Abstract: A device for axially holding a ring spacer sector of a high-pressure turbine of a turbomachine, said spacer sector comprising an upstream radial wall provided with an outer upstream tongue for being engaged in a corresponding upstream groove of a casing of the turbine, and with an inner upstream tongue for being engaged in a corresponding upstream groove of a ring sector, said device comprising a piece of longitudinally-extending sheet metal that constitutes a heat shield disposed upstream from the upstream radial wall of the spacer sector against an inner surface of the casing, the piece of longitudinally-extending sheet metal bearing axially against the upstream radial wall of the spacer sector and being secured to the casing so as to hold the spacer sector axially.

Abstract: A turbine blade gap control system configured to move a vane carrier and attached ring segments of a turbine engine axially relative to a turbine blade assembly to reduce the gaps between the tips of the turbine blades and the ring segments to increase the efficiency of the turbine engine once operating in a steady state condition. The turbine blade assembly may be formed from a plurality rows of turbine blades extending radially from a rotor, wherein each row may be formed from a plurality of turbine blades and wherein the turbine blades may have tips positioned at an acute angle relative to a rotational axis of the turbine blade assembly. A vane carrier may be positioned concentric with the rotor and positioned radially outward from the turbine blades, and a plurality of ring segments may be attached to the vane carrier.

Abstract: The interturbine duct (ITD) directs hot combustion gases from a high pressure turbine stage outlet to a low pressure turbine stage inlet in a gas turbine engine. The ITD comprises an outer wall and a heat resistant non-metallic resilient annular seal provided around an end of the outer wall that is adjacent to the high pressure turbine stage outlet.

Abstract: One embodiment of a transition-to-turbine seal (300) comprises a first, flattened section (302) adapted to be received in a peripheral axial slot (320) of a transition (325), and a second, generally C-shaped section (301). The generally C-shaped section (301) comprises a flattened portion (305) near the first, flattened section (302), and a curved portion (306) extending to a free edge (307). A fiber metal strip component (309) may be attached to the flattened portion (305) to define a first engagement surface adapted to engage an upstream side (336) of an outer vane seal rail (337), and a second engagement surface 308, adjacent the free edge (307), provides an opposed wear surface adapted to engage a downstream side (338) of the outer vane seal rail (337). System embodiments also are described, in which such transition-to-turbine seal (300) is isolated from a hot gas path (350) by provision of a plurality of cooling apertures (327) in the transition (325).

Abstract: A vane assembly for a gas turbine engine is disclosed having lower thermally induced stresses resulting in improved component durability. The stresses in the vane assembly airfoils are lowered by increasing the flexibility of the vane platform and reducing its resistance to thermal deflection. This is accomplished by placing an opening along the innermost vane assembly rail that reduces the effective stiffness of the platform, thereby lowering the operating stresses in the airfoils of the vane assembly. A removable seal is then placed in the opening in order to prevent undesired leakages, while maintaining the benefit of the increased platform flexibility.

Abstract: A sealing device for sealing between two static portions of a turbomachine includes a brush gasket formed by a plurality of resilient bristles which are, at least in part, held in a housing. Each bristle has a main portion and a secondary portion interconnected by a bent portion. The secondary portion has a free end ensuring leaktight contact with a surface that is to be sealed of the other static portion. The sealing device further includes a flexible washer of bent section which has a free end bearing both against the free ends of the secondary portions of the bristles of the brush gasket, and against the surface that is to be sealed so that the free ends of the secondary portions of the bristles of the brush gasket are permanently in contact with the surface that is to be sealed.

Abstract: A vane assembly for a gas turbine engine is disclosed having lower thermally induced stresses resulting in improved component durability. The stresses in the vane assembly airfoils are lowered by increasing the flexibility of the vane platform and reducing their resistance to thermal deflection. This is accomplished by placing an opening along the vane assembly rail that reduces the effective stiffness of the platform, thereby lowering the operating stresses in the airfoils of the vane assembly. A removable seal is then placed in the opening in order to prevent undesired leakages, while maintaining the benefit of the increased platform flexibility.

Abstract: A stator blade segment for a compressor includes an inner ring segment and an outer ring segment and a plurality of stator blades extending radially between the inner and outer ring segments, the stator blade secured to the outer ring at a shank portion of the blade and loosely held in a slot in the inner ring segment at a tip portion of the blade, the slot formed to substantially match a cross-sectional profile of the tip portion of the blade but sized to create a clearance between the tip portion and the slot.

Abstract: A sealing device for use between a gas turbine combustor transition duct aft frame and a turbine inlet having improved durability, reduced wear on the mating turbine vane, and reduced manufacturing costs, is disclosed. The sealing device has a circumferential length, an axial width, and a radial height and contains a plurality of channels extending axially along the seal inner surface for passing a controlled amount of cooling air to a turbine inlet. The sealing device is formed of abradable honeycomb having a plurality of honeycomb cells with the honeycomb cells oriented to ensure maximum control against cooling air leakage while also providing maximum flexibility during assembly. The sealing device is captured between the transition duct aft frame, bulkhead, and turbine vane platform, thereby allowing easy replacement of the seal without requiring major disassembly of the transition duct aft frame section.

Abstract: A seal arrangement for reducing the seal gaps within a rotary flow machine, preferably an axial turbomachine, has rotor blades and guide vanes, which are respectively arranged in at least one rotor blade row and guide vane row and have respective blade/vane roots (2,3) which protrude into fastening contours within the rotor blade and guide vane rows. A sealing element (4) of plastically deformable material is provided between at least two adjacent blade/vane roots (2,3) along a rotor blade row or guide vane row or between a blade/vane root (2,3) of a rotor blade or guide vane and a rotary flow machine component directly adjoining the blade/vane root.

Abstract: A mechanical arrangement for protection against catastrophic nozzle failures includes a turbine nozzle segment including an inner platform rail, a turbine nozzle inner support ring in part in axial registration with the rail on one side, an inner retainer segment secured to the inner support ring and in part in axially spaced registration relative to the rail on an axial side of the rail opposite from the support ring, a first inclined conical surface on the inner retainer segment, and a second inclined conical surface on the inner platform rail of the turbine nozzle, the second inclined conical surface opposing the first inclined conical surface, so as to bind the inner platform rail to the turbine nozzle between the inner retainer segment and the inner support ring, resulting in a wedge lock that prevents the inner platform from being lost downstream into rotating hardware of the turbine.

Abstract: It is disclosed a seal assembly (5) which comprises a layered structure. It consists of a first layer (9, 12) of a base material, a second layer (10) of thermal insulating material on top of the base material and a third layer (11, 12) of base material or an oxide resistant material on top of the second layer (10) of thermal insulation. The inventive seal assembly (5) described herein can be arranged between combustor liner segments (2).

Abstract: Primary objective of the present invention is to provide a seal device wherein the seal surfaces of the device are always brought into close contact against the side walls of the mounting slot regardless of fluid conditions. The seal device has two seal portions located to both sides of a bight shaped elastic portion which has a “U”-shaped cross section wherein the respective seal portions have a first side portion and a second side portion. The second side portions of the seal portions abut against each other in a symmetric manner in order to form a “W”-shaped seal body element. The end portions of the both second side portions of the seal body element are joined and seal surfaces are defined on the outer side surfaces of the both first side portions.

Abstract: A sealing arrangement that can be employed to maintain a seal in any suitable pressure vessel including but not limited to a gas turbine engine. The sealing arrangement is forg sealing a leakage gap between at least two relatively moveable parts that are adjacent to each other in a flow path between a region of high fluid pressure and a region of low fluid pressure. The sealing arrangement comprises at least two resilient sealing strips, each strip having a portion formed along at least part of its width to locate in a groove provided along adjacent faces of the relatively moveable parts. The remaining portion of each of the sealing strips has a substantially flat surface which, in operation, abuts a corresponding flat surface of an adjacent sealing strip, thereby forming a seal.

Abstract: Primary objective of the present invention is to effect a seal against assembly clearances formed between components which admit flows of high temperature fluids or are subjected to vibrations.

Abstract: A blower unit of an air conditioning unit including a scroll casing including a first engaging portion, a fan casing including a first engaging portion, and a controller for controlling an electric motor to move at least one of the scroll casing and the fan casing such that the movement of at least one of the scroll casing and the fan casing stops when the first engaging portion of the scroll casing abuts the first engaging portion of the fan casing. The precision of the position of the scroll casing with respect to the fan casing, the precision of the temperature of air discharged into the compartment, and/or the volume of the blower unit may be improved.

Abstract: The invention concerns a casing aggregate for the turbine of an exhaust turbocharger. The invention is identified by the following characteristics: a spiral casing adapted for surround the running wheel of the turbine; a tongue-like wall part (tongue) in the inside of the spiral casing; an inlet connection; an outlet connection; a flange to connect to a bearing casing; wherein the casing aggregate is manufactured of thin-walled precision casting; wherein the casing aggregate is made of at least two parts, so that at least one separation joint is present; and the separation joint is arranged as follows: it runs in an axially perpendicular level; it runs along the apex line of the spiral casing; it extends over an arc of a circle of approximately 270 degrees; it lies outside of the area of the tongue.

Abstract: There is disclosed a gas turbine shroud structure comprising: a shroud support component 12 attached to an inner surface of a gas turbine casing 3; a shroud segment 14 divided in a peripheral direction and supported by the inner surface of the shroud support component; and a heat-resistant restricting spring 18 held between the shroud segment and shroud support component to urge the shroud segment inwards in a radial direction. The shroud segment 14 is formed of a ceramic composite material, and includes a coating layer 15 having heat insulation and impact absorption effect on an inner surface of the segment.

Abstract: The clearances between an array of turbine blades and its surrounding blade track may be controlled by an expansion control material system supporting the blade tracks. The blade track support hoop is placed in tension by the expansion control material placed therein and the expansion control material is placed in compression.

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 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: 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 suspension of an annular secondary structure on a primary structure, in the form of a spoke-type centering device, has at least three sliding guides distributed uniformly over the structure circumference. Each sliding guide allows at least a linear relative movement of the structures transversely to their axial direction. The linear direction of movement of each sliding guide runs, in relation to the structure-related radial direction at the location of the sliding guide, at an angle having radial and tangential direction components.

Abstract: A biased, wear resistant seal assembly for sealing gas flow paths in a turbine system includes a planar shim having transverse legs at each end, a flattened leaf spring to urge the shim into sealing contact with the sealing surfaces of the turbine parts, and a shim protection material to avoid wear to the shim on the contact side. The shim protection material can include one or more metal fiber cloth layers secured to the shim, for example, by seam welding to a pair of rails or to the legs of the shim.

Abstract: The support spacer sector (14) minimizes functional clearances (J) between the end of the blades (3) and the ring (12) of the high-pressure turbine and the assembly clearances of the support spacer sectors (14) on the casing of the high-pressure turbine (1). Each support spacer sector (14) has a tab (20) on the upstream side one end (21) of which is supported on the inside wall (1I) of the casing of the high-pressure turbine (1) thus forming an intimate contact between the attachment parts of this support spacer sector (14) with the corresponding parts of the casing of the high-pressure turbine (1). This invention applies to turbo-machines fitted on an aircraft.

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 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: Upper and lower halves of the inner barrel of a compressor for a turbine are bolted to one another along a horizontal midline joint with metal-to-metal surface contact. Elongated grooves are provided in the engaging surfaces, with the upper groove having a depth greater than the depth of the lower groove. Elongated seals are provided in the grooves, including an elongated shim with a metallic woven covering on opposite sides, and margins having sealing edges. The depth of the seal is equal to or less than the depth of the upper groove and an adherent material maintains the seal in the upper groove during installation. In operation, the adherent material dissolves at temperatures above ambient and the seal drops by gravity to bear against the walls of the grooves in response to high pressure air leaking through any gap at the horizontal midline joint.

Abstract: Spline seals are formed at the joints between packing casing shells in a steam turbine having high pressure and intermediate pressure turbine sections sealing between the high pressure and intermediate pressure regions on opposite axial sides of the packing casing. The spline seals extend in registering grooves from adjacent the seal segments of the rotor radially outwardly to adjacent axial load surfaces cooperable between the outer shell and the packing casing. The spline seals minimize or eliminate steam leakage paths through the horizontal joint of the packing casing shells.

Abstract: A transition duct with a thermally free aft frame for use in a gas turbine engine is disclosed. The transition duct includes an aft frame that is thermally free through the use of a plurality of retention lugs, bushings, and bulkhead assemblies. The aft frame is allowed to adjust from thermal changes as a result of relative sizing between the bushings and retention lugs of the aft frame. An additional feature of this invention is the use of radially extending ribs along the sidewalls of the aft frame, to form an interlocking sealing means with adjacent transition ducts to reduce the amount compressor air leakage into the turbine inlet.

Abstract: In the gas turbine split ring, on an outer peripheral surface 1b between two cabin attachment flanges, a circumferential rib which extends in the circumferential direction and an axial rib which extends in the axial direction and has a height taller than that of the circumferential rib are, respectively, formed in plural lines, so that it is possible to suppress heat deformation in the axial direction which largely contributes to reduction of the tip clearance compared to head deformation in the circumferential direction more efficiently.

Abstract: A segmented vane support structure for use in a gas turbine engine having an engine casing, includes a single piece inner ring and separated front outer and rear outer rings. The vane segments circumferentially abut to form a stator ring which is clamped at the respective opposed outer ends by the front outer and rear outer rings therebetween onto the inner ring. The front outer and rear outer rings are axially restrained to the engine casing by a retaining ring which is fitted in an inner annular groove of the engine casing. Each vane segment has a lug member at its outer diameter which radially and slidably engages in a slot of the engine casing to provide angular positioning of the vane segments within the engine casing and to transmit circumferential vane load into the engine casing. This support structure arrangement transmits circumferential loading to the engine casing and isolates radial loading from the engine casing caused by thermal growth changes of the vane segments.

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: 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: An integral aft seal for sealing the interface between a gas turbine transition piece and first stage nozzle exhibits sufficient variable thickness to adjust for relative component movement. The seal comprises a contiguous material having opposing planar faces and stiffeners attached at each of the faces. The contiguous material has corrugations to provide variable thickness. The stiffeners are attached to the planar faces of the contiguous material and generally have the same planar shape as the seal. Flexibility in the thickness dimension due to the corrugations biases the stiffeners away from one another and into contact with the turbine components. The aft seal is affixed to the transition piece by an attachment structure that eliminates relative movement between the seal and the transition piece to thereby eliminate wear.

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 device and a method are described for the controlled setting of the gap between the shroud arrangement and rotor arrangement of a turbomachine, the shroud arrangement of which has a shroud carrier and at least one shroud segment connectable to the shroud carrier via at least two holding webs and the rotor arrangement of which provides at least one moving blade row rotatable about an axis of rotation and having a plurality of individual moving blades, the movable blade ends of which are located opposite the shroud segment and with the latter enclose a radial gap.

Abstract: During a standard fired shutdown of a turbine, a loaded rabbet joint between the fourth stage wheel and the aft shaft of the machine can become unloaded causing a gap to occur due to a thermal mismatch at the rabbet joint with the bearing blower turned on. An open or unloaded rabbet could cause the parts to move relative to each other and therefore cause the rotor to lose balance. If the bearing blower is turned off during a shutdown, the forward air/oil seal temperature may exceed maximum design practice criterion due to “soak-back.” An air/oil seal temperature above the established maximum design limits could cause a bearing fire to occur, with catastrophic consequences to the machine. By controlling the bearing blower according to an optimized blower profile, the rabbet load can be maintained, and the air/oil seal temperature can be maintained below the established limits. A blower profile is determined according to a thermodynamic model of the system.

Abstract: A tip clearance control device for a gas turbine engine having a shroud surrounding a stage of rotor blades. The tip clearance control device comprises a one-piece split ring having opposed overlapping end portions. The split ring is directly supported onto the inner surface of the shroud and is adapted to automatically adjust for thermal growth of the shroud during engine operation.

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 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 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: A seal assembly that seals a gap formed by a groove comprises a seal body, a biasing element, and a connection that connects the seal body to the biasing element to form the seal assembly. The seal assembly further comprises a concave-shaped center section and convex-shaped contact portions at each end of the seal body. The biasing element is formed from an elastic material and comprises a convex-shaped center section and concave-shaped biasing zones that are opposed to the convex-shaped contact portions. The biasing element is adapted to be compressed to change a width of the seal assembly from a first width to a second width that is smaller than the first width. In the compressed state, the seal assembly can be disposed in the groove. After release of the compressing force, the seal assembly expands. The contact portions will move toward a surface of the groove and the biasing zones will move into contact with another surface of the groove.

Abstract: The invention provides a novel sealing device and plenum assembly particularly adapted for forming a plenum inwardly of a circumferential stator blade array in a gas turbine engine. A Tangential On Board Injection plenum (T.O.B.I.) is defined by assembly of: a forward plate; a rearward plate; and a segmented inner shroud disposed between the forward and rearward plates. Releasable elongate fasteners are disposed in joints between a series of inner shroud segments with sealing sheaths laterally surrounding the fasteners. While clamping the shroud segments between the forward and rearward plates, the fasteners and sheaths simultaneously seal the joints and secure the inner shroud segments with forward and rearward plates in a rigid structural assembly.

Abstract: A ceramic turbine inlet vane(s) is resiliently mounted to stator portions (10) and (15) of a gas turbine engine by outer and inner resilient mounts (20) and (25). The resilient mounts each include springs, which accommodate varying rates of radial and axial thermal expansion between the vane and adjacent metallic stator structure.

Abstract: A turbine spline seal includes an elongated turbine seal member. The seal member has an elongated, imperforate, and manually-flexible first portion and a manually-rigid second portion lengthwise adjoining the first portion. The turbine assembly includes the turbine spline seal and also includes first and second turbine members. The first and second turbine members are spaced apart to define a fluid-path leakage gap therebetween, and the seal member is placed in the gap. Each lengthwise edge of the seal member engages a respective one of two opposing surface grooves of the first and second turbine members.

Abstract: A gas turbine engine blade tip clearance control system and method is described. An annular housing is formed about an engine casing to which an annular shroud segment assembly is secured and closely spaced about blade tips of a stage of blades. The annular housing forms an air passage means communicating with the casing for directing a cooling air stream to the casing. A thermally operable passive ring valve is formed by two overlapped metal ring segments having a dissimilar coefficient of thermal expansion selected whereby to produce a radial gap between the ring segments when the valve temperature reaches a predetermined value. The radial gap admits a cooling air flow into the housing for cooling the casing and its associated shroud segment assembly to control radial growth and thereby prevent blade tip pinching.

Abstract: A turbine assembly having a turbine stator, a ceramic inner shroud, and a first spring. The stator has a longitudinal axis and an outer shroud block with opposing and longitudinally outward facing first and second sides. The first side has a longitudinally outward projecting first ledge and has a first side portion located radially outward of the first ledge. The ceramic inner shroud has a first hook portion longitudinally and radially surrounding the first ledge. The first spring is attached to one of the first side portion and the first hook portion and unattachedly and resiliently contacts the other of the first side portion and the first hook portion.