Patent Publication Number: US-6340047-B1

Title: Core tied cast airfoil

Description:
The U.S. Government may have certain rights in this invention in accordance with Contract No. F33657-83-C-0281 awarded by the Department of the Air Force. 
    
    
     BACKGROUND OF THE INVENTION 
     The present invention relates generally to gas turbine engines, and, more specifically, to casting of turbine airfoils therein. 
     In a gas turbine engine air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases which flow downstream through multiple turbine stages that extract energy therefrom. Since the turbine stages are heated by the hot combustion gases, they are typically internally cooled by using a portion of the pressurized air bled from the compressor. 
     A typical turbine stage includes an annular turbine stator or nozzle having a plurality of circumferentially spaced apart nozzle vanes extending radially between outer and inner bands. Disposed downstream from the nozzle is a row of circumferentially spaced apart turbine rotor blades extending radially outwardly from a supporting rotor disk. 
     The vanes and blades define airfoils having respective aerodynamic geometries for maximizing efficiency of energy extraction from the combustion gases. A typical airfoil includes a generally concave, pressure side and an opposite, generally convex, suction side extending axially between leading and trailing edges, and radially between a root and a tip. 
     In a nozzle vane, the airfoil extends radially between the outer and inner bands and is typically formed in a one-piece casting. In a rotor blade, the airfoil tip is spaced from a surrounding turbine shroud, with the root of the airfoil being integrally formed with a dovetail which mounts the blade in a complementary dovetail slot formed in the perimeter of the rotor disk. 
     Since turbine blades rotate during operation they are subject to considerable centrifugal force and corresponding stress, with the force increasing the complexity of cooling the blade. A typical blade includes an internal cooling circuit formed by multiple, radially extending flow passages or channels through which the cooling air is channeled. The blade airfoil is initially internally cooled by the air which is then discharged through various holes extending though the walls of the airfoil. 
     Due to the aerodynamic profile of the airfoil, the heat transfer coefficient between the hot combustion gases and the airfoil varies over the pressure and suction sides between the leading and trailing edges and between the root to tip. Accordingly, the internal cooling circuit varies in complexity for best utilizing the limited cooling air to cool the different portions of the airfoil differently in response to the varying heat influx from the combustion gases. Many compromises must be made in defining the internal cooling circuit due to the aerodynamic limitations of channeling the cooling air therethrough, and while balancing the centrifugal and thermal stress experienced by the blade during operation. 
     A high pressure turbine rotor blade typically includes a dedicated cooling passage or channel behind its leading edge, a dedicated cooling passage behind its trailing edge, and a multi-pass serpentine cooling passage disposed axially therebetween and extending radially between the root and tip of the blade airfoil. The flow passages typically also include turbulators in the form of small ribs extending from the inside surface of the airfoil which trip a portion of the cooling air as it flows radially through the cooling passages for enhancing cooling air heat transfer. The airfoil typically includes several radial rows of film cooling holes extending through the walls thereof for discharging the internal cooling air in corresponding films along the outer surface of the airfoil for providing film cooling thereof. 
     In order to precisely form the external and internal features of the airfoil, turbine rotor blades are typically cast using high-strength superalloys. In the lost wax method of casting, a ceramic casting core is initially molded to precisely define the internal cooling circuit, including any turbulators or other features desired. The core is then surrounded by wax to define the desired metal portions of the blade, and the wax is then surrounded by a ceramic outer shell. 
     The wax is removed, and molten metal is injected into the space previously occupied by the wax. The metal solidifies, the shell is removed, and the core is leached away leaving behind the cast blade, including its airfoil and dovetail having the desired precise configurations thereof, both externally and internally. The various holes in the airfoil, such as the film cooling holes, may then be suitably drilled therein. 
     Some turbine blades, such as stage two blades, have relatively long airfoils which require relatively long casting cores. Since the typical casting core includes multiple legs for matching the multiple internal flow channels of the airfoil, the legs are slender and subject to movement and breakage during the casting process. Misaligned core legs correspondingly change the dimensions of the resulting airfoil, and can lead to out-of-specification locally thick or thin regions for which the airfoil may be rejected. And, core breakage during the casting process also may result in rejection of the cast blade. 
     As a solution to this problem, it is known to provide one or more core ties between adjacent legs to fixedly join together the legs for reducing undesirable movement therebetween during the casting process and reducing the likelihood of core breakage. However, the ties necessarily define a corresponding tie hole in the intermediate airfoil rib through which a portion of the cooling air being channeled through the flow channels is short circuited. Cooling air short circuits in the complex internal flow channels reduce the cooling efficiency of the available air and correspondingly adversely affect the useful life of the blade during operation. 
     Accordingly, it is desired to provide an improved method of casting turbine airfoils which reduces the adverse effects of core ties used in the casting thereof. 
     BRIEF SUMMARY OF THE INVENTION 
     A gas turbine engine airfoil is cast around a core having a plurality of legs to form matching flow channels in the airfoil. The legs have a tie extending therebetween to maintain alignment. And, the tie is relocated along the core span to reduce differential static pressure of the cooling air across the resulting tie hole formed by the core tie. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention, in accordance with preferred and exemplary embodiments, together with further objects and advantages thereof, is more particularly described in the following detailed description taken in conjunction with the accompanying drawings in which: 
     FIG. 1 is an isometric view of an exemplary turbine rotor blade for a gas turbine engine in accordance with an exemplary embodiment of the present invention. 
     FIG. 2 is a radial sectional view through a portion of the blade airfoil illustrated in FIG.  2  and taken along line  2 — 2 . 
     FIG. 3 is an elevational sectional view through the airfoil illustrated in FIG.  2  and taken along line  3 — 3 . 
     FIG. 4 is an isometric view of an exemplary casting core for casting the turbine blade illustrated in FIGS. 1-3 in accordance with an exemplary method, also shown in flowchart form in the several figures. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Illustrated in FIG. 1 is an exemplary turbine rotor blade  10  for a gas turbine engine (not shown). The blade is configured as a second stage turbine blade and is therefore relatively long along its radial or span axis as compared to a first stage turbine blade which is shorter. 
     The blade includes an airfoil  12  and an integral axial-entry dovetail  14  formed in a unitary one-piece casting in accordance with the present invention. The airfoil is configured for extracting energy from hot combustion gases  16  which flow downstream thereover, with the dovetail being disposed in a complementary dovetail slot in a rotor disk (not shown) which is rotated during operation. 
     The airfoil  12  is specifically configured for each engine application by defining an aerodynamic geometry or outer profile thereof specific to the flowfield of the combustion gases  16  channeled thereover. The airfoil includes a generally concave, pressure side  18 , and an opposite generally convex, suction side  20  which extend axially between opposite leading and trailing edges  22 , 24 , and radially along the longitudinal or span axis of the airfoil from a root  26  to a tip  28 . A typical radial section through the airfoil is illustrated in FIG.  2  and includes the typical crescent-shaped aerodynamic profile thereof. 
     Since the turbine blade is heated during operation by the combustion gases  16  which flow over the airfoil thereof, the blade is further specified by defining an internal cooling circuit  30  which extends radially through the dovetail  14  and the airfoil  12  to its tip. The cooling circuit receives pressurized cooling air  32  bled from a compressor (not shown) of the engine. The cooling circuit  30  may take any conventional form for preferentially channeling the cooling air through the different portions of the airfoil for providing corresponding cooling thereof against the varying heat affect of the combustion gases  16 . 
     The air enters the dovetail  14  at its lower end and is discharged from the airfoil through various outlet holes  34  typically in the form of radial rows of film cooling holes which discharge the air in a protective film over the outer surface of the airfoil as a barrier against the hot combustion gases flowing thereover. 
     An exemplary embodiment of the internal cooling circuit  30  is illustrated in more detail in FIG.  3 . The circuit typically includes a plurality of cooling flow channels  36  extending longitudinally or radially between the root and tip of the airfoil as well as radially through the dovetail. The flow channels  36  extend generally along the radial span of the airfoil and are separated axially from each other by corresponding bridges or ribs  38  which are laterally or circumferentially formed integrally with the pressure and suction sides of the airfoil. 
     In the exemplary embodiment illustrated in FIG. 3, the cooling circuit  30  includes a dedicated or lone flow channel  36  inside the airfoil behind the leading edge  22 , and another dedicated or lone flow channel  36  inside the airfoil behind the trailing edge  24 . And, additional ones of the flow channels  36  define a five-pass serpentine flow channel having a first pass behind the leading edge channel and subsequent passes axially therebehind. The five flow channels defining the serpentine are disposed end-to-end with suitable reverse bends near the root and tip of the airfoil so that the last or fifth channel extends outwardly to the airfoil tip immediately adjacent to the trailing edge channel. 
     These three sub-circuits each include a separate inlet through the dovetail for receiving in parallel the cooling air  32  at the base of the dovetail. The cooling air  32  flows radially through the separate flow channels and loses pressure awhile gaining heat as the airfoil is cooled thereby. 
     Internal airfoil cooling may be further enhanced by providing corresponding rows of turbulators  40  on either or both sides of the airfoil along the separate flow channels  36 . The turbulators trip the cooling air as it flows and further reduce the pressure thereof along the length of the channels. 
     The turbine blade as above described is conventional in configuration and operation. The outer profile of the airfoil is suitably defined analytically and adjusted as desired during testing thereof for maximizing aerodynamic performance. The cooling circuit  30  may also be defined analytically and modified as desired by testing for maximizing cooling performance thereof. The so defined turbine blade requires mass production with precise reproduction of the outer and inner features thereof. Mass production is typically effected by casting individual blades using the lost wax method, with the wax representing the metallic features of the blade as molten metal replaces the volume previously occupied by the wax. 
     FIG. 4 illustrates schematically a method of making the exemplary turbine blade  10  illustrated in FIGS. 1-3 in accordance with a preferred embodiment of the present invention. After the aerodynamic geometry of the blade and the internal cooling circuit  30  are suitably initially defined as shown in FIGS. 1 and 3, a corresponding ceramic casting core  42  is then initially defined or formed to match the internal cooling circuit  30  in any conventional manner. 
     The core  42  has a plurality of branches or legs  44  which are configured to match respective ones of the flow channels  36  illustrated in FIG.  3 . Each of the core legs  44  is axially separated from its neighbor by a corresponding gap  46  which matches the corresponding ribs  38  of the resulting cast blade. Each of the core legs  44  includes corresponding cavities or depressions  40   c  which match respective ones of the turbulators  40 . The depressions  40   c  thusly define the respective turbulators  40  when metal is cast therein. 
     The core  42  has a longitudinal or span axis which corresponds with that of the resulting blade  10  illustrated in phantom outline in FIG.  4 . The legs  44  and the intervening gaps  46  extend along the span axis of the core, with the legs being cantilevered from a common support base  48 . The individual legs  44  require precise alignment for precisely forming the internal flow channels  36 . The common base  48  is formed integrally with the several legs  44  in a unitary casting itself. The base  48  supports the radially inner ends of the several leas  44 , and a ceramic cap  50  is suitably attached to the radially outer ends of two or more of the legs  44 . The cap  50  defines a corresponding recess in the airfoil tip illustrated in FIG. 3, for example, and defines the bottom of the tip floor which closes the top of the cooling circuit  30 . 
     The core  42  illustrated in FIG. 4 is thusly configured to extend through both the blade airfoil  12  and dovetail  14 , with the core base  48  being disposed below the dovetail. For the relatively long stage two turbine blade  10 , the corresponding core  42  requires long and slender legs  44  which may be subject to movement and misalignment during the casting process, as well as breakage, in vies of the brittle nature of the ceramic used. 
     Accordingly, the process of casting the blade also includes locating or defining at least one core tie  52  between two adjacent ones of the core legs  44  to maintain fixed alignment therebetween for ensuring proper size of the gap  46  and the resulting proper thickness of the corresponding ribs  38 , as well as correct wall thickness of the airfoil. One or more of the core ties  52  may be used as required to maintain alignment of the legs  44  and reduce the likelihood of core breakage during casting. 
     The number and position of the core ties  52  may be determined in any conventional manner for maintaining precision and integrity of the core  42  itself during the casting process. Manufacture of the core and its ties, and blade casting are typically accomplished by vendor companies specializing therein. For example, the casting of superalloy turbine blades may be performed by Howmet Corporation, Whitehall, Mich. which has proven experience developed over many years of commercial production in this country. 
     In the lost wax method of casting, wax (not illustrated) is cast around the core  42  using a master mold (not shown) to define the outer profile of the blade, including its airfoil and dovetail. The mold is removed and a ceramic shell  54 , shown in part in FIG. 4, is built around the wax. The wax is then removed by melting for leaving a void or gap between the shell  54  and the core  42  suitably mounted therein. 
     Molten metal  56  is then poured or injected into the casting void to completely surround the core as bounded by the shell. The metal is then solidified followed by removal of the shell  54  and leaching away of the core  42  for leaving behind the cast blade  10  illustrated in FIGS. 1-3. The various holes  34  may then be conventionally drilled through the outer surface of the airfoil for providing outlets for the cooling air channeled therethrough during operation. 
     Although the core ties  52  may be desirable for maintaining alignment of the core legs  42  and reducing the likelihood of core breakage during casting, they correspondingly form undesirable tie holes  58  as shown in FIGS. 2 and 3. But for the tie holes  58 , the corresponding ribs  38  are preferably imperforate in the preferred embodiment, with the tie holes being a necessary consequence of using the core ties. 
     As shown in the exemplary configuration illustrated in FIG. 3, there are four tie holes  58  formed in the intermediate ribs  38  corresponding to the four core ties  52  illustrated in FIG.  4 . The number of core ties and their initial positions are initially determined solely by the mechanical requirements for maintaining alignment of the core legs and reducing core breakage during the casting process. 
     The resulting tie holes  58  accordingly provide short circuits in the predefined internal cooling circuit  30  which adversely affects cooling performance thereof. In the hostile operating environment of a gas turbine engine, the small adverse affect created by the tie holes  58  can significantly adversely affect the useful life of the blade during operation. Reduced cooling performance can occur from the tie holes  58  subjecting the airfoil to additional thermal stress during operation and reducing the cycle life thereof. 
     However, and in accordance with the present invention, the tie holes  58  may be preferentially relocated along the span of the airfoil to minimize their adverse affect on airfoil cooling. More specifically, and as shown in FIG. 3, an improved process of making the blade includes additionally determining the internal static pressure distribution of the cooling air  32  across each of the intermediate ribs  38  in which a corresponding tie hole  58  is located. The static pressure distribution inside the airfoil may be determined in any conventional manner, such as using a one-dimensional mathematical analysis given the internal geometry of the cooling circuit  30  and the typical cooling parameters of the cooling air  32  channeled through the blade. The static pressure distribution is determined preferably without including the tie holes  58 , with the intermediate ribs being otherwise imperforate. 
     In this way, the adverse affect of including the tie holes  58  in the intermediate ribs may be determined based on the expected effect of the short circuits provided by the tie holes. The internal static pressure distribution in the airfoil is affected by the specific configuration and lengths of the several flow channels  36 . As shown in FIG. 3, the leading and trailing edge flow channels have a single pass and perform differently than the five-pass serpentine flow channels therebetween. 
     All three sub-circuits receive respective portions of the common cooling air  32  at the base of the dovetail, with the air losing pressure differently and absorbing heat differently in each of the three circuits. Furthermore, since the blade rotates during operation, the cooling air is subject to centrifugal force which locally pumps the air for increasing its pressure greater near the tip of the airfoil than near its root. 
     Accordingly, for each of the desired locations of the core ties  52  which create the tie holes  58 , the differential static pressure across the respective tie holes  58  may be determined. If that differential pressure or pressure drop is near zero, the tie hole will have little adverse affect on blade cooling. If the pressure drop is large, cooling air will short circuit through the tie hole and adversely affect blade cooling in the corresponding flow channel deprived of its full complement of cooling air. 
     In accordance with the present invention, each of the initially defined core ties  52  may be relocated along the core span to reduce the differential static pressure across the corresponding tie hole  58 . As shown in FIG. 4, each of the core ties  52  has a span position or height A measured from the common base  48 . The span height of the individual core ties  52  is initially determined by the mechanical requirements to maintain precise alignment between the slender core legs  44  and reduce core breakage. 
     The span heights of the respective core ties  52  may then be adjusted following determination of the pressure distribution inside the airfoil for reducing the differential pressure across the tie holes. In this way, the core ties  52  may be repositioned to reduce their adverse affect on airfoil cooling in a compromise with alignment of the legs and core breakage during casting. 
     The final casting core  42  is therefore preferably formed with the relocated core ties  52  for improving the location of the resulting tie holes  58  for increasing cooling performance and life of the airfoil. The blade and its airfoil is then normally cast using the reconfigured core  42  in a conventional manner using the lost wax method. 
     In the exemplary embodiment illustrated in FIG. 4, the core legs  44  are cantilevered at their lower base ends from the common support base  48 , and are tied together at their outer ends by the cap  50 . Since the legs  44  are long and slender, misalignment between the five-pass serpentine legs and the lone leading and trailing edge legs is a concern. One or more of the core ties  52  is therefore preferably located near the upper ends of the legs opposite to their base ends. And, one or more of the core ties  52  is preferably relocated further from the base  48  and closer to the outer ends of the legs for reducing the pressure drop across the corresponding tie holes  58 . 
     As the cooling air flows radially outwardly through the several flow channels illustrated in FIG. 3, it is subject to friction losses, heat gain, and centrifugal pumping. The five-pass serpentine flow channels illustrated in FIG. 3 alternately channel the cooling air radially outwardly in the direction of centrifugal pumping and radially inwardly against the direction of centrifugal pumping. When the cooling air reaches the last pass of the serpentine flow channel directly adjacent the trailing edge flow channel, it has lost significant pressure and has absorbed heat. 
     A significant pressure differential will therefore exist between the last pass serpentine channel and the trailing edge channel from root to tip of the airfoil. And, by relocating the tie hole  58 , and its corresponding core tie  52 , closer to the airfoil tip, differential pressure across the tie hole may be reduced due to the significant centrifugal pumping of the cooling air. Correspondingly, at other locations of the tie holes, they may be relocated radially inwardly closer to the airfoil root than they would otherwise be without considering the differential pressure thereacross. 
     Accordingly, for the serpentine flow channels  36  illustrated in FIG. 3, the corresponding casting core  42  illustrated in FIG. 4 includes matching legs  44  disposed end-to-end in a serpentine configuration from the base  48 , with the lone trailing edge leg also extending from the base to adjoin the last serpentine leg at the corresponding core tie  52 . 
     In the exemplary embodiment illustrated in FIG. 4, the core  42  includes an additional core tie  52  disposed between the adjacent second and third legs of the serpentine configuration for maintaining alignment therebetween. And that core tie  52  may be suitably relocated for reducing the differential pressure acting across the corresponding tie hole  58  between the second and third flow channels  36  of the serpentine configuration illustrated in FIG.  3 . 
     In the specific embodiment illustrated in FIG. 4, four of the core ties  52  are used to adjoin respective core legs  44 , with each of the core ties  52  being staggered from each other along the core span. Correspondingly, the resulting tie holes  58  illustrated in FIG. 3 are also staggered along the airfoil span. Since the internal pressure distribution from channel to channel in FIG. 3 will vary, the individual tie holes, and corresponding core ties, may be relocated either radially outwardly or radially inwardly as the specific pressure distribution dictates for reducing the corresponding pressure drops thereacross. 
     Accordingly, the resulting turbine blade  10  has tie holes  58  which are differently located along the airfoil span for reducing air short circuits, than they would otherwise be located based on maintaining alignment and integrity of the casting core. The relocated core ties  52  and corresponding tie holes  58  enjoy the benefit of accurate casting with reduced core breakage, with the additional advantage of decreasing the adverse affect of the cooling air short circuits provided by the tie holes  58 . The airfoil therefore enjoys improved cooling which can lead to an improved useful life thereof not previously available for the same design without relocated tie holes. 
     While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein, and it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.