US20180135638A1

US20180135638A1 - Ceramic coating composition for compressor casing and methods for forming the same

Info

Publication number: US20180135638A1
Application number: US15/353,182
Authority: US
Inventors: Neelesh Nandkumar Sarawate; Luc Stephane Leblanc; Sathyanarayanan Raghavan; Changjie Sun; Ananda Barua
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2016-11-16
Filing date: 2016-11-16
Publication date: 2018-05-17

Abstract

Coating systems for components of a gas turbine engine, such as a compressor casing, are provided. The coating system can include a ceramic material disposed along the compressor casing on a surface to be adjacent to a rotating compressor blade. The coating system is harder than the compressor blade and can reduce the rub ratio between the casing and blade. The coating system can thereby increase the lifetime of the compressor casing and blades. Methods are also provided for applying the coating system onto a compressor casing.

Description

FIELD

Embodiments of the present invention generally relate to ceramic coating systems for metallic components, particularly for use on a compressor casing in a gas turbine engine.

BACKGROUND

Gas turbine engines typically include a compressor for compressing air. The compressor includes a series of stages of blades rotating around a shaft. The compressed air is mixed with a fuel and channeled to a combustor, where the mixture is ignited within a combustion chamber to generate hot combustion gases. The combustion gases are channeled to a turbine. The turbine section of a gas turbine engine contains a rotor shaft and one or more turbine stages, each having a turbine disk (or rotor) mounted or otherwise carried by the shaft and turbine blades mounted to and radially extending from the periphery of the disk. A turbine assembly typically generates rotating shaft power by expanding hot compressed gas produced by the combustion of a fuel. Gas turbine buckets or blades generally have an airfoil shape designed to convert the thermal and kinetic energy of the flow path gases into mechanical rotation of the rotor.
In a compressor, as well as in a turbine, engine performance and efficiency may be enhanced by reducing the space between the tip of the rotating blades and the respective casing to limit the flow of air over or around the top of the blade that would otherwise bypass the blade. For example, a compressor blade may be configured so that its tip fits close to the compressor casing during engine operation. During engine operation, however, blade tips may rub against the casing, thereby increasing the gap and resulting in a loss of efficiency, or in some cases, damaging or destroying the blade set. Blade material may be transferred to the compressor case creating scabs on the casing that extend into the clearance between the blades and casing, further aggravating any rubbing against the blade tip. In addition, the high speeds and high contact forces increase the local temperature at the blade tip such that the metal blade tip may melt or soften. The melting or softening of the blade tip may then lead to additional removal of the blade tip material when rubbed against the compressor case. These interactions result in a reduced lifetime of the compressor components.
Thus, an improved design of a compressor casing and a compressor blade and case assembly is desirable in the art.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.
A coated compressor casing is generally provided, the coated compressor casing comprising a compressor casing defining an inner surface of the compressor casing and comprising a base material, and a coating system disposed along the inner surface of the compressor casing, wherein the compressor casing is configured to have a compressor blade positioned within the casing, and wherein the coating system comprises a ceramic material that is harder than the compressor blade. For instance, in some embodiments, the coating system has a hardness about 10% to about 50% higher than a hardness of the compressor blade, and in some embodiments, the coating system has a modulus about 10% to about 50% higher than a modulus of the compressor blade.
In some embodiments, the ceramic material comprises yttria stabilized zirconia, mullite, alumina, ceria, rare-earth zirconates, rare-earth oxides, metal-glass composites, zirconia stabilized with an oxide, silicate, chromium oxide, chrome carbide, or combinations thereof. In certain embodiments, the coating system has a uniform thickness across the inner surface of the compressor casing, and in some embodiments, the inner surface of the compressor casing is configured to be adjacent to a rotating compressor blade. In some embodiments, the coating system has a thickness of about 127 microns to about 254 microns. In some embodiments, the coating system includes a bond coat. The coating system is generally non-abradable, and in some embodiments, the compressor case is configured to be positioned in a turbofan engine.
Aspects of the present disclosure are also drawn to a gas turbine engine comprising a compressor comprising a compressor case having an inner surface, wherein the compressor case comprises a base material, and a compressor blade having a blade tip, wherein the compressor blade comprises a base material, and a coating system disposed along the inner surface of the compressor case, wherein the coating system has a higher hardness than a hardness of the compressor blade base material. For instance, in some embodiments, the coating system has a hardness about 10% to about 50% higher than a hardness of the compressor blade base material. In some embodiments, the coating system includes a bond coat and in some embodiments, the compressor blade base material is uncoated. In certain embodiments, the compressor blade base material comprises a nickel superalloy, while in some embodiments, the compressor blade comprises a curved airfoil.
Aspects of the present disclosure are also drawn to a method of preparing a coated compressor casing, the method comprising forming a coating system comprising a ceramic material along an inner surface of a compressor casing, wherein the compressor casing is configured to have a compressor blade positioned within the casing, and wherein the ceramic material is harder than the compressor blade. For instance, in some embodiments, the coating system has a hardness of about 10% to about 50% higher than a hardness of the rotating compressor blade. In some embodiments, forming the coating system along the surface of the compressor casing comprises forming the ceramic material along a surface configured to be adjacent to a rotating compressor blade. In some embodiments, forming the coating system along the surface of the blade tip comprises forming the ceramic material along the surface of the compressor casing to a thickness of about 127 microns to about 254 microns.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended FIGS., in which:

FIG. 1 is a schematic cross-section of an exemplary compressor casing comprising a coating system in accordance with one embodiment of the present disclosure;

FIGS. 2a and 2b are schematic views of exemplary compressor casings and blade assemblies comprising a coating system in accordance with one embodiment of the present disclosure;

FIG. 3 is a schematic cross-sectional view of an exemplary gas turbine engine in accordance with one embodiment of the present disclosure;

FIG. 4 is an exemplary method of preparing a coating system in accordance with one embodiment of the present disclosure;

FIGS. 5a, 5b, and 5c illustrate the rubbing of a conventional compressor blade and casing;

FIGS. 6a, 6b, and 6c illustrate the rubbing of a compressor blade and casing in accordance with one embodiment of the present invention; and

FIGS. 7a and 7b compare the rubbing of a conventional compressor blade and casing to that of a compressor blade and casing in accordance with one embodiment of the present invention.

Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.
The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.
In the present disclosure, when a layer is being described as “on” or “over” another layer or substrate, it is to be understood that the layers can either be directly contacting each other or have another layer or feature between the layers, unless expressly stated to the contrary. Thus, these terms are simply describing the relative position of the layers to each other and do not necessarily mean “on top of” since the relative position above or below depends upon the orientation of the device to the viewer.
Chemical elements are discussed in the present disclosure using their common chemical abbreviation, such as commonly found on a periodic table of elements. For example, hydrogen is represented by its common chemical abbreviation H; helium is represented by its common chemical abbreviation He; and so forth.
A coating system for a compressor casing is generally provided herein, along with methods of forming such coating system. As used herein, “compressor case” and “compressor casing” may be used interchangeably. The coating system may be used to protect the compressor casing against rubbing of a moving part such as a compressor blade. The composition of the coating system and the methods of applying the coating system to the compressor casing reduce the overall wear of the compressor blade and casing during high-speed rubs and may thereby increase the lifetime of the compressor blade and casing. The coating system includes a ceramic coating disposed along the compressor casing that is harder than the material with which the compressor blade is formed. The coating system is non-abradable in that the coating system is designed to be harder than the material of the rotating compressor blade such that the blade material tends to be removed rather than the coating system.
Without intending to be limited by theory, the difference in hardness of the coating system and the compressor blade may reduce the overall amount of material that is rubbed off of the blade. As used herein, “rub ratio” refers to the amount of clearance opening divided by the amount of rubbing incursion. The coating system can (a) reduce damage to the blade tip during a rubbing event between stator and rotor, (b) achieve a tighter clearance between the stator and rotor during engine operations, and (c) solve high rub ratio problems.
The coating system has a higher hardness (or Young's Modulus) compared to the blade material. Without intending to be bound by theory, the coating system affects the contact interaction between the blade and the casing during a rub event. Because the casing has a harder coating on it, when the blade first rubs against the casing, there is initially more wear on the blade. However, because the blade loses material readily in the initial rub, less contact forces are generated. Consequently, the blade has less deflection and lower amplitudes of vibration, leading to less radial growth of the blade and lower overall wear of the blade during the whole rub event.
The lower contact forces at the blade tip also result in lower frictional energy dissipation at the blade tip. Heat generated from frictional energy created with conventional blades rubbing against conventional casings may lead to softening of the blade tip or melting of the blade tip. Such softening or melting of the blade material increases the amount of material removed during subsequent rubbing events. With less frictional dissipation due to the coating system, a smaller temperature increase may be observed in the blade. The coating system may thereby reduce softening or melting of the blade and thus reduce further wear of the blade.
In addition, the coating system prevents or reduces scab (deposition of blade material on the casing) build-up during a rub event. When removed, the coating system may wear out cleanly—that is, without building any material deposition on the casing. When a bare blade and a conventional compressor casing rub against each other, the rubbing creates a scab, or deposition of the blade material on the casing. Scab build-up during a rub event can lead to a reduction in clearance between the blade tip and the casing, increasing the amount of rubbing between the blade tip and casing, and thus, leading to enhanced material loss from the blade tip during a rub event. The scab may act as a cutting tool to remove more material from the tip of the blade. The presence of the hard coating system can help reduce the blade wear, by reducing scab build up. In some embodiments, the amount of scab build-up is significantly reduced. When the scab is eliminated or reduced, the overall wear of the blade is also reduced.
With certain blades the amount of material loss at the blade tip may be equivalent to the incursion or interference depth. Turbine blades typically have a 1:1 rub ratio (the ratio of blade material lost to interference). However, compressor blades, particularly aft compressor blades, can have a high rub ratio due to their design and geometry, such as a curved airfoil. When running at high speeds, the airfoil may be pushed radially up to an almost standing position (which may be referred to as “radial growth”), thereby rubbing more against the compressor case. Rub ratios significantly exceeding 1:1 have been previously observed for high pressure compressors. The compressor blades can rub on the casing during certain transients, and upon rub, the blades can lose a substantially higher amount of material than the magnitude of the interference. This high rub ratio leads to high blade wear, thereby opening the clearance between the blade tip and the casing, which results in loss of flow that does useful work. High rub ratios have a significant impact on engine performance and operability. Thus, reducing the rub ratio may improve the compressor performance and operability. The coating system incorporates a ceramic material with a higher hardness than that of the compressor blade. The harder coating system may reduce overall blade loss during a rub event and may result in reduced clearance between the stator and rotor during engine operating conditions. The coating system may thereby improve the specific fuel consumption (SFC) of the engine, resulting in increased fuel economy.
The coated compressor casing can be utilized as a component for a gas turbine engine. In particular, the coated compressor casing can be positioned within a gas flow path of a gas turbine engine such that the coating system protects the compressor casing within the gas turbine engine. The coating system may be applicable to casings in a high pressure compressor (HPC), fan, booster, high pressure turbine (HPT), and low pressure turbine (LPT) of both airborne and land-based gas turbine engines.
FIG. 1 is a schematic cross-section of an exemplary compressor casing comprising a coating system in accordance with one embodiment of the present disclosure. In particular, FIG. 1a is a cross-sectional schematic view of a compressor casing 10 comprising a base material 12 and a surface 14. In the embodiment illustrated in FIG. 1, a coating system 20 comprising a ceramic material 16 is disposed along the surface 14 of the compressor casing 10. The coating system 20 has a surface 18 that is generally exposed to a compressor blade.
In certain embodiments, the compressor casing 10 comprises a base material 12. In some embodiments, the base material 12 may include a metal such as steel or superalloys (e.g., nickel-based superalloys, cobalt-based superalloys, or iron-based superalloys), or combinations thereof.
The coating system 20 is configured such that rubbing of the compressor casing 10 and the associated compressor blade may be reduced. The coating system incorporates components that have a higher hardness than the compressor blade and thereby protect the underlying metal of the base material 12 of the compressor casing 10 from rubbing events and reduce the overall rub ratio of the compressor blade. For instance, in certain embodiments, the coating system 20 may comprise a ceramic material 16 with a higher hardness than the compressor blade with which the compressor casing is to be used.
Various ceramic materials may be suitable in the coating system 20. For instance, the ceramic material may be one or more of yttria stabilized zirconia (YSZ), mullite (3Al₂O₃-2SiO₂), alumina (Al₂O₃), ceria (CeO₂), rare-earth zirconates (e.g., La₂Zr₂O₇), rare-earth oxides (e.g., La₂O₃, Nb₂O₅, Pr₂O₃, CeO₂), and metal-glass composites, and combinations thereof (e.g., alumina and YSZ or ceria and YSZ). Zirconia may be stabilized with a multitude of other oxides including CaO, MgO, CeO2, and also many rare earth oxides such as Gd₂O₃, Yb₂O₃, La₂O₃to name a few. In addition to stabilized zirconates, several alumina based compounds with TiO₂, ZrO₂, SiO₂, Y₂O₃in various ratios are suitable. Silicate coatings including Zircon (ZrSiO₄) and hafnon may be used. Chromium oxide and mixtures of chromium oxide with TiO₂, SiO₂, and Al₂O₃may also be used. In addition, cermet such as chrome carbide in a high temperature resistant matrix such as superalloys may be especially suitable. In some embodiments, the coating system may include a matrix of the above mentioned materials (one or more of these materials) with an embedded hard fugitive material such as boron nitride, tungsten carbide, or combinations thereof. At certain temperatures boron nitride or tungsten carbide may eventually degrade leaving the hard matrix material and leaving the thickness of the coating system substantially intact. The hard fugitive material would participate in the initial rubbing events.
The coating system may be formed by any suitable process. For instance, the coating system may be deposited on the compressor casing by air-plasma spray (APS), electron beam physical vapor deposition (EBPVD), high velocity oxygen fuel (HVOF), electrostatic spray assisted vapor deposition (ESAVD), direct vapor deposition, high velocity air fuel (HVAF), or combinations thereof. For the thermal spray processes (e.g., plasma spray, HVOF, HVAF, etc.), the type of feedstock material may be a powder, a suspension, a solution, or a mixture thereof. Tip grinding may occur before or after application of the coating system 20.
In some embodiments, the ceramic material 16 may be applied to the compressor casing 10 to form one or more layers of ceramic material 16. In certain embodiments, the ceramic material 16 may be applied to the compressor casing 10 such that the ceramic material 16 becomes dispersed throughout another layer, such as dispersed throughout a matrix of another component along the compressor casing 10, while in some embodiments the ceramic material 16 may be a matrix phase with another material dispersed throughout the ceramic material 16. The ceramic material 16 can be a discontinuous phase within the matrix or a continuous phase within the matrix.
One or more ceramic materials 16 may be used along the compressor casing 10. For instance, a plurality of ceramic materials 16 may be applied to the compressor casing 10 and may form one or more ceramic materials 16 along the compressor casing 10. Various alternative configurations are possible without deviating from the intent of the present disclosure.
The coating system 20 may have a thickness of about 1 mils (about 25 microns) to about 20 mils (about 508 microns), such as about 2 mils (about 50 microns) to about 15 mils (about 381 microns), about 3 mils (about 76 microns) to about 12 mils (about 305 microns), or about 5 mils (about 127 microns) to about 10 mils (about 254 microns) along the compressor case 10.
As shown in FIG. 1, in this embodiment, the compressor casing 10 is coated with a coating system 20. The coating system 20 is disposed along the surface 14 of the compressor casing 10. The coating system 20 may cover at least a portion of the compressor casing 10, and in some cases, the coating system 20 may uniformly cover the compressor casing 10. For instance, the coating system 20 may uniformly cover the portion of the compressor casing 10 most immediately adjacent to the compressor blades when positioned in the compressor section of the engine (see e.g., FIG. 3).
FIGS. 2a and 2b are schematic views of exemplary compressor casings and blade assemblies comprising a coating system in accordance with one embodiment of the present disclosure. As shown in FIGS. 2a and 2b , in these embodiments, the coating system 20 is disposed with a uniform thickness along the compressor casing 10. For instance, the coating system 20, in some embodiments, may have a thickness of about 5 mils (about 127 microns) to about 10 mils (about 254 microns) along the compressor casing 10. In some embodiments, the coating system 20 may be disposed along certain areas of the compressor casing 10 with different thicknesses.
In FIGS. 2a and 2b , a compressor casing 10 is shown comprising a base material 12 and having a surface 14. The compressor casing 10 is coated with a coating system 20 which has a surface 18. FIG. 2a also illustrates a compressor blade 22 having a blade tip 24 that rotates around the inner surface of the compressor case (the inner surface in this embodiment is surface 14 of the compressor case). FIG. 2b illustrates a compressor blade 22 a having a blade tip 24 that rotates around the inner surface of the compressor case. The blades 22, 22 a may comprise a base material such as a metal such as steel or superalloys (e.g., nickel-based superalloys, cobalt-based superalloys, or iron-based superalloys), or combinations thereof. As shown in FIGS. 2a and 2b , the blades 22, 22 a are uncoated. As used herein, “uncoated” or “bare” refers to the absence of a coating or additional layer applied to the base material of the component. No abradable coating or additional protective coating is needed for the compressor blades in these embodiments.
In the embodiments illustrated in FIGS. 2a and 2b , the compressor casing 10 includes a coating system 20 comprising a ceramic material 16 disposed along the surface 14 of the compressor casing 10. The coating system 20 has a higher hardness than the base material of the blades 22, 22 a. For instance, the coating system 20 may have a hardness of about 15 Rockwell C to about 80 Rockwell C, such as about 20 Rc to about 75 Rc, or about 25 Rc to about 70 Rc or a Young's modulus of greater than about 27,000 ksi (about 186,158 MPa), such as about 27,000 ksi (about 186,158 MPa) to about 40,000 ksi (about 275,790 MPa), or about 32,000 ksi (about 220,632 MPa) to about 35,000 ksi (about 241,316 MPa). The coating system 20 may be at least about 5% harder, such as about 10% to about 50%, or about 10% to about 45% harder than the base material of the blades 22, 22 a. The coating system 20 may have a Young's modulus of about 5% greater than, such as about 10% to about 50%, or about 10% to about 45% greater than the Young's modulus of the base material of the blades 22, 22 a.
In the embodiments illustrated in FIGS. 2a and 2b , the blades 22, 22 a are generally represented as being adapted for mounting to a disk or rotor within the compressor section of an aircraft gas turbine engine (illustrated in FIG. 3). For this reason, the blades 22, 22 a are represented as including a dovetail 34 for anchoring the blades 22, 22 a to a compressor disk by interlocking with a complementary dovetail slot formed in the circumference of the disk. As represented in FIGS. 2a and 2b , the interlocking features comprise protrusions referred to as tangs 32 that engage recesses defined by the dovetail slot. The blades 22, 22 a are further shown as having a platform 28 that separates an airfoil 26, 26 a from a shank 30 on which the dovetail 34 is defined.
The blades 22, 22 a include a blade tip 24 disposed opposite the platform 28. As such, the blade tip 24 generally defines the radially outermost portion of the blades 22, 22 a and, thus, may be configured to be positioned adjacent to the stationary casing 10 of the compressor. The blade tip 24 may be referred to as the interface between the blade and the casing and may be referred to as the rubbing area between the blade and the casing.
In the embodiment illustrated in FIG. 2a , the airfoil 26 is generally straight. In some embodiments, it may be beneficial to have a generally straight airfoil 26. In certain embodiments, it may be beneficial to have a curved airfoil, shown for example in FIG. 2b as airfoil 26 a. As shown in FIG. 2b , the airfoil 26 a of the compressor blade 22 a is a generally curved body in that a portion of the airfoil 26 a bends out away from the blade tip 24. While curved airfoils provide certain benefits to the compressor, during use, force applied to the compressor blade 22 a may push the generally curved body into a more straightened position (which may be referred to as “radial growth”) forcing the blade tip 24 to contact the casing 10, increasing the occurrence or magnitude of a rub event between the blade tip 24 and the casing 10. Since the blade wears cleanly with the coating system 20 disposed on the casing 10, the blade 22 a may experience less vibrational forces and thus, less radial growth than in systems without the coating system 20. The blade 22 a may thereby have less blade wear while still obtaining the benefits of curved airfoils.
FIG. 3 is a schematic cross-sectional view of a gas turbine engine in accordance with one embodiment of the present disclosure. Although further described below generally with reference to a turbofan engine 100, the present disclosure is also applicable to turbomachinery in general, including turbojet, turboprop and turboshaft gas turbine engines, including industrial and marine gas turbine engines and auxiliary power units.
As shown in FIG. 3, the turbofan 100 has a longitudinal or axial centerline axis 102 that extends therethrough for reference purposes. In general, the turbofan 100 may include a core turbine or gas turbine engine 104 disposed downstream from a fan section 106.
The gas turbine engine 104 may generally include a substantially tubular outer casing 108 that defines an annular inlet 120. The outer casing 108 may be formed from multiple casings. The outer casing 108 encases, in serial flow relationship, a compressor section having a booster or low pressure (LP) compressor 122, a high pressure (HP) compressor 124, a combustion section 126, a turbine section including a high pressure (HP) turbine 128, a low pressure (LP) turbine 130, and a jet exhaust nozzle section 132. A high pressure (HP) shaft or spool 134 drivingly connects the HP turbine 128 to the HP compressor 124. A low pressure (LP) shaft or spool 136 drivingly connects the LP turbine 130 to the LP compressor 122. The LP spool 136 may also be connected to a fan spool or shaft 138 of the fan section 106. In particular embodiments, the LP spool 136 may be connected directly to the fan spool 138 such as in a direct-drive configuration. In alternative configurations, the LP spool 136 may be connected to the fan spool 138 via a speed reduction device 137 such as a reduction gear gearbox in an indirect-drive or geared-drive configuration. Such speed reduction devices may be included between any suitable shafts/spools within engine 100 as desired or required.
As shown in FIG. 3, the fan section 106 includes a plurality of fan blades 140 that are coupled to and that extend radially outwardly from the fan spool 138. An annular fan casing or nacelle 142 circumferentially surrounds the fan section 106 and/or at least a portion of the gas turbine engine 104. It should be appreciated by those of ordinary skill in the art that the nacelle 142 may be configured to be supported relative to the gas turbine engine 104 by a plurality of circumferentially-spaced outlet guide vanes 144. Moreover, a downstream section 146 of the nacelle 142 (downstream of the guide vanes 144) may extend over an outer portion of the gas turbine engine 104 so as to define a bypass airflow passage 148 therebetween. The HP turbine 128 includes, in serial flow relationship, a first stage of stator vanes 154 (only one shown) axially spaced from turbine rotor blades 158 (only one shown) (also referred to as “turbine blades”) and a second stage of stator vanes 164 (only one shown) axially spaced from turbine rotor blades 168 (only one shown) (also referred to as “turbine blades”).
While not illustrated in FIG. 3, the high pressure (HP) compressor 124 may comprise a coating system as described herein. The coating system may reduce the rub ratio of the compressor blades and casing and thereby increase the life and efficiency of the compressor and engine.
FIG. 4 is a method of preparing a coating system in accordance with one embodiment of the present disclosure. In the embodiment illustrated in FIG. 4, the method of preparing a coated compressor casing 400, particularly a coated compressor casing configured for use with a compressor blade in a gas turbine engine, comprises the step of applying a coating system to a surface of a compressor casing 410. The coating system comprises a ceramic material and has a higher hardness than the hardness of the compressor blade with which the compressor casing will be used. The coating system may be applied by any suitable method as described herein. The method may comprise other treatments to the compressor casing between each application of coating to further improve blade wear. In some embodiments, a bond coating may be applied to the compressor casing to improve adhesion of the ceramic material while in certain embodiments, a bond coat may not be needed. The bond coat may help accommodate differences in coefficients of thermal expansion, to provide a rougher surface for adhesion, or combinations thereof.
While the present application is discussed in relation to compressor cases, the disclosure may be applied in other applications such as where a coating with a harder material may protect the underlying metal of a stator in a rubbing event (e.g., a rotor and a stator).

EXAMPLES

The feasibility of this concept has been demonstrated with experiments using blade and casing coupons. FIGS. 5a, 5b, and 5c illustrate the rubbing of a conventional compressor blade and casing. In particular, FIG. 5a illustrates the casing shoe and blade coupon after a rubbing event, FIG. 5b is another image of the casing shoe after a rubbing event, and FIG. 5c shows the topography of the casing shoe after a rubbing event. A nickel superalloy blade coupon and casing shoe was used to illustrate the rubbing experienced with conventional compressor systems. FIGS. 6a, 6b, and 6c illustrate the rubbing of a compressor blade and casing in accordance with one embodiment of the present invention. In particular, FIG. 6a illustrates the casing shoe and blade coupon after a rubbing event, FIG. 6b is another image of the casing shoe after a rubbing event, and FIG. 6c shows the topography of the casing shoe after a rubbing event. In this embodiment of the coating system, a nickel superalloy blade coupon was rubbed against a nickel superalloy casing shoe coated with aluminum oxide (Al₂O₃). FIGS. 7a and 7b compare the rubbing of a conventional compressor blade and casing to that of a compressor blade and casing in accordance with one embodiment of the present invention. Besides the addition of the aluminum oxide coating, all other parameters were kept constant in the two tests.
As shown in FIG. 5c , with the conventional compressor system, a scab formed along the casing due to material transfer from the blade tip to the casing. FIG. 5c illustrates the build-up of scab on the casing shoe. In sharp contrast, with the incorporation of the coating system, much less build-up was seen as less material from the blade tip was transferred to the casing. FIG. 6c illustrates the lack of scab build-up on the casing shoe when coated with the coating system. The coating system had a generally more jagged topography, which is illustrated in FIG. 6c ; however, as shown in this graph, a scab was not generally formed on the coating system surface.
The amount of force seen on the blade tip is generally less with the incorporation of the coating system. FIG. 7a shows the maximum force seen during one rubbing event. As shown in FIG. 7a , the normal forces for the coated casing were lower in comparison to the uncoated casing. Less resistance is seen from the blade tip, so the maximum force is less.
FIG. 7b illustrates the blade wear value obtained from the two experiments. The wear for the blade rubbing against the aluminum oxide coated casing was lower showing that the coating did reduce the rub ratio (about 15%). The blade coupons used in the experiment were flat blade coupons. In certain embodiments, the compressor blades have a curved body or airfoil (see e.g., FIG. 2b ). In such embodiments, the improvement in rub ratio is expected to be significantly more due to the effect on radial growth as discussed previously. In certain embodiments, the rub ratio is expected to be reduced by more than half with the incorporation of the coating system.
While the invention has been described in terms of one or more particular embodiments, it is apparent that other forms could be adopted by one skilled in the art. It is to be understood that the use of “comprising” in conjunction with the coating compositions described herein specifically discloses and includes the embodiments wherein the coating compositions “consist essentially of” the named components (i.e., contain the named components and no other components that significantly adversely affect the basic and novel features disclosed), and embodiments wherein the coating compositions “consist of” the named components (i.e., contain only the named components except for contaminants which are naturally and inevitably present in each of the named components).
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Claims

What is claimed is:

1. A coated compressor casing, the coated compressor casing comprising:

a compressor casing defining an inner surface of the compressor casing and comprising a base material, and

a coating system disposed along the inner surface of the compressor casing, wherein the compressor casing is configured to have a compressor blade positioned within the casing, and wherein the coating system comprises a ceramic material that is harder than the compressor blade.

2. The coated compressor casing according to claim 1, wherein the ceramic material comprises yttria stabilized zirconia, mullite, alumina, ceria, rare-earth zirconates, rare-earth oxides, metal-glass composites, zirconia stabilized with an oxide, silicate, chromium oxide, chrome carbide, or combinations thereof.

3. The coated compressor casing according to claim 1, wherein the coating system has a uniform thickness across the inner surface of the compressor casing.

4. The coated compressor casing according to claim 1, wherein the inner surface of the compressor casing is configured to be adjacent to a rotating compressor blade.

5. The coated compressor casing according to claim 1, wherein the coating system has a hardness about 10% to about 50% higher than a hardness of the compressor blade.

6. The coated compressor casing according to claim 1, wherein the coating system has a modulus about 10% to about 50% higher than a modulus of the compressor blade.

7. The coated compressor casing according to claim 1, wherein the coating system has a thickness of about 127 microns to about 254 microns.

8. The coated compressor casing according to claim 1, wherein the coating system includes a bond coat.

9. The coated compressor casing according to claim 1, wherein the coating system is non-abradable.

10. The coated compressor casing according to claim 1, wherein the compressor casing is configured to be positioned in a turbofan engine.

11. A gas turbine engine comprising:

a compressor comprising a compressor casing having an inner surface, wherein the compressor casing comprises a base material, and a compressor blade having a blade tip, wherein the compressor blade comprises a base material, and

a coating system disposed along the inner surface of the compressor casing, wherein the coating system has a higher hardness than a hardness of the compressor blade base material.

12. The system according to claim 11, wherein the coating system includes a bond coat.

13. The system according to claim 11, wherein the compressor blade base material is uncoated.

14. The system according to claim 11, wherein the coating system has a hardness about 10% to about 50% higher than a hardness of the compressor blade base material.

15. The system according to claim 11, wherein the compressor blade base material comprises a nickel superalloy.

16. The system according to claim 11, wherein the compressor blade comprises a curved airfoil.

17. A method of preparing a coated compressor casing, the method comprising:

forming a coating system comprising a ceramic material along an inner surface of a compressor casing, wherein the compressor casing is configured to have a compressor blade positioned within the casing, and wherein the ceramic material is harder than the compressor blade.

18. The method according to claim 16, wherein forming the coating system along the surface of the compressor casing comprises forming the ceramic material along a surface configured to be adjacent to a rotating compressor blade.

19. The method according to claim 16, wherein forming the coating system along the surface of the compressor casing comprises forming the ceramic material along the surface of the compressor casing to a thickness of about 127 microns to about 254 microns.

20. The method according to claim 18, wherein the coating system has a hardness of about 10% to about 50% higher than a hardness of the rotating compressor blade.