Abstract:
In a featured embodiment, a gas turbine engine has a compressor section having a downstream rotor and a diffuser downstream of the compressor section. A combustor receives air downstream of the diffuser. A turbine section has at least one component to be cooled. A conduit is spaced from the diffuser and defines a cooling airflow path. The cooling airflow path is separate from an airflow downstream the diffuser, and passing to the combustor. The conduit passes cooling air to the component to be cooled.

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 61/847,104, filed Jul. 17, 2013. 
     
    
     BACKGROUND 
       [0002]    This application relates to a supply duct for supplying cooling air with minimal pressure loss. 
         [0003]    Gas turbine engines are known and, typically, include a fan delivering air into a compressor. The air is compressed and delivered into a combustor section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. 
         [0004]    The products of combustion are quite hot and cooling air is typically provided to a number of locations within the gas turbine engine. 
         [0005]    In addition, the flow of air to the combustor section is closely controlled. Often, a diffuser is positioned immediately upstream of the combustor section and serves to prepare the air for delivery into the combustor section. Due to various packaging realities, the airflow downstream of the diffuser is turned through an approximately 90 degree angle and then back into an inlet through another 90 degree angle. 
         [0006]    In the prior art, this same airflow is utilized as a source of cooling air. 
       SUMMARY 
       [0007]    In a featured embodiment, a gas turbine engine has a compressor section having a downstream rotor and a diffuser downstream of the compressor section. A combustor receives air downstream of the diffuser. A turbine section has at least one component to be cooled. A conduit is spaced from the diffuser and defines a cooling airflow path. The cooling airflow path is separate from an airflow downstream the diffuser, and passing to the combustor. The conduit passes cooling air to the component to be cooled. 
         [0008]    In another embodiment according to the previous embodiment, the cooling airflow path is tapped from a location downstream of the downstream rotor, and upstream of the diffuser. 
         [0009]    In another embodiment according to any of the previous embodiments, the conduit is provided by a pair of radially spaced shells. 
         [0010]    In another embodiment according to any of the previous embodiments, the shells are positioned radially inwardly of the diffuser and the combustor. 
         [0011]    In another embodiment according to any of the previous embodiments, shells are also positioned radially outwardly of the diffuser and the combustor section to provide a second cooling airflow path. 
         [0012]    In another embodiment according to any of the previous embodiments, the component to be cooled includes at least one of a turbine vane, a turbine rotor, and a blade outer air seal. 
         [0013]    In another embodiment according to any of the previous embodiments, one of the shells has a downstream end secured to a base of the turbine vane to provide cooling air to the turbine vane. 
         [0014]    In another embodiment according to any of the previous embodiments, the cooling airflow path, downstream of the shells, passes into an injector tube for supplying cooling air to the turbine rotor. 
         [0015]    In another embodiment according to any of the previous embodiments, one of the shells has an upstream end positioned downstream of an upstream end of a second of the shells to provide an open inlet into the cooling airflow path. 
         [0016]    In another embodiment according to any of the previous embodiments, one of the shells is positioned closer to an outer surface of the diffuser than the second of the shells. 
         [0017]    In another embodiment according to any of the previous embodiments, the shells and the cooling air path are positioned radially outwardly of the diffuser and the combustor section. 
         [0018]    In another embodiment according to any of the previous embodiments, the component to be cooled includes a blade outer air seal. 
         [0019]    In another embodiment according to any of the previous embodiments, the diffuser is mounted by a mount structure to an inner housing. 
         [0020]    In another embodiment according to any of the previous embodiments, at least one of the shells has a slot to be received on the mount structure. 
         [0021]    In another embodiment according to any of the previous embodiments, one of the shells has an upstream end positioned downstream of an upstream end of a second of the shells to provide an open inlet into the cooling airflow path. 
         [0022]    In another embodiment according to any of the previous embodiments, the component to be cooled includes at least one of a turbine vane, a turbine rotor, and a blade outer air seal. 
         [0023]    In another embodiment according to any of the previous embodiments, the conduit is provided by a pair of radially spaced shells. 
         [0024]    In another embodiment according to any of the previous embodiments, one of the shells has a downstream end secured to a base of the turbine vane to provide cooling air to the turbine vane. 
         [0025]    In another embodiment according to any of the previous embodiments, the cooling airflow path, downstream of the shells, passes into an injector tube for supplying cooling air to the turbine rotor. 
         [0026]    In another embodiment according to any of the previous embodiments, a combustor housing is positioned downstream of an outlet of the diffuser, such that air downstream of the diffuser bends through an approximately ninety degree angle in one radial direction, then moves back through an approximately ninety degree angle through an inlet port into a combustion chamber. 
         [0027]    These and other features may be best understood from the following drawings and specification. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0028]      FIG. 1  schematically shows a gas turbine engine. 
           [0029]      FIG. 2  shows a portion of the gas turbine engine of  FIG. 1 . 
           [0030]      FIG. 3  shows one mechanical feature of the  FIG. 2  structure. 
           [0031]      FIG. 4  shows an alternative embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0032]      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. 
         [0033]    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. 
         [0034]    The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a 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 high pressure compressor  52  and 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. 
         [0035]    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 . 
         [0036]    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. 
         [0037]    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 1 bm of fuel being burned divided by 1 bf 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. 
         [0038]      FIG. 2  shows a downstream-most compressor rotor  179  and a downstream most compressor vane  80 . This may be part of an engine such as shown in  FIG. 1 . Downstream of the compressor vane  80  is a diffuser  82 . As known, the diffuser has an upstream end  95  and a downstream exit  81  that is typically of a larger cross-sectional area than the upstream end  95 . 
         [0039]    A plurality of circumferentially spaced mount structures  79  mount the diffuser  82  to a radially inner housing  21 . 
         [0040]    Downstream of the diffuser exit  81  is a portion  86  of a housing for a combustion section  56 . As shown by arrows, part of the air leaving the exit  81  bends through a radially inward direction (approximately through a ninety degree angle), then flows axially along an outer surface of the housing  86 , then radially outwardly (again, approximately through a ninety degree angle) into ports  88  and into a combustion chamber  15 . Fuel is injected through elements  89  and an igniter  91  ignites the fuel and air within the combustion chamber  15 . Products of this combustion pass downstream over a vane  104  and a turbine rotor  102 . 
         [0041]    As is known, the turbine vane  104  and turbine rotor  102  will become quite hot due to the products of combustion. Thus, cooling air is provided. In the past, part of the air flowing to ports  88  was diverted as cooling air. 
         [0042]    In this disclosure, a conduit is formed of a radially inner shell  90  and a radially outer shell  92  to provide a flow path  198  from an inlet  93 . As shown, an upstream end  94  of the inner shell  90  is more upstream than an upstream end  96  of the outer shell  92 . As can be seen, the upstream end  96 , which is downstream of upstream end  94 , is on the outer shell  92 , which is closer to an outer surface  23  of diffuser  82  than is shell  90 . The forward facing inlet provided by this positioning results in a reduced pressure drop across the inlet  93 . As can be appreciated, the shells extend for 360° about a center axis (A) of the engine. 
         [0043]    Air flows through the path  198  and exits through exit port  106  and injector tube  100 . The air exiting port  106  cools the turbine vane  104 , while the air through the injector tube  100  is aimed at the inner bore of the turbine rotor  102 . 
         [0044]    By utilizing the separate cooling air flow path  198 , pressure losses across the diffuser  82 , and through the bending of the air on the way to the inlet  88  do not occur to the cooling air being delivered to the vane  104  and the rotor  102 . As such, more efficient use of the cooling air is achieved. 
         [0045]    As also shown in  FIG. 2 , an inner end  201  of the outer shell  92  abuts against an inner surface  203  or base of the turbine vane  104  such that the air is delivered into the inner surface  203  of the turbine vane  104 . 
         [0046]      FIG. 3  shows a feature that may be found in both the inner and outer shells  90  and  92 , but is illustrated at the inner shell  90 . As shown, the mount structures  79  may be received within slots  101  in the shell  90 . Thus, there are effective vane structures within the cooling air path  198 . 
         [0047]      FIG. 4  schematically shows an alternative embodiment wherein there are shells  190  and  192  at a radially inner end delivering air to the uses  298  which may be schematically a vane, such as turbine vane  104  and a rotor, such as turbine rotor  102 . An outer flow path  288  is provided radially outwardly of the diffuser  82  by two shells  290  and  292  and delivers air, such as to a use  396 , which may be radially outward of the combustor  56 . As an example, the use  396  may be a blade outer air seal  296 , such as shown in  FIG. 2 . 
         [0048]    The use of the dedicated shells to provide the cooling air path result in very efficient use of the cooling airflow. While shells are shown as a conduit defining a cooling air passage, any other method of providing a conduit to define a cooling airflow path separate from the combustion flow path can be utilized. As an example, the shells could be split into several circumferentially spaced pieces, and bolted together. Alternatively, axial ribs can extend the length of the shells and tie them together structurally. Alternatively, there could be individual tubes that carry the airflow from aft of the compressor to the components to be cooled. Again any number of other ways of defining a separate flow path would come within the scope of this disclosure. 
         [0049]    Since the inlet to the cooling air passages faces axially forwardly, or toward an upstream end, the air delivered into the passage sees a total pressure, rather than just static pressure. As can be appreciated from  FIG. 2 , the shape of the cooling air path is smooth, and has no sharp bends which could reduce the pressure of the air. 
         [0050]    Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.