.

Abstract: According to one aspect of the invention, an assembly for preventing fluid flow between turbine components includes a shim and a first woven wire mesh layer that includes a first surface coupled to a first side of the shim and a second surface of the woven wire mesh layer opposite the first surface. The assembly also includes a first outer layer coupled to the second surface of the woven wire mesh layer, where the first outer layer includes a high temperature non-metallic material.

Abstract: A method for controlling turbine blade tip seal clearance on a non-cowled gas turbine engine includes the step of directing a flow of relatively hot, pressurized air in contact with an exterior surface of the engine core casing circumscribing the turbine. The flow of relatively hot, pressurized air may be selectively directed in contact with the exterior surface of the engine core casing circumscribing the turbine when it is desired to expand the turbine shroud radially outward, such as during transient operation of the gas turbine engine.

Abstract: A turbine vane or blade segment includes at least one airfoil extending radially outwardly from a radially inner band. A plurality of cooling passages are formed in the radially inner band in fluid communication with an internal plenum in the airfoil and exiting the inner band via a plurality of exit holes in the one of the axially-extending side edges of the inner band. The plurality of exit holes are confined to a region along the one of the axially-extending side edges where static pressure PS along the one of the substantially axially-extending side edges lies in a pressure range substantially between the stage inlet total pressure PT and a pressure that is substantially about 1.5 times the dynamic pressure range ?PD, below the stage inlet total pressure.

Abstract: Shroud assembly for use in a turbine, having an outer shroud having a side proximate to flow of operative fluid of the turbine, the side having a boss and an outer shroud aperture extending through the boss. An inner shroud is located between the side of the outer shroud proximate to the flow of operative fluid and the operative fluid flow of the turbine, the inner shroud having an inner shroud aperture which is dimensioned to fit over the boss.

Abstract: A turbine shroud (30) of a gas turbine engine comprises a plurality of arcuate shroud segments (31) combined into an annular configuration, each shroud segment including a main body (32) defining an inner circumferential surface opposing the tips of the turbine rotor blades (11a) at a small clearance and an engagement feature including an axial wall (33a, 34a) having a prescribed circumferential length and a prescribed axial length, the turbine casing including an axial slot (51, 52) extending coaxially around the center line of the engine and configured to receive the axially extending wall of each shroud segment. A clearance defined between each circumferential end part (E) of the axial wall and an opposing inner circumferential surface of the turbine casing is greater than that defined between a circumferentially middle part (M) of the axial wall and an opposing inner circumferential surface of the turbine casing under a cool condition of the engine.

Abstract: A ferrule of a stator with blades of an axial turbomachine consists of at least two segments having the general form of an arc of a circle and intended to be arranged end to end, so as to form the ferrule, where each segment contains on its outside a plurality of openings arranged at regular distances along the general arc of the circle of the segment, into each of which fits the end of a blade of the stator. Each end of a segment possesses a section of which the profile essentially corresponds with that of the end of the corresponding blade when the segment is placed on the blades of the stator, such that the joint between two adjacent segments corresponds with the profile of the corresponding blade.

Abstract: A blade outer air seal (BOAS) of a gas turbine engine has a segmented support ring to support a segmented turbine shroud. The support ring has a cooling air distribution system which includes a plurality of inlet cavities extending axially and inwardly to communicate with an inner cooling air passage within the respective support segments. The inlet cavities each are formed with two recesses defined in respective adjacent two support segments.

Abstract: An article of manufacture has a body formed in part of a first metal alloy and in part of a second metal alloy, the second metal alloy having a thermal coefficient of expansion that is less than the thermal coefficient of expansion of the first metal alloy. A BOAS segment for a gas turbine engine is disclosed wherein the formation of cracks due to thermal mechanical fatigue in the body of the disclosed BOAS segment is minimized, if not eliminated, through a unique construction of the disclosed BOAS segment, whether original equipment manufacture or a repaired blade outer air seal. A method for manufacture of a BOAS segment and a method for modifying a BOAS segment are disclosed.

Abstract: An assembly for a turbomachine comprising: a plurality of components (16, 26) which, in use, are assembled into a circumferentially extending annular array; each component having a radially outer platform (20, 28), the outer platforms (20, 28) of the components (16, 26) forming a ring in the assembled array; wherein adjacent components (16, 26) interface at adjacent surfaces (22, 30) of their outer platforms (20, 28); wherein the outer platforms (20, 28) each comprise at least one lug (24) which extends from the outer platform (20) over a portion of the outer platform (28) of an adjacent component (26), and in use prevents relative radial displacement of the outer platform (20, 28) of the adjacent component (16, 26).

Abstract: A blade outer air seal with ring segments forming a mate face gap in which a riffle seal is placed to seal the axial gap. The riffle seal includes a horizontal plate and a vertical plate extending from a bottom surface of the horizontal plate and occupies the axial gap space. The bottom end of the vertical extending plate is angled in a direction of rotation of rotor blades and includes ribs that form open slots in the riffle seal. The ring segments include metering holes that discharge cooling air into the slots to discharge film cooling air onto the hot gas surface of the BOAS and cool the ring segments.

Abstract: A system for recirculating a hot gas flowing through a gas turbine generally includes a transition piece having a downstream surface, a stationary nozzle having a leading edge surface adjacent to the transition piece downstream surface, a gap defined between the transition piece downstream surface and the stationary nozzle leading edge surface, and a projection having an inner arcuate surface radially separated from an outer surface. The projection generally extends from the stationary nozzle leading edge surface towards the transition piece downstream surface and at least partially decreases the gap. A seal extends across the gap and may be in contact with the transition piece and the stationary nozzle leading edge surface. A recirculation zone may be at least partially defined between the projection inner arcuate surface, the stationary nozzle leading edge surface and the transition piece downstream surface.

Abstract: Gas turbine engines and related systems involving blade outer air seals are provided. In this regard, a representative blade outer air seal segment for a set of rotatable blades includes: a blade arrival end; and a blade departure end; each of the blade arrival end and the blade departure end being angularly offset with respect to a longitudinal axis about which the blades rotate.

Abstract: In an embodiment, a system includes a rotary machine. The rotary machine includes a rotor having a plurality of blades and a segmented stator having a plurality of stator segments arranged circumferentially about the plurality of blades. Each adjacent pair of first and second segments of the plurality of stator segments includes a recess extending circumferentially across an intermediate joint between the first and second segments. Furthermore, each recess includes at least one eccentricity control insert configured to mitigate eccentricity of an inner circumference of the segmented stator due to thermal expansion or thermal contraction of the segmented stator.

Abstract: Systems, methods, and apparatus for linking machine stators are provided. A bracket may be utilized to link adjacent stators within a machine, for example, a turbine. The bracket may include a body portion, a first plurality of extensions, and a second plurality of extensions. The body portion may include a first edge operable to align with a base of a first stator and a second edge opposite the first edge and operable to align with a base of a second stator. The first plurality of extensions may extend from the first edge and may be operable to align with at least one corresponding groove on the base of the first stator. The second plurality of extensions may extend from the second edge and may be operable to align with at least one corresponding groove on the base of the second stator. When utilized to link the first stator and the second stator, the bracket may facilitate a reduction of vibrations in the first stator and the second stator.

Abstract: A seal member for use in a channel between a transition seal structure and a vane seal structure in a gas turbine engine. The seal member includes a spring member and a sheathing assembly. A first end of the spring member is affixed to either the transition seal structure or the vane seal structure. The second end is free to move within the channel. The sheathing assembly includes a main body and a plate portion. The main body surrounds the spring member and is affixed to the second end thereof. The plate portion extends from the main body and is adapted to extend toward the other of the transition seal structure and the vane seal structure. The spring member provides a bias on the sheathing assembly such that the plate portion engages the other of the transition seal structure and the vane seal structure to limit leakage through the channel.

Abstract: A seal housing is provided to substantially cover at least one duct wall of vane array duct of a gas turbine engine, and one example arrangement is employed in a mid-turbine frame. The arrangement provides improved sealing of the vane array duct through the provision of a plurality of cavities extending along the duct wall. The arrangement may also include insulation tubes to assist in sealing around load transfer spokes passing through the vane array.

Abstract: A stator vane assembly with a mate face gap and a seal slot to receive a seal pin. The seal pin includes a row of axial cooling air channels opening on a top side of the seal pin and extending toward a forward end of the seal pin, and a row of metering holes that supply cooling air from the gap below the seal pin to each of the axial cooling channels. Cooling air flows through the metering holes and along the axial cooling channels to provide cooling for the endwall mate face surfaces and the top of the seal pin exposed to a hot gas flow in the gap. Vortex chambers are formed on the forward ends of the seal pin mate face slots, and cooling air holes discharge cooling air from the vortex chambers downward from the vane leading edge corner.

Abstract: In exemplary embodiments, a nozzle can include a first flow wall, a second flow wall and a vane disposed between the first and second flow walls, wherein the vane is mechanically coupled to the first flow wall and in contact with the second flow wall.

Abstract: A turbine nozzle is constructed from a plurality of associated circumferential turbine nozzle segments each of which incorporates at least one pre-formed annular-segment passage bounded at a first azimuthal boundary by a first portion of a first bounding nozzle vane, and at a second azimuthal boundary by a second portion of a second bounding nozzle vane, with the at least one pre-formed annular-segment passage therebetween, wherein when assembled in the turbine nozzle, a first portion of a first bounding nozzle vane of one circumferential turbine nozzle segment is joined to a second portion of a second bounding nozzle vane of another adjacent circumferential turbine nozzle segment so as to form therebetween a nozzle vane of the turbine nozzle, wherein the minimum through-flow area of the associated at least one pre-formed annular-segment passage is substantially unaffected by the assembly of the circumferential turbine nozzle segments to form the turbine nozzle.

Abstract: A spline seal for a hot gas path component is provided. The spline seal may include a first metal layer and a second metal layer. The first metal layer may have a first volumetric thermal expansion coefficient. The second metal layer may be disposed adjacent the first metal layer and have a second volumetric thermal expansion coefficient. The second volumetric thermal expansion coefficient may be higher than the first volumetric thermal expansion coefficient. When the spline seal is exposed to a heat source, the first and second metal layers may deform to provide a seal between the hot gas path component and an adjacent hot gas path component.

Abstract: A stator vane with a mate-face cooling and sealing arrangement, the vane end rails include an upper axial U-shaped seal slot and a lower axial U-shaped seal slot with an upper U-shaped riffle seal and a lower U-shaped riffle seal secured within the two axial U-shaped seal slots. The two U-shaped seals form a box-like arrangement with riffle teeth formed on the top sides of a horizontal plate for each of the U-shaped seals.

Abstract: A segment of a component for use in a gas turbine includes a leading edge; a trailing edge; a pair of opposed lateral sides between the leading and trailing edges; and a seal slot provided in each lateral side. The seal slot includes a surface having a channel extending in an axial direction defined from the leading edge to the trailing edge, at least one inlet to the channel, and at least one outlet from the channel. The at least one outlet is spaced downstream from the at least one inlet in the axial direction. The segment may be an inner shroud segment or a nozzle segment.

Abstract: A gas turbine is provided that includes a rotor which is rotatable around an axis and equipped with rotor blades, and which is concentrically enclosed at a distance by a casing, which is equipped with stator blades, forming an annular hot gas passage. Rings with stator blades and rotor blades are arranged in a manner alternating in the axial direction. Between adjacent stator blades, heat shield segments are arranged, which delimit the hot gas passage on the outside in a region of the rotor blades and are cooled by impingement cooling where a cooling medium from an outer annular cavity flows into the heat shield segment.

Abstract: The invention relates to a compressor, particularly a high pressure compressor, of a gas turbine, particularly of a gas turbine aircraft engine, having at least one rotating blade ring on the rotor side, and having at least one guide blade ring on the stator side, wherein the, or each, guide blade ring is formed of a plurality of guide blade segments, and wherein each guide blade segment is formed of a plurality of individual blades. According to the invention, adjacent individual blades (11, 12; 12, 13; 13, 14) are permanently connected to each other within each guide blade segment (10) of at least one guide blade ring on opposing surfaces positioned radially outward, whereas the same are not connected to each other on opposing surfaces positioned radially inward.

Abstract: A seal structure is provided for preventing leakage of gases across a gap between first and second components in a turbine engine. The seal structure is adapted to be received in first and second adjacent slots provided in the first and second components. The seal structure may comprise: a wear resistant layer; and a deformable layer defined by a material having one of a varying density and a varying porosity.

Abstract: A seal assembly includes a body having two circumferential sides, a leading end, and a trailing end. At least one of the circumferential sides includes a first channel sidewall, a second channel sidewall, and a channel bottom wall that together define a seal channel for receiving a seal. The seal channel includes a slot that cooperates with a tab on the seal to facilitate securing the seal within the seal channel.

Abstract: An arrangement between blade elements in a blade row in a turbine is described. Each blade element has at least one shroud element and a blade airfoil which abuts on, and is connected to, the shroud element, and essentially extends radially with regard to a principal axis of the blade row. When installed, the shroud element sides, which extend circumferentially, abut on the respectively adjacent shroud element of the respectively adjacent blade element, each forming an essentially radial gap. At least one blade element has a projection which projects into the shroud element of the abutting blade element and extends in the circumferential direction, and at least one blade element has a recess which accommodates such a projection. In the region of the projection or recess there is a stepped region of the radial gap, and the guiding of the radial gap in this stepped region is a labyrinth seal.

Abstract: A blade outer air seal segment assembly for a turbine in which adjacent segments form a curved metering and diffusion cooling slot for the discharge of film cooling air onto an inner surface of the segments. The segments include a first diffusion cavity connected to the backside surfaces of the segments through metering holes. The curved slot is connected to the first diffusion cavity. spend cooling air for cooling of the backside surfaces of the segments is metered into the first diffusion cavity and then discharged as film cooling air from the curved slot in a direction of rotation of the rotor blades to provide both sealing and cooling for the mate face of the segments.

Abstract: The gas turbine engine shroud comprises a plurality of circumferentially-disposed and concentric shroud segments. Each shroud segment has an arc-shaped platform with opposite ends, each end comprising an inter-segment seal slot, at least one slot extending substantially across each corresponding end and having a lengthwise-variable depth.

Abstract: A bi-metallic strip seal used to seal the gap between adjacent shroud segments in a gas turbine engine. The bi-metallic strip seal is formed from one or more layers of bi-metallic materials with different coefficients of thermal expansion. In a cold state the bi metallic strip seal will fit within the slot easily, while in the hot state the strip seal will expand to fit tightly within the slot and block leakage flow. The strip seal can be formed with the bi-metallic layers oriented similarly or oppositely in order to control the seal force and deflection properties. In one embodiment, the strip seal is formed from four bi-metallic layers in which the middle is bonded while the adjacent layers are free from shear.

Abstract: A sealing arrangement for use in a turbine engine having ceramic components. The sealing arrangement is retained in a seal gap formed between adjacent segments and is compliant to accommodate variations in the size of the seal gap as the adjacent segments move relative to one another.

Abstract: In one embodiment, a system includes a turbine engine that includes a rotor including multiple blades. The turbine engine also includes a shroud disposed about the blades. The shroud includes multiple segments engaged with one another via mating teeth. The mating teeth are oriented in an axial direction along a longitudinal axis of the turbine engine.

Abstract: Systems involving feather seals are provided. A representative vane assembly for a gas turbine engine includes: a first mounting platform having a first slot; a first airfoil extending from the first mounting platform; and a feather seal having opposing faces, a first side extending between the faces, and a first tab, the first tab extending outwardly beyond the first side; the first slot being sized and shaped to receive the feather seal including the first tab.

Abstract: A locking member for a device for fastening ring sectors to an aircraft turbine engine casing is disclosed. The member has first and second clamping branches, which are connected together by curving joining portions, via a connecting branch. The joining portion of center thickness E1 has an inside surface of mean radius Rm1, and the joining portion of center thickness E2 has an inside surface of mean radius Rm2. Thickness E1 is strictly smaller than thickness E2, and mean radius Rm1 is strictly smaller than mean radius Rm2.

Abstract: A vane carrier assembly is provided for supporting vanes within a main engine casing of a gas turbine engine. The vane carrier assembly comprises a plurality of vane support panels positioned adjacent to one another so as to define a vane support assembly. The support panels are assembled such that the support panels expand circumferentially to minimize radial expansion of the vane support assembly during operation of the gas turbine engine. A control ring is coupled to the main engine casing, and the vane support assembly is coupled to the control ring. The control ring may be formed from a material having a coefficient of thermal expansion less than that of the material from which the vane support assembly is formed.

Abstract: A gap seal (9) radially seals a gap (8) which extends axially and radially between two blades (1, 2) of a turbomachine which are adjacent in the circumferential direction (3). The two blades (1, 2) have, in each case, an axially extending longitudinal slot (10, 11), which is open towards the gap (8), on its respective blade root (6, 7). A band-form or strip-form sealing element (12) engages with its longitudinal sides (13, 14) in the two longitudinal slots (10, 11) and bridges the gap (8). The one blade (1) has a projection (15) on its blade root (6), which projects from the blade root (6) in the circumferential direction (3) and extends in the circumferential direction (3) and radially, at least in the region of the respective longitudinal slot (10), and bridges an axial longitudinal end of the gap (8) in the process. The other blade (2) has a step-shaped recess (16) on its blade root (7), complementary to the projection (15) of the one blade (1) and in which the projection engages.

Abstract: A compressor has a rotor (13), which rotor (13) has a number of rows of rotor blades (14, 14?) which are at a distance from one another, one behind the other in the axial direction, with a number of thermal barrier segments (10), which are detachably attached to the rotor (13) and are mounted such that they can move in the circumferential direction, being arranged one behind the other in the circumferential direction on the circumference of the rotor (13), and with securing devices (15) being provided on the thermal barrier segments (10), which secure the thermal barrier segments (10) against being moved in the circumferential direction. In the case of such a rotor (13), production and assembly are simplified in that only some of the thermal barrier segments (10) are equipped with the securing devices (15).