Abstract:
A component includes at least one thermal riser that extends from an exterior surface of the component. At least one cooling passage extends through a wall and adjoins an interior cooling passage and provides an exterior surface. At least one cooling passage is configured to direct cooling fluid through the wall adjacent to at least one thermocouple.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to U.S. Provisional Application No. 62/068,354 which was filed on Oct. 24, 2014 and is incorporated herein by reference. 
     
    
     BACKGROUND 
       [0002]    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 combustion 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. 
         [0003]    In the pursuit of ever higher efficiencies, gas turbine manufacturers have long relied on high and higher turbine inlet temperatures to provide boosts to overall engine performance. In typical modem gas turbine engine applications, gas path temperatures within the turbine exceed the melting point of the component constituent materials. Due to this, dedicated cooling air must be extracted from the compressor and used to cool gas path components in the turbine. This incurs significant cycle penalties especially when cooling air is utilized in the low pressure turbine (sometimes referred to as the power turbine). 
         [0004]    A significant driver of turbine cooling is the spatially varying temperature distribution of the gas path. The nature of the combustion method and geometry produces a variance is gas temperature in both the radial and circumferential directions of flow. This variation in gas temperature also shifts and changes in magnitude and shape as the hot gases progress through the turbine. 
         [0005]    To help minimize the amount of cooling air needed for the gas path turbine components, an understanding of the spatially resolved temperature of the gas stream is desired to tailor cooling to areas of high heat load. This understanding is achieved through temperature measurements through the gas path. These are typically achieved through direct measurement on thermocouple probes, such as kiel head probes, protruding into the gas path situated. The thermocouple is typically attached to the leading edge of a static structure or rake and protrudes into the gas path allowing direct measurement of the total gas temperature. 
         [0006]    However, since the probe is mounted to a static structure which generally needs to be several hundred degrees cooler than the gas path, film cooling on the static structure is generally needed to cool the static structure. The film cooling, which is typically machined into the static structure, generally directs the cooling air at the thermocouple. This perturbs the flow temperature measurements by the thermocouple and causes large discrepancies in actual values. Therefore, there is a need for improved temperature measurements in the gas path of a gas turbine engine. 
       SUMMARY 
       [0007]    In one exemplary embodiment, a component includes at least one thermal riser that extends from an exterior surface of the component. At least one cooling passage extends through a wall and adjoins an interior cooling passage and provides an exterior surface. At least one cooling passage is configured to direct cooling fluid through the wall adjacent to at least one thermocouple. 
         [0008]    In a further embodiment of the above, at least one thermocouple includes a plurality of thermocouples that extend along a leading edge of the component. 
         [0009]    In a further embodiment of any of the above, a plurality of trailing edge cooling passages extend along a trialing edge of the component and a plurality of side cooling passages located on circumferentially facing sides of the component. 
         [0010]    In a further embodiment of any of the above, at least one cooling passage includes a curvature in a radial and a curvature in an axial direction. 
         [0011]    In a further embodiment of any of the above, at least one cooling passage at least partially circumscribes at least one thermal riser. 
         [0012]    In a further embodiment of any of the above, at least one cooling passage includes a first cooling passage that extends radially inward and a second cooling passage extends radially outward. 
         [0013]    In a further embodiment of any of the above, the first cooling passage includes a first film cooling hole and the second cooling passage includes a second cooling film hole. The first film cooling hole intersects the second film cooling hole. 
         [0014]    In a further embodiment of any of the above, at least one cooling passage includes an inlet portion, an intermediate portion, and an outlet portion. The intermediate portion is misaligned relative to the outlet portion. 
         [0015]    In a further embodiment of any of the above, the inlet portion extends in a generally perpendicular direction from the interior cooling passage. 
         [0016]    In a further embodiment of any of the above, the component is at least one of an airfoil or a rake. 
         [0017]    In another exemplary embodiment, a gas turbine engine includes a combustor section. A component is located downstream of the combustor section and includes at least one thermal riser with an embedded temperature measurement device that extends from a leading edge of the component. At least one cooling passage extends through a wall and adjoins an interior cooling passage and provides an exterior surface At least one cooling passage is configured to direct cooling fluid through the wall adjacent at least one thermal riser. 
         [0018]    In a further embodiment of the above, at least one cooling passage includes a curvature in a radial and a curvature in an axial direction and at least partially circumscribes at least one thermal riser. 
         [0019]    In a further embodiment of any of the above, at least one cooling passage includes a first cooling passage that extends radially inward and a second cooling passage that extends radially outward. 
         [0020]    In a further embodiment of any of the above, the first cooling passage includes a first film cooling hole and the second cooling passage includes a second cooling film hole. The first film cooling hole intersects the second film cooling hole. 
         [0021]    In a further embodiment of any of the above, at least one cooling passage includes an inlet portion, an intermediate portion, and an outlet portion. The intermediate portion is misaligned relative to the outlet portion. 
         [0022]    In a further embodiment of any of the above, the inlet portion extends in a generally perpendicular direction from the interior cooling passage. 
         [0023]    In another exemplary embodiment, a method of cooling a component includes directing a cooling fluid through at least one non-linear cooling passage in an exterior wall of a component adjacent a thermal riser. The cooling fluid is directed through a film cooling hole of at least one non-linear cooling passage away from the thermal riser. 
         [0024]    In a further embodiment of the above, at least one non-linear cooling passage includes a first cooling passage that extends radially inward and a second cooling passage that extends radially outward. 
         [0025]    In a further embodiment of any of the above, the first cooling passage and the second cooling passage at least partially circumscribe the thermal riser. 
         [0026]    In a further embodiment of any of the above, at least one cooling passage includes an inlet portion, an intermediate portion, and an outlet portion. The intermediate portion is misaligned relative to the outlet portion. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0027]      FIG. 1  is a schematic view of an example gas turbine engine. 
           [0028]      FIG. 2  illustrates a perspective view of an example gas turbine engine component. 
           [0029]      FIG. 3  illustrates a side view of the example gas turbine engine component of  FIG. 2 . 
           [0030]      FIG. 4  illustrates internal passages in the component of  FIG. 2 . 
           [0031]      FIG. 5  illustrates an enlarged view of  FIG. 4  surrounding a leading edge. 
           [0032]      FIG. 6  illustrates film cooling holes and cooling passages in more detail. 
           [0033]      FIG. 7  illustrates another example gas turbine engine component. 
           [0034]      FIG. 8  illustrates cooling passages in the example component of  FIG. 7 . 
           [0035]      FIG. 9  illustrates an enlarged view of the cooling passages surrounding a leading edge of the component of  FIG. 7 . 
           [0036]      FIG. 10  illustrates film cooling holes for the cooling passages of the component of  FIG. 7  in more detail. 
       
    
    
     DETAILED DESCRIPTION 
       [0037]      FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
         [0038]    The exemplary 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, and the location of bearing systems  38  may be varied as appropriate to the application. 
         [0039]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated 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 second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . 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 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
         [0040]    The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
         [0041]    The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
         [0042]    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 (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by 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 1.45. “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 (350.5 meters/second). 
         [0043]      FIGS. 2 and 3  illustrate an example gas turbine engine component  60 . In the illustrated example, the component  60  is a rake for supporting a plurality of thermal risers  70 . The thermal risers  70  are designed to be heated by the external gas path such that the bulk temperature of the thermal risers  70  is equal to that of the flow path total temperature. The thermal risers  70  are typically made of extremely high temperature capable material, such as platinum, where operating at such temperatures have little impact on the thermal riser  70  durability. However, the component  60  could also be an airfoil or another component in a hot gas path airflow of the gas turbine engine  20 . The component  60  includes a leading edge  62 , a trialing edge  64 , a radially inner end  66 , and a radially outer end  68 . In the illustrated embodiment, the radially outer end  68  of the component is fixed to a portion of the engine static structure  36  downstream of the combustor section  26 . 
         [0044]    The thermal risers  70  are located along the leading edge  62  of the component  60  and are spaced radially from each other. In this disclosure, radial or radial direction is relative to the engine axis A of the gas turbine engine  20  unless otherwise specified. The thermal risers  70  have an imbedded temperature measurement device within them a sensor that measures the bulk temperature of the thermal riser  70 , such as a thermocouple or thermoresistor. The thermal risers  70  are generally cylindrical in shape and extend axially forward of the leading edge  62 . Although the thermal risers  70  are cylindrical in the illustrated example, other shapes, such as hexagons, octagons, rectangles, or ellipses, could also be used. The thermal risers  70  measure a temperature of the hot gas path airflow exiting the combustor section  26  at varying radial positions for monitoring the operation of the combustor section  26  and the temperature of the hot gas path airflow entering the turbine section  28 . The length of the thermal risers  70  are designed to thermally isolate the measurement portion of the riser at its tip which allows the component  60  to be made of a lower capable material that is actively cooled. 
         [0045]    Multiple components  60  can also be spaced circumferentially around the engine axis A in the hot gas airflow path to measure variations in temperature at multiple circumferential and radial locations. In this disclosure, circumferential or circumferentially spaced is relative to a circumference surrounding the engine axis A of the gas turbine engine  20  unless otherwise specified. 
         [0046]    Because the component  60  is located immediately downstream of the combustor section  26 , the component  60  is subjected to substantial temperatures and requires cooling. The component  60  includes trailing edge cooling passages  72 , side cooling passages  74 , and leading edge cooling passages  76  to properly cool the component  60 . The trailing edge cooling passages  72 , the side cooling passages  74 , the leading edge cooling passages  76  are in fluid communication with an internal cooling passage  78  for ( FIGS. 4-6 ) and can be formed with additive manufacturing. In the illustrated example, the leading edge cooling passages  76  are non-linear. The internal cooling passage  78  is in fluid communication with a pressurized cooling fluid source, such as the compressor section  24 , to feed cooling fluid through the trailing edge cooling passages  72 , the side cooling passages  74 , the leading edge cooling passages  76  to cool the component  60 . 
         [0047]    As shown in  FIGS. 4-6 , the leading edge cooling passages  76  each include an inlet portion  80 , an intermediate portion  82 , and an outlet portion  84  extending through an outer wall  86  of the component  60 . The inlet portion  80  is generally perpendicular to an interior surface  88  of the outer wall  86 . The intermediate portion  82  is generally perpendicular to the inlet portion  80  and parallel to the outer wall  86 . The outlet portion  84  extends in a direction radially inward from the intermediate portion  82  and is generally tangential to an outer surface  90  of the outer wall  86 . However, the outlet portion  84  could also extend in a direction radially outward relative to the intermediate portion  82 . 
         [0048]    The intermediate portion  82  and the outlet portion  84  define an angle a there between as shown in  FIG. 5 . In one example, the angle a is between 45 and 90 degrees and in another example, the angle a is between 45 and 50 degrees. 
         [0049]    A diameter of the inlet portion  80 , the intermediate portion  82 , and the outlet portion  84  is approximately  0 . 08  inches ( 2 . 0  mm) The small diameter of the inlet portion  80 , the intermediate portion  82 , and the outlet portion  84  allows for less cooling fluid to be supplied from the pressurized air source to prevent ingesting hot gases from the hot gas path airflow during operation of the gas turbine engine  20 . 
         [0050]    As shown in  FIGS. 5 and 6 , the leading edge cooling passage  76  extends behind the thermal riser  70  such that a portion of the leading edge cooling passage  76  is radially and circumferentially aligned with the thermal riser  70  in order to cool the leading edge  62  of the component  60 . Because the leading edge cooling passage  76  extends behind the thermal riser  70  and a film cooling hole  92  exiting the leading edge cooling passage  76  is directed away from the thermocouple  70 , the temperature measured by the thermal riser  70  is not influenced by impingement cooling from the leading edge cooling passages  76 . 
         [0051]    The changes in direction of the cooling fluid traveling through the inlet portion  80 , the intermediate portion  82 , and the outlet portion  84  increase convective cooling to reduce the temperature of the leading edge  62  of the component  60 . The leading edge cooling passages  76  direct the cooling fluid through the leading edge  62  of the component  60  without directing impingement cooling fluid in the direction of the thermocouples  70 , which could influence the accuracy of the temperature measurements from the thermocouple  70 . The cooling fluid exiting the leading edge cooling passages  76  through the film cooling holes  92  creates vortices that are shed axially downstream towards the trailing edge  64  during operation of the gas turbine engine  20  when the hot gas path airflow is flowing over the component  60 . 
         [0052]      FIG. 7  illustrates another example component  160 . The component  160  is similar to the component  60  except where shown in the Figures or described below. The component  160  includes the leading edge  62 , the trialing edge  64 , the radially inner end  66 , and the radially outer end  68  similar to the component  60 . 
         [0053]    Because the component  160  is located immediately downstream of the combustor section  26 , the component  160  is subjected to substantial temperatures and requires cooling. The component  160  includes the trailing edge cooling passages  72  and the side cooling passages  74  that were used in the component  60  in addition to leading edge cooling passages  176 . The trailing edge cooling passages  72 , the side cooling passages  74 , the leading edge cooling passages  176  are in communication with the internal cooling passage  78  ( FIGS. 8-9 ). 
         [0054]    In the illustrated example shown in  FIGS. 8 and 9 , the leading edge cooling passages  176  include sets of three cooling passages that are generally symmetric about a circumferential centerline CL extending radially through the component  160 . The cooling passages  176  branch off from the internal cooling passage  78  at a central location  176   a . The set of three cooling passages include a first cooling passage  180  that extends radially inward, a second cooling passage  182  that extends radially outward, and a third cooling passage  184  that extends axially upstream. However, the third cooling passage  184  could be eliminated from the leading edge cooling passages  176  such that only the first cooling passage  180  and the second cooling passage  182  cool the leading edge  62 . 
         [0055]    The first and second cooling passages  180  and  182  include a curvature in a radial direction as shown in  FIG. 9  and in an axial direction as shown in  FIG. 10 . As shown in  FIG. 9 , the curvature in the axial direction of each of the first and second cooling passages  180  and  182  circumscribe approximately 25% of an adjacent one of the thermal risers  70 , such that the first and second cooling passages  180  and  182  do not extend behind the thermal risers  70 . The curvature of the first and second cooling passages  180  and  182  increases the surface area for the cooling fluid to absorb heat from the leading edge  62  of the component  160  without impingement cooling the thermal riser  70 . 
         [0056]    As shown in  FIG. 9 , the curvature of the first and second cooling passages  180  and  182  allow the cooling fluid exiting from first and second film cooling holes  192   a  and  192   b , respectively, to flow in a direction is away from the thermal riser  70  that the first and second cooling passages  180  and  182  are circumscribing. This ensures that the stagnation temperature read by the thermal riser  70  is not influence by impingement cooling from the first and second cooling passages  180  and  182 . The third cooling passage  184  also includes a third film cooling hole  193 . 
         [0057]    The cooling fluid is deflected from impinging on one of the thermal riser s  70  radially spaced from the thermal riser  70  that either the first or the second cooling passages  180  and  182  circumscribe. The cooling fluid is deflected at the intersection of the cooling fluid exiting the first and second cooling film holes  192   a  and  192   b  because the cooling fluid from the first cooling film hole  192   a  intersects the cooling fluid from the second cooling film hole  192   b  at approximately the same angle and deflects the cooling fluid downstream. 
         [0058]    Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
         [0059]    It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure. 
         [0060]    The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claim should be studied to determine the true scope and content of this disclosure.