CERAMIC MATRIX COMPOSITE COMPONENT, TURBINE SYSTEM AND FABRICATION PROCESS

Abstract

A ceramic matrix composite component, turbine system and fabrication process are disclosed. The ceramic matrix composite (CMC) component includes a CMC material, an environmental barrier coating (EBC) on the CMC material, and a hard wear coating applied over the EBC. The turbine system includes a rotatable CMC component having a hard wear coating, and a stationary turbine component, the stationary turbine component having an abradable coating arranged and disposed to be cut by the silicon carbide material. The fabrication process includes positioning the rotatable CMC a pre-determined distance from the stationary turbine component and rotating the rotatable CMC component. The hard wear coating on the rotatable CMC component cuts the abradable coating on the stationary turbine component.

Description

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under contract number DE-FC26-05NT42643 awarded by the Department of Energy. The government has certain rights in the invention.

FIELD OF THE INVENTION

The present invention is directed to manufactured components and a process of using manufactured components. More particularly, the present invention is directed to a wear coating applied to ceramic matrix composite (CMC) components.

BACKGROUND OF THE INVENTION

Gas turbines are continuously being modified to provide increased efficiency and performance. These modifications include the ability to operate at higher temperatures and under harsher conditions, which often requires material modifications and/or coatings to protect components from such temperatures and conditions. As more modifications are introduced, additional challenges are realized.

One modification to increase performance and efficiency involves minimizing a gap between a turbine bucket tip and a turbine shroud. The minimized gaps lead to rubs between the bucket tip and the shroud during certain power transients. An environmental barrier coating (EBC) used on CMC gas turbine bucket tips can be damaged by these rubs.

Damage to the EBC on a turbine bucket may expose an underlying CMC to high temperature combustion gasses leading to increased volatilization. Increasing the EBC thickness on the bucket increases the weight of the bucket and decreases the efficiency and performance of the turbine. Furthermore, opening the gap to minimize rubs also decreases efficiency and performance.

A manufactured component and process of using manufactured components not suffering from the above drawbacks would be desirable in the art.

BRIEF DESCRIPTION OF THE INVENTION

In an exemplary embodiment, a ceramic matrix composite (CMC) component includes a CMC material, an environmental barrier coating applied over the CMC material, and a hard wear coating applied over the environmental barrier coating.

In another exemplary embodiment, a turbine system includes a rotatable CMC component having a CMC material, an environmental barrier coating (EBC) applied over the CMC material, and a hard wear coating applied over the EBC. The turbine system further includes a stationary turbine component having an abradable coating arranged and disposed adjacent to the rotatable CMC component, to be cut by the hard wear coating of the rotatable turbine component.

In another exemplary embodiment, a fabrication process includes positioning a rotatable CMC component a pre-determined distance from a stationary turbine component, the stationary turbine component comprising an abradable material. The fabrication process further includes rotating the rotatable CMC component, wherein, a hard wear coating on the rotatable CMC component cuts the abradable material of the stationary turbine component.

Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a turbine bucket.

FIG. 2 is a schematic view of a turbine system.

FIG. 3 is a schematic view of a rotating part to stationary part interference pattern.

Wherever possible, the same reference numbers will be used throughout the drawings to represent the same parts.

DETAILED DESCRIPTION OF THE INVENTION

Provided is an exemplary ceramic matrix composite component, turbine system and fabrication process. Embodiments of the present disclosure, in comparison to methods and products not utilizing one or more features disclosed herein, utilize decreased environmental barrier coating (EBC) material, decrease damage to the EBC, decrease repairs, decrease ceramic matrix composite (CMC) volatilization, decrease tip clearance, increase turbine efficiency, or a combination thereof.

Referring to FIG. 1, in one embodiment, a rotatable CMC component 101 includes a CMC material 102, an EBC 104 applied over the CMC material 102, and a hard wear coating 106 applied over the EBC 104. The rotatable CMC component 101 is any suitable component that may experience volatilization and/or rub wear such as, but not limited to, a blade or bucket. The bucket may be either an unshrouded bucket or a shrouded bucket. The hard wear coating 106 is a material such as, but not limited to silicon carbide (SiC), SiO₂, cubic boron nitride (CBN), or combinations thereof. Examples of CMC material 102 include, but are not limited to, carbon-fiber-reinforced carbon (C/C), carbon-fiber-reinforced silicon carbide (C/SiC), silicon-carbide-fiber-reinforced silicon carbide (SiC/SiC), and alumina-fiber-reinforced alumina (Al₂O₃/Al₂O₃). The CMC's have increased elongation, fracture toughness, thermal shock, dynamical load capability, and anisotropic properties as compared to a monolithic ceramic structure. However, without an environmental coating, the CMC's may volatilize during gas turbine operation.

For example, at temperatures above 1500° F. water vapor chemically reacts with the CMC material 102. The water vapor reacts with silicon and carbon in the CMC material 102 to produce silicon hydroxide (SiOH) and carbon dioxide (CO₂), respectively. The SiOH and CO₂ formed by the reaction between the water vapor and the CMC material 102 slowly gasses off, or volatilizes. Over many hours of operation above 1500° F. the CMC material 102 is removed from the outside in.

The EBC 104 protects the CMC material 102 from water vapor, heat, and other combustion gases which may cause the volatilization or deterioration of the CMC material 102. In one embodiment, the EBC 104 reduces or eliminates the occurrence of the chemical reaction between the water vapor and the CMC material 102. The EBC 104 may be any suitable material for protecting the CMC material 102 from the hot gases of combustion. Suitable EBC materials include, but are not limited to, barium strontium alumino silicate (BSAS), mullite, yttria-stabilized zirconia, and combinations thereof.

Referring to FIG. 1, in one embodiment, the hard wear coating 106 forms an outermost layer of the rotatable CMC component 101, being applied over one or more EBCs 104. The hard wear coating 106 is applied over any suitable portion of the rotatable CMC component 101 that experiences wear during a rub event. Suitable portions of the rotatable CMC component 101 include, but are not limited to, a tip portion 103, a platform interface, the contact point between an airfoil and a shroud, the contact point between the airfoil and shroud rails, or a combination thereof. Additionally, the hard wear coating 106 is applied with any suitable thickness to reduce or eliminate a loss of the EBC 104 during deep rubs. Suitable thicknesses include, but are not limited to, between about 0.1 mils and about 4 mils, between about 0.5 and about 3 mils, between about 1 mil and about 2.5 mils, or any combination, sub-combination, range, or sub-range thereof. For example, in one embodiment, the hard wear coating 106 extends across the entire width and length of the tip portion 103. In another example, the hard wear coating 106 is applied around an edge of the tip portion 103 and down to a radially extending surface of the rotatable CMC component 101 to provide increased protection from severe conditions.

The hard wear coating 106 is applied over the EBC 104 through any suitable deposition process. For example, one suitable deposition process is physical vapor deposition (PVD). PVD condenses a vaporized form of the hard wear coating 106 on the EBC 104 of the rotatable CMC component 101 to form a hard thin hard wear coating 106. Other suitable deposition process include, but are not limited to, chemical vapor deposition (CVD), air plasma spray (APS), combustion spraying with powder or rod, slurry coating, sol gel, electrophoretic deposition, tape casting, or a combination thereof. In one embodiment, SiC is used as the hard wear coating 106. The SiC adheres well because a coefficient of thermal expansion (CTE) of the SiC is well matched to the CTE for both the CMC material 102 and the EBC 104. The EBC 104 may be applied to and in contact with the CMC material 102. The hard wear coating 106 may be applied to and contact the EBC 104.

Referring to FIG. 2, in one embodiment, a turbine system 100 includes the rotatable CMC component 101 such as a turbine bucket attached to a rotatable member 107 such as a turbine disc. The CMC component 101 and the rotatable member 107 extend away from a central point 131 such as a rotatable shaft. In a further embodiment, a plurality of rotatable CMC components 101 are positioned around the rotatable member 107, and extend away from the central point 131. In one embodiment, a stationary turbine component 120 forms a perimeter around the rotatable member 107, the CMC component 101 being between the rotatable member 107 and the stationary turbine component 120. The stationary turbine component 120 is centered around the central portion 131, sharing a common central point with the rotatable member 107. The tip portion 103 of the rotatable CMC component 101 ideally forms a seal with the stationary turbine component 120.

An abradable coating 122 on the stationary turbine component 120 is arranged and disposed to be cut by the rotatable CMC component 101. The abradable coating 122 is any suitable coating based upon stationary turbine component 120 material, and operating temperature. Suitable abradable coatings 122 for a metallic stationary turbine compound 120 at lower temperatures (up to approximately 2200° F.) include, but are not limited to, metal abradables having the general form MCrAlY, such as NiCrAlY, CoCrAlY, FeCrAlY, or combinations thereof. In one embodiment, the metal abradable is applied by air plasma spray, wire arc and combustion spraying, high velocity oxy-fuel (HVOF), or combinations thereof.

Suitable abradable coatings 122 for the metallic stationary turbine component 120 at higher temperatures (at least approximately 2200° F.) include, but are not limited to, ceramic abradables such as partially stabilized zirconia, with yttria as the stabilizer (7 or 8 YSZ). In one embodiment, the metal abradables for lower temperature applications and/or the ceramic abradables for higher temperature applications, initially contain a polyester component during application. The polyester component is burned out by heating in air to leave behind porosity. The porosity makes the abradable coating 122 compliant with a rotating CMC component 101 during a rub. In one embodiment, at least one of the metal abradable coating and the ceramic abradable coating further includes boron nitride as a solid lubricant.

Suitable abradable coatings 122 for a ceramic stationary turbine component 120 include, but are not limited to, silicates. For example, in one embodiment, the suitable abradable coating 122 for the ceramic stationary turbine component 120 operating at lower temperatures (up to approximately 2200° F.) is barium strontium alumino silicate (BSAS). As another example, in one embodiment, the suitable abradable coatings 122 for the ceramic stationary turbine component 120 operating at higher temperatures (at least approximately 2201° F.) include high rare earth content silicates, such as Yb₂O₃Si₂O₇.

In one embodiment, the silicate abradable coatings 122 include a polyester component during application. The polyester component is burned out by heating in air to leave behind porosity which contributes to compliance. In one embodiment, the silicate abradable coating 122 is applied by air plasma spray, slurry coating, sol gel, electrophoretic deposition, tape casting, or a combination thereof.

Referring to FIG. 3, in one embodiment, the fabrication process 200 includes positioning the rotatable member 107 a pre-determined distance 105 from the stationary turbine component 120. The rotatable member 107 is rotated around the central point 131, rotating the CMC component 101. During power transients the pre-determined distance 105 fluctuates, resulting in contact between the tip portion 103 and the abradable coating 122 of the stationary turbine component 120. The hard wear coating 106 cuts into the abradable coating 122, forming a pathway 201 to reduce contact during further power transients. In one embodiment, the direction of rotation 203 is clockwise. In another embodiment, the direction of rotation 203 is counter-clockwise (not shown).

Referring to FIG. 1, in one embodiment, a hardness value of the hard wear coating 106 is higher than the hardness value of the EBC 104. The hardness value represents an ability of a material to withstand contact without being damaged. For example, the hard wear coating 106 having the higher hardness value than the EBC 104 is able to withstand deeper rubs than the EBC 104.

In one embodiment, the hard wear coating 106 having the higher hardness value is applied over the EBC 104 to protect the EBC 104 from being damaged when energy transients create contact between the rotating CMC component 101 and the stationary turbine component 120. In one embodiment, the hard wear coating 106 is SiC, the SiC having a knoop value (hardness) of 2480 and the EBC 104 having a knoop value of less than 2480. In one embodiment, the EBC 104 has a knoop value of between approximately 500 to 1800, depending on the architecture of the EBC 104. Architectural elements include the number of layers, the thickness of each layer, layer composition, the application process and the heat treat used.

For example, in one embodiment, the EBC 104 is only able to survive contact between the tip portion 103 and the abradable coating 122 of up to about 10 mils. Gas turbine rotating parts incur contact against static parts well in excess of 10 mils. Contact above 10 mils deteriorates the EBC 104 structure and function while possibly subjecting the base material to environmental attack. To protect the EBC 104, the hard wear coating 106 is applied over the EBC 104, decreasing or eliminating damage to the EBC 104 on deep rubs.

Typical turbine bucket running conditions include water vapor, hot gases of combustion at temperatures in the range of approximately 2,000° F. to approximately 3,000° F. The hard wear coating 106 gradually volatizes under these conditions, exposing the EBC 104. Applying a minimal thickness of the hard wear coating 106 minimizes a gap between the EBC 104 on the rotatable CMC component 101 and the abradable coating 122 of the stationary turbine component 120. Decreasing the gap increases the efficiency of a system by preventing increased air flow between the rotatable CMC component 101 and the stationary turbine component 120.

The time it takes for the sacrificial hard wear coating 106 to volatize under operating conditions is a break-in period. In one embodiment, the break-in period is up to about 100 hours. In one embodiment, the break-in period is at least about 100 hours. The break-in period during which the hard wear coating 106 remains on the EBC 104 in the tip portion 103 is sufficient so that the hard wear coating 106 is present to abrade during power transients resulting in deep rubs.

In one embodiment, the break-in period includes a constant pre-determined speed of rotation for a predetermined time. In one embodiment, the break-in period includes a pre-determined fluctuation and/or series of fluctuations in speed of rotation. The pre-determined speed of rotation or fluctuations in speed of rotation are set to induce power transients during the break-in period. During the break-in period, the power transients of the fabrication process 200 cause the hard wear coating 106 to contact, and cut, the abradable material 122. The cut by the hard wear coating 106 forms a pathway 201 in the abradable material 122.

The hard wear coating 106 gradually volatilizes throughout the break-in period, creating the pathway 201 between the stationary abradable material and the tip portion 103. The break-in period minimizes the pathway 201 between the stationary component and the tip portion 103 around which gas may leak, without damaging the EBC 104. Gas leakage is minimized due to the cutting character of the hard wear coating 106 and its minimal thickness. Forming the minimized space between the pathway 201 and the tip portion 103, increases the efficiency of the turbine system 100. Additionally, the pathways 201 formed in the stationary abradable material by the cuts from the hard wear coating 106 decrease damage to the EBC 104 due to rubs after the hard wear coating 106 has volatilized. The undamaged EBC 104 provides the CMC material 102 with increased volatilization protection.

While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.

Claims

1. A ceramic matrix composite component, comprising:

a ceramic matrix composite material;

an environmental barrier coating applied over the ceramic matrix composite material; and

a hard wear coating applied over the environmental barrier coating.

2. The ceramic matrix composite component of claim 1, wherein the hard wear coating further comprises silicon carbide.

3. The ceramic matrix composite component of claim 1, wherein the hard wear coating is the outermost layer.

4. The ceramic matrix composite component of claim 1, wherein the hard wear coating coats an outer surface of the component.

5. The ceramic matrix composite component of claim 1, wherein the hard wear coating is applied by physical vapor deposition.

6. The ceramic matrix composite component of claim 1, wherein the hard wear coating is applied by chemical vapor deposition.

7. A turbine system, comprising:

a rotatable ceramic matrix composite bucket attached to a disc, the rotatable ceramic matrix composite bucket having a ceramic matrix composite material, an environmental barrier coating applied over the ceramic matrix composite material, and a hard wear coating applied over the environmental barrier coating; and

a stationary turbine shroud, the stationary turbine shroud having an abradable coating arranged and disposed to be cut by the hard wear coating of the rotatable ceramic matrix composite bucket.

8. The turbine system of claim 7, wherein the stationary turbine shroud forms a perimeter around the rotatable bucket.

9. The turbine system of claim 7, wherein a plurality of rotatable ceramic matrix composite buckets are attached circumferentially to the turbine disc.

10. A fabrication process, comprising:

positioning a rotatable ceramic matrix composite component a pre-determined distance from a stationary turbine component, the stationary turbine component having an abradable coating and the rotatable component having a hard wear layer applied over an environmental layer; and

rotating the rotatable ceramic matrix composite component;

wherein, the hard wear layer on the rotatable ceramic matrix composite component cuts the abradable coating on the stationary turbine component.

11. The fabrication process of claim 10, wherein power transients cause the rotatable ceramic matrix composite component to contact the abradable coating.

12. The fabrication process of claim 10, wherein the cut by the rotatable ceramic matrix composite component forms a pathway in the abradable coating.

13. The fabrication process of claim 10, wherein the process further comprises a break-in period.

14. The fabrication process of claim 13, wherein the break-in period is up to about 100 hours.

15. The fabrication process of claim 13, wherein the break-in period is at least 100 hours.

16. The fabrication process of claim 10, wherein the hard wear layer volatilizes after predetermined exposure to elevated temperature.

17. The fabrication process of claim 16, wherein the elevated temperature comprises at least 1,500° F.

18. The fabrication process of claim 16, wherein the elevated temperature is at least 2,900° F.

19. The fabrication process of claim 16, wherein the hard wear coating volatilization exposes the environmental barrier coating.

20. The fabrication process of claim 10, wherein the environmental barrier coating has a hardness value of less than that of the hard wear coating.