Abstract:
A method of machining cooling holes in a component includes the steps of inserting an electro discharge machining guide that houses an electrode into an internal cavity of a component, and machining a cooling hole into a wall of the component with the electrode. A gas turbine engine component includes first and second spaced apart walls providing an internal cavity. The first wall has outer and inner surfaces. The inner surface faces the internal cavity. A cooling hole extends through the first wall from the inner surface to the outer surface. The cooling hole includes entry and exit openings respectively provided in the inner and outer surfaces. The exit opening includes a cross-sectional area that is smaller than a cross-sectional area of the entry opening.

Description:
BACKGROUND 
     This disclosure relates to components for a gas turbine engine, such as airfoils. More particularly, the disclosure relates airfoils that are electro discharge machined. 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     Turbine vanes used in the turbine section are manufactured as either single vanes, vane doublets, or multiple vanes combining quantities of more than two vanes. Multiple combined vanes such as this may have cooling holes that are not in a location that provides gun barrel line of sight from the machine tool to the hole location. Without line of sight access one way to machine these cooling holes is to use complicated tooling and programming to gain access to the intended location on the external airfoil. The obstructed hole is machined externally. 
     Turbine airfoil cooling hole position is typically inspected at the external hole breakout to gain some level of confidence that the hole is breaking into the intended internal cavity. However on multiple combined vanes the feature location cannot be inspected since the hole cannot be viewed from the gun barrel axis of the hole. Since the location of the holes cannot be accurately measured from the external surface, there exists some risk that the hole may not be drilled into the intended cavity. 
     The two manufacturing methods typically utilized for the machining of cooling holes in turbine airfoils are electro discharge machining (EDM) and laser. Many turbine airfoils have a thermal barrier coating applied to the airfoil surfaces and sometimes this is done prior to the installation of cooling holes. The thermal barrier coating is non-conductive so this prevents the use of the EDM process when machining the cooling holes from the external part surface. In these instances the only option is laser hole drilling, which does not have as much dimensional control when compared to EDM and also is not capable of non-line of sight machining. 
     SUMMARY 
     In one exemplary embodiment, a method of machining cooling holes in a component includes the steps of inserting an electro discharge machining guide that houses an electrode into an internal cavity of a component, and machining a cooling hole into a wall of the component with the electrode. 
     In a further embodiment of any of the above, the guide includes first and second portions that are non-colinear with respect to one another. 
     In a further embodiment of any of the above, the component includes an airfoil. The internal cavity is a cooling passage within the airfoil. 
     In a further embodiment of any of the above, the component is a turbine stator vane. 
     In a further embodiment of any of the above, the turbine stator vane is a doublet. 
     In a further embodiment of any of the above, the method includes the step of applying a thermal barrier coating on the wall on a surface opposite the internal cavity. The applying step is performed before the machining step. 
     In a further embodiment of any of the above, the wall is obstructed by a structure on a side opposite the internal cavity. 
     In another exemplary embodiment, a gas turbine engine component includes first and second spaced apart walls providing an internal cavity. The first wall has outer and inner surfaces. The inner surface faces the internal cavity. A cooling hole extends through the first wall from the inner surface to the outer surface. The cooling hole includes entry and exit openings respectively provided in the inner and outer surfaces. The exit opening includes a cross-sectional area that is smaller than a cross-sectional area of the entry opening. 
     In a further embodiment of any of the above, the component includes an airfoil and the internal cavity is a cooling passage within the airfoil. 
     In a further embodiment of any of the above, the component is a turbine stator vane. 
     In a further embodiment of any of the above, the turbine stator vane is a doublet. 
     In a further embodiment of any of the above, the gas turbine engine component includes a thermal barrier coating provided on the outer surface. The cooling hole extends through the thermal barrier coating. The thermal barrier coating remains intact surrounding the exit opening. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
         FIG. 1  schematically illustrates a gas turbine engine embodiment. 
         FIG. 2  is a schematic plan view of a stator vane doublet. 
         FIG. 3  schematically illustrates walls of adjacent vanes having film cooling holes electrode discharge machined in one of the walls. 
         FIG. 4  is a schematic view of an example EDM system configured to machine film cooling holes in a wall with limited access. 
         FIG. 5  is a schematic view of film cooling holes machined in a wall from an internal cavity of a component, such as an airfoil. 
         FIG. 6  illustrates a manifold providing passageways used to guide an electrode to a desired location adjacent to a wall. 
         FIG. 7  is an enlarged cross-sectional view of a film cooling hole machined by an electrode. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
     Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
     The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , to drive the fan  42  at a lower speed than the low speed spool  30 . The high-speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
     A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The example low pressure turbine  46  has a pressure ratio that is greater than about five (5). The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
     A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
     The core airflow C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes vanes  59 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  59  of the mid-turbine frame  57  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  57 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
     The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
     In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
     “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
     “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
       FIG. 2  illustrates a stator vane  62 , which may be used between stages in the turbine section, such as the high pressure turbine section  54 . In the example illustrated, the stator vane  62  is a “doublet” having a pair of airfoils  68  that extend radially between inner and outer platforms  64 ,  66 . Although a stator vane is illustrated as the component in which holes are drilled using an EDM process, other components may benefit from the disclosed system and process. 
     In some applications, sufficient room between adjacent airfoils exist to machine film cooling holes  82  using an EDM electrode  86  fed through a straight guide  84 , as best shown in  FIG. 3 . First and second walls  70 ,  72  are spaced apart from one another. First and second walls  70 ,  72  may correspond to adjoining walls of a pair of airfoils  68 . A space  78  is provided between the first and second walls  70 ,  72 . The first wall  70  has spaced apart first and second surfaces  74 ,  76 . The guide  84  is inserted into the space  78  from an end  80  into the space between the airfoils to a location adjacent to the first surface  74 . The electrode  86 , which is consumable brass, for example, is fed through the guide  84  as current is provided to the electrode  86 , which removes material from the first wall  70  to provide the film cooling hole  82 . 
     A schematic of an example EDM system  89  is illustrated in  FIG. 4 . The system  89  includes a non-linear guide  184  that may be used to feed a tip  88  of the electrode  86  in areas with much more limited space or conventional guides cannot be used, for example, area obstructed by external structures. In one example, the guide  184  is constructed from stainless steel with a zirconia tip. The guide  184  includes first and second portions  85 ,  87  that are not co-linear with respect to one another. The first and second portions  85 ,  87  are canted at an angle relative to one another that enables the guide  184  to be inserted in tight spaces, such as the cooling passage  178  of the airfoil  68  (shown in  FIG. 2 ). 
     The system  89  includes a guide positioning device  90  that moves the guide  184  in A, B and W directions. The guide may also made movable in additional directions to provide more complicated film hole cooling geometries. The electrode  86  is advanced in a U direction using an electrode feed device  92 , which provides current to the electrode  86 . 
     The stator vane  62  is mounted to a table  94  by a fixture  96 . The table  94  is movable in X and Y directions. The controller  98  communicates with the guide positioning device  90 , electrode feed device  92  and table  94  to position the guide  184  and electrode  86  in desired locations to machine film cooling holes  182 , as shown in  FIG. 5 . 
     With continuing reference to  FIG. 5 , the guide  184  with its electrode  86  is inserted into ends  180  of the cavity  178 . In the example, the cavity  178  corresponds to an internal cooling passage of the airfoil  68  between pressure and suction sides of the airfoil  68 . The first and second walls  170 ,  172  are relatively close to one another, such that access to the cavity  178  is limited. 
     A thermal barrier coating (TBC)  100  is provided on an outer surface  176  of the first wall  170 . The electrode  86  is positioned by the guide  184  in a desired position adjacent to the inner wall  174 . The current is applied to the electrode  86  and advanced as the electrode  86  is consumed to machine the film cooling holes  182 . The TBC  100  is not conductive. However, the electrical and thermal energy that is built up from the initiation of the EDM and through the EDM drilling is sufficient to liberate the TBC in the area around the exit of the film cooling hole  182  at the external breakout location in the outer surface  176 . Removing the TBC  100  in this manner will not cause any further damage to the TBC  100  surrounding the film cooling hole  182 . That is, the TBC  100  will remain intact surrounding the film cooling hole  182  at the outer surface  176 . As a result, the TBC  100  can be applied to the wall  170  prior to machining the film cooling holes  182 . 
     In another example illustrated in  FIG. 6 , a manifold  102 , which provides the guide, may be placed within the cavity  178 . The manifold  102  conforms to the internal cavity shape of the part being machined. The manifold  102  is undersized relative to the size of the cavity  178 . The manifold  102  may include one or more locators  104  to facilitate insertion of the manifold  102  into the cavity  178  and locate the manifold  102  in a desired position with respect to the first wall  170 . 
     The manifold  102  includes multiple passages  106 , which are non-linear enabling the manifold  102  to guide the electrode  86  to the position desired with respect to the first wall  170 . A conventional EDM electrode guide may be used to feed the electrodes through the manifold passages  106  to machine the film cooling holes  182  from the cavity  178 . 
     Referring to  FIG. 7 , the film cooling hole  182  is shown in more detail. The electrode  86  is provided within the cavity  178 . The probe  86  begins forming an entry opening  190  in the inner surface  174  of the wall  170 . The electrode  86  continues to remove material from the wall  170  until an exit opening  192  in the outer wall  176  is formed. The exit opening  192  has a smaller cross-sectional area than then the entry opening  190 . As a result, the flow of cooling air will be more restricted at the outer surface  176 . 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.