WO2020046359A1

WO2020046359A1 - Radiatively cooled hybrid components

Info

Publication number: WO2020046359A1
Application number: PCT/US2018/049015
Authority: WO
Inventors: Jan H. Marsh; Evan C. LANDRUM
Original assignee: Siemens Aktiengesellschaft; Siemens Energy, Inc.
Priority date: 2018-08-31
Filing date: 2018-08-31
Publication date: 2020-03-05

Abstract

There is provided a hybrid component including a metal core (102); a shell (104) comprising a ceramic matrix composite material about the metal core (104); and one or more cavities (106) defined between the metal core (102) and the shell (104). In addition, the component (100) includes an extended surface (110) disposed on the metal core (102). The extended surface (110) is effective to increase an amount of radiative heat transfer through the one or more cavities (106) from the shell (104) to the metal core (102) (relative to a like component without the extended surface (110).

Description

RADIATIVELY COOLED HYBRID COMPONENTS FIELD OF THE INVENTION

The present invention relates to components for use in high temperature environments, such as in gas turbines, and more particularly hybrid components which promote radiative cooling from a ceramic matrix composite shell to a metal core of the hybrid component.

BACKGROUND OF THE INVENTION

Gas turbines 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 a high temperature and high pressure (working) gas. This working gas is then ejected past the combustor transition and injected into the turbine section of the turbine.

The turbine section 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 the rotor. The rotor is also attached to the compressor section, thereby turning the compressor and also an electrical generator for producing electricity. A high efficiency of a combustion turbine is achieved by heating the gas flowing through the combustion section to as high a temperature as is practical. The hot gas, however, may degrade the various metal turbine components, such as the combustor, transition ducts, vanes, ring segments and turbine blades that it passes when flowing through the turbine.

For this reason, strategies have been developed to protect such components from extreme temperatures including the development and selection of high

temperature materials adapted to withstand these extreme temperatures, and cooling strategies to keep the components adequately cooled during operation. For one, ceramic matrix composite (CMC) materials have been developed with high temperature resistance. CMC materials include a ceramic matrix reinforced with ceramic fibers. Despite the advancement in materials, CMC turbine components typically still utilize active cooling regimens to maintain the CMC body temperature below a predetermined temperature limit, such as 1200° C in many instances. Above 1200° C, the grain size of the CMC’s fibers typically coarsen and lose their strength. Further, since CMC is a low thermal conductivity material, a thermal gradient across the CMC body results in a strain gradient across its wall thickness, which can result in damage during thermal cycling.

The aforementioned active cooling regimens include distributing a pressurized gas through radially extending cavities or through cooling channels (e.g., serpentine cooling channels) formed in the component (e.g., airfoil) body. In some instances, the pressurized gas is additionally fed through cooling holes that lead to an exterior of the component and provide a thin protective gas layer over the component surface. While there are some cooling benefits to such active cooling regimens, several drawbacks are associated with the use of pressurized cooling air as the primary cooling strategy. For one, the use thereof adds significant material and operational costs to the associated gas turbine engine. In addition, the use of pressurized air as such inevitably results in air leakage through any gaps in the component, thereby reducing cooling efficiency, engine efficiency, and increasing capital costs. Further, the use of pressurized air may result in delamination of the individual fiber layers (e.g., plies) of the CMC material or splitting of the airfoil at its trailing edge due to the weak interlaminar strength of the CMC material and the significant pressures involved. Even minimal delamination may affect the temperature resistance and lifetime of the component, as well as reduce cooling effectiveness due to leakage. Accordingly, there is a need for novel CMC airfoil cooling strategies that eliminate the drawbacks of prior solutions.

SUMMARY

In accordance with an aspect, the present inventors have developed a hybrid component which utilizes an outer ceramic matrix composite (CMC) shell for its high temperature resistance (relative to superalloys or other high temperature materials) at the interface of the component and with hot gas. The component further includes a metal core which is spaced from the CMC shell by defined cavities within an interior the component body. The metal core comprises an extended surface on an exterior of the metal core which increases the radiative cooling ability of the component (relative to the same component without the extended surface). Importantly, the extended surface increases the surface area about the metal core and increases a heat transfer coefficient of the metal core, thereby promoting radiative heat transfer from the CMC portion to the metal core.

In certain embodiments, a coating is provided on or about the CMC shell, the metal core, and/or the extended surface to further enhance the radiative cooling of the component. Increased radiative cooling provides benefits for the hybrid CMC/metal component. For example component lifetime may be extended via reduction of the temperatures to which the hybrid component is subjected over time. Alternatively, the increased thermal capability of the radiatively cooled hybrid component may allow for use of the hybrid component in higher temperature environments, thereby leading to increased turbine efficiency when utilized within a turbine engine.

In accordance with a particular aspect, there is disclosed a hybrid component comprising a metal core and a shell comprising a ceramic matrix composite material about the metal core, the shell and the core defining one or more cavities therebetween. The hybrid component further comprises an extended surface on the metal core, the extended surface effective to increase an amount of radiative heat transfer through the one or more cavities from the ceramic matrix composite (CMC) shell to the metal core (relative to a like component without the extended surface).

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic of a gas turbine engine which incorporates a hybrid component in accordance with an aspect of the present invention.

FIG. 2 illustrates a hybrid component in accordance with an aspect of the present invention.

FIG. 3 illustrates a cross section of the airfoil of FIG. 2 illustrating a hybrid component comprising a ceramic matrix composite shell and a metal core with an extended surface in accordance with an aspect of the present invention.

FIG. 4 illustrates an example of an extended surface in accordance with an aspect of the present invention. FIG. 5 illustrates a hybrid component having a coating that increases radiative cooling on the shell, the metal core, and/or the extended surface in accordance with an aspect of the present invention.

DETAILED DESCRIPTION

Referring now to the figures, FIG. 1 illustrates a gas turbine engine 2 which includes one or more hybrid components formed from a ceramic matrix composite material and a metal material as described herein. The gas turbine engine 2 includes a compressor section 4, a combustor section 6, and a turbine section 8. In the turbine section 8, there are alternating rows of stationary airfoils 18 (commonly referred to as "vanes") and rotating airfoils 16 (commonly referred to as "blades"). Each row of blades 16 is formed by a circular array of airfoils connected to an attachment disc 14 disposed on a rotor 10 having a rotor axis 12. The blades 16 extend radially outward from the rotor 10 and terminate in blades tips. The vanes 18 extend radially inward from an inner surface of vane carriers 22, 24 which are attached to an outer casing 26 of the engine 2. Between the rows of vanes 18 a ring seal 20 is attached to the inner surface of the vane carrier 22. The ring seal 20 is a stationary component that acts as a hot gas path guide between the rows of vanes 18 at the locations of the rotating blades 16. The ring seal 20 is commonly formed by a plurality of ring segments that are attached either directly to the vane carriers 22, 24 or indirectly such as by attachment to metal isolation rings (not shown) attached to the vane carriers 22, 24. During engine operation, high- temperature/high-velocity gases 28 flow primarily axially with respect to the rotor axis 12 through the rows of vanes 18 and blades 16 in the turbine section 8.

Referring now to FIG. 2, there is shown an exemplary hybrid component 100, which may comprise a gas turbine component illustrated in FIG. 1. In the embodiment of FIG. 2, the component 100 comprises an elongated airfoil 32, the airfoil 32 having a body 34 which extends in a radial direction (R). The body 34 is defined between a leading edge 54 and a trailing edge 56, and further includes an outer wall 36. The outer wall 36 may have a generally concave-shaped portion 38 defining a pressure side 40 and a generally convex shaped portion (opposite side) 42 defining the suction side 44.

In an embodiment, the comprises a vane 18 (FIG. 1 ) comprising the airfoil 32, wherein the airfoil 32 is disposed between an outer platform 46 at a first end 48 of the vane 18 and an inner platform 50 at a second end 52 of the vane 18. Although a vane 18 is shown, it is appreciated that the component 100 is not limited to a vane 18, but may include any component for high temperature use, such as another component (e.g., blades 16) of a turbine engine 2 shown in FIG. 1.

Referring to FIG. 3, there is shown a cross-section taken at line A-A of the component 100 of FIG. 2. As shown, the hybrid component 100 comprises a metal core 102 and a shell 104 comprising a ceramic matrix composite (CMC) material about the metal core 102. One or more cavities 106 are defined between the metal core 102 and the shell 104. An extended surfacel 10 is disposed on the metal core 102 and is effective to increase an amount of radiative heat transfer from the shell 104 to the metal core 102 (relative to a like component without the extended surface). To accomplish the heat transfer, energy may be radiated through the one or more cavities 106 from the CMC shell 104 to the metal core 102. In certain embodiments, the CMC shell 104 comprises a plurality of ribs 112 as shown in FIG. 3 that span across the body from the pressure side 40 to the suction side 44 to define multiple cavities 106 in the component 100. It is appreciated, however, that the present invention is not so limited and that no ribs 112 or a greater number of such ribs 1 12 may be provided.

The CMC shell 104 comprises any suitable ceramic matrix material which hosts a plurality of reinforcing fibers as may be known or later developed in the art, thereby defining a ceramic matrix composite (CMC) material. If a fiber reinforced material is used, the fibers may comprise oxide ceramics, non-oxide ceramics, or a combination thereof. For example, the oxide ceramic fibers may include those commercially available from the Minnesota Mining and Manufacturing Company under the trademark Nextel, including Nextel 720 (alumino-silicate), Nextel 610 (alumina), and Nextel 650 (alumina and zirconia). In an embodiment, the fibers are comprised of alumino-silicate fibers. For another example, the non-oxide ceramic fibers can include those

commercially available from the COI Ceramics Company under the trademark Sylramic (silicon carbide), and from the Nippon Carbon Corporation, Limited under the trademark Nicalon (silicon carbide).

The matrix material composition that surrounds the fibers may be made of an oxide or non-oxide material, such as alumina, mullite, aluminosilicate, ytrria alumina garnet, silicon carbide, silicon nitride, silicon carbonitride, and the like. The CMC material may combine a matrix composition with a reinforcing phase of a different composition (such as mullite/silica), or may be of the same composition

(alumina/alumina or silicon carbide/silicon carbide). In an embodiment, the CMC material comprises an oxide-oxide (Ox-Ox) CMC material. The fibers may be continuous or long discontinuous fibers, and may be oriented in a direction generally parallel, perpendicular, or otherwise disposed relative to the major length of the CMC material. The matrix composition may further contain whiskers, platelets, particulates, or fugitives, or the like. The reinforcing fibers may be disposed in the matrix material in layers, with the plies of adjacent layers being directionally oriented to achieve a desired mechanical strength.

The metal core 102 may comprise any suitable metal (e.g., alloy) material. In an embodiment, the metal material comprises a superalloy material, such as a nickel- based or a cobalt-based superalloy material as is well known in the art. The term "superalloy" may be understood to refer to a highly corrosion-resistant and oxidation- resistant alloy that exhibits excellent mechanical strength and resistance to creep - even at high temperatures. Exemplary superalloy materials are commercially available and are sold under the trademarks and brand names Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g. Rene N5, Rene 41 , Rene 80, Rene 108, Rene 142, Rene 220), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-750, ECY 768, 262, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys, GTD 111 , GTD 222, MGA 1400, MGA 2400, PSM 116, CMSX-8, CMSX-10, PWA 1484, IN 713C, Mar- M-200, PWA 1480, IN 100, IN 700, Udimet 600, Udimet 500 and titanium aluminide, for example.

In certain embodiments, the metal core 102 defines a cavity that also extends in the radial direction (R) through the airfoil 32 such that a cooling fluid, e.g., gas, may be passed through the core 102 to carry away heat absorbed by the metal core 102. In the case of a vane 18 or the like, the cooling fluid may exit the component 100 from one of the platforms 46, 50 to an exterior of the component 100.

To further enhance the degree of radiative heat transfer from the shell 104 to the core 102, the component 100 comprises an extended surface 110 disposed on the metal core 102. As used herein, the term“extended surface or surfaces” thus refers generally to an additional structure associated with a primary heat absorbing structure (e.g., metal core 102 in this instance) so as to increase a rate of heat transfer to the primary heat absorbing structure, e.g., metal core 102. The extended surface 110 enhances the heat transfer, namely by increasing an area available for heat transfer. In this way, the extended surface 1 10 provides an increased surface area to the metal core 102, acts as a heat sink for the metal core 102, and assists the metal core 102 in absorbing thermal energy stemming from the CMC shell 104. Typically, the extended surface 110 is disposed about an entire circumference of the metal core 102, and as with the metal core 102 and shell 104, extends radially (direction (R)) through the component 100. The extended surface 110 may be disposed/provided on the metal core 102 by any suitable structure or process, such as by additive manufacturing, spot welding, a mechanical attachment, or by any other suitable structure or process.

In certain embodiments, the extended surface 110 may be made fully integral with the metal core 102. For example, in an embodiment, the extended surface 110 is additively manufactured on the metal core 102, and thus is fully integrated with the metal core 102. As used herein, the term“additively manufactured” refers to any additive manufacturing process which utilizes layer-by-layer construction, including, but not limited to, selective laser sintering, selective laser melting (SLM), direct metal deposition, direct metal laser sintering (DMLS), direct metal laser melting, electron beam melting, electron beam wire melting, and others known in the art. In other embodiments, the extended surface 110 may be spot welded to the metal core 102 at a plurality of spaced apart locations, and thus only periodically integrated or connected with the metal core 102. Alternatively, any other suitable process may be utilized for fixedly connecting the extended surface 1 10 to the metal core 102.

The extended surface 110 may comprise any suitable material which may withstand the expected temperatures within the component 100. In an embodiment, the extended surface 100 comprises a metal material as described above for the metal core 102, such as a superalloy material. In certain embodiments, the extended surface 110 comprises the same material utilized for the metal core 102; however, it is understood the invention is not so limited. As mentioned, the extended surface 110 comprises any suitable structure or configuration which will increase a degree of radiative heat transfer from the shell 104 to the core 102 (relative to the same component without the extended surface). In an embodiment, the extended surface 110 comprises a non- continuous solid surface about a circumference thereof. By way of example only and as shown in FIG. 4, the extended surface 110 may comprise a circumferential ring 114 which extends about the metal core 102 (FIG. 3) and which includes a plurality of spaced apart dendritic branches 116 defining channels 1 18 therebetween.

Alternatively, the extended surface 110 may comprise any other suitable structure which increases a surface area of the metal core 102 to allow for an even greater amount of radiative heat transfer from the shell 104 to the metal core 102 (relative to the same component without the extended surface 110), such as fins or the like.

In accordance with another aspect, the component 100 further comprises a coating on a component surface thereof which further increases radiative heat transfer from the shell 104 to the core 102. In an embodiment, one or more of the core 102, the shell 104, or the extended surface 110 comprises a coating on a surface thereof which is effective to further increase an amount of radiative heat transfer from the shell 104 to the core 102. In certain embodiments, each of the core 102, shell 104, and extended surface 110 includes the coating thereon.

Referring to FIG. 5, there is shown a cross-section of an airfoil 32 comprising a coating (“first coating”) 120 disposed on an inner surface of the ceramic matrix composite shell 104. The first coating 120 is effective to increase an amount of surface emissivity of the CMC shell 104. In an embodiment, the surface emissivity is increased upon being subjected to a temperature of about 600° C or more. A surface’s emissivity generally refers to a ratio of energy radiated from a material's surface to that radiated from a blackbody (a perfect emitter) at the same temperature and wavelength and under the same viewing conditions. It is a dimensionless number between 0 (for a perfect reflector) and 1 (for a perfect emitter). In an embodiment, the first coating 120 comprises an emissivity value of greater than .85, and in a particular embodiment greater than .95. The first coating 120 may further comprise any suitable material and thickness to provide the desired emissivity value for the shell 104.

In an embodiment, the first coating 120 comprises a refractory pigment, a high emissivity additive, and a binder/suspension agent. Exemplary refractory pigments include zirconia, zirconia silicate, aluminum oxide, aluminum silicate, silicon oxide, and the like. The high emissivity additive may comprise a transition metal oxide, such as chromium oxide (C^C ), cobalt oxide (CoO, CO₂O3, Co30₄), ferrous oxide (FeaC ), and nickel oxide (NiO), or a rare earth metal, such as cerium oxide (Ce0₂). Cerium oxide has a melting point of 2750K (2477° C) and exhibits an emissivity of approximately 0.9 from 1000° C to 2000° C. In certain embodiment, the refractory pigment and the high emissivity additive may comprise the same material. The binder/suspension agent may further comprise any suitable material which enables the high emissivity coating to be applied to the CMC, such as a binder/suspension agent comprising 40-70 wt % aluminum phosphate solution, 25-45 wt % peptized aluminum oxide monohydrate, and 5-15 wt % ethyl alcohol.

Further, one or both of the metal core 102 and the extended surface 110 may instead or additionally comprise a second coating 122 thereon. The second coating 122 is effective to increase a surface absorptivity of the associated surface of the metal core 102 and/or the extended surface 110. Absorptivity generally refers to a measure of the ability of a body to absorb the incident radiant energy. Quantitatively, absorbance is defined as a ratio of radiant energy absorbed by a body to incident radiant energy. In certain embodiments, the second coating 122 works collectively with the first coating 120 to optimize the radiative heat transfer from the metal core 102 to the extended surface 110. Thus, in certain embodiments, the shell 104 comprises the first coating 120 thereon and the core 102 and/or extended surface 110 comprise the second coating 122 thereon as shown in FIG. 5. Flowever, it is understood that the present invention is not so limited. In other embodiments, only one of the first coating 120 and the second coating 122 is provided in the component 100.

When provided, the second coating 122 may comprise any suitable material and thickness to provide the desired absorptivity value for the respective surface. In an embodiment, the second coating 122 comprises a metal oxide or mixture of metal oxides (along with any suitable additives) that provides the desired degree of absorptivity while having a sufficiently high melting point to maintain its integrity. In other embodiments, the second coating 122 may comprise a high temperature black paint. In addition, the first and second coatings 120, 122 may be of any suitable thickness. In an embodiment, the coatings 120, 122 for the CMC shell 104 and metal core 102 (and extended surface 110), respectively, may be made so thin so as to only influence surface radiation and not provide any conduction resistances. Further, it is appreciated that one or more layers of each coating 120, 122 may be provided to provide the respective layers of the same..

While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.

Claims

CLAIMS What we claim is:

1. A hybrid component (100) comprising:

a metal core (102);

a shell (104) comprising a ceramic matrix composite material about the metal core (102);

one or more cavities (106) defined between the metal core (102) and the shell (104); and

an extended surface (110) disposed on the metal core (102), the extended surface (110) effective to increase an amount of radiative heat transfer through the one or more cavities (106) from the shell (104) to the metal core (102) relative to a like component without the extended surface (110).

2. The component (100) of claim 1 , wherein the extended surface comprises a non-continuous solid surface about a circumference thereof.

3. The component (100) of claim 1 , wherein the extended surface (110) comprises a circumferential ring (114) about the metal core (102) comprising a plurality of spaced apart dendritic branches (116) defining channels (118) therebetween.

4. The component (100) of claim 1 , wherein at least one of the core (102), the shell (104), and the extended surface (110) comprises a coating (120, 122) on a surface thereof effective to further increase radiative heat transfer from the shell (104) to the core (102).

5. The component (100) of claim 4, further comprising a first coating (120) disposed on an inner surface of the shell (104) facing the metal core (102), and wherein the first coating (120) is effective to increase an amount of surface emissivity of the shell (104).

6. The component (100) of claim 5, wherein the first coating (120) comprises an emissivity value of greater than .85.

7. The component (100) of claim 6, wherein the first coating (120) comprises an emissivity value of greater than .95.

8. The component (100) of claim 7, wherein the first coating (120) comprises a ceramic coating having an emissivity value greater than .85.

9. The component (100) of claim 1 to 8, further comprising a second coating (122) on a surface of the metal core (102) effective to increase a surface absorptivity of the metal core (102).

10. The component (100) of claims 1 to 9, further comprising a second coating (122) on a surface of the extended surface (110) effective to increase a surface absorptivity of the extended surface (110).

11. The component (100) of claims 9 to 10, wherein the second coating (122) comprises a metal oxide.

12. The component (100) of claims 1 to 11 , wherein the shell (104) comprises an oxide-oxide ceramic matrix composite material.

13. The component (100) of claims 1 to 12, wherein the component (100) comprises an airfoil (32).

14. The component (100) of claim 13, wherein the component (100) comprises a turbine blade (16) or vane (18) comprising the airfoil (32).