Patent Abstract:
A turbine engine includes a frame assembly including an outer cavity and an inner cavity with the outer cavity including at least one opening configured and adapted to communicate cooling air to the turbine case. A baffle within the outer cavity includes a plurality of openings for directing cooling airflow into the outer cavity for preventing impingement on a radially inner wall of the outer cavity for maintaining a desired temperature of the cooling air within the outer cavity.

Full Description:
REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 61/939,950 filed on Feb. 14, 2014. 
    
    
     BACKGROUND 
     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. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     A mid-turbine frame is sometimes provided between the high pressure turbine and the low pressure turbine to aid in supporting bearing assemblies. The low pressure turbine case requires cooling air to maintain temperatures within a desired limit. Cooling air is extracted from the compressor section and routed to a cavity within the mid-turbine frame. Cooling air from the cavity within the mid-turbine frame is then routed to cool the low pressure turbine case. In some applications, the mid-turbine frame is at a temperature such that cooling air within the cavity is heated above a temperature capable of sufficiently cooling the low pressure turbine case. 
     Accordingly, it is desirable to design and develop cooling features and systems for maintaining desired temperatures within the turbine case. 
     SUMMARY 
     A turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a turbine section including a turbine case disposed about an axis. A frame assembly defines an outer cavity. The outer cavity includes radially outer wall, a radially inner wall and at least one opening configured and adapted to communicate cooling air to the turbine case. A baffle is configured to receive cooling air through the radially outer wall and direct cooling airflow within the outer cavity to prevent impingement on the inner wall. 
     In a further embodiment of any of the foregoing turbine engines, the baffle includes a plurality of openings for directing cooling air transverse to the radially inner wall of the outer cavity. 
     In a further embodiment of any of the foregoing turbine engines, the baffle is disposed within the outer cavity. 
     In a further embodiment of any of the foregoing turbine engines, the plurality of openings are disposed about an outer periphery of the baffle for directing cooling airflow forward, aft and circumferentially within the outer cavity. 
     In a further embodiment of any of the foregoing turbine engines, the plurality of openings includes holes. 
     In a further embodiment of any of the foregoing turbine engines, the plurality of openings includes slots. 
     In a further embodiment of any of the foregoing turbine engines, includes a compressor section in communication with a supply tube for supplying cooling air to the baffle. 
     In a further embodiment of any of the foregoing turbine engines, the compressor section includes a high pressure compressor. 
     In a further embodiment of any of the foregoing turbine engines, the turbine section includes a high pressure turbine and a low pressure turbine and the frame is a mid-turbine frame which defines a flow path between the high pressure turbine and the low pressure turbine. 
     A frame assembly for a turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a plurality of vane struts extending radially outward relative to an axis, an outer cavity which includes an opening for communicating cooling air to a turbine section of the turbine engine, and a baffle within the outer cavity configured and adapted to receive cooling air. The baffle includes a plurality of openings for directing cooling airflow into the outer cavity for preventing impingement on a radially inner wall of the outer cavity for maintaining a desired temperature of the cooling air within the outer cavity. 
     In a further embodiment of any of the foregoing frame assemblies, the plurality of openings direct cooling airflow forward, aft and circumferentially within the outer cavity. 
     In a further embodiment of any of the foregoing frame assemblies, the plurality of openings includes a plurality of holes. 
     In a further embodiment of any of the foregoing frame assemblies, the plurality of openings includes a plurality of slots. 
     In a further embodiment of any of the foregoing frame assemblies, the plurality of openings define an total opening area for metering cooling airflow into the outer cavity. 
     In a further embodiment of any of the foregoing frame assemblies, includes an inner cavity radially inward of the plurality of vane struts. The inner cavity is in communication with the outer cavity. 
     In a further embodiment of any of the foregoing frame assemblies, the opening for communicating cooling air to the turbine section include a plurality of openings disposed circumferentially within the outer cavity. 
     In a further embodiment of any of the foregoing frame assemblies, the baffle includes at least two baffles directing cooling air within the outer cavity. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic view of an example gas turbine engine. 
         FIG. 2  is an axial section view of an example mid-turbine frame assembly. 
         FIG. 3  is a sectional view of a portion of the example mid-turbine frame assembly. 
         FIG. 4  is a perspective view of a portion of an outer cavity of the mid-turbine frame assembly. 
         FIG. 5  is a schematic view of the outer cavity and example baffle. 
         FIG. 6  is a top schematic view of the example baffle. 
         FIG. 7  is a sectional view of cooling airflow within the example mid-turbine frame assembly. 
     
    
    
     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 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 assembly  58  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 assembly  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
     Airflow through the core airflow path 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 assembly  58  includes vanes  60 , which are in the core airflow path C and function as an inlet guide vane for the low pressure turbine  46 . Temperatures of the exhaust gases are such that cooling of the mid-turbine frame assembly  58  may be required. A low temperature cooling air flow (LTCA) supply tube  66  communicates relatively cool air from the compressor section  24  to the turbine section  28 . In this example, the supply tube  66  communicates relatively low temperature cooling air  18  from one of the initial stages of the high pressure compressor  52  to the mid-turbine frame assembly  58 . 
     Utilizing the vane  60  of the mid-turbine frame assembly  58  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 assembly  58 . 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 airflow through the bypass flow path 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 (10.67 km). The flight condition of 0.8 Mach and 35,000 ft (10.67 km), 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 (350 meters/second). 
     The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section  22  includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about six (6) turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about three (3) turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     Referring to  FIGS. 2, 3 and 4  an example mid-turbine frame assembly  58  includes an outer cavity  62  and an inner cavity  64 . The outer cavity  62  is disposed radially outward of the airfoils  60  and the inner cavity  64  is disposed radially inward of the airfoils  60 . Several LTCA supply pipes  66  deliver cooling air from the compressor section  24  to the outer cavity  62 . In this example, four (4) supply tubes  66  are arranged ninety (90) degrees apart about the circumference of the mid-turbine vane assembly  58 . As appreciated, different numbers of supply tubes  66  could be utilized in different locations about the mid-turbine vane assembly  58 . In this example, cooling air  18  is extracted from an initial stage of the high pressure compressor  52 . As appreciated, cooling air may be obtained from other portions of the engine  20  that include air at appropriate pressures and temperatures. 
     The mid-turbine frame assembly  58  includes a plurality of airfoils  60  and vane struts  76  arranged circumferentially about the engine axis A. The airfoils  60  define passages between the high pressure turbine  54  and the low pressure turbine  46 . The vane struts  76  provide support for structures such as bearings supported radially inward of the airfoils  60 . The outer cavity  62  and inner cavity  64  are provided with cooling air  18  that is circulated from the outer cavity  62  to the inner cavity  64  through openings between the airfoils  60  and vane struts  76 . 
     The outer cavity  62  is defined between a radially outer wall  80  and a radially inner wall  78 . The radially inner wall  78  is exposed to high temperature gas flow  82  and it therefore operates at a substantially higher temperature than the radially outer wall  80 . 
     Cooling air  18  is communicated to the outer cavity  62  to cool the mid-turbine frame  58 . The cooling air  18  is also communicated through the outer cavity  62  to a low pressure turbine (LPT) cavity  86  defined within a turbine case  74  ( FIG. 3 ) through a plurality of supply holes  72 . Cooling air  18  may also be communicated to the LPT cavity  86  through a feather seal  72  defined at an aft portion of the outer cavity  62 . 
     The mid-turbine frame assembly  58  is very hot and therefore the temperature of the cooling air  18  provided to cool the low pressure turbine case  74  may require additional cooling features to provide a flow of a desired temperature determined to provide the desired cooling of the low pressure turbine  46 . Cooling air  18  that directly impinges on the radially inner wall  78  is heated and can reach temperatures above desired threshold values for fooling the turbine case  74 . Additionally, direct impingement of cooling air onto the inner wall  78  can result in non-uniform temperatures of the inner wall  78  that can increase thermal stresses. 
     Accordingly, the example mid-turbine frame assembly  58  includes features that prevent direct impingement and provide a more uniform temperature distribution within the inner wall  78 . 
     Referring to  FIGS. 5, 6, and 7 , the supply pipe  66 , communicates cooling air flow  18  to a baffle  68 . The baffle  68  is disposed within the outer cavity  62  and includes a plurality of openings  86 . In the disclosed example, the baffle  68  directs incoming cooling air outward in a direction transverse to the inner radial wall  78  to prevent direct impingement of cooling air on the inner radial wall  78 . 
     The example baffle  68  receives cooling air flow  18  and distributes the cooling airflow as indicated by arrows  84  forward, aft, and circumferentially within the outer cavity  62  such that the cooling air flow  84  is directed transverse relative to the incoming airflow  18 . The transverse direction can include components in the forward and aft direction parallel with the axis A and also include a circumferential component within the outer cavity  62 . 
     In this example, the baffle  68  is cylindrical and includes openings disposed about an outer periphery to distribute cooling airflow  84  into the outer cavity  62 . It should be understood that although a cylindrical shape is disclosed, the baffle  68  may comprise any shapes desired to direct airflow within the outer cavity  62 . Moreover, the openings  86  are holes that provide a desired flow area for the cooling airflow  84 . The openings  86  may be holes, slots, or any other shape that provides a desired direction of cooling airflow into the outer cavity  62 . 
     The openings  86  combine to provide a desired flow area for the cooling airflow  84 . The flow area provided by the plurality of openings  86  can be tailored to provide a desired metering of cooling airflow as is desired for cooling of both the mid-turbine frame and the turbine case  74 . 
     The directed airflow  84  does not directly impinge on the inner radial wall  78  and therefore does not become heated above desired threshold limits. Moreover, the baffle directs cooling airflow  84  to provide a substantially uniform temperature of the radially inner wall  78 . The reduction in heating of the cooling airflow  84  within the outer cavity  64  provides a uniform flow of cooling air into through the openings  72  into the cavity  88  of the turbine case  74 . 
     Accordingly, the disclosed baffle  68  prevents impingement of cooling airflow on the radially inner wall  78  of the cavity  62  to generate a more uniform temperature. Additionally, the baffle  68  directs cooling air transverse to the radially inner wall  78  such that cooling air within the cavity  62  may be maintained at a lower temperature within a desired threshold temperature range for cooling of a turbine case  74 . 
     The example mid-turbine frame  58  includes baffles  68  at each inlet for cooling airflow  18  ( FIG. 7 ) such that airflow is directed circumferentially about the axis A. In this example, inlets  66  are spaced evenly apart about the axis A and provide cooling air to a corresponding baffle  68 . The baffle  68  distributes the cooling airflow  84  transverse to incoming airflow  18  and to the inner radial wall  78  to prevent absorption of excessive heat in any one location. The distribution provided by the baffles  68  generate a more uniform temperature distribution in both the radial wall  78  and the cooling air  84  circulating though the outer cavity  62 . 
     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 this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.

Technology Classification (CPC): 5