Abstract:
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.

Description:
U.S. GOVERNMENT RIGHTS 
       [0001]    The invention was made with U.S. Government support under contract N00019-02-C-3003 awarded by the U.S. Navy. The U.S. Government has certain rights in the invention. 
     
     BACKGROUND OF THE INVENTION 
       [0002]    The invention relates to cooling of high temperature components. More particularly, the invention relates to coated gas turbine engine vane clusters. 
         [0003]    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. 
         [0004]    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. 
         [0005]    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. 
         [0006]    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 
       [0007]    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. 
         [0008]    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. 
         [0009]    The details of one or more embodiments of the invention are set forth in the accompanying drawings and the description below. Other features, objects, and advantages of the invention will be apparent from the description and drawings, and from the claims. 
     
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]      FIG. 1  is a schematic view of a gas turbine engine. 
           [0011]      FIG. 2  is a view of a vane ring of the engine of  FIG. 1 . 
           [0012]      FIG. 3  is a first view of a vane cluster of the ring of  FIG. 2 . 
           [0013]      FIG. 4  is a second view of the vane cluster of  FIG. 3 . 
           [0014]      FIG. 5  sectional view of airfoils of a prior art/baseline cluster. 
           [0015]      FIG. 6  is a sectional view of airfoils of a first reengineered cluster. 
           [0016]      FIG. 7  is a sectional view of airfoils of a second reengineered cluster. 
           [0017]      FIG. 8  is a schematic plan view of an exemplary non-serpentine suction side wall cooling circuit of the cluster of  FIG. 7 . 
           [0018]      FIG. 9  is a schematic plan view of an exemplary non-serpentine pressure side wall cooling circuit of the cluster of  FIG. 7 . 
           [0019]      FIG. 10  is a schematic plan view of an exemplary interconnected circuit of the cluster of  FIG. 7 . 
           [0020]      FIG. 11  is a schematic plan view of an exemplary serpentine circuit of the cluster of  FIG. 7 . 
       
    
    
       [0021]    Like reference numbers and designations in the various drawings indicate like elements. 
       DETAILED DESCRIPTION 
       [0022]      FIG. 1  shows a gas turbine engine  20  having a central longitudinal axis  500  and extending from an upstream inlet  22  to a downstream outlet  24 . 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 combustor  34 , a high speed/pressure turbine (HPT)  36 , a low speed/pressure turbine (LPT)  38 , an augmentor  40 , and an exhaust duct/nozzle  42 . 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 axis  500 . The blades of the HPC and HPT are mounted on a high speed spool for such rotation. 
         [0023]    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 stage  50  of the HPT  36  is schematically shown in  FIG. 1 .  FIG. 2  shows further details of the exemplary vane stage  50 . The ring includes an inboard platform  52  and an outboard shroud  54 . 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 ring  56  and at the shrouds by an engine case. 
         [0024]      FIGS. 3 and 4  show an exemplary two-airfoil cluster (doublet)  60 . Each exemplary cluster includes a first airfoil  62  and a second airfoil  64 . Each of the airfoils extends from an associated inboard end  66  at a platform segment  68  to an associated outboard end  70  ( FIG. 4 ) at a shroud segment  72 . The exemplary platform segment has an outboard surface  74  along the inboard extreme of the core flowpath. The shroud segment has an inboard surface  76  along an outboard extreme of the core flowpath. 
         [0025]    An underside  80  of the platform segment may include features for mounting each platform segment to its adjacent segments (e.g., by bolting to the ring  56 ). The platform segment has a forward/upstream end  82 , a rear/downstream end  84 , and first and second circumferential ends or matefaces  86  and  88 . Similarly, the shroud segment  72  has an upstream end  92 , a downstream end  94 , and first and second circumferential ends  96  and  98 . Each of the platform circumferential ends  86  and  88  and a shroud circumferential ends  96  and  98  may include a groove or channel  108  for receiving a seal (not shown). A given such seal spans the gap between the adjacent grooves of each adjacent pair of clusters. 
         [0026]    The cluster  60  has cooling passageways. An exemplary passageway network may include one or more inlet ports.  FIG. 3  shows exemplary inlet ports  110 ,  111 ,  112 ,  113 ,  115 ,  116 ,  117 , and  118  (discussed below) in the shroud segment  72 . The inlet ports direct cooling air (e.g., bleed air) through one or more spanwise passageway segments in the airfoils  62  and  64 . 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 platform  68 .  FIG. 4  shows exemplary outlets  120 ,  121 , and  122 . The air discharged through the outlets  120 - 122  may pass downstream to the adjacent blade stage to, in turn, pass through cooling passageways of those blades to cool the blades. 
         [0027]    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). 
         [0028]      FIG. 5  is a sectional view of the airfoils of a baseline version of a cluster from which the inventive clusters may represent reengineerings. The first airfoil  62  is shown having a leading edge  140 , a trailing edge  142 , a pressure side  144 , and a suction side  146 . Pressure and suction side walls are shown as  148  and  149 , respectively. Similarly, the second airfoil  64  has a leading edge  150 , a trailing edge  152 , a pressure side  154 , a suction side  156 , a pressure side wall  158 , and a suction side wall  159 . The airfoils also have passageways described below. 
         [0029]    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 side  146  and the second airfoil pressure side  154 , essentially normal line-of-sight flow access is available. However, along portions of the first airfoil pressure side  144  and second airfoil suction side  156  the other airfoil will block normal line-of-sight access. This blocking/occlusion mandates off-normal application with attendant reduction in coating thickness. 
         [0030]      FIG. 5  shows series of line-of-sight spray directions  510  positioned at boundaries of occlusion by the airfoils.  FIG. 5  also shows a local surface normal  520 . Along a leading region  160  of the first airfoil pressure side, there is essentially normal or near-normal line-of-sight access. Thus, along this region  160 , the coating is full thickness. Downstream thereof, the off-normal angle θ increases. There may be progressive degradation of coating thickness. For example, in a region  162  to an angle θ of about 30°, the coating may be deemed marginal. In a region  164  downstream thereof, and with greater θ, the coating may be deemed poor. 
         [0031]    Similarly, along a trailing region  168  of the second airfoil suction side  156 , the coating may be full-thickness. Along a region  170  thereahead, the coating may be marginal. Along a region  172  yet thereahead, the coating may be poor. Along a region  174  yet thereahead, the coating may be marginal. Along a leading region  176 , 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 thinning is disadvantageous. In such regions, it is desirable to add supplemental cooling. 
         [0032]    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. 
         [0033]      FIG. 6  shows a reengineered cluster (e.g., reengineered from the  FIG. 5  baseline) to add supplemental wall cooling. In the exemplary cluster  60 , each of the airfoils includes a streamwise array of spanwise-elongate passageway legs: a leading edge feed cavity  200 ; a first through-flow leg  202 ; a second through-flow leg  204 ; a third through-flow leg  206 ; and a trailing edge feed cavity  208 . In the exemplary cluster  60 , the leading edge cavity  200  has a closed inboard end and discharges air through spanwise arrays of leading edge outlet holes  210 . Similarly, the cavity  208  may 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. 
         [0034]    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 passageways  214 ,  216 , and  218  in the suction side wall  159  of 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 region  164 . For example,  FIG. 7  shows passageways  220 ,  222 , and  224  adjacent a feed passageway  208 ′ (thinned relative to  208 ) from which an array of outlet holes (or a slot)  212 ′ extends. 
         [0035]      FIG. 8  shows an implementation of the added passageways  214 ,  216 , and  218  as discrete, non-interconnnected, and non-serpentine upstream-to-downstream arrayed legs. Airflow  230  passes 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. 9  shows the passageways  220 ,  222 , and  224 . 
         [0036]    Returning to  FIG. 3 , the exemplary inlet ports are shown in one exemplary combination corresponding to the passageway positions of  FIG. 7 . In the exemplary implementation, inlet ports  110  and  115 , respectively, feed the lead passageways  200  of the first and second airfoils  62  and  64 . The inlet port  111  feeds the next three through-passageways of the first airfoil  62 . For the second airfoil  64 , the port  111  is replaced with two ports  116  and  117 . The port  116  feeds passageways  202 ,  204 , and  206  whereas the port  117  feeds the passageways  214 ,  216 , and  218 . Conversely, for the second airfoil  64 , the port  118  feeds the trailing feed passageway  208 . For the first airfoil  62 , the port  118  is replaced by ports  112  and  113  feeding the feed passageway  208 ′ on the one hand and the through-passageways  220 ,  222 , and  224  on the other hand. In the exemplary platform of  FIG. 4 , the port  122  is positioned to receive the combined flow from the passageways  202 ,  204 ,  206 ,  214 ,  216 , and  218  for the second airfoil  64 . For the first airfoil  62 , however, the port  120  discharges the flow from the three main through-passageways whereas the port  121  discharges the flow from the added passageways  220 ,  222 , and  224 . Where multiple passageways are fed by or feed a single port, an associated plenum structure is defined within the shroud or platform. 
         [0037]      FIG. 10  shows a circuit having legs  232 ,  233 ,  234 , and  235  interconnected by gaps  236  in the walls  237  separating adjacent legs. 
         [0038]      FIG. 11  shows an exemplary two-circuit single serpentine arrangement. A first circuit  240  passes a flow  242  and a second circuit  244  passes a flow  246 . The two circuits each have a first downpass leg  248 ;  250  receiving the flow from one or more inlets (e.g., the inlet  112 ). Therefrom, the circuits each have an inboard turn  252 ;  254 . Therefrom the circuits each have a backpass leg  256 ;  258 . Therefrom, the circuits each have an outboard turn  260 ;  262 . Therefrom, the two circuits have a final downpass leg  264 ,  266  discharging the associated flow  242 ;  246  from an associated outlet in the platform. Relative to the direction of flow over the airfoil, the exemplary direction of the flow  242  is downstream to upstream (e.g., toward the leading edge) while the direction of the airflow  246  is downstream (e.g., toward the trailing edge). 
         [0039]    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.