Patent Application
20190271234
This application incorporates by reference in its entirety International Application entitled “CERAMIC-MATRIX-COMPOSITE (CMC) TURBINE ENGINE BLADE WITH PIN ATTACHMENT, AND METHOD FOR MANUFACTURE”, filed concurrently herewith, and assigned application number (unknown).
The invention relates to components for combustion turbine engines. More particularly, the invention relates to pin attachment mounts for ceramic-matrix-composite (CMC) blades, which during engine operation, accommodate varying thermal expansion rates between ceramic blades and metallic components that mate the blade to a corresponding rotor disc of the engine, and methods for making such components.
Ceramic matrix composite (“CMC”) structures are being incorporated into gas turbine engine components as insulation layers and/or structural elements of such components, such as insulating sleeves, ring segments, vanes, and turbine blades. These CMCs provide better oxidation resistance, and higher temperature capability, in the range of approximately 1150 degrees Celsius (“C”) for oxide/oxide (“Ox/Ox”) based ceramic matrix composites, and up to around 1350 degrees C. for Silicon Carbide fiber-Silicon Carbide core (“SiC—SiC”) based ceramic matrix composites, whereas nickel or cobalt based superalloys are generally limited to approximately 950 to 1000 degrees Celsius under similar operating conditions within engines. While 1150 degrees C. (1350 degrees C. for SiC—SiC based CMCs) operating capability is an improvement over traditional superalloy temperature limits, mechanical strength (e.g., load bearing capacity) of CMCs is also limited by grain growth and reaction processes with the matrix and/or the environment at 1150/1350 degrees C. and higher. Therefore, some combustion-turbine engine components, such as blades and ring segments, utilized hybrid combinations of CMC and superalloy or other metals structures, which include the benefits higher temperature resistance of the CMC material and mechanical strength of the metals. However, inclusion of mating CMC and superalloy substrates in combustion turbine engines presents new and different thermal expansion mismatch challenges. During gas turbine engine operation, superalloy and CMC materials have different thermal expansion properties. Superalloy material expands more than the CMC material, which in extreme cases leads to crack formation and/or delamination in the CMC material. In some cases expansion rates between the CMC material and the superalloy or other metallic material are affected by the local ambient temperatures of the respective components.
By way of background, CMC structures typically comprise a solidified ceramic substrate, in which are embedded ceramic fibers. The embedded ceramic fibers within the ceramic substrate of the CMC improve elongation rupture resistance, fracture toughness, thermal shock resistance, and dynamic load capabilities, compared to ceramic structures that do not incorporate the embedded fibers. The CMC embedded fiber orientation also facilitates selective anisotropic alteration of the component's structural properties. CMC structures are fabricated by laying-up or otherwise orienting ceramic fibers, also known as “rovings”, into fabrics, filament windings, tows, or braids. Fiber-reinforced ceramic substrate fabrication for CMCs is comparable to what is done to form fiber-reinforced polymer structural components for aircraft wings or boat hulls. Unless the ceramic fibers are pre-impregnated with a resin containing ceramic material, they are subsequently impregnated with ceramic material by such techniques as gas deposition, melt infiltration, preceramic polymer pyrolysis, chemical reactions, sintering, or electrophoretic deposition of ceramic powders, creating a solid ceramic structure with embedded, oriented ceramic fibers.
Ox/Ox CMCs are being evaluated to replace nickel superalloys as rotating components in combustion or gas turbine engines. An important strength criterion for a rotting blade is the specific strength (strength over density) of the material. Nickel-based superalloys have high strength and high density (i.e., approximately 336 MPa and 8.1 g/cm3), while Ox/Ox CMS have lower strength and density (i.e., approximately 81 MPa and 2.7 g/cm3). Based on these physical property measurements, specific strength of superalloys is about 38% higher than that of Ox/Ox CMCs. Given the lower specific strength of CMC materials, design challenges for construction of CMC blade bodies include meeting rotating centrifugal force (“CF”) loads imparted on material forming the walls of the airfoil portions of such blade bodies, and attachment of the CMC blade bodies to turbine rotor discs in a way that can carry the centrifugal load of the entire composite blade. Another CMC blade attachment challenge is compatibility of blade mounting with fir-tree shaped attachment systems that have been utilized historically for mounting metallic blades to rotor discs of combustion turbine engines.
Typically, known metallic turbine blade bodies meet CF load requirements by use of tapered wall construction along the radial length or stand of the airfoil. Airfoil walls are thicker at the blade platform portion than at the blade tip. Such blade wall taper also facilitates easier casting of metallic blades. Fir-tree blade root to rotor disc attachment structures efficiently spread tensile CF forces imparted on the blade across faces of the mating blade root and rotor disc recess teeth, while necked portions within the blade root metal substrates have sufficient strength to absorb and distribute both tensile and compressive forces imparted therein. In contrast, CMC materials do not have strength properties needed to fabricate a fir-tree compatible blade root. A toothed and necked blade root formed of CMC material cannot meet the tensile load requirements in its neck portion, compared to a comparable superalloy material having approximately 38% higher specific strength. Moreover, total compressive loads imparted in the necked portion the CMC material, if left unchecked, are sufficiently high enough to cause delamination of the embedded ceramic fibers in that region. Traditional fir-tree blade mounting systems also beneficially offer flexibility to accommodate thermal growth mismatch between blades and their disc rotor. The attachment system decouples thermal growth of the disc and the blade root by allowing relative sliding along the fir-tree mating teeth. Given the greater thermal mismatch between CMC blade bodies and their rotor disc, it is desirable to decouple them through use of an attachment system that is as effective as a fir-tree type blade mounting system.
U.S. Pat. No. 8,231,354, issued Jul. 31, 2012, entitled “TURBINE ENGINE AIRFOIL AND PLATFORM ASSEMBLY”, is incorporated by reference in its entirety herein, and describes a composite-construction turbine blade with a metallic airfoil having a shank portion that is attached to two flanking platform portions, which are constructed of a different alloy, by clevis pins formed in the platform portions. The platform portions include integrally formed fir-tree blade root profiles, for attachment to mating recesses formed within a rotor disc of a combustion turbine engine.
As previously discussed, ring segments of combustion turbine engines also are susceptible to thermal expansion mismatch with adjoining metallic support structures, such as turbine vane casings and their support rings. Both U.S. Pat. No. 7,278,820, issued Oct. 9, 2007, entitled RING SEAL SYSTEM WITH REDUCED COOLING REQUIREMENTS”, and U.S. Pat. No. 7,950,234, issued May 31, 2011, entitled “CERAMIC MATRIX COMPOSITE TURBINE ENGINE COMPONENTS WITH UNITARY STIFFENING FRAME” (which are both incorporated by reference in their respective entireties herein) incorporate cantilevered pins, supported by the turbine casing support structure, whose free ends slidably engage apertures formed in the ring segment supporting structure. The sliding engagement facilitates relative sliding motion between the engine casing support structure and the ring segment as the ring segment material expands and contracts at different rates than its mating metallic support structure. Factors other than material composition differences that affect relative thermal expansion of components include component locations with the turbine engine. For example, some portions of a ring segment or its support structure have more cooling air exposure than portions that are directly exposed to combustion gasses.
In exemplary embodiments described herein, components for combustion turbine engines, such as ring segments, CMC blades and CMC vanes, are coupled to turbine engines by mounting systems, which comprise clevis pin-type attachments. The component is slidably coupled to one or more pins. Both ends of the respective pins project from respective sides of a through-aperture formed in the component. Both pin ends are also coupled opposed support structure within the engine, forming the clevis-like structure. The pinned component slides along the pin, as needed to accommodate thermal expansion, within limits of a gap established between the supported ends of the pin. In some embodiments, the defined gap is established by utilizing a pin, which bottoms out in surrounding support structure, so that the minimum gap is defined by length of the pin. In some blade embodiments, the pin is shorter than depth of pin-retaining apertures in the pin support structure. In such a short-pin construction, the gap formed between the ends of the pin and their respective aperture maximum depths limits compression of the portion of the component incorporating the through-aperture, once the pin ends bottom out in the receiving apertures. Such a short pin construction is useful for applying a maximally permitted compression on a CMC-material mounting shank of a turbine blade, which is captured between opposing portions of the support structure, such as a pair of opposed clevis attachment pieces. In some ring segment embodiments, the sliding pin structure accommodates thermal expansion in the axial direction of the turbine engine, which is generally aligned along axial length of the engine's rotor shaft. In some ring segment embodiments, the sliding pin structure is additionally incorporated as part of another sliding joint structure, such as a dovetail mount, that in turn accommodates thermal expansion in the circumferential direction of the engine casing, such as between adjoining ring segments forming the ring structure, or any other desired orientation, within the engine.
In some embodiments, the component is a turbine blade or vane constructed from CMC material, having a shank portion that is also formed from CMC material. In some embodiments, CMC material forming the shank portion of the blade or vane body defines a two-dimensional array of pin-receiving apertures, in order to spread tensile loads relatively equally throughout the shank. In some embodiments, the CMC blade or vane has a pair of shank portions, to split tensile loads. In some embodiments, a CMC blade is constructed of multiple plies of ceramic fabric having different axial lengths from the blade shank to the blade tip, such that the wall thickness tapers, from thicker at the shank portion to thinner at the blade tip.
In some embodiments, the mounting system is a ring segment mounting system. In such systems, the ring segment includes a first flange portion that is constrained between second and third flanges of supporting components of the turbine engine casing, in clam shell-like fashion. A first aperture, through-aperture in the first flange of the ring segment is axially aligned with respective second and third apertures formed in the second and third flanges. A mounting pin is slidably engaged within the first through-aperture. Ends of the mounting pin are retained within the second and third apertures, forming a clevis pin-type mounting structure, which accommodates thermal mismatch expansion between the ring segment and the corresponding support structures of the turbine engine casing. In some embodiments, the second flange is incorporated within a forward isolation ring that supports the ring segment. In some embodiments, the third flange is incorporated within vane blocks that support the ring segment. The ring segment mounting pins in some embodiments are retained within support structures that permit sliding motion in a second (e.g., circumferential) direction, which is complimentary to the sliding direction of the clevis pin-type mounting pin structure (e.g., an axial sliding direction).
Exemplary embodiments of the invention feature a ceramic-matrix-composite (CMC) blade for a combustion turbine engine, including a fiber-reinforced, ceramic blade body. The blade body includes an airfoil portion with a tapered blade wall defined between an outer wall surface and an inner wall surface. The outer wall surface defines respective concave pressure and convex suction sides joined by leading and trailing edges; a first end defining at least one blade shank. The at least one blade shank has a shank first portion proximate the airfoil portion, a shank tip distal the airfoil portion, and first and second shank sides between the first and tip distal portions thereof. The blade body has a second end coupled to a blade tip; with blade wall thickness in the airfoil portion between the outer and inner wall surfaces decreasing from the first end to the second end of the blade body. The blade body includes a layered structure of laid-up ceramic fibers embedded within cured ceramic material, including at least one inner layer, which delimits the inner wall surface. The inner layer has a length extending from the at least one blade shank distal tip of the first end of the blade body to the second end of the blade body, and successively shorter length extending layers, applied over previously laid-up layers. Each successively shorter layer has a length extending from the at least one blade shank distal tip of the first end toward the second end thereof, so that thickness of the composite, laid-up, successive fiber layers decreases from the first end to the second end. A two-dimensional array of rows of apertures is formed in the at least one blade shank. Each of the apertures extends through the at least one blade shank between the first and second shank sides, for insertion and receipt of corresponding load-carrying pins. Rows of apertures formed proximate the distal tip have larger diameter than rows of apertures formed closer to the blade tip, so that when any axial tensile load is applied to the at least one blade shank while corresponding load-carrying pins are inserted into their respective apertures, the applied tensile load stress is split between successive rows of apertures from proximate the at least one blade shank distal tip toward the blade tip, so that each row of apertures carries its own tensile load plus aggregate tensile load of all other rows of apertures in the blade shank portion that are closer to the blade tip.
The respective features of the exemplary embodiments that are described herein may be applied jointly or severally in any combination or sub-combination.
The exemplary embodiments are further described in the following detailed description, in conjunction with the accompanying drawings, in which:
FIG.18 is an elevational view of an alternative embodiment of a slidable ring segment mounting system, which incorporates a mounting pin retained within a pair of separate clevis pieces, which are separately dovetail mounted to the engine casing;
To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures. The figures are not drawn to scale.
Exemplary embodiments described herein are utilized to couple or otherwise affix components, including by way of example CMC blades, CMC vanes, and ring segments within combustion turbine engines. Those components are coupled to the turbine engine's casing or its rotor discs by clevis-type attachment pins. The pins are slidably engaged within corresponding through-apertures that are formed within the component, with respective ends of the pins projecting outwardly from the component. Both projecting pin ends engage structural supports within the engine, such as a turbine vane carrier-supporting ring or a rotor disc. The slidably mounted component is movable along the corresponding attachment pin, within a gap formed between the engaged ends of the pins and outer facing surfaces of the component through-aperture. The component freedom to move along the pin gap distance advantageously accommodates thermal expansion mismatch between the component and its supporting structure. By way of example, blades are pin-mounted to rotor discs, whereas vanes or ring segments are pin-mounted to turbine vane carriers or other engine casing supporting rings. In some embodiments, pin-mounted CMC blades and vanes are structurally self-supporting, relying on internally embedded fibers to provide additional strength to its fiber-reinforced, ceramic substrate.
Some embodiments of the CMC blade and vane components have a solidified, fiber-reinforced ceramic substrate, with ceramic fibers embedded therein. In accordance with method embodiments of the invention, exemplary CMC turbine blades are made by laying-up ceramic fibers into a layered, tapered structure. In some CMC blade manufacture embodiments, innermost fabric layers extend in length from a distal end of a blade shank to the blade tip. Subsequently applied, outboard fabric layers extend from the blade shank distal end toward the blade tip in progressively shorter lengths. In this way, the blade wall structure is thicker proximate the blade shank, where it is attached to a corresponding rotor disc, and tapers to a thinner wall structure proximate the blade tip. Some CMC blade embodiments incorporate a two-dimensional array of attachment pins within the blade shank, in order to distribute centrifugal loads imparted on the blade uniformly (e.g., within plus or minus five percent) throughout the blade shank. When constructing CMC blades in accordance with the methods described herein, if the ceramic fibers forming the blade body are not already pre-impregnated with ceramic material prior to their laying-up, they are subsequently infiltrated with ceramic material, forming a solidified, fiber-reinforced ceramic substrate. In some embodiments, the two-dimensional array of attachment pins within an attachment shank is incorporated into CMC vanes for turbine engines. Typically a CMC turbine vane will have shanks at both ends of the vane body, i.e., outboard of the vane airfoil.
Some embodiments of ring segments, and their mounting system, utilize a clevis pin-type mounting system on at least one axial end of the ring segment, e.g., the forward axial end that is closest to the engine combustion section. A forward axial end of the ring segment includes a first flange or a lug that projects outwardly in a generally radial direction, away from the engine's combustion path, and incorporates one or more first through-apertures. The first lug is flanked by opposed second and third flanges, in clam shell-like fashion, which project inwardly toward the combustion path. In some embodiments, the second and third flanges are incorporated within isolation rings and vane blocks, which are in turn coupled to the engine casing of the combustion turbine engine. The second and third flanges incorporate respective second and third apertures, which are coaxial with the first through-aperture of the first flange. A mounting pin is captured within the first through-aperture, with its ends in turn captured and supported within the second and third apertures of the respective second and third flanges. The first flange is slidable along the mounting pin, within the gap formed between the second and third flanges, which accommodates thermal expansion. In some ring segment embodiments, the second and third flanges incorporate sliding joints around their second and third apertures, such as pin retaining pieces, dovetail joints, or circumferential grooves, which facilitate thermal expansion in another direction within the engine (e.g., a circumferential direction about the engine casing).
Referring to
The CMC blade 50 wall taper angle θ, including the number of reinforcing ceramic fiber layers and varying horizontal, cross-sectional thickness, is selected so that sufficient material tensile strength is provided to resist the centrifugal load CF imparted on the spinning blade. Generally, blade taper angle θ for a CMC blade will be about double comparable to the taper angle used in a superalloy blade. The taper angle θ for the CMC blade 50 is five degrees or greater, whereas a comparable taper angle for a superalloy blade is two-three degrees. Use of a pair of blade shanks 70 splits the total tensile load, that must be carried by each shank, safely within the material properties of the CMC material. One blade shank is on the pressure side 62 of the blade and the other blade shank is on the suction side 64 of the blade, as shown in
As shown in
Each load-carrying pin (e.g., pin 132) respectively has an axial length LP between its first and second pin ends that is shorter than combined axial depth D1+D2=D3 of corresponding apertures (e.g., apertures 108, 124 and 110) that are formed in the blade shank 70 and the inner sides 104, 106 of the first and second clevis attachment pieces 100, 102. Because of the defined carrying-pin length and aperture depth dimensions, clearance gaps GA are formed pin ends and the corresponding bottom depths of the partial-depth apertures. See for example, the ends of load carrying pin 132 and their interface with the corresponding partial-depth apertures 108, 110. When a compressive load force FC is applied to the clevis attachment pieces 100, 102, the pin ends of the respective load-carrying pins (e.g., 132) bottom out against the corresponding partial-depth apertures (e.g., 108, 110), closing the gap GA while concurrently compressing the CMC material in the blade shank 70. Generally, a compressive load on the CMC material in the blade shank 70 enhances ability of the material to carry tensile loads, so long as the compression force is below that which causes delamination of the ceramic fibers embedded within the blade shank. Advantageously, the gap GA is chosen to be smaller than that which would enable application of a delamination compressive force on the blade shank 70.
Referring to
Some CMC blade 50 embodiments, such as those shown in
As shown in
In some embodiments, provisions are made for mismatched thermal expansion in the axial direction of the blade shank 70, between the blade 50 leading edge 66 and trailing edge 68, through use of separate, parallel clevis attachment pieces 100, 102, as shown in
In the embodiments of
In accordance with method embodiments of the invention, exemplary CMC turbine blades 50, (as well as similar construction CMC vane components) are made by laying-up the ceramic fiber layers 84, 86, 88, 90, 92 into the layered, tapered structure, to form the blade body 52. In some CMC blade manufacture embodiments, innermost fabric layers 84 of the laid-up fibers extend in length from a distal end 74 of the to-be-formed blade shank 70 of the blade body 52 to the juncture with the blade tip 82. Subsequently applied, outboard fabric layers 86, 88, 90, 92 are laid-up to extend from the blade shank distal end 74 in progressively shorter lengths toward the blade tip 82. In this way, the blade wall structure is thicker proximate the blade shank 70, where it is attached to a corresponding rotor disc 46, and tapers to a thinner wall structure proximate the blade tip 82. After laying up the ceramic fibers (e.g., the fabric layers 84, 86, 88, 90, 92) they are impregnated with ceramic slurry material, if those fibers were not previously impregnated with ceramic material prior to their lay-up. The impregnated ceramic fibers are cured, thereby solidifying the ceramic CMC material 94, which forms the blade body 52. In some embodiments, a thermal barrier coating (“TBC”) outer insulative layer 95 is applied over the solidified CMC material 94 of the blade body 52. In some embodiments, the TBC layer is applied to the blade shanks 70. In exemplary embodiments, the TBC layer 95 is thermally sprayed, vapor deposited, or solution/suspension plasma sprayed over the outer wall surface 58 of the blade body 52.
In some embodiments, a clevis pin-like attachment system is incorporated into CMC vanes for turbine engines. The CMC vane mounting system is constructed, in the alternative, with or without the two-dimensional array of attachment pins within an attachment shank that was previously described for application with the CMC turbine blade 50 embodiments. Typically a CMC turbine vane will have shanks at both ends of the vane body, outboard of the vane airfoil. In this way, a CMC structure vane body is mated to metallic support structure of a corresponding turbine vane cavity.
Ring segment mounting system embodiments, which incorporate clevis pin-type mounting structures, are shown in
An array of first apertures 218 pass entirely through forward 214 and aft 216 axial sides of the first flange 212; i.e., they are through-apertures. The respective first apertures 218 slidably receive corresponding ones of a plurality of mounting pins 220. Each of the mounting pins 220 has a forward 224 and an aft 226 distal end or tip, which tips respectively extend or project outwardly from both forward 214 and aft 216 axial sides of the first flange 212. The ring segment mounting system 200 includes a ring-segment support ring 230 that is coupled to the combustion-turbine engine casing 22. The ring-segment support ring 230 has a forward isolation ring 232, an aft isolation ring 233, and a plurality of circumferentially aligned vane blocks 234. The forward isolation ring 232 forms a second flange, and the vane blocks 234 collectively form a third flange. The second flange i.e., the forward isolation ring 232 and the third flange (i.e., the vane blocks 234) are circumferentially aligned with the first flange 212 of the ring segment 202. Both the second 232 and the third 234 flanges radially project inwardly toward the outer circumferential surface 206 of the ring segment 202, and respectively are rigidly oriented in spaced, axially opposed, circumferentially flanking relationship with corresponding forward 214 and aft 216 axial sides of the first flange 212, in clam shell-like fashion, establishing axial spacing gaps “g” there between.
The second 232 and third 234 flanges of the ring-segment support ring 230 having respective second 236 and third 238 partial-depth apertures, which are coaxially aligned in opposing corresponding relationship with the first through-apertures 218 of the first flange 212, and which respectively receive corresponding forward 224 and aft 226 ends of the mounting pins 220. The second 236 and third 238 partial-depth apertures constrain radial movement of the mounting pins 220, as is done in a clevis pin-type mounting system. Radial constraint of the respective mounting pins 220 in turn radially constrains each pin's corresponding ring segment 202 within the combustion-turbine engine casing 22, by way of the slidable engagement with a corresponding first through-aperture 218. While the slidable engagement between the mounting pin 220 and the first flange 212 constrains radial movement of the ring segment 202, it allows axial movement of the ring segment 202 within the engine casing 22, by relative sliding motion of the respective first apertures 218 along their respective mounting pins 220. The axial movement accommodates thermal mismatch relative expansion between the ring segment 202 and the engine casing 22 in the general direction of the engine 20 from the compressor section 24 to the turbine section 28.
In the ring segment mounting system embodiment of
The ring segment mounting system embodiment of
The first pin-retaining piece 240 is captured within the second flange formed within the forward isolation ring 232, and interposed between aft side 216 of the first flange 212 of the ring segment 202 and said forward isolation ring. The first pin-retaining piece 240 has an outboard side 244 that defines a second portion 246 of the second dovetail mating joint 245, which is in slidable, mating engagement with a circumferential groove 248 formed within the second flange formed within the forward isolation ring 232. The circumferential groove 248 forms a first portion of the second dovetail mating joint 245. An inboard side 250 defines the second, partial-depth apertures 236, which receive the aft-projecting pin end 226 of the corresponding mounting pin 220. Circumferential sides 264 defined by each of the respective first pin-retaining pieces 240 are laterally/circumferentially spaced from neighboring first pin-retaining pieces by the gap “g′“.
Similarly, the second pin-retaining piece 242 is captured within the third flange formed within the vane blocks 234, and interposed between the forward side 214 of the first flange 212 of the ring segment 202 and said vane blocks. The second pin-retaining piece 242 has an outboard side 252 that defines a second portion 254 of the third dovetail mating joint 258, which is in slidable, mating engagement with a circumferential groove 256 formed within the third flange formed within the vane blocks 234. The circumferential groove 256 forms a first portion of the third dovetail mating joint 258. An inboard side 260 defines the third, partial-depth apertures 238, which receive the forward-projecting pin end 224 of the corresponding mounting pin 220. Circumferential sides 262 defined by each of the respective second pin-retaining pieces 242 are laterally/circumferentially spaced from neighboring second pin-retaining pieces by the gap “g′“.
Alternative embodiments of clevis pin-type, mounting systems for ring segments are shown in
The ring-segment mounting system 310 of
The ring-segment mounting system 340 of
The ring-segment mounting system 360 of
Although various embodiments that incorporate the invention have been shown and described in detail herein, others can readily devise many other varied embodiments that still incorporate the claimed invention. The invention is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways.
By way of non-limiting examples, while cross-sectional profiles of mounting and retaining pins and their corresponding receiving apertures, in different types of clevis pin-like mounting systems shown in the figures, are circular, other cross sectional profiles can be substituted for the circular profiles. A threaded fastener, such as a cap screw, can be substituted for one or more of the dowel-like, cylindrical profile, mounting and retaining pins. Similarly, non-oval, elongated apertures can be substituted for oval profile apertures. Flanges of blade or vane shanks, as well as flanges on casing rings, ring segments, and ring segment supports can have continuous circumferential profiles, as shown, or such flanges can comprise sub arrays of segmented or split sub-flanges. While some blade and vane embodiments are described as having CMC material construction, the clevis pin-type mounting systems shown and claimed herein can be utilized with metallic vanes or blade bodies.
In some embodiments, if a threaded fastener is utilized for a mounting or retaining pin, one or more of the pin-receiving apertures in the clevis attachment pieces or the second or third flanges of a ring segment support can be constructed with corresponding female threads, for engagement of the fastener's male threads. Concomitantly, the blade shank or ring segment flange aperture slides over the outer diameter, thread profile of the male threads of the fastener.
In some embodiments, the outer profile of the clevis-attachment pieces forms the outer profile of a blade platform. In some other embodiments, the blade platform defined by the clevis-attachment pieces is subsequently coated with a thermal barrier coating.
In addition, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of “including,” “comprising,” or “having” and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms “mounted”, “connected”, “supported”, and “coupled”, and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, “connected” and “coupled” are not restricted to physical, mechanical, or electrical connections or couplings.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US2016/058369 | 10/24/2016 | WO | 00 |
