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1. (WO2019046022) HYBRID CMC COMPONENT HAVING INTEGRAL AND MODULAR SHROUDS
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HYBRID CMC COMPONENT HAVING INTEGRAL AND MODULAR

SHROUDS

CROSS REFERENCE TO RELATED APPLICATIONS

[0001] This application claims priority to and the benefit of the filing date of U.S. Provisional Application No. 62/550,820, filed August 28, 2017, the entirety of each of which is hereby incorporated by reference.

FIELD

[0002] Disclosed embodiments are generally related to turbine engines and in particular to blades and vanes within the turbine engines.

BACKGROUND

[0003] Gas turbine engines typically comprise a casing or cylinder for housing a compressor section, a combustion section, and a turbine section. A supply of air is compressed in the compressor section and directed into the combustion section. The compressed air enters the combustion inlet and is mixed with fuel. The air/fuel mixture is then combusted to produce high temperature and high pressure (working) gas. This working gas then travels through the transition and into the turbine section of the turbine.

[0004] The turbine section typically comprises rows of vanes which direct the working gas to the airfoil portions of the turbine blades. The working gas travels through the turbine section, causing the turbine blades to rotate, thereby turning a rotor attached thereto. The rotor is also attached to the compressor section, thereby turning the compressor and also operatively connected to an electrical generator for producing electricity.

[0005] High efficiency of a combustion turbine is improved by heating the gas flowing through the combustion section to as high a temperature as is practical. However, the hot gas may degrade various metal turbine components, such as the combustor, transition ducts, vanes, ring segments, and turbine blades as it flows through the turbine.

[0006] High temperature resistant ceramic matrix composite (CMC) materials have been developed and increasingly utilized in gas turbine engines. Typically, CMC materials include a ceramic or a ceramic matrix material, either of which hosts a plurality of reinforcing fibers. The fibers may have predetermined orientation(s) to provide the CMC materials with additional mechanical strength. Generally, fiber reinforced ceramic matrix composites are manufactured by the infiltration of a matrix slurry (e.g., alumina, mullite, silicon-containing polymers, molten silicon, or the like) into a fiber preform, but can be manufactured in a variety of other ways.

[0007] As more metallic components are replaced by CMC steps should be taken to ensure that cooling air savings are met while ensuring the life of the component. One section of the gas turbine engine where this is pursued is the turbine section.

SUMMARY

[0008] Briefly described, aspects of the present disclosure relate to an apparatus for attaching gas turbine engine vanes.

[0009] An aspect of the present disclosure may be a gas turbine engine. The gas turbine engine may have an outer metallic shroud substructure forming part of an outer shroud; an inner metallic shroud substructure forming part of an inner shroud; a metallic core airfoil structure extending between the outer metallic shroud substructure and the inner metallic shroud substructure, wherein the metallic core airfoil structure forms part of an airfoil; a first CMC structure covering the metallic core airfoil structure and the outer metallic shroud substructure, wherein the first CMC structure forms part of the outer shroud and the airfoil; and a second CMC structure covering the inner metallic shroud substructure, wherein the second CMC structure forms part of the inner shroud.

[0010] Another aspect of the present disclosure may be a turbine section for a gas turbine engine. The turbine section may have an outer shroud; an inner shroud; an airfoil extending between the outer shroud and the inner shroud; wherein the outer shroud is formed by an outer metallic shroud substructure and a first CMC structure; wherein the inner shroud is formed by an inner metallic shroud substructure and a second CMC structure; and wherein the airfoil is formed by a metallic core airfoil structure and by the first CMC structure.

[0011] Yet another aspect of the present disclosure may be an airfoil for a gas turbine engine. The airfoil may have a metallic core airfoil structure adapted to extend between an outer metallic shroud substructure and an inner metallic shroud

substructure; a first CMC structure covering the metallic core airfoil structure, wherein the first CMC structure also forms part of an outer shroud for a gas turbine engine; and wherein the metallic core airfoil structure is adapted to be fixedly attached to the outer metallic shroud substructure and the inner metallic substructure.

BRIEF DESCRIPTION OF THE DRAWINGS

[0012] Fig. 1 shows a gas turbine engine.

[0013] Fig. 2 is an isometric view of a portion of the turbine section having a vane of modular construction.

[0014] Fig. 3 is a schematic view of an embodiment of an airfoil that has integral inner and outer shroud CMC components.

[0015] Fig. 4 is a schematic view of an embodiment of an airfoil where the CMC material is used with separate inner and outer shrouds.

[0016] Fig. 5 is a schematic view of a hybrid airfoil that uses an integral CMC component and separate shroud.

[0017] Fig. 6 is a schematic view of the plies in the integral CMC component shown in Fig. 5.

DETAILED DESCRIPTION

[0018] To facilitate an understanding of embodiments, principles, and features of the present disclosure, they are disclosed hereinafter with reference to implementation in illustrative embodiments. Embodiments of the present disclosure, however, are not limited to use in the described systems or methods and may be utilized in other systems and methods as will be understood by those skilled in the art.

[0019] The components described hereinafter as making up the various embodiments are intended to be illustrative and not restrictive. Many suitable components that would perform the same or a similar function as the components described herein are intended to be embraced within the scope of embodiments of the present disclosure.

[0020] Fig. 1 shows an overview of a gas turbine engine 100. Located within the gas turbine engine 100 is a turbine section 50. The turbine section 50 typically comprises rows of vanes which direct the working gas to the airfoil portions of the turbine blades. The working gas travels through the turbine section 50, causing the turbine blades to rotate, thereby turning a rotor attached thereto. The rotor is also attached to the compressor section, thereby turning the compressor. The rotor is also operatively connected to an electrical generator for producing electricity.

[0021] Turning to Fig. 2, an isometric view of a portion of the turbine section 50 is shown. The turbine section 50 comprises a component 25 (e.g., which may, for example, comprise a blade or vane as described above. In the case of a vane, the component 25 comprises an airfoil 10 located between an outer shroud 12 and an inner shroud 14. The airfoil 10 is secured between the outer shroud 12 and the inner shroud 14. The airfoil 10 and the accompanying outer shroud 12 and inner shroud 14 can be made from metallic materials, such as IN939 and CM247LC.

[0022] Fig. 3 shows a schematic view of an airfoil 10 between an outer shroud 12 and an inner shroud 14. Fig. 3 shows how the airfoil 10 is connected and secured between the outer shroud 12 and the inner shroud 14. In this embodiment the airfoil

10 is made of a first CMC structure 20 that surrounds a metallic core airfoil structure 9. The CMC material may be an oxide-oxide CMC material, which can provide better thermal insulation. It should be understood that in the discussion of the CMC structures disclosed herein with respect to Figs. 2-6 they may be of conventional CMC material or oxide-oxide CMC material. The oxide-oxide CMC material can provide some advantages with respect to its application in the gas turbine engine.

[0023] The metallic core airfoil structure 9 is secured between an outer metallic shroud substructure 11 and an inner metallic shroud substructure 13. The metallic components of the metallic core airfoil structure 9, outer metallic shroud substructure

11 and the inner metallic shroud substructure 13 may be made of a superalloy metals such as IN738, IN939 or CM247LC.

[0024] The metallic coil airfoil structure 9 is attached to the inner metallic shroud substructure 13 and the outer metallic shroud substructure 11 at attachment points 15. The attachment points 15 may be where the metallic core airfoil structure 9 is welded, brazed and/or bolted. The metallic core airfoil structure 9 may also be manufactured or cast integrally.

[0025] Surrounding the metallic core airfoil structure 9, the inner metallic shroud substructure 13 and the outer metallic shroud substructure 11 is the first CMC structure 20. The first CMC substructure 20 is used in the integrated embodiment for forming the airfoil 10 and the outer shroud 12 and inner shroud 14. The first CMC structure 20 in this embodiment surrounds the metallic core airfoil structure 9, and covers the hot gas path surface of the inner metallic shroud substructure 13 and the outer metallic shroud substructure 11 and is formed as an integral structure in this embodiment.

[0026] In the embodiment shown in Fig. 3, to deal with loads that occur during the operation of the gas turbine engine 100, side rails 8, or secondary attachment features (e,g, hooks or clips) are used in order to assist in the securing of the airfoil 10.

[0027] The side rails 8 are separate components that add cost and complexity to the structure of the gas turbine engine 100. The side rails 8 are located at the distal ends of the outer shroud 12 and the inner shroud 14. The side rails 8 may be constructed from metal or an alloy and attached so as to reinforce the structural integrity of the airfoil 10. The side rails 8 are used to secure the outer edges of the outer shroud 12 and the inner shroud 14 to the outer metallic shroud substructure 11 and the inner metallic shroud substructure 13. The metallic side rails 8 may be protected with a protective thermal or environmental barrier coating.

[0028] The embodiment shown in Fig. 3 prevents leakage paths from existing between the airfoil 10, the outer shroud 12, and the inner shroud 14 due to the integral nature of the CMC structure. However, there is difficulty in manufacturing this embodiment. Furthermore, the thermal expansion between the connections between the CMC material and the metal material are hard to accommodate since they may expand at different rates. Furthermore, the load from the operation of the gas turbine engine 100 has to travel through the side rails 8 and may be a point of weakness.

[0029] Fig. 4 shows an embodiment of a component 25, wherein the CMC material is applied in a modular fashion. That is to say that instead of having one integral CMC material, there are multiple applications of CMC material.

[0030] In Fig. 4 a metallic core airfoil structure 9 extends between the outer metallic shroud substructure 11 and the inner metallic shroud substructure 13. The metallic core airfoil structure 9 is attached via attachment points 15 to the outer metallic shroud substructure 11 and the inner metallic shroud substructure 13.

[0031] The embodiment shown in Fig. 4 has three different CMC structures applied to it. The first CMC structure 20 is applied to the metallic core airfoil structure 9. The second CMC structure 21 is applied to the outer metallic shroud substructure 11. The third CMC structure 22 is applied to the inner metallic shroud substructure 13.

[0032] The embodiment shown in Fig. 4 permits the airfoil 10 to be directly supported by outer shroud 12 and the inner shroud 14. The arrangement shown in Fig. 4 also avoids issues with thermal expansion since the CMC inner shroud 14 and the CMC outer shroud 12 are attached to the inner metallic shroud substructure 13 and outer metallic shroud substructure 11, respectively. The attachment of the CMC outer shroud 12 and the CMC inner shroud 14 in this manner permits them to move independently of the airfoil 10.

[0033] The embodiment shown in Fig. 4 is easy to manufacture. However, two leakage paths are present in the embodiment shown in Fig. 4. Those leakage paths are present in the area where the airfoil 10 is connected to the inner shroud 14 and the outer shroud 12.

[0034] Now turning to Fig. 5, wherein an embodiment of the component 25 is shown that improves upon the features shown and discussed above and combines positive features discussed above with respect to Figs. 3 and 4.

[0035] Metallic core airfoil structure 9 extends between the inner metallic shroud substructure 13 and the outer metallic shroud substructure 11. The metallic core airfoil shroud structure 9 is secured to the inner metal shroud substructure 13 and outer metal shroud substructure 11 via attachment points 15. The connections may be made via welding or brazing the metallic core airfoil structure 9. The metallic core airfoil structure 9 may additionally be attached via bolts. Alternatively, the metallic core airfoil structure 9 may also be manufactured or cast integrally. Connecting the metallic core airfoil structure 9 in this manner provides good structural integrity.

[0036] The metallic core airfoil structure 9 also provides radiative cooling to the first CMC structure 20. The metallic core airfoil structure 9 forms part of the airfoil 10 along with the first CMC structure 20.

[0037] The outer metallic shroud substructure 11 forms part of the outer shroud 12. The metallic core airfoil substructure 9 and the outer metallic shroud substructure 11 are covered by a first CMC structure 20. The outer metallic shroud substructure 11 and the first CMC structure 20 form the outer shroud 12.

[0038] Inner metallic shroud substructure 13 forms part of the inner shroud 14. Inner metallic shroud substructure 13 is covered by a second CMC structure 21. The second CMC structure 21 along with the inner metallic shroud substructure 13 forms part of the inner shroud 14.

[0039] The second CMC structure 21 may be attached to the inner metallic shroud substructure 13 through the use of a side rail, or via the use of bolts.

[0040] In the embodiment shown in Fig. 5 the airfoil 10 is directly supported at the inner metallic shroud substructure 13 and the outer metallic shroud substructure 11. There are no thermal expansion issues since one of the CMC shrouds can move independently with respect to the CMC airfoil 10. In the embodiment shown in Fig. 5, the airfoil 10 is fixed to the outer metallic shroud substructure 11 and is permitted to have radial expansion within the inner metallic shroud substructure 13.

[0041] This embodiment also minimizes the leakage that can occur at the shroud locations. This is because the first CMC structure 20 forms part of the outer shroud 12. Furthermore the outer shroud 12 is self-supporting and there is no need for separate side rails, such as illustrated in the embodiment in Fig. 3.

[0042] In the embodiment shown in Fig. 5 CMC fillets 23 are formed from the first CMC structure 20 and provide further support for the airfoil 10. The CMC fillets 23 are better at handling load transfer than the embodiment shown in Fig. 3.

[0043] Fig. 6 shows how the CMC fillets 23 are formed from the first CMC structure 20. The first CMC structure (20) is formed with a plurality of sets of plies 30, 31, 32. The first set of plies 30 runs along the length of the airfoil 10 and directly into the outer metallic shroud substructure 11. A second set of plies 31 extends along the length of the airfoil 10 and curves outwardly forming part of the outer shroud 12 and the CMC fillet 23. A third set of plies 32 extends along the length of the airfoil 10 and forms part of the outer shroud 12 and the CMC fillet 23. Having both the second set of plies 31 and the third set of plies 32 provides for stronger interaction in the formation of the outer shroud 12. This increases the overall strength and integrity of the airfoil 10 and the outer shroud 12.

[0044] By allowing portions of the airfoil 10 to extend into the outer shroud 12 and directly load-up with the outer metallic shroud substructure 11 most of the gas load is transferred directly into the outer metallic shroud substructure 11 through the airfoil 10 rather than the outer shroud 12. Furthermore, by decoupling the inner shroud 14 from the airfoil 10, the thermal expansion that occurs can occur freely without damage to the connections. This decreases the stress experienced by airfoil 10.

[0045] Through the integration outer shroud 12 with the airfoil 10 and the

modularity of the inner shroud 14 it is possible to combine the benefits of both the integral and modular concept while only creating one leak-path. Furthermore, the airfoil geometry supports a layup process that can be performed in a way that provides a much stronger fillet 23 than employed on the integral design.

[0046] While embodiments of the present disclosure have been disclosed in exemplary forms, it will be apparent to those skilled in the art that many modifications, additions, and deletions can be made therein without departing from the spirit and scope of the invention and its equivalents, as set forth in the following claims.