Patent Publication Number: US-2019176228-A1

Title: Gas turbine engine component cooling passage and space eating core

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This disclosure is a divisional of U.S. patent application Ser. No. 14/766,061 filed Aug. 5, 2015, which is a U.S. National Phase of PCT/US2014/011833 filed Jan. 16, 2014, which claims priority to Provisional Application No. 61/763,661 filed Feb. 12, 2013. 
    
    
     BACKGROUND 
     This disclosure relates to a gas turbine engine component, such as an airfoil. More particularly, the disclosure relates to a component cooling passage having an obstruction and a core for making the component cooling passage. 
     Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. 
     Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades. 
     Many blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages. In some applications, such as vanes, a baffle is provided in the cooling passage to reduce the cross-sectional area of the cooling passage and increase cooling fluid velocity. Typically, the baffles are constructed from sheet metal inserted into the vane after the vane has been cast. 
     SUMMARY 
     In one exemplary embodiment, a gas turbine engine includes a structure that has walls that provide a cooling passage and a cooling surface. A non-ferrous obstruction is relative to the walls. The obstruction includes a portion spaced from the cooling surface to provide a gap which is configured to receive a cooling fluid. 
     In a further embodiment of any of the above, the structure is an airfoil. 
     In a further embodiment of any of the above, the structure is a blade outer air seal. 
     In a further embodiment of any of the above, the structure is a platform that supports an airfoil. 
     In a further embodiment of any of the above, the structure is a combustor line 
     In a further embodiment of any of the above, the structure is an exhaust liner. 
     In a further embodiment of any of the above, the obstruction is arranged to block a flow of the cooling fluid through the cooling passage. 
     In a further embodiment of any of the above, the obstruction includes a refractory metal. 
     In a further embodiment of any of the above, the obstruction provides multiple gaps adjacent to the cooling surface. 
     In a further embodiment of any of the above, the structure includes a nickel alloy, and the obstruction is provided by a material that is different than the nickel alloy. 
     In another exemplary embodiment, a core structure for providing a gas turbine engine a cooling passage includes a first material provided on a second material and together provides a perimeter. The first material is provided along multiple portions of the perimeter with the second material separating the multiple portions. 
     In a further embodiment of any of the above, the first and second materials are different than one another. 
     In a further embodiment of any of the above, the first material is ceramic. 
     In a further embodiment of any of the above, the second material includes a refractory metal. 
    
    
     
       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. 2A  is a perspective view of the airfoil having the disclosed cooling passage. 
         FIG. 2B  is a plan view of the airfoil illustrating directional references. 
         FIG. 3  is an end view of a core structure. 
         FIG. 4  is a cross-sectional view of the air foil taken along line  4 - 4  in  FIG. 2A . 
         FIG. 5  is an example flow chart depicting an example method of manufacturing a gas turbine engine component, such as an air foil. 
     
    
    
     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 . 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 X 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 X. 
     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 (′TSFC)”—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. 
     The disclosed cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade  64  is described. It should be understood that the cooling passage may also be used in vanes, blade outer air seals, and turbine platforms, for example. 
     Referring to  FIGS. 2A and 2B , a root  74  of each turbine blade  64  is mounted to the rotor disk. The turbine blade  64  includes a platform  76 , which provides the inner flow path, supported by the root  74 . An airfoil  78  extends in a radial direction R from the platform  76  to a tip  80 . It should be understood that the turbine blades may be integrally formed with the rotor such that the roots are eliminated. In such a configuration, the platform is provided by the outer diameter of the rotor. The airfoil  78  provides leading and trailing edges  82 ,  84 . The tip  80  is arranged adjacent to a blade outer air seal (not shown). 
     The airfoil  78  of  FIG. 2B  somewhat schematically illustrates exterior airfoil surface extending in a chord-wise direction C from a leading edge  82  to a trailing edge  84 . The airfoil  78  is provided between pressure (typically concave) and suction (typically convex) wall  86 ,  88  in an airfoil thickness direction T, which is generally perpendicular to the chord-wise direction C. Multiple turbine blades  64  are arranged circumferentially in a circumferential direction A. The airfoil  78  extends from the platform  76  in the radial direction R, or spanwise, to the tip  80 . 
     The airfoil  78  includes a cooling passage  90  provided between the pressure and suction walls  86 ,  88 . The exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage  90 . Example cooling passages  90  illustrated in  FIG. 2A  are shown in more detail in  FIG. 4 . 
     A cooling source  92  is in fluid communication with the cooling passage  90 . In the example, the cooling passage  90  is illustrated as having an obstruction  94  arranged within the cooling passage to restrict flow from the cooling source  92 , which increases the velocity of the cooling fluid within the cooling passage  90 . 
     An example cross-sectioned through the airfoil  78  is illustrated in  FIG. 4 . Although the obstruction  94  is illustrated as being arranged in only one of the cooling passages, the obstruction may be arranged in multiple cooling passages or portions thereof. The cooling passage  90  includes cooling passages  90   a - 90   e . In the example, the obstruction  94  is arranged in the cooling passage  90   c  which is surrounded by walls  96  that provide a cooling surface  98 . 
     In the example, the obstruction  94  is non-ferrous and engages the cooling surface  98  to locate the obstruction  94  relative to the walls  96 . The obstruction  94  includes portions spaced from the cooling surface  98  to provide one or more gaps  102 . The gaps  102  receive the cooling fluid from the cooling source  92  ( FIG. 2A ). 
     A multi-material core is used to produce the obstruction  94 . Referring to  FIG. 3 , the core structure  104  includes first and second materials  106 ,  108 , the second material  108  corresponds to the obstruction  94  in the finished gas turbine engine component. In one example, the first material  106  is a ceramic, the second material  108  is a refractory metal, such as molybdenum and/or niobium. 
     Generally, to provide the obstruction  94  within the cooling passage  90   c , the core structure  104  is arranged within a wax mold, as is known in conventional investment casting processes, to produce a wax structure in the shape of the component to be cast. A perimeter  105  of the core structure  104  provides the cooling surface  98  within a space  100  of the cast component, as shown in  FIG. 4 . That is, the cast component engages the perimeter  105 , which is provided by both the first and second materials  106 ,  108 . In this manner, the second material  108  will locate the obstruction  94  relative to the cooling surface  98  once the first material  106  has been removed. The wax structure is then coated in ceramic slurry, which is permitted to harden into a ceramic mold. Molten metal, such as a nickel alloy, is poured into the ceramic mold, which removes the wax. 
     Referring to the method  110  shown in  FIG. 5 , the core structure is provided having multiple, different materials in the example, as indicated at block  112 . After the wax structure is produced, as described above, the gas turbine engine component is cast about the core structure, as indicated in block  114 . A portion of the core structure is removed from the cast component. In one example, the first material  106 , ceramic, is chemically removed, for example, as indicated at block  116 . With the first material  106  removed, which leaves the gaps  102 , the second portion  108  is left within the cast component to provide a reduced cross-section cooling passage  90 , as indicated at block  118 . 
     Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For example, the obstruction may be provided in passages other than cooling passages. The passages may be formed in other structures, such as blade outer air seals, platforms, combustor liner and exhaust liners. For that and other reasons, the following claims should be studied to determine their true scope and content.