Abstract:
A cooling system for a gas turbine engine turbine section includes a rotor supporting a blade having a cooling passage. A disc is secured relative to the rotor and it forms a cavity between the rotor and the disc. A bleed air source is in fluid communication with the cavity. An impeller is arranged in the cavity. The impeller is configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage.

Description:
BACKGROUND 
       [0001]    This disclosure relates to a gas turbine engine, and more particularly to an impeller used in a high pressure turbine section to increase the pressure of a cooling fluid. 
         [0002]    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. 
         [0003]    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. 
         [0004]    Typically bleed air from a compressor stage is used to cool the turbine blades in the turbine section. The cooling fluid is routed to the turbine blades by a variety of structures and then fed to internal cooling passages in the blade through a space in a rotor slot within which the turbine blade&#39;s root is mounted. Sufficiently high pressures must be provided to ensure desired flow through the cooling passages to achieve desired cooling. 
       SUMMARY 
       [0005]    In one exemplary embodiment, a cooling system for a gas turbine engine turbine section includes a rotor supporting a blade having a cooling passage. A disc is secured relative to the rotor and it forms a cavity between the rotor and the disc. A bleed air source is in fluid communication with the cavity. An impeller is arranged in the cavity. The impeller is configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage. 
         [0006]    In a further embodiment of any of the above, the blade is in a last stage of a high pressure turbine section. 
         [0007]    In a further embodiment of any of the above, the rotor and the disc are affixed to a spool for rotation therewith. 
         [0008]    In a further embodiment of any of the above, the bleed air source is a stage of a high pressure compressor section. 
         [0009]    In a further embodiment of any of the above, the high pressure compressor section includes an aft hub having an aft hub leak path. The aft hub leak path is in fluid communication with the cavity and is configured to provide aft hub fluid to the cavity. 
         [0010]    In a further embodiment of any of the above, a tangential on board injector has a TOBI leak path. The TOBI leak path is in fluid communication with the cavity and is configured to provide a TOBI fluid to the cavity. 
         [0011]    In a further embodiment of any of the above, the impeller is mounted on the disc. 
         [0012]    In a further embodiment of any of the above, the impeller includes circumferentially spaced paddles integral with disc. 
         [0013]    In a further embodiment of any of the above, the cooling system for a gas turbine engine turbine section includes the static structure. The disc includes a seal configured to seal relative to the static structure. 
         [0014]    In one exemplary embodiment, a turbine stage for a gas turbine engine includes a rotor. A disc is secured relative to the rotor to provide a cavity there between. An impeller is arranged in the cavity. 
         [0015]    In a further embodiment of any of the above, the impeller includes a set of first paddles and a set of second paddles. The first and second paddles are interleaved relative to one another. The first paddles are larger than the second paddles. 
         [0016]    In a further embodiment of any of the above, a rotor supports turbine blades that have a cooling passage in fluid communication with the cavity. The disc includes a seal in engagement with turbine blades. 
         [0017]    In a further embodiment of any of the above, the disc includes an annular flange that provides the seal. The annular flange extends in an axial direction and is spaced radially from the sets of the first and second paddles. The annular flange provides an annular channel radially between the annular flange and the sets of first and second paddles. 
         [0018]    In another exemplary embodiment, a disc for a turbine stage includes a disc-shaped wall supporting paddles that extend from an inlet radially outward to an outlet. An annular flange extends axially from the wall to provide an annular channel arranged radially between the outlet and the annular wall. 
         [0019]    In a further embodiment of any of the above, the paddles include a set of first paddles and a set of second paddles. The first and second paddles are interleaved relative to one another. The first paddles are larger than the second paddles. 
         [0020]    In a further embodiment of any of the above, the annular flange includes a first seal. 
         [0021]    In a further embodiment of any of the above, a second seal is supported by the wall on a side opposite the paddles. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0022]    The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: 
           [0023]      FIG. 1  schematically illustrates a gas turbine engine embodiment. 
           [0024]      FIG. 2  is a cross-sectional view through a high pressure turbine section including an impeller. 
           [0025]      FIG. 3  is a partial perspective view of the impeller. 
       
    
    
     DETAILED DESCRIPTION 
       [0026]      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 . 
         [0027]    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. 
         [0028]    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. 
         [0029]    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. 
         [0030]    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. 
         [0031]    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. 
         [0032]    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 . 
         [0033]    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. 
         [0034]    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. 
         [0035]    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. 
         [0036]    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. 
         [0037]    “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. 
         [0038]    “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) 0.5]. The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
         [0039]    Referring to  FIG. 2 , a cross-sectional view through a high pressure turbine section  54  is illustrated. In the example high pressure turbine section  54 , first and second arrays  54   a,    54   c  of circumferentially spaced fixed vanes  60 ,  62  are axially spaced apart from one another. A first stage array  54   b  of circumferentially spaced turbine blades  64  is arranged axially between the first and second fixed vane arrays  54   a,    54   c.  A second stage array  54   d  of circumferentially spaced turbine blades  66  is arranged aft of the second array  54   c  of fixed vanes  62 . The last stage of blades  66  are mounted to a rotor  68 . The blades  66  include a cooling passage  114 . 
         [0040]    The turbine blades each include a tip adjacent to a blade outer air seal  70  of a case structure  72 . The first and second stage arrays  54   a,    54   c  of turbine vanes and first and second stage arrays  54   b,    54   d  of turbine blades are arranged within a core flow path C and are operatively connected to the shaft  32 . 
         [0041]    A disc  74  is secured relative to the rotor  68  at the aft of the high pressure turbine  54 . In the example, a fixing device, such as a fastening element  76  secures the disc  74  for rotation with the shaft  32 . 
         [0042]    In one example, a seal assembly  78  is provided to seal the disc  74  relative to the static structure  36 . The seal assembly  78  includes a seal  80 , such as knife edge seals, to seal relative to a land  82  supported by the static structure  36 . 
         [0043]    A cavity  84  is provided between the disc  74  and rotor  68 . Fluid F from a space  86  provided radially between the shaft  32  and rotor  68  is communicated to the cavity  84  for cooling the turbine blade  66 . In one example, a tangential on-board injector (TOBI)  88  communicates a first fluid  90  to the space  86 . Second fluid  94  from a high pressure compressor aft hub  92  is also provided to the space  86 . The first and second fluids  90 ,  92  are fluid leaked past various seals through leak paths, and are insufficient to cool the turbine blade  66 . 
         [0044]    A bleed air source  96  provides a third fluid  98  that mixes with the first and second fluids  90 ,  94  in the space  86 . The bleed air source  96  is typically the lowest pressure bleed air sufficient to deliver sufficient cooling fluid F to the turbine blade  66 . In the example, the bleed air is provided by one of the latter stages in the high pressure compressor  52 . 
         [0045]    To enable a lower pressure bleed air source to be used, an impeller  100  is provided on the disc  74  within the cavity  84 . The impeller  100  includes a wall  102  on which the first and second radially extending paddles  104 ,  106  are arranged. The first paddles  104  extend from an inlet  108  to an outlet  110 . The first paddles  104  are larger than the second paddles  106 , which are arranged circumferentially between the first paddles  104  in an interleaved relationship. The first and second paddles  104 ,  106  are shaped to fill the cavity  84  and provide an aerodynamic shape that increases the pressure of the fluid F. The impeller  100  increases the pressure of the fluid F and reduces the pressure loss of the fluid F as it reaches the cooling passage  114 . In one example, the paddles are cast as a unitary structure with the wall  102 . 
         [0046]    An annular flange  112  extends axially forward from the wall  102  to provide an annular channel  113  between the outlet  110  and the annular flange  112 . A seal  116  is provided on the annular flange  112  and engages the turbine blades  66 . The fluid F is delivered from the annular channel  113  after its pressure has been increased relative to the pressure of the fluid within the space  86  and delivered to the cooling passage  114  within the turbine blade  66 . 
         [0047]    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 the claims. For that reason, the following claims should be studied to determine their true scope and content.