Vane with enhanced heat transfer

A vane cluster has a coated metallic substrate. The cluster includes a platform and a shroud. At least first and second airfoils extend between an outer face of the platform and an inner face of the shroud. Each airfoil has a pressure side and a suction side. The pressure side of the first airfoil faces the suction side of the second airfoil. The cluster includes a cooling passageway system including one or more first feed passageways in the first airfoil and one or more second feed passageways in the second airfoil. At least a first side selected from the pressure side of the first airfoil and the suction side of the second airfoil includes a first region with a local thinning or gap in the coating. Along the first side, the cooling passageway system includes means for locally cooling said first region.

BACKGROUND OF THE INVENTION

The invention relates to cooling of high temperature components. More particularly, the invention relates to coated gas turbine engine vane clusters.

In the aerospace industry, a well-developed art exists regarding the cooling of components such as gas turbine engine components. Exemplary components are gas turbine engine blades and vanes. Exemplary blades and vanes airfoils are cooled by airflow directed through the airfoil to be discharged from cooling holes in the airfoil surface. Also, there may be cooling holes along the vane shroud or vane or blade platform. The cooling mechanisms may include both direct cooling as the airflow passes through the component and film cooling after the airflow has been discharged from the component but passes downstream close to the component exterior surface.

By way of example, cooled vanes are found in U.S. Pat. Nos. 5,413,458 and 5,344,283 and U.S. Application Publication 20050135923. Vane clustering may have several advantages. The reduced engine part count may ease manufacturing and reduce weight. The reduction in the number of platform and shroud gaps (e.g., a halving with doublets) may have performance advantages. First, intergap leakage may correspondingly be reduced. Second, diversion of cooling air to cool gap seals may also be reduced.

Exemplary cooled vanes are formed by an investment casting process. A sacrificial material (e.g., wax) is molded over one or more cores (e.g., refractory metal cores and/or ceramic cores) to form a pattern. The pattern is shelled. The shell is dewaxed. Alloy (e.g., nickel- or cobalt-based superalloy) is cast in the shell. The shell and core(s) may be destructively removed (e.g., by mechanical means and chemical means, respectively). The casting may be finish machined (including surface machining and drilling of holes/passageways). The casting may be coated with a thermal and/or erosion-resistant coating.

Exemplary thermal barrier coatings include two-layer thermal barrier coating systems An exemplary system includes an NiCoCrAlY bond coat (e.g., low pressure plasma sprayed (LPPS)) and a yttria-stabilized zirconia (YSZ) barrier coat (e.g., air plasma sprayed (APS) or electron beam physical vapor deposited (EBPVD)). With vane clusters (e.g., doublets), each airfoil may interfere with the application of the coating to the adjacent airfoil(s). This may cause local thinning of the applied coating or even gaps.

SUMMARY OF THE INVENTION

One aspect of the invention involves a vane cluster having a coated metallic substrate. The cluster includes a platform and a shroud. At least first and second airfoils extend between an outer face of the platform and an inner face of the shroud. Each airfoil has a pressure side and a suction side. The pressure side of the first airfoil faces the suction side of the second airfoil. The cluster includes a cooling passageway system including one or more first feed passageways in the first airfoil and one or more second feed passageways in the second airfoil. At least a first side selected from the pressure side of the first airfoil and the suction side of the second airfoil includes a first region with a local thinning or gap in the coating. Along the first side, the cooling passageway system includes means for locally cooling said first region.

In various implementations, the means may be provided in a reengineering of an existing cluster configuration. The means may include an in-wall circuit. This circuit may direct flow from the shroud to the platform.

DETAILED DESCRIPTION

FIG. 1shows a gas turbine engine20having a central longitudinal axis500and extending from an upstream inlet22to a downstream outlet24. From upstream to downstream, the engine may have a number of sections along a core flowpath. From upstream to downstream, the sections may include a low speed/pressure compressor (LPC)30, a high speed/pressure compressor (HPC)32, a combustor34, a high speed/pressure turbine (HPT)36, a low speed/pressure turbine (LPT)38, an augmentor40, and an exhaust duct/nozzle42. Each of the compressor and turbine sections may include a number of blade stages interspersed with a number of vane stages. The blades of the LPC and LPT are mounted on a low speed spool for rotation about the axis500. The blades of the HPC and HPT are mounted on a high speed spool for such rotation.

As is discussed in further detail below, one or more of the vane stages may be formed as a cluster ring. For example, a second vane stage50of the HPT36is schematically shown inFIG. 1.FIG. 2shows further details of the exemplary vane stage50. The ring includes an inboard platform52and an outboard shroud54. A circumferential array of airfoils (discussed below) span between the platform and shroud. As is discussed in further detail below, the ring may be segmented into a plurality of separately-formed clusters interlocked at the platforms by a structural ring56and at the shrouds by an engine case.

FIGS. 3 and 4show an exemplary two-airfoil cluster (doublet)60. Each exemplary cluster includes a first airfoil62and a second airfoil64. Each of the airfoils extends from an associated inboard end66at a platform segment68to an associated outboard end70(FIG. 4) at a shroud segment72. The exemplary platform segment has an outboard surface74along the inboard extreme of the core flowpath. The shroud segment has an inboard surface76along an outboard extreme of the core flowpath.

An underside80of the platform segment may include features for mounting each platform segment to its adjacent segments (e.g., by bolting to the ring56). The platform segment has a forward/upstream end82, a rear/downstream end84, and first and second circumferential ends or matefaces86and88. Similarly, the shroud segment72has an upstream end92, a downstream end94, and first and second circumferential ends96and98. Each of the platform circumferential ends86and88and a shroud circumferential ends96and98may include a groove or channel108for receiving a seal (not shown). A given such seal spans the gap between the adjacent grooves of each adjacent pair of clusters.

The cluster60has cooling passageways. An exemplary passageway network may include one or more inlet ports.FIG. 3shows exemplary inlet ports110,111,112,113,115,116,117, and118(discussed below) in the shroud segment72. The inlet ports direct cooling air (e.g., bleed air) through one or more spanwise passageway segments in the airfoils62and64. Some of this airflow may exit cooling holes (discussed below) along the airfoils. In the exemplary doublet, a majority of the mass flow of air is discharged thought one or more outlets in the underside of the platform68.FIG. 4shows exemplary outlets120,121, and122. The air discharged through the outlets120-122may pass downstream to the adjacent blade stage to, in turn, pass through cooling passageways of those blades to cool the blades.

Some of the airflow, however, may be directed to exit the platform through one or more cooling outlet holes (e.g., along the platform outboard surface and the platform circumferential ends).

FIG. 5is a sectional view of the airfoils of a baseline version of a cluster from which the inventive clusters may represent reengineerings. The first airfoil62is shown having a leading edge140, a trailing edge142, a pressure side144, and a suction side146. Pressure and suction side walls are shown as148and149, respectively. Similarly, the second airfoil64has a leading edge150, a trailing edge152, a pressure side154, a suction side156, a pressure side wall158, and a suction side wall159. The airfoils also have passageways described below.

After casting, a coating is applied along the airfoils. Exemplary coating techniques are line-of-sight spray techniques (e.g., air plasma spray (APS) and electron beam physical vapor deposition (EBPVD)). Advantageous coating applications are achieved when the spray direction is near normal to the surface being coated. For the first airfoil suction side146and the second airfoil pressure side154, essentially normal line-of-sight flow access is available. However, along portions of the first airfoil pressure side144and second airfoil suction side156the other airfoil will block normal line-of-sight access. This blocking/occlusion mandates off-normal application with attendant reduction in coating thickness.

FIG. 5shows series of line-of-sight spray directions510positioned at boundaries of occlusion by the airfoils.FIG. 5also shows a local surface normal520. Along a leading region160of the first airfoil pressure side, there is essentially normal or near-normal line-of-sight access. Thus, along this region160, the coating161(FIGS. 12&13) is full thickness. Downstream thereof, the off-normal angle θ increases. There may be progressive degradation of coating thickness producing a local thinning163or gap165. For example, in a region162to an angle θ of about 30°, the coating may be deemed marginal. In a region164downstream thereof, and with greater θ, the coating may be deemed poor.

Similarly, along a trailing region168of the second airfoil suction side156, the coating may be full-thickness. Along a region170thereahead, the coating may be marginal. Along a region172yet thereahead, the coating may be poor. Along a region174yet thereahead, the coating may be marginal. Along a leading region176, the coating may be full. The exact distribution of coating quality will be highly dependent upon the particular cluster geometry. The presence of regions of relatively thin coating may locally increase thermal damage. In addition to being affected by coating thickness, the locations of possible thermal damage are influenced by the locations of aerodynamic heating. Thus, a combination of high local aerodynamic heating and local coating thinning163and/or gap165is disadvantageous. In such regions, it is desirable to add supplemental cooling.

One possible avenue for supplemental cooling would be to add outlets from the existing passageways to the airfoil surface (e.g., film cooling holes). However, the dilution associated with such discharge of air would impact the thermodynamic performance of the engine and counter the advantage that doublets have in reduced intergap air discharge relative to singlets. Furthermore, discharge along the suction side affects aerodynamic performance of the airfoil particularly significantly, thereby impeding turbine performance.

FIG. 6shows a reengineered cluster (e.g., reengineered from theFIG. 5baseline) to add supplemental wall cooling (e.g., via adding in-wall circuits). In the exemplary cluster60, each of the airfoils includes a streamwise array of spanwise-elongate passageway legs: a leading edge feed cavity200; a first through-flow leg202; a second through-flow leg204; a third through-flow leg206; and a trailing edge feed cavity208. In the exemplary cluster60, the leading edge cavity200has a closed inboard end and discharges air through spanwise arrays of leading edge outlet holes210. Similarly, the cavity208may discharge through an array of trailing edge outlet holes (or a slot)212. The through-flow legs discharge through the associated platform outlets to a plenum (not shown) for feeding blade cooling.

The basic arrangement of such passageways may be preserved in the reengineering. However, local wall thickening to accommodate added passageways may correspondingly narrow the adjacent legs/cavities. The exemplary reengineering adds cooling passageways214,216, and218in the suction side wall159of the second airfoil. To permit use of identical casting cores, similar passageways may be added to the first airfoil. In some asymmetric alternatives, the first airfoil could be left unchanged relative to the baseline. In other asymmetric alternatives, the first airfoil (and not the second) could include similar cooling along its region164. For example,FIG. 7shows passageways220,222, and224adjacent a feed passageway208′ (thinned relative to208) from which an array of outlet holes (or a slot)212′ extends.

FIG. 8shows an implementation of the added passageways214,216, and218as discrete, non-interconnected, and non-serpentine upstream-to-downstream arrayed legs. Airflow230passes inboard from outboard inlets of the legs and is discharged through the platform outlets without diversion (e.g., via film cooling holes to the suction surface). In an alternative (not shown) with interconnected legs, the overall flow may also enter from the shroud and discharge from the platform. In variations of either embodiment, there may, however, be diversions from this flow (e.g., for film cooling). Similarly,FIG. 9shows the passageways220,222, and224.

Returning toFIG. 3, the exemplary inlet ports are shown in one exemplary combination corresponding to the passageway positions ofFIG. 7. In the exemplary implementation, inlet ports110and115, respectively, feed the lead passageways200of the first and second airfoils62and64. The inlet port111feeds the next three through-passageways of the first airfoil62. For the second airfoil64, the port111is replaced with two ports116and117. The port116feeds passageways202,204, and206whereas the port117feeds the passageways214,216, and218. Conversely, for the second airfoil64, the port118feeds the trailing feed passageway208. For the first airfoil62, the port118is replaced by ports112and113feeding the feed passageway208′ on the one hand and the through-passageways220,222, and224on the other hand. In the exemplary platform ofFIG. 4, the port122is positioned to receive the combined flow from the passageways202,204,206,214,216, and218for the second airfoil64. For the first airfoil62, however, the port120discharges the flow from the three main through-passageways whereas the port121discharges the flow from the added passageways220,222, and224. Where multiple passageways are fed by or feed a single port, an associated plenum structure is defined within the shroud or platform.

FIG. 11shows an exemplary two-circuit single serpentine arrangement. A first circuit240passes a flow242and a second circuit244passes a flow246. The two circuits each have a first downpass leg248;250receiving the flow from one or more inlets (e.g., the inlet112). Therefrom, the circuits each have an inboard turn252;254. Therefrom the circuits each have a backpass leg256;258. Therefrom, the circuits each have an outboard turn260;262. Therefrom, the two circuits have a final downpass leg264,266discharging the associated flow242;246from an associated outlet in the platform. Relative to the direction of flow over the airfoil, the exemplary direction of the flow242is downstream to upstream (e.g., toward the leading edge) while the direction of the airflow246is downstream (e.g., toward the trailing edge).

One or more embodiments of the present invention have been described. Nevertheless, it will be understood that various modifications may be made without departing from the spirit and scope of the invention. For example, the principles may be applied in the remanufacturing of an existing engine or the reengineering of an existing baseline engine configuration. In such a remanufacturing or reengineering situation, details of the baseline configuration may influence details of the particular implementation. Accordingly, other embodiments are within the scope of the following claims.