Patent Abstract:
This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Full Description:
FIELD OF THE INVENTION 
     This invention is directed generally to abradable coating systems, and more particularly to abradable coating systems useful for creating individualized seals between turbine blades and corresponding ring segment shrouds. 
     BACKGROUND 
     Axial gas turbines typically contain rows of turbine blades, referred to as stages, coupled to disks that rotate on a rotor assembly. The turbine blades extend radially and terminate in turbine blade tips. Ring seal segments are positioned radially outward from the turbine blade tips, but in close proximity to the tips of the turbine blades to limit gases from passing through the gap created between the turbine blade tips and the inner surfaces of the ring seal segments. The gaps between the turbine blade tips and the ring seal segments are designed to be as small as possible between the blade tips and the surrounding segment because the larger that gap, the more inefficient the turbine engine. 
     The size of the gap between the tips of the turbine blades and the ring seal segments must account for the turbine blades and the ring seal segments being formed from materials having different coefficients of thermal expansion. As a turbine engine begins to heat up during startup procedures, the length of the turbine blades increases radially outward while the ring seal segments move radially outward as well. The gap may change during the thermal growth. Thus, the gap is sized such that at steady state operating conditions in which the turbine blades are heated to an operating temperature, the gap is a small as possible without risking significant damage from the tips contacting the ring seal segments. However, as the gap is reduced, the incidences of rubbing between the turbine blade tips and the outer ring seal increases. 
     Attempts have been made to minimize the clearance gap to improve efficiency while avoiding excessive wear on the turbine blade tips. For instance, some conventional turbine engines include thermal barrier coatings (TBCs) on the ring seal segments that are designed to abrade when contacted by the blade tips. The TBCs also insulate the underlying turbine components from the hot gases present during operation, which may be approximately 2500 degrees Fahrenheit. Use of the TBCs can keep the underlying turbine component generally at temperature of less than approximately 1800 degrees Fahrenheit. 
     While the gap between the tips of the turbine blade and the ring seal segments may be designed to enable smooth startup from a cold engine, problems are typically encountered during a warm restart. In particular, a warm restart occurs when a turbine engine running at steady state operating temperatures is shut down, allowed to cool for two to three hours, and then restarted. During the restart, the turbine blade tips often contact the abradable coating on the ring seal segments because during the shut down period turbine disks remain hot and thermally expanded radially, while the thermally insulated turbine shroud ring has cooled and retracted somewhat, thereby reducing the gap. With the gap reduced, the turbine blade tips often contact the abradable coating. 
     Abradable coatings are designed such that when contacted by a turbine blade, a portion of the coating will break away to prevent damage to the turbine blade. A problem that is widespread with abradable coatings is that the coatings generally sinter after exposure to turbine engine operating temperatures of about 2,500 degrees Fahrenheit after about 50 to 100 hours. Sintering of the abradable coating significantly reduces the abradable coatings ability to shear when contacted by tips of turbine blades. For instance, as shown in  FIG. 1 , abradable coatings greatly lose their ability to shear when contacted by tips of turbine blades with greater and greater exposure to turbine engine operating temperatures. In particular,  FIG. 1  illustrates the impact of sintering on the abradability of a conventional abradable coating, 79% dense 8YSZ, 8YSZ refers to 8 weight percent yttria stabilized zirconia, which is a common TBC in both aero and IGT engines. The coating exhibited an abradability volume wear ratio (VWR) of 34 (coating wear/blade wear, where larger values are better) prior to exposure to elevated temperatures. After the same coating was exposed to approximately 2000 degrees Fahrenheit for 200 hours, the VWR declined to nine. The VWR declined to seven when exposed to approximately 2200 degrees Fahrenheit for 200 hours. Finally, the VWR was two after exposure to approximately 2375 degrees Fahrenheit for 200 hours. Thus, the usefulness of an abradable coating is nearly negated once sintered. Therefore, a need exists for an abradable coating system capable of shearing when contacted by turbine blade tips even if a portion of the abradable coating has sintered. 
     SUMMARY OF THE INVENTION 
     This invention relates to an abradable coating system for use in axial turbine engines. In particular, the abradable coating system may include an abradable coating formed from a plurality of columns that limit sintering of the coating to outermost portions of the coating, thereby enabling the columns forming the abradable coating to shear off near the base of the columns. Shearing in the unsintered area near the base of the column creates for a smooth break with reduced losses relative to the prior art. 
     The abradable coating system may include an abradable coating attachable to an outer surface of a turbine component, such as but not limited to, a ring seal segment, also known as a blade outer air seal (BOAS). The abradable coating may be formed from any ceramic powder capable of being thermally sprayed, such as, but not limited to, 8YSZ, compositions of ceria-stabilized zirconia, materials that are capable of withstanding higher temperatures and are not based on yttria, ceria or zirconia, and other appropriate materials. The abradable coating system may also include a forming matrix supported on the outer surface of the turbine component. The forming matrix may be formed from a plurality of walls that are coupled together to form a plurality of cells having at least one opening opposite the outer surface for receiving the abradable coating. The forming matrix may be formed from a material having a melting point less than about 2,500 degrees Fahrenheit such that the forming matrix melts during operation of a turbine engine in which the coating system is positioned, thereby leaving the first abradable coating attached to the turbine component and forming a plurality of columns from the abradable coating. The forming matrix may be a fugitive material such as, but not limited to plastics, molybdenum, and other appropriate materials. The choice of fugitive materials is based more upon convenience than on composition, since any material that can be formed into the desired “forming matrix” shape (herein termed “honeycomb”) that will burn off at turbine temperatures will be a suitable choice. Polymer materials such as common plastics may be used and, unless very high temperature thermal spraying is required, have been shown to function well. For higher temperature spray requirements, metal “honeycomb” or metalized plastics may be used. Molybdenum and moly alloys are suitable choices since they tend to form volatile oxides rather than melting when heated in oxidizing atmospheres. Fugitive materials are materials that occupy a physical area and burn off when exposed to temperatures above a threshold temperature, leaving a void absent of the fugitive materials where the materials once existed. 
     The forming matrix may have a wall thickness of less than about five mils (0.005 inches), with typical thicknesses being approximately one mil. The cells of the forming matrix may have a cross-sectional area in a plane generally aligned with the outer surface of the turbine component that is less than about two mm 2  and typically will be less than one mm 2 . At least one cell of the plurality of cells forming the forming matrix may have a cross-sectional shape that is selected from the group consisting of a circle, an ellipse, a triangle, a rectangle, a hexagon, and a diamond. 
     The abradable coating system may also include a second coating deposited between the first abradable coating and the outer surface of a turbine component and below the first abradable coating such that said forming matrix is attached to an outer surface of the second coating. The second coating may be a thermal barrier coating or a bond coating, or other appropriate material. In one embodiment, a bond coating may be deposited on the outer surface of the turbine component, and the second coating may be a thermal coating deposited on the bond coating. 
     The abradable coating system may include an alarm system for identifying whether a turbine blade tip has contacted the first abradable coating. The alarm system may be formed from a metalized layer positioned between an outer surface of the turbine component and a tip of the columns of the abradable coating, wherein the metalized layer may be coupled to the alarm system that is usable for actuating an alarm when a tip of a turbine blade contacts the metalized layer indicating the tip has worn through a predetermined distance of the abradable coating. The abradable coating system may also include a temperature sensor on the first abradable coating. The temperature sensor may be formed from at least two metals. 
     During use, a turbine engine is ramped up to a steady state operating temperature. At the steady state operating condition, the abradable coating system is typically exposed to gases having temperatures of about 2,500 degrees Fahrenheit. Exposure of the forming matrix to these gases causes the forming matrix to burn, thereby leaving the inter-columnar channels and forming columns of the abradable coating. The width of the inter-columnar channels  46  may be between about 0.25 mm and about 1.5 mm. After prolonged exposure to the exhaust gases, the tips of the columns of the abradable coating may become sintered; however, the bases of the columns are either unsintered or sintered to a much lesser degree than the tips. Thus, should a tip of a turbine blade contact the abradable coating, such as during a warm restart, the columns of the abradable coating may shear at the base, thereby breaking free and protecting the tip of the turbine blade from damage. The columns may also provide the abradable coating with an increased resistance to spallation due to the inter-columnar channels that enable the columns to expand. 
     An advantage of this invention is that the columnar structure of the abradable coating system allows columns to break near the base, resulting in reduced blade wear compared to the conventional systems. This configuration is particularly advantageous after the tips of the columns of the abradable coating become sintered, in part, because the base of the columns may not be sintered. 
     Another advantage of the invention is that the abradable coating reduces or eliminates thermal barrier coating (TBC) spallation due to thermal cycling since the columnar structure naturally relieves thermally-induced strains caused by the contraction and expansion of the underlying metal substrate. 
     Yet another advantage of the invention is that the abradable coating may include an alarm system and thermocouples for monitoring the performance and condition of the abradable coating system and the turbine engine. 
     These and other embodiments are described in more detail below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The accompanying drawings, which are incorporated herein and form a part of the specification, illustrate embodiments of the presently disclosed invention and, together with the description, disclose the principles of the invention. 
         FIG. 1  is a chart showing the impact of high temperatures on the abradability of a conventional abradable coating. 
         FIG. 2  is a cross-sectional view of turbine engine with a rotor assembly and including aspects of this invention. 
         FIG. 3  is a detailed view taken at detail  3 - 3  in  FIG. 2  of the abradable coating system. 
         FIG. 4  is a cross-sectional view of the abradable coating system of this invention with the forming matrix intact. 
         FIG. 5  is a cross-sectional view of the abradable coating system of this invention after the forming matrix has been burned off due to exposure to turbine engine steady state operating temperatures. 
         FIG. 6  is a cross-sectional view of a tip of a turbine blade contacting and shearing the abradable coating at a base of a column of abradable material forming the abradable coating. 
         FIG. 7  is a cross-sectional view of an alternative embodiment of the abradable coating system of this invention with the forming matrix intact. 
         FIG. 8  is a cross-sectional view of the alternative embodiment of the abradable coating system shown in  FIG. 7  after the forming matrix has been burned off due to exposure to turbine engine steady state operating temperatures. 
         FIG. 9  is a perspective view of portion of a forming matrix of this invention. 
         FIG. 10  is a perspective view of a portion of a forming matrix of this invention having an alternative configuration. 
         FIG. 11  is a perspective view of a portion of a forming matrix of this invention having an alternative configuration. 
         FIG. 12  is a perspective view of a portion of a forming matrix of this invention having an alternative configuration. 
         FIG. 13  is a perspective view of a portion of a forming matrix of this invention having an alternative configuration. 
         FIG. 14  is a perspective view of a portion of a forming matrix of this invention having an alternative configuration. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     As shown in  FIGS. 2-14 , this invention is directed to an abradable coating system  10  for use in turbine engines  12 . In particular, the abradable coating system  10  may include an abradable coating  14  formed from a plurality of columns  16  that limit sintering of the coating  14  to outermost portions of the coating  14 , thereby enabling the columns  16  forming the abradable coating  14  to shear off near the base  18  of the columns  16 . The abradable coating  14  may be applied to an outer surface  17  of a turbine component  19 , such as, but not limited to, one or more turbine ring seal segments  20 . The turbine ring seal segments  20  may be positioned radially outward from tips  22  of turbine blades  24  to create a seal between the turbine blades  24  and the surrounding ring seal segments  20 . The abradable coating system  10  may be formed an abradable material and may have a columnar configuration that prevents bases  18  of the columns  16  from sintering, thereby enabling the columns  16  to break at the base  18  if struck by a turbine blade  24 . The abradable columnar coating material be composed of a substance that is abradable and thermally resistant, such as, but not limited to 8YSZ, ceria stabilized zirconia, and other coatings not based on yttria, ceria, or zirconia. The abradable coating system  10  may reduce blade wear and spalling of the abradable coating  14  in comparison with conventional coatings. 
     As shown in  FIG. 2 , the abradable coating system  10  may be used together with a turbine engine  12 . For instance, the turbine engine  12  may include a plurality of turbine blades  12  extending radially outward from a rotor assembly  26  and positioned into a plurality of rows forming stages. The turbine blades  12  may be formed from a material capable of withstanding the high temperature exhaust gases in the turbine engine  12 . Stationary turbine vanes  28  may extend radially inward from an outer casing and be positioned in rows between adjacent turbine vanes  28 . A plurality of ring seal segments  20  may be positioned radially outward from the tips  22  of the turbine blades  24 . The ring seal segments  20  may be offset radially from the tips  22  of the turbine blades  24  forming a gap  32  such that the turbine blades  24  may rotate without contacting the ring seal segments  20 . 
     The abradable coating system  10  may include an abradable coating  14  applied to an outer surface  17  of a turbine component  19 , which may be, but is not limited to, ring seal segments  20 . The abradable coating  14  is configured to minimize the gap  32  while preventing excessive wear and damage to the turbine blade tip  22  that may occur while the turbine components are in different states of expansion, such as during a warm restart. The abradable coating system  14  may be formed from a forming matrix  36 , as shown in  FIGS. 9-14 , covered with the abradable coating  14 . The forming matrix  36  may be formed from a plurality of walls  38  that are coupled together to form a plurality of cells  40  having at least one opening  42  opposite to the ring seal segment  20 . The opening  42  enables the abradable coating  14  to be applied into the cells  40  during the formation process. The cells  40  may have any appropriate configuration, such as, but not limited to, a hexagon, as shown in  FIG. 9 , an ellipse, as shown in  FIG. 10 , a circle, as shown in  FIG. 11 , a triangle, as shown in  FIG. 12 , a rectangle, as shown in  FIG. 13 , a diamond, as shown in  FIG. 14 , and other appropriate configurations. A single side wall  38  may be used to form a portion of one or more adjacent cells  40 . 
     The forming matrix  36  may be made from any material having a melting point less than a steady state operating temperature of a turbine engine  12 . In at least one embodiment, a steady state operating temperature of the turbine engine  12  may be about 2,500 degrees Fahrenheit. In at least one embodiment, the forming matrix  36  may be formed from materials such as, but not limited to, a material having a melting point less than a steady state operating temperature of a turbine engine or a fugitive material such as plastics, molybdenum, and other appropriate materials. A fugitive material is a material that occupies a physical area and burns off when exposed to temperatures above a threshold temperature, leaving a void absent of the fugitive material where the material once existed. In the abradable coating system  10 , it is preferred that the material forming the forming matrix  36  have a melting point less than the steady state operating temperature of the turbine engine  12 , which may be about 2,500 degrees Fahrenheit. 
     The forming matrix  36  may have any appropriate height. In at least one embodiment, the height of the cells  40  forming the forming matrix  36  as indicated by distance A in  FIGS. 4 and 8  may be between about 0.005 and about 0.060 inches, and may be between about 0.020 and about 0.040 inches. The height of the cells  40  may vary depending on the gap  32  desired in a particular turbine engine  12 . In at least one embodiment, a width of the cells, as indicated by distance B in  FIG. 9  may be between about 0.125 millimeters and about 1.5 millimeters. 
     The abradable coating system  10  may be formed by positioning the forming matrix  36  onto a ring seal segment  20 . The forming matrix  36  may be attached directly to an outer surface  17  of the ring seal segment  20  or to one or more bond coatings  44  positioned between the outer surface  17  of the ring seal segment  20  and the forming matrix  36 . The bond coatings  44  may be formed from materials such as, but not limited to, powders such as CoCrAlY, NiCrAlY, CoNiCrAlY, and rhenium containing versions and other appropriate materials. In another embodiment, as shown in  FIG. 7 , the abradable coating  14  may not be formed from columns  16  across the entire thickness. Rather, an abradable coating intermediate layer  48  may be applied to the ring seal segment  20  and then, the forming matrix  36  and abradable coating  14  may be applied to an outer surface of the abradable coating intermediate layer  48 . The abradable coating intermediate layer  48  may provide additional thermal protection for the underlying turbine blade  24 . In addition, since the inter-columnar channel  46  does not extend to the bond coating  44 , overfracture may be limited to the intersection of the abradable coating intermediate layer  48  and the abradable coating  14  formed from the columns  16 , as shown in  FIG. 8 . The abradable coating intermediate layer  48  may also be a thermal barrier coating (TBC), such as, but not limited to, 8YSZ, ceria stabilized zirconia, and other coating compositions not based on yttria, ceria, or zirconia. 
     During use, a turbine engine  12  is ramped up to a steady state operating temperature. At the steady state operating condition, the abradable coating system  10  is typically exposed to gases having temperatures of about 2,500 degrees Fahrenheit. Exposure of the forming matrix  36  to these gases causes the forming matrix  36  to burn or melt, thereby leaving the inter-columnar channels  46  and forming columns  16  of the abradable coating  14 . The width of the inter-columnar channels  46  may be between about 0.5 mils and about 5.0 mils. After prolonged exposure to the exhaust gases, the tips  50  of the columns  16  of the abradable coating  14  may become sintered; however, the bases  18  of the columns  16  do not sinter. Thus, should a tip  22  of a turbine blade  24  contact the abradable coating  14 , such as during a warm restart, the columns  16  of the abradable coating  14  may shear at the base  18 , thereby protecting the tip  22  of the turbine blade  24  from damage. The columns  16  may also provide the abradable coating  14  with an increased resistance to spallation due to the inter-columnar channels  46  that enable the columns  16  to expand. In addition, the inter-columnar channels  46  may relieve stress on the abradable coating  14  that is imparted onto the abradable coating  14  from thermal expansion of the turbine blade  24 . 
     The cells  40  of the forming matrix  36  may be configured to minimize the amount of force exerted on the blade tip  22  when contacting the abradable coating  14  during operation of the turbine engine  12 , yet create as small a gap  32  as possible within safety parameters between the blade tips  22  and the abradable coating  14  on the ring seal segment  20 . In particular, the abradable coating  14  may be formed with columns  16  having relatively small cross-sectional areas, such as less than about two mm 2  and, in one embodiment between about two mm 2  and about one mm 2 , thereby resulting in a relatively high number of columns  16  per unit area. The cross-sectional area may be generally aligned with the outer surface  17  of the turbine component  19 . This configuration may create a more efficient seal between the tips  22  of the turbine blades  24  and the abradable coating  14  on the ring seal segments  20  because the amount of unnecessary columns broken off at the outer edges of the seal will be reduced. In addition, as the cross-sectional area of the columns  16  decreases, the amount of force exerted on the blade tips  22  during the abrasion of the blade tips  22  with the abradable coating  14  decreases. 
     In another embodiment, the abradable coating system  10  may include an alarm system  54 , as shown in  FIG. 8 , for indicating when a turbine blade tip  22  contacts the abradable coating  14 . In at least one embodiment, the alarm system  54  may be formed from a metallic layer  56 , such as, but not limited to, a thin metal foil. The alarm system  54  may be configured such that when a tip  22  of a turbine blade  24  contacts and cuts the metallic foil, a circuit is broken and an alarm is actuated. The metallic layer  56  may be deposited in a calibrated manner such that the alarm is triggered when the columnar abradable coating layer is worn to a specified depth by placing the metal layer  56  between the tip  50  and the base  18  of the column  16 . 
     In another embodiment, as shown in  FIG. 8 , the abradable coating system  10  may include a temperature sensor  58 . For instance, the temperature sensor  58  may be formed from two or more metals used to generate an EMF to determine temperature. 
     The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.

Technology Classification (CPC): 8