Adjusting airflow in turbine component by depositing overlay metallic coating

A method for adjusting the airflow in a turbine component having a plurality of airflow holes. The method comprises the step of depositing an overlay metallic coating on the surface of the turbine component in a manner such that at least some of the airflow holes are partially filled such that the volume of the partially filled airflow holes is changed so as to adjust the airflow through the turbine component. Also provided is a turbine component having a plurality of airflow holes, at least some of the airflow holes being partially filled with the overlay metallic coating to change the volume thereof so as to adjust the airflow through the turbine component.

BACKGROUND OF THE INVENTION

This invention relates to adjusting the airflow in turbine components, e.g., gas turbine components, having airflow holes by depositing an overlay metallic coating. This invention further relates to the turbine component whose airflow has been adjusted by depositing such an overlay metallic coating.

Airflow holes are formed in many gas turbine components, such as combustor liners, for transporting film air through the component to typically cool the component and to form a fluid barrier between the component and hot gases traveling in the main flowpath of the engine. In addition to flowing air over the hot surfaces, combustor liner cooling is also provided by a thin layer of cooling air along the inner, combustion side of the liner by directing airflow through an array of very small airflow holes formed in the liner, typically having a diameter of from about 0.02 to about 0.03 inches (from about 508 to about 762 microns). This film cooling is also induced through “nugget” holes that are typically in one row along the forward edge of the liner, and a much greater number of “transpiration” holes that are typically arranged in a plurality of rows across the entire surface of the liner to induce a more uniform airflow. These “transpiration” holes are typically angled or slanted from the “cold” or air supply side, to the “hot” or combustion side of the liner in a downstream direction, and typically have a circumferential orientation. See, for example, FIG. 2 of commonly assigned U.S. Pat. No. 6,655,149 (Farmer et al), issued Dec. 2, 2003. This arrangement, commonly referred to as “multi-hole film cooling,” reduces the overall liner cooling airflow requirement because the mass flow through the airflow holes dilutes the hot combustion gas next to the liner surface, with the flow through the airflow holes providing convective cooling of the liner walls. In addition to these smaller diameter “nugget” and “transpiration” airflow holes, larger diameter holes (commonly referred to as “dilution holes”) to introduce dilution air into the combustion zone are also provided at spaced intervals. See, for example, commonly assigned FIG. 2 of U.S. Pat. No. 6,408,629 (Harris et al), issued Jun. 25, 2002 and FIG. 3 of U.S. Pat. No. 6,655,149 (Farmer et al), issued Dec. 2, 2003.

The combustion side of these combustion liners can be coated with a thermal barrier coating to help protect the liner from thermal fatigue caused by the hot gas radiation and conduction to the combustor liner or liners. See commonly assigned U.S. Pat. No. 6,620,457 (Farmer et al), issued Sep. 16, 2003, which discloses a physical vapor deposition process to thermally insulate the combustor liner. After a period of service, these combustor liners are typically removed from the engine for replacement, repair, cleaning and/or removal of contaminants (e.g., oxidative deposits and residual combustion products), cracking and other thermally induced stresses that the liners have been subjected to. At least some of these cracks run through the various holes, including the smaller “nugget” and “transpiration” holes.

To repair a part or component, or a portion of a part or component, adequate cleaning of the surface thereof is typically required. During cleaning, the thermal barrier coating and contaminants are typically removed from the combustor liners by chemical and/or mechanical processes, for example, a conventional acid strip process. Repair of the combustor liners, and in particular repair of the cracks that typically form in the liners during operation and use, can cause at least some of the very small “nugget” and “transpiration” holes to become obstructed, occluded, plugged, or otherwise blocked which can then require chemical and/or mechanical processes to reopen the holes. The reopening of these airflow holes, as well as the chemical stripping process that removes the coatings and contaminants, can also remove some of metal substrate of the combustor liner where these holes are located, resulting in enlarging of these holes. These enlarged airflow holes can significantly and undesirably increase the airflow of these liners. Indeed, after several cycles of such cleaning and repair, the airflow can be increased to the extent that the combustor liner is no longer usable.

Problems in controlling airflow can also occur during the original manufacture of the combustor liners. To form typically thousands of these very small “nugget” and “transpiration” holes, and especially at a slanted angle, the liner is typically drilled using special machining processes such as laser beam or electrical discharge machining (EDM) processes. While a certain amount of control can be exercised over the pattern and size of the drilled holes, it is still extremely difficult to provide combustor liners that have a consistent pattern and size of holes such that the airflow rate is within desired limits. In addition, laser or EDM drilling forms a recast layer along the surface of the hole as it is generated. During subsequent stripping and cleaning of the combustor liner, this recast layer can also be removed, thus enlarging the hole and increasing the airflow.

A method for adjusting airflow in a turbine component having such airflow holes is disclosed in commonly assigned U.S. Pat. No. 6,408,610 (Caldwell et al), issued Jun. 25, 2002. This method involves depositing a thermal barrier coating by a physical vapor deposition (PVD) process (e.g., electron beam PVD) on the exterior and/or interior surfaces of the component to at least partially obstruct the airflow through the airflow holes. While this method provides the ability to adjust the airflow through the airflow holes, the physical vapor deposition apparatus, because of its size, may not provide the flexibility needed to use it with some turbine components. In addition, thermal barrier coatings typically comprise ceramic materials that may or may not adhere adequately to the metal surface of the liner over time without an overlay metallic bond coat layer. See U.S. Pat. No. 6,620,457, supra, which discloses spraying NiCrAlY as a bond coat layer110of from about 4 to about 10 mils on the inner combustion surface40of the combustor liner14before depositing the thermal barrier coating120.

Accordingly, it would be desirable to be able to economically and evenly adjust the airflow in turbine components (e.g., gas turbine components), such as combustor liners, having a large quantity of smaller diameter airflow holes where the air volume of the respective holes has been changed, and especially enlarged, during subsequent repair, replacement, cleaning, and/or removal processes. It would also be desirable to be able to adjust the airflow through turbine components after final manufacture without removing, or substantially removing, previously applied thermal barrier coatings. It would additionally be desirable to be able to adjust the airflow in turbine components, such as combustor liners, having airflow holes that, when originally manufactured (i.e., an OEM component), require airflow adjustment to be within acceptable limits. It would be further desirable to be able to have the flexibility to adjust the airflow in a variety of turbine components having airflow holes and especially a component having many thousands of airflow holes as, for example, a combustor liner.

BRIEF DESCRIPTION OF THE INVENTION

An embodiment of this invention relates to a method for adjusting the airflow in a turbine component (e.g., a gas turbine component) having a plurality of airflow holes. This method comprises the step of depositing an overlay metallic coating on the surface of the turbine component in a manner such that at least some of the airflow holes are partially filled such that the volume of the partially filled airflow holes is changed so as to adjust the airflow through the turbine component.

Another embodiment of this invention relates to a turbine component having a plurality of airflow holes, at least some of the airflow holes being partially filled with the overlay metallic coating to change the volume thereof so as to adjust the airflow through the turbine component.

The method and turbine component of this invention provides a number of benefits and advantages. These include but are not limited to: (1) providing the ability to repair or correct the airflow in turbine components, such as combustor liners, having a plurality of airflow holes so as to come within desired airflow specification or limits; (2) providing the ability to locally correct or change the airflow within specific regions or areas having such airflow holes as needed or desired, including controlling such airflow within desired specifications or limits; (3) recovering and repairing previously used or operational combustor liners and other turbine components having a plurality of airflow holes in a cost effective and efficient manner; (4) performing such recovery or repair of the turbine component without having to remove original or restored thermal barrier coatings or otherwise undesirably altering the physical or chemical properties, or physical dimensions of the component; and (5) recovering or salvaging turbine components manufactured with airflow greater than desired operational specifications or limits.

DETAILED DESCRIPTION OF THE INVENTION

As used herein, the term “turbine component having a plurality of airflow holes” refers to any turbine engine component (e.g., gas turbine engine component) having at least two and typically a multiplicity of airflow holes through which air is circulated, typically for the purpose of cooling the component. Representative examples of such turbine components include but are not limited to turbine shrouds, such as, for example, those disclosed in commonly assigned U.S. Pat. No. 5,127,793 (Walker et al), issued Jul. 7, 1992; U.S. Pat. No. 5,169,287 (Proctor et al), issued Dec. 8, 1992; U.S. Pat. No. 6,340,285 (Gonyou et al), issued Jan. 22, 2002); and U.S. Pat. No. 6,354,795 (White et al), issued Mar. 12, 2002, all of which are incorporated by reference; combustor liners, such as, for example, those disclosed in commonly assigned U.S. Pat. No. 6,408,629 (Harris et al), issued Jun. 25, 2002; U.S. Pat. No. 6,655,149 (Farmer et al), issued Dec. 2, 2003; and U.S. Pat. No. 6,620,457 (Farmer et al), issued Sep. 16, 2003, all of which are incorporated by reference, heat shields, vanes, impingement rings, nozzles, etc. This invention is particularly directed at combustor liners and other turbine components having a plurality, or more typically a multiplicity, of such airflow holes, and more particularly combustor liners having a plurality, or more typically a multiplicity, of transpiration cooling airflow holes to provide what is commonly referred to as “multi-hole film cooling.” See, for example, U.S. Pat. No. 6,408,629, supra; U.S. Pat. No. 6,655,149, supra; and U.S. Pat. No. 6,620,457, supra, all of which are incorporated by reference.

As used herein, the term “airflow hole(s)” refers to holes through which air passes or otherwise circulates to adjust, change or otherwise control the airflow in the turbine component, and which typically provide or induce, directly or indirectly, cooling of the turbine components, i.e., are cooling holes, such as nugget holes or transpiration holes.

As used herein, the term “overlay metallic coating” refers to any metallic coating that can be sprayed, applied or otherwise deposited by thermal spray deposition techniques and are often used to deposit additive metallic coatings on a metal substrate before thermal barrier coating materials are deposited. These overlay metallic coatings include MCrAlY coatings wherein M is a metal such as iron, nickel, platinum, cobalt or alloys thereof, etc. Typically, the overlay metallic coating used in this invention is a NiCrAlY coating.

As used herein, the term “thermal spray deposition methods” refers to any method for spraying, applying or otherwise depositing the overlay metallic coating that involves heating and typically at least partial or complete thermal melting of the overlay coating material and depositing of the heated/melted material, typically by entrainment in a heated gas stream, onto the metal substrate to be coated. Suitable thermal spray deposition techniques include plasma spray, such as air plasma spray (APS) and vacuum plasma spray (VPS), high velocity oxy-fuel (HVOF) spray, detonation spray, wire spray, etc., as well as combinations of these techniques. A particularly suitable thermal spray deposition technique for use herein is plasma spray. Suitable plasma spray techniques are well known to those skilled in the art. See, for example, Kirk-Othmer Encyclopedia of Chemical Technology, 3rd Ed., Vol. 15, page 255, and references noted therein, as well as U.S. Pat. No. 5,332,598 (Kawasaki et al), issued Jul. 26, 1994; U.S. Pat. No. 5,047,612 (Savkar et al) issued Sep. 10, 1991; and U.S. Pat. No. 4,741,286 (Itoh et al), issued May 3, 1998 (herein incorporated by reference) which are instructive in regard to various aspects of plasma spraying suitable for use herein. In general, typical plasma spray techniques involve the formation of a high-temperature plasma, which produces a thermal plume. The overlay metallic coating material, e.g., as a powder, is fed into the plume, and the plume is directed toward the base metal substrate to be coated. Various details of such plasma spray coating techniques are also well-known to those skilled in the art, including various relevant steps and process parameters such as cleaning of the surface of the base metal substrate prior to deposition, plasma spray parameters such as spray distances (gun-to-substrate), selection of the number of spray-passes, powder feed rates, particle velocity, torch power, plasma gas selection, oxidation control to adjust oxide stoichiometry, angle-of-deposition, post-treatment of the applied coating, and the like. Suitable plasma spray systems are described in, for example, U.S. Pat. No. 5,047,612 (Savkar et al), supra, which is incorporated by reference.

As used herein, the term “comprising” means various compositions, components, materials, layers, steps, etc., can be conjointly employed in this invention. Accordingly, the term “comprising” encompasses the more restrictive terms “consisting essentially of” and “consisting of.”

All amounts, parts, ratios and percentages used herein are by weight unless otherwise specified.

The various embodiments of the method and turbine component of this invention are further illustrated by reference to the drawings as described hereafter. Referring to the drawings,FIG. 1shows a gas turbine engine combustor indicated generally as10which comprises at least one component having a plurality, and more typically a multiplicity, of airflow holes for which the method of this invention is useful. As shown inFIG. 1, combustor10includes a cowl assembly indicated as14, an outer combustor liner indicated as22and an inner combustor liner indicated generally as26. The outer and inner liners22and26are disposed between an outer combustor casing30and an inner combustor casing34. Outer and inner liners22and26are generally annular in form about a centerline axis (not shown) and are radially spaced from each other to define a combustion chamber38therebetween. The outer liner22and the outer casing30form an outer passage therebetween indicated by arrow42, while inner liner assembly26and the inner casing34form an inner passage therebetween indicated by arrow46.

Cowl assembly14is mounted to the upstream ends of outer and inner liners22and26. An annular opening50is formed in the cowl assembly14for the introduction of compressed air into combustor10. The compressed air is supplied from a compressor (not shown) and is channeled in a direction generally indicated by arrows54ofFIG. 1. The compressed air passes principally through the opening50to support combustion and partially into the outer and inner passages42and46where it is used to at least partially provide cooling airflow for outer and inner liners22and26. Disposed between and interconnecting liners22and26near their respective upstream ends is an annular dome plate58. A plurality of circumferentially spaced swirler assemblies62are mounted in the dome plate58. Each swirler assembly62receives compressed air from the opening50and fuel from a corresponding fuel tube66. The fuel and air are swirled and mixed by swirler assemblies62, and the resulting fuel/air mixture is discharged into the combustion chamber38. It should be noted that althoughFIG. 1illustrates a single annular combustor, the method of this invention is equally applicable to any type of combustor, including double and triple annular combustors, that use liners with multi-hole film cooling. One such representative double annular combustor is shown inFIG. 2as having two rows of swirler assemblies62and corresponding fuel tubes66, as well as a centershield68.

The outer and inner liners22and26each comprise a single wall, metal shell having a generally annular and axially extending configuration. As shown inFIGS. 1 and 2, outer liner22has a hot or combustion surface or side70facing the hot combustion gases in the combustion chamber38and a cold or air supply surface or side74in contact with the relatively cooler air in outer passage42. Similarly, the inner liner26has a hot or combustion surface or side78facing the hot combustion gases in the combustion chamber38and a cold or air supply surface or side82in contact with the relatively cooler air in inner passage46. As shown particularly inFIG. 2, each of liners22and26also has respective forward bands indicated generally as84and86, as well as aft leaf seal lips indicated generally88and90, for securing liners22and26within combustor10.

As shown inFIG. 1, both outer and inner liners22and26have formed therein a large number of airflow holes. Referring toFIG. 3which shows, for illustrative purposes, inner liner26having a row of relatively small nugget cooling holes indicated generally as94at the forward end of the liner, several arranged arrays comprising a plurality of rows of smaller transpiration cooling holes indicated generally as98, and at least one row of larger dilution holes indicated generally as102that extend from the cold side82(or74in the case of outer liner22) to the hot side78(or70in the case of outer liner22). Dilution holes102introduce air into combustor chamber38, are generally far smaller in number than nugget holes94, and especially transpiration holes98, and have a cross-sectional area that is substantially greater than the cross-sectional area of holes94and98. The transpiration holes98are typically axially slanted from the cold side74,82to the respective hot side70,78in the downstream direction. Thus, air from the outer and inner passages42and46passing through the nugget and transpiration holes94and98is directed downstream so as to form a cooling film on the hot side70,78of each liner22and26.

Referring toFIG. 4, and viewed from the hot or combustion side78, a representative arrangement of nugget holes94, transpiration holes98and dilution holes102is shown for inner liner26. As shown inFIG. 4, and proceeding from the upstream end to the downstream end of liner26, a circumferentially extending row indicated as106of nugget holes94is formed at the upstream end of the liner, with many circumferentially extending rows of laterally spaced transpiration holes98indicated generally as110being formed and arrayed after nugget hole row106. The nugget holes94in row106and the transpiration holes98in rows110typically have a diameter of from about 0.01 to about 0.05 inches (from about 254 to about 1270 microns), more typically from about 0.02 to about 0.03 inches (from about 508 to about 762 microns), and a spacing therebetween of from about 3 to about 10 hole diameters.

As also shown inFIG. 4, a circumferentially extending row indicated as114of primary dilution holes102is located towards the upstream end of liner26, with a circumferentially extending row indicated as118of secondary dilution holes located downstream of the primary dilution holes. Primary dilution holes114and secondary dilution holes118typically have a diameter of from about 0.1 to about 0.75 inches (from about 2.5 to about 19 mm.), with the primary dilution holes114typically being larger in diameter than the secondary dilution holes118. Depending upon the diameter and spacing, there are typically from about 20 to about 60 primary dilution holes114and from about 20 to about 120 secondary dilution holes118formed in the liner.

Referring toFIG. 5, a cross-section of one such transpiration cooling hole formed in outer liner22that is indicated as122, and extends at a slanted angle downwardly towards the downstream end from cold side74to hot side70of liner22. (The cross-section for such a transpiration cooling hole formed in inner liner26would be similar, i.e., extending at a slanted angle upwardly towards the downstream end from cold side82to hot side78of liner26.) As shown inFIG. 5, hot side70typically has formed thereon an additive thermal barrier coating (TBC) indicated generally as126. To adjust the airflow through hole122, an overlay metallic coating indicated as130is deposited (e.g., sprayed) by an apparatus (not shown) using thermal spray deposition techniques on the cold or air supply side70. The particular spray angle indicated as134for depositing coating130can depend on a variety of factors, including the slanted angle of hole122, etc. For example, and as shown inFIG. 5, the spray angle134for depositing coating130can be the same or similar to that of the slanted angle of hole122, or can be different from that of the slanted angle of hole122.

Coating130can be deposited to any desired thickness, but is typically deposited to a thickness in the range of from about 1 to about 10 mils (from about 25 to about 254 microns), more typically from about 2 to about 7 mils (from about 51 to about 178 microns), i.e., is relatively thin. Depending upon the spray angle134at which the overlay metallic coating is deposited, a portion of coating130enters hole122at the cold side74end thereof, and partially fills hole122, as indicated by138. Because the lip end142of hole122is partially filled with the portion138of coating130, this effectively reduces the diameter of hole122, and thus partially obstructs or restricts the effective area or volume of hole122, thus reducing the volume of airflow that can enter such holes122from passage42(or from passage46adjacent to liner26) for a given pressure drop. As a result, the airflow from passage42(or46adjacent to liner26) into holes122is reduced. The coating130can be applied in such a manner so as to restrict the airflow through a larger or smaller quantity or number of holes122, so as to restrict the airflow through a specified, defined or otherwise selected pattern (e.g., row(s), array(s), area(s), etc.) of holes122, etc., or any combination thereof. In this manner, the airflow of liner22(or26) can be controlled and adjusted within desired limits or specifications depending upon the number of holes122that are partially filled with the portion138of coating130, the pattern of such holes122that are filled with the portion138of coating130, and/or the degree to which holes122are partially filled with the portion138of coating130.

In gas turbine components such as liners22and26, airflow hole sizes and airflow passage areas are typically controlled tightly with regard to both upper and lower airflow limits or specifications. In the combustion chamber38, insufficient cooling airflow or area could result in reduced component durability. In addition, excessive cooling airflow or area could result in detrimental effect on engine performance, operability and ignition. This condition is usually the result of at least some of the airflow holes122being larger than desired. This can occur as a result of two factors: (1) holes122being larger than the specified or desired size or diameter after original manufacture of the turbine component (i.e., an OEM component) such that desired airflow specifications or limits are exceeded; or (2) holes122becoming enlarged after original manufacture and during operation of the turbine component such that desired airflow specifications or limits are exceeded. The various embodiments of the method of this invention are intended to be applicable to turbine components (e.g., combustor liners22and26) that have larger than desired holes122due to either of factors 1 or 2.

The diameter of holes122can become enlarged (i.e., factor 2) for a variety of reasons, thus increasing the airflow through outer and inner liners22and26. For example, enlargement of holes122can result from the various processes that are used to perform repairs on outer and inner liners22and26, as well as the manner in which holes122are formed. This is exemplified byFIGS. 6 and 7.FIG. 6shows a hole122formed by a special machining process, such as laser or electrical discharge machine (EDM) drilling. As a result of laser or EDM drilling, a recast layer of material indicated generally as146is formed on the surface of hole122. During the process of cleaning a portion of a component, for example, TBC126, liners22,26are typically subjected to dry or wet grit blasting with an abrasive media such as alumina particles. In wet grit blasting, the abrasive media is suspended in a liquid, typically water, and then sprayed at high velocity to remove TBC126. As a result, and as shown inFIG. 7, recast layer146is typically partially or completely removed, thus increasing or enlarging the effective diameter of hole122from that indicated by double headed arrow150(seeFIG. 6), to that indicated by double headed arrow154(seeFIG. 7). By using an embodiment of the method of this invention previously described with respect toFIG. 5, diameter154of these enlarged holes122can be reduced to adjust and control airflow therethrough.

An embodiment of the method of this invention for adjusting and controlling the airflow through holes122, and thus adjusting and controlling the airflow through liners22and26, includes the steps of: (a) initially determining the amount of airflow through holes122prior to applying coating130; (b) applying coating130to the cold side74,82of liner22and/or26, as previously described; (c) determining the amount of airflow through holes122after coating130has been applied; and (d) repeating steps (b) and (c) as needed until the desired degree of airflow is achieved. An embodiment of such a method can be carried out by appropriate modification of the technique disclosed in commonly assigned U.S. Pat. No. 6,408,610 (Caldwell et al), issued Jun. 28, 2002, which is incorporated by reference. The Caldwell et al method, as modified for use in the embodiment of the method of this invention, includes the steps of: (1) developing a predetermined pressure drop across holes122prior to applying coating130; (2) calculating the airflow through holes122resulting from the predetermined pressure; (3) depositing a selected thickness of coating130based on the measured airflow through holes122so that airflow through holes122after depositing the coating130is within a preselected range of airflows. The minimum airflow of the preselected range of desired airflows is typically selected to provide sufficient airflow through holes122to maintain liners22and26below a selected maximum temperature during engine operation. This maximum temperature is calculated to provide an environment in which liners22and26life requirements will be met. The maximum airflow of the range is typically selected to ensure sufficient airflow through other components within the gas turbine engine to maintain the other components below maximum temperatures at which their respective life requirements are met. Local airflow changes could also cause a detrimental effect on component life as a result of excessive gradient temperatures and can also be corrected according to this embodiment of this invention.

In using the Caldwell et al method, as modified for use in the method of this invention (see FIG. 3 of Caldwell et al patent), the turbine component, such as combustor liner22,26is checked by a pressure flow stand, before and after coating130is deposited on cold side74,82. After pressurization, and because of the difference in pressure between the upstream and downstream ends of the pressure flow stand, a pressure drop develops across holes122of the combustor liner22,26. Because the upstream and downstream pressures are known, airflow through holes122can be calculated. For a turbine component or portion thereof, this airflow is compared to a preselected range of desired airflows. If the airflow is within the preselected range of desired airflows and the combustor liner22,26otherwise meets desired component specifications or limits, it is then ready (i.e., acceptable) for use.

If, however, the airflow is above the preselected range of desired airflows, additional layers of coating130are then deposited on the combustor liner22,26. The coated combustor liner22,26is again measured for airflow through holes122as before and the measured airflow is again compared to the preselected range. In one embodiment of this alternative method, these steps can be repeated until the measured airflow is within the preselected range of desired airflows. Typically, the step of depositing coating130need be repeated no more than once in this alternative method.

The minimum airflow of the preselected range of desired airflows is selected to provide sufficient airflow through holes122to maintain the turbine component (e.g., combustor10) below a selected maximum temperature during engine operation. This maximum temperature is calculated to provide an environment in which component life requirements will be met. The maximum airflow of the range is selected to ensure that there is sufficient airflow through other components of the gas turbine engine and that the other components are maintained below maximum temperatures at which their respective life requirements are met.

While specific embodiments of this invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of this invention as defined in the appended claims.