Abstract: A turbine engine shroud segment is provided with a radially outer surface including a pair of spaced apart, opposed first and second edge portion surface depressions, for example spaced circumferentially, having a seal surface shaped to receive, in a surface depression formed between assembled adjacent segments in a shroud assembly, or axially assembled adjacent segments and engine members, a radially outer fluid seal member. The depression portions of a shroud segment are joined with the radially outer surface of the shroud segment through an arcuate transition surface. Stresses generated during engine operation in the shroud material are reduced, enabling practical use of a low ductility material, for example a ceramic matrix composite. The edge portion surface depressions are provided with a first shape and the fluid seal member, disposed at the depression, is provided with a surface of a second shape matched in shape with the first shape.

Abstract: A turbine seal closure mechanism for sealing an opening in a turbine component or between one or more turbine components. The turbine seal may include a main body having a longitudinal axis positioned transverse to the axial direction of the turbine and having at least one arm for sealing the main body to a surface of a turbine component and configured to fit within a recess in a turbine component. The turbine seal may also have one or more arms coupled to a side of the main body opposite to the first side having arms attached thereto for sealing the turbine seal to another turbine surface.

Abstract: At least one shroud segment, floating axially independently of adjacent turbine engine shroud assembly members, includes a segment body comprising a radially outer surface and a radially outwardly projecting segment support that includes an axial support wall surface therein. The assembly includes a shroud hanger in axial juxtaposition with the segment support, and at least one axial support projection from the shroud hanger into the segment support at the support wall surface. The support projection supports the shroud segment releasably at the support wall surface sufficiently to enable relative axial movement of the shroud segment on the support projection independently of the shroud hanger and adjacent engine members.

Abstract: A gap seal in a gas turbine for sealing off a cooling space, formed between a heat shield and a heat shield carrier, from a hot gas space of the gas turbine. A sealing body is retained in a movable manner on the heat shield, on the one hand, and on a blade carrier, on the other hand, adjacent to the heat shield carrier.

Abstract: Nozzle segments mounting vanes are received in circumferentially extending, generally dovetail-shaped grooves in an outer casing of a steam turbine, the nozzle segments forming part of a stage with rotating buckets of the steam turbine. The inclined slashfaces of the adjoining bases of the nozzle segments are provided with circumferentially opening slots to receive spline seals. The spline seals preclude or minimize steam leakage flow past the gap between the adjoining nozzle segments thereby enhancing the steam flow through the partitions of the nozzles.

Abstract: Spline seals are provided in registering end faces of upper and lower diaphragm segments in a steam turbine at each of the horizontal joints. The diaphragm segments each include inner and outer semi-annular rings with circumferentially spaced partitions extending therebetween. The registering end faces of each of the inner and outer semi-annular rings of the segments are spanned by spline seals which may comprise a sheet metal plate, a sheet metal body having reversely-curved ends, a solid body having bulbous ends or a sheet metal plate covered by a metallic or ceramic woven cloth. Leakage paths, both in axial and radial directions, are minimized or precluded.

Abstract: A stator shroud segment is provided that includes an outer shroud having a leading edge groove and a trailing edge groove, both grooves of the outer shroud opening in a first, axial direction; and a plurality of inner shrouds each having a leading edge hook and a trailing edge hook. The hooks of the inner shrouds project in a second, axial direction, diametrically opposite the first axial direction and the leading and trailing hooks of each of the inner shrouds are respectively engaged with the leading and trailing edge grooves of the outer shroud so as to axially and radially lock the inner shrouds to the outer shroud. The assembly simplifies access to and removal of the inner shroud(s) without added complexity.

Abstract: A turbine engine shroud segment having a body including a circumferentially arcuate radially inner surface defining a circumferential arc and a radially outer surface is provided, at least at one axially spaced apart outer surface edge portion surface, with a surface depression extending circumferentially across the outer edge portion and including a planar seal surface. The planar seal surface is spaced apart radially outwardly from the circumferential arc defining a spaced apart chord of the arc. The planar seal surface is joined with the segment body radially outer surface through an arcuate transition surface. In a circumferential assembly of a plurality of the shroud segments into a turbine engine shroud assembly, at least one of the outer surface edge portions and its respective depression portion and fluid seal surface is distinct axially from an axially juxtaposed engine member by a separation therebetween.

Abstract: A turbine engine shroud segment having a body including a circumferentially arcuate radially inner surface defining a circumferential arc and a radially outer surface is provided, at least at one axially spaced apart outer surface edge portion surface, with a surface depression extending circumferentially across the outer edge portion and including a planar seal surface. The planar seal surface is spaced apart radially outwardly from the circumferential arc defining a spaced apart chord of the arc. The planar seal surface is joined with the segment body radially outer surface through an arcuate transition surface. In a circumferential assembly of a plurality of the shroud segments into a turbine engine shroud assembly, at least one of the outer surface edge portions and its respective depression portion and fluid seal surface is distinct axially from an axially juxtaposed engine member by a separation therebetween.

Abstract: In a gas turbine having a chordal hinge seal between an inner rail of each nozzle segment and an annular axially facing sealing surface of a nozzle support ring, a supplemental seal is disposed between the support ring and inner bands of the nozzle segments on a lower pressure side of the chordal hinge seals. To minimize or prevent leakage flow across the chordal hinge seals, a generally U-shaped supplemental seal having reversely folded U-shaped marginal portions is received in a cavity formed in the axially extending sealing surface of the inner rail of the nozzle segment. At operating conditions, the marginal portions seal against the base of the cavity and the annular sealing surface of the nozzle support ring to prevent leakage flow past the chordal hinge seal from entering the hot gas path.

Abstract: A shroud assembly for a turbine portion of a gas turbine engine, the shroud assembly comprising an annular ceramic shroud ring, a plurality of arcuate shroud support segments, and a plurality of inwardly biased resilient members. The annular ceramic shroud ring is circumferentially disposed about radially extending blades of a turbine rotor and partially defines an annular hot gas passage of the turbine portion. The plurality of arcuate shroud support segments are radially disposed outwardly of the ceramic shroud ring and are contiguous therewith. The plurality of inwardly biased resilient members are each engaged between one of the shroud support segments and an outer annular turbine support case composed of a material having a different thermal expansion coefficient than the ceramic shroud ring. The resilient members maintain contact between the shroud support segments and the ceramic shroud ring.

Abstract: A gas turbine having a metallic outer shroud and a ceramic inner shroud. The ceramic inner shroud being secured to the outer shroud by hooks carried on the outer shroud. The hooks on the outer shroud being coated with TBC. A pin and spring system being provided to hold the ceramic inner shroud against the forward hook of the outer shroud. An anti-rotation pin being provided to trap the aft bend of the ceramic inner shroud against the aft hook of the outer shroud. The gas turbine further including a damping spring and pin system, disposed between a heat shield, within the outer shroud, and the ceramic inner shroud, to provide damping of the inner shroud.

Abstract: Spline seals are disposed in circumferentially registering slots of adjacent arcuate seal strip segments disposed in grooves of diaphragm assemblies of a steam turbine in radial opposition to the rotating tips of the buckets. The spline seals extend in the gaps between the endfaces of the segments and minimize or preclude steam leakage past the endfaces. The spline seals are disposed in slots which may be formed as part of original equipment manufacture or may be machined in segments with the spline seals provided as retrofits.

Abstract: Spline seals are disposed in circumferentially registering slots of adjacent arcuate sealing segments disposed in grooves of the casing of a steam turbine about a rotor. The spline seals extend in the gap between the endfaces of the segments and minimize or preclude steam leakage past the endfaces. The spline seals may be oriented axially or inclined radially to minimize leakage paths in the radial inclined and axial directions, respectively. The spline seals are disposed in the slots which may be formed as part of original equipment manufacture or may be machined in segments with the spline seals provided as retrofits.

Abstract: A method of producing a laminated iron core by punching and laminating rotor and stator core pieces. The rotor core pieces are punched from a metal sheet and laminated. Magnetic pole teeth are roughly punched in the metal sheet with the rotor core pieces removed. Each of a plurality of magnetic pole teeth are pressed at plural different portions to form thin parts so as to develop the plurality of magnetic pole teeth toward the punched side of the rotor core piece. Thereafter, the stator core piece is punched to form front tips on the magnetic pole teeth defining an internal form on the stator core piece. An external form of the stator core piece is punched. The punched stator core pieces are laminated in succession.

Abstract: A vane cluster for a turbine engine compressor or turbine includes a shroud 12 with a nonlinear slot 26 extending therethrough to divide the shroud into thermally independent shroud segments 24. The slot is bordered by matching nonlinear surfaces 28 that are easy and inexpensive to produce with conventional wire EDM equipment. The nonlinear profile of the slots effectively resists fluid leakage.

Abstract: A process for producing a groove-like recess which is part of a strip seal between two components which enclose a gap, uses a material-removing process technology to create the groove-like recess in the two components in such a manner that the groove-like recess in each of the two components, after orientation with respect to one another, lie substantially opposite one another in an end position in which they enclose the gap between them, and that a strip element can be introduced into the groove-like recesses in such a manner that the strip element at least partially projects into both groove-like recesses and, under the action of a force, forms a substantially gastight joint with the groove-like recesses in the two components. Furthermore, a groove-like recess of this type is part of a strip seal.

Abstract: A turbine includes a sealing element with a receiving area for sealing the guide blade vanes which are adjacent to each other in the peripheral direction of the turbine. The foot plates of the guide blade vanes extend into the receiving area. The edge area of the foot plates does not have to be reinforced compared to a conventional seal, which enables the entire foot plate to be cooled homogeneously. A closed cooling system can therefore be used for cooling, especially with steam.

Abstract: A turbine installation, especially a gas turbine installation, includes foot plates of the guide blades of adjacent turbine stages being interconnected with a clip-type sealing element on their rear sides facing away from the gas area. This provides a simple seal between adjacent foot plates which is effective regardless of the thermal expansion of the foot plates. The clip-type sealing element is also suitable for sealing the tiles of a combustor of the turbine installation together.

Abstract: A turbine engine shroud segment, preferably a low ductility material, for example a ceramic matrix composite, comprises a segment body extending between body circumferential ends. The body includes a body radially inner surface arcuate at least circumferentially, and a body radially outer surface from which a plurality of hooks extend generally radially outwardly. Each hook comprises a generally radially outwardly extending hook arm with a generally axially extending hook end having a generally radially inner surface in spaced apart juxtaposition with a portion of the body radially outer surface. Such body outer surface includes at least two body contact surfaces each matched in shape and in juxtaposition at least at body circumferential ends with a cooperating hanger contact surface. The radially inner surface of each hook end includes a hook end contact surface matched in shape with a cooperating hanger contact surface at least in a circumferential middle portion of such surface.

Abstract: Spline seals are disposed in circumferentially registering slots of adjacent arcuate seal strip segments disposed in grooves of diaphragm assemblies of a steam turbine in radial opposition to the rotating tips of the buckets. The spline seals extend in the gaps between the endfaces of the segments and minimize or preclude steam leakage past the endfaces. The spline seals are disposed in slots which may be formed as part of original equipment manufacture or may be machined in segments with the spline seals provided as retrofits.

Abstract: The invention pertains to a seal element (1) for sealing a gas-path leakage-gap (5) between spaced apart first and second components (2,3) of a turbo machinery (22), which first and second components (2,3) having opposing inner and outer surfaces (9, 10) situated outside the leakage-gap (5). The seal element (1) comprises a generally gas impervious sealing member (4) and a layer (6) comprising ceramic fibers. The layer (6) covers at least partially said sealing element (1) and defines a sealing surface (21) for contacting outside the gap (5) the outer surfaces (10) of the first and second components (2,3) of the turbo machinery (22).

Abstract: Spline seals are disposed in circumferentially registering slots of adjacent arcuate sealing segments disposed in grooves of the casing of a steam turbine about a rotor. The spline seals extend in the gap between the endfaces of the segments and minimize or preclude steam leakage past the endfaces. The spline seals may be oriented axially or inclined radially to minimize leakage paths in the radial inclined and axial directions, respectively. The spline seals are disposed in the slots which may be formed as part of original equipment manufacture or may be machined in segments with the spline seals provided as retrofits.

Abstract: The stator shroud is made up of components (2) juxtaposed so as to form circumferences, which are successive and assembled separately to a common housing which surrounds them. Characteristically, the assembly is made with only one single fixing point (7) per component (2), while being completed by a support point (21) sliding along the circumference. The result is a more flexible component assembly and a more even distortion, which allows better control of the clearances between the shroud and the rotor blades it surrounds.

Abstract: A turbine engine shroud segment, preferably a low ductility material for example a ceramic matrix composite, comprises a segment body extending between body circumferential ends. The body includes a body radially inner surface arcuate at least circumferentially, and a body radially outer surface from which a plurality of hooks extend generally radially outwardly. Each hook comprises a generally radially outwardly extending hook arm with a generally axially extending hook end having a generally radially inner surface in spaced apart juxtaposition with a portion of the body radially outer surface. Such body outer surface includes at least two body contact surfaces each matched in shape and in juxtaposition at least at body circumferential ends with a cooperating hanger contact surface. The radially inner surface of each hook end includes a hook end contact surface matched in shape with a cooperating hanger contact surface at least in a circumferential middle portion of such surface.

Abstract: An inner shroud assembly for a turbine comprising a plurality of part-annular segments combining to form an inner, annular shroud adapted to surround rotating components of a turbine, each segment having a pair of end faces that are juxtaposed similar end faces on adjacent segments with gaps therebetween; at least one convection cooling hole in the segment, opening along at least one of the pair of end faces. The cooling hole opens specifically into a diffuser recess formed in one of the pair of end faces for diffusing the flow of cooling air into the gap.

Abstract: An inner shroud assembly for a turbine comprising a plurality of part-annular segments combining to form an inner, annular shroud adapted to surround rotating components of a turbine, each segment having a pair of end faces that are juxtaposed similar end faces on adjacent segments with gaps therebetween; at least one convection cooling hole in the segment, opening along at least one of the pair of end faces. The cooling hole opens specifically into a diffuser recess formed in one of the pair of end faces for diffusing the flow of cooling air into the gap.

Abstract: A split ring is disposed on an inner wall of a gas turbine casing. The split ring is composed of a plurality of split segments that are arranged on the inner wall of the casing in circumferential direction. A predetermined clearance is formed between the inner face of a split segment and the rotor blade tips. The split segments are arranged so that a predetermined circumferential clearance is formed between the split segments in order to allow the thermal expansion of the segments. A circumferential end face located upstream side of the segments with respect to the direction of the rotor blade rotation is connected to the inner face by a transition face having a surface formed as an inclined plane. The inclined plane prevents the swirl flow caused by the rotating rotor blade from impinging the upstream end face and, thereby, suppresses a temperature rise of the split segment at the upstream end face.

Abstract: The gas-path leakage seal is provided for generally sealing a gas-path leakage-gap between spaced-apart first and second members of a turbine. The seal includes an elongated seal member having an elongated, imperforate, and flexible first portion and a rigid second portion that defines a mounting bracket. The first portion has first and second surfaces and raised longitudinal edges that extend a generally identical distance above the first surface and generally the aforesaid identical distance below the second surface. A cloth layer having a thickness generally equal to or slightly greater than the aforesaid identical distance is superimposed on the second surface between the raised edges.

Abstract: A gas turbine segmental ring has an increased rigidity to suppress a thermal deformation and enables less cooling air leakage by less number of connecting portions of segment structures. Cooling air (70) from a compressor flows through cooling holes (61) of an impingement plate (60) to enter a cavity (62) and to impinge on a segmental ring (1) for cooling thereof. The cooling air (70) further flows into cooling passages (64) from openings (63) of the cavity (62) for cooling an interior of the segmental ring (1) and is discharged into a gas path from openings of a rear end of the segmental ring (1). Waffle pattern (10) of ribs arranged in a lattice shape is formed on an upper surface of the segmental ring (1) to thereby increase the rigidity. A plurality of slits (6) are formed in flanges (4, 5) extending in the turbine circumferential direction to thereby absorb the deformation and thermal deformation of the segmental ring (1) is suppressed.

Abstract: A seal is provided between a pair of members by a flexible seal element fixed to one member and having a turned edge. The seal element overlies a contact surface of the other member to effect the seal, the seal element spanning a gap between the members and being pressed into the contact surface by a high pressure region on the same side of the gap as the sealing element. In other forms, a second sealing element is employed in conjunction with a spline seal forming a tortuous sealing path between the members.

Abstract: In a heat shield (10), in particular for the stator of gas turbines, which heat shield (10) is composed of a plurality of individual segments (10a, b; 20a, b), whose end surfaces (15a, b) respectively abut one another so as to form a gap (12), and which have cooling holes (13a, b) for cooling purposes in the region of the end surfaces (15a, b), through which cooling holes (13a, b) a cooling fluid is blown out into the gap (12), cooling is ensured, even when the gap is closed, by a chamber (11), which is widened relative to the gap (12) and into which the cooling holes (13a, b) open, being arranged in the region of the gap (12).

Abstract: A gas turbine engine component including a nozzle outer band, a plurality of nozzle vanes extending inward from the outer band, and an inner band extending circumferentially around inner ends of the vanes. Further, the component has a shroud integral with the outer band adapted for surrounding a plurality of blades mounted in the engine for rotation about a centerline thereof.

Abstract: A gas turbine engine arcuate segment has an arcuate band segment, at least one tab and an arcuate rail segment spaced axially apart from the tab and located along an axial end of the band segment. The tab and the arcuate rail segment extend radially away from the band segment and a space is located between the tab and the rail segment. A tab aperture extends entirely through the tab. The rail segment has a rail aperture entirely therethrough. A leaf seal and a spring are disposed within the space to bias the leaf seal against the aft rail. In an exemplary embodiment, a spring aperture extends entirely through the spring, a seal aperture extends entirely through the leaf seal, and a headless pin is disposed in the apertures such that the leaf seal and the spring are radially restrained by the pin.

Abstract: Circumferentially spaced arrays of support pins are disposed through access openings in an outer turbine shell and have projections received in recesses in forward and aft sections of an inner turbine shell supported from the outer shell. The projections have arcuate sides in a circumferential direction affording line contacts with the side walls of the recesses and are spaced from end faces of the recesses, enabling radial and axial expansion and contraction of the inner shell relative to the outer shell. All loads are taken up in a tangential direction by the outer shell with the support pins taking no radial loadings.

Abstract: In a method of producing an iron core by punching and laminating stator core pieces and rotor core pieces from the same metal sheet, it is an object of the invention that magnetic pole teeth of the roughly punched stator core pieces are caused to stably develop by pressing toward space sides made by previously punching the rotor core pieces so as to sufficiently secure punching margins for pressing out the front tips of the magnetic pole teeth of the internal forms of the stator core pieces and the punching can be performed without obstacles even if a gap with the rotor core pieces is small, and further ruggedness formed on a surface of the metal sheet by pressing does not cause disturbances in a flow of magnetic flux.

Abstract: Disengagement of the C-clip in turbine shroud assemblies is prevented by providing a C-clip retainer. The shroud assembly includes a shroud support having a hook and at least one shroud having a mounting flange. A C-clip overlaps the hook and the mounting flange to clamp the shroud to the shroud support, and the retainer is secured to the shroud support and located so as to engage the C-clip. The retainer engages the C-clip in such a manner so as to limit aft axial movement of the C-clip, thereby eliminating C-clip back-off.

Abstract: A split ring is disposed on an inner wall of a gas turbine casing. The split ring is composed of a plurality of split segments that are arranged on the inner wall of the casing in circumferential direction. A predetermined clearance is formed between the inner face of a split segment and the rotor blade tips. The split segments are arranged so that a predetermined circumferential clearance is formed between the split segments in order to allow the thermal expansion of the segments. A circumferential end face located upstream side of the segments with respect to the direction of the rotor blade rotation is connected to the inner face by a transition face having a surface formed as an inclined plane. The inclined plane prevents the swirl flow caused by the rotating rotor blade from impinging the upstream end face and, thereby, suppresses a temperature rise of the split segment at the upstream end face.

Abstract: A gas turbine nozzle segment has outer and inner bands and vanes therebetween. Each band includes a side wall, a cover and an impingement plate between the cover and nozzle wall defining two cavities on opposite sides of the impingement plate. Cooling steam is supplied to one cavity for flow through apertures of the impingement plate to cool the nozzle wall. The side wall of the band and inturned flange define with the nozzle wall an undercut region. Slots are formed through the inturned flange along the nozzle side wall. A plate having through-apertures extending between opposite edges thereof is disposed in each slot, the slots and plates being angled such that the cooling medium exiting the apertures in the second cavity lie close to the side wall for focusing and targeting cooling medium onto the side wall.

Abstract: A leaf seal assembly is secured to the trailing edge of a shroud segment for sealing between the shroud segment and the leading edge side wall of a nozzle outer band. The leaf seal includes a circumferentially elongated seal plate biased by a pair of spring clips disposed in a groove along the trailing edge of the shroud segment to maintain the seal plate in engagement with the flange on the leading edge side wall of the nozzle outer band. The leaf seal plate and spring clips receive pins tack-welded to the shroud segment to secure the leaf seal assembly in place.

Abstract: A stator suitable for high-pressure compressors in gas turbine engines includes two points of ventilation at different temperatures. It is recommended that the upstream section of the stator be constructed with a casing and a shroud, both of which are unbroken around a circumference and made of a material with a low degree of thermal expansion. Downstream the shroud is, however, constructed using angular sectors in a material with a higher degree of thermal expansion.

Abstract: A gas turbine nozzle segment has outer and inner bands and a vane therebetween. Each band includes a nozzle wall, a side wall, a cover and an impingement plate between the cover and the nozzle wall defining two cavities on opposite sides of the impingement plate. Cooling steam is supplied to one cavity for flow through apertures of the impingement plate to cool the nozzle wall. The side wall of the band and inturned flange define with the nozzle wall an undercut region. The impingement plate has a turned flange welded to the inturned flange. A backing plate overlies the turned flange and aligned apertures are formed through the backing plate and turned flange to direct and focus cooling flow onto the side wall of the nozzle segment.

Abstract: A stator vane frame assembly including an outer structure ring, an inner structure ring, a set of discrete vanes each connecting the outer structure ring and the inner structure ring forming an inner and outer platforms with neighboring vanes defining a set of flow paths, a set of sealing members contoured to a set of gaps between the set of vanes disposed to sealing the gap. A method for applying a sealing member to the gaps between a platform formed resulting from the coupling between vanes including affixing a set of stator vanes on an inner frame and an outer frame, and sealing a set of sealing members on the backside surface of a stator vane frame. The discrete nature of individual vane blade sealed by the sealing members renders the stator vane frame assembly to add a damping effect which increases the fatigue life of the vane, and permits the use of lower strength, lighter, and less expensive material.

Abstract: A nozzle vane segment includes outer and inner band portions with a vane extending therebetween and defining first and second cavities separated by an impingement plate for flowing cooling medium for impingement cooling of nozzle side walls. The side wall of each nozzle segment has an undercut region. The impingement plate has an inturned flange with a plurality of openings. Cooling inserts or receptacles having an open end are received in the openings and the base and side walls of the receptacles have apertures for receiving cooling medium from the first cavity and directing the cooling medium for impingement cooling of the side wall of the nozzle segment and a portion of the nozzle wall.

Abstract: A segmented turbine cooling component, such as shroud segment for the high pressure and low pressure turbine sections of a gas turbine engine, useful in providing preferential cooling to the side rails or panels of the turbine component through the combination of cooling air fed into the intermediate pressure cavity between the discourager and primary seals and then through recesses in the bottom wall of the lower discourager seal slot that impinges air on the portion of the side panel below the discourager seal. The side panels of the turbine component have an upper primary seal slot and a lower discourager seal slot with a bottom wall having a plurality of alternating lands and recesses along the length thereof and a plurality of cooling air passages having outlets exiting into the lower slot above the lands.

Abstract: A casing structure of metal construction for the rotor blade area of axial flow compressor and turbine stages includes a closed, stable outside wall, a segmented inside wall that is interrupted by multiple expansion joints, and a load-transmitting connecting structure between inside and outside wall. A multiple sub-divided hollow chamber structure with a plurality of thin wall elements standing upright on the inside wall and outside wall, at least the majority of which are directly connected together, is arranged as connecting structure. The hollow chamber structure is connected to the inside wall and/or the outside wall by brazing.

Abstract: A turbine inner shroud and a turbine assembly. The turbine assembly includes a turbine stator having a longitudinal axis and having an outer shroud block with opposing and longitudinally outward facing first and second sides having open slots. A ceramic inner shroud has longitudinally inward facing hook portions which can longitudinally and radially surround a portion of the sides of the outer shroud block. In one attachment, the hook portions are engageable with, and are positioned within, the open slots.

Abstract: Cooling is facilitated in stationary flowpath components, such as turbine nozzle segments and shrouds, by providing circumferential edges defining a non-linear curvature that matches the curvature of the flow line defined by the gas flow past the stationary flowpath component. A flexible seal member is provided to accommodate the non-linear edges.

Abstract: In gas turbine split rings, end faces having bent surfaces are formed in flanges. Adjoining split rings are coupled together with a groove therebetween to form a cylindrical split ring. Notches are formed in the flanges. These notches are sealed by inserting a seal plate into the notches of adjoining split rings. A hole for passing cooling air is drilled obliquely in the flange. Cooling air is allowed to flow out along the direction of rotation (of the turbine). This cooling air cools the outlet of the groove due to the effect of film cooling. Because of such cooling, high temperature gas is prevented from staying in this area, cooling effect is enhanced, and hence burning of the end portions can be prevented.

Abstract: A method for cooling the trailing edge 126, 146 region of a stator vane platform 48, 54 and for forming a cooling passage 198, 214 using platform slots in the side of a stator vane is disclosed. Various steps are developed which provide for effective cooling with minimal intrusion into the working medium flowpath 18. In one particular embodiment, the method includes extending a feather seal slot rearwardly with an extension 88, 96 in facing sides 116, 118, or 136, 138 of a pair of vane platforms 48, 54 and for flowing cooling fluid laterally prior to discharging the cooling air for impingement cooling of an adjacent stator vane.