Abstract:
An example method of cooling a compressor section of a gas turbine engine includes diverting a flow from a compressor through a heat exchanger, the flow moving from the compressor in a first direction, and moving the flow from the heat exchanger back to the compressor in a second direction. An example spacer for a compressor of a gas turbine engine includes a first side portion, a second side portion spaced apart from the first side portion, and a middle web arranged between the first and second side portions. At least one of the first and second side portions and the middle web include at least one orifice to communicate flow in a direction that is different from a core flowpath flow direction. An example compressor including the spacer is also disclosed.

Description:
BACKGROUND 
       [0001]    This disclosure generally relates to a cooling arrangement for a compressor. 
         [0002]    A gas turbine engine typically includes 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 a fan section or other engine loads. The compressor section may include low and high pressure compressors. 
         [0003]    The compressor section, and especially the high pressure compressor, is subject to high temperatures during engine operation. This affects the lifetime of the compressor section. In order to achieve a desired service lifetime, the compressor section temperature, and thus pressure are limited. However, higher operating pressures may improve the efficiency of the compressor section and overall efficiency of the engine. Some compressor sections may thus employ various cooling arrangements to reduce the temperatures of certain components while still operating at relatively high temperatures. 
       SUMMARY 
       [0004]    A method of cooling a compressor section of a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, diverting a flow from a compressor through a heat exchanger, the flow moving from the compressor in a first direction, and moving the flow from the heat exchanger back to the compressor in a second direction. 
         [0005]    In a further non-limiting embodiment of the foregoing method of cooling a compressor section, the method further comprises the step of removing a first amount of thermal energy from the flow by the heat exchanger. 
         [0006]    In a further non-limiting embodiment of any of the foregoing methods of cooling a compressor section, the method further comprises the step of removing a second amount of thermal energy from the flow by the heat exchanger, the second amount different from the first amount. 
         [0007]    In a further non-limiting embodiment of any of the foregoing methods of cooling a compressor section, the flow moves from the heat exchanger to a rim of an aftmost stage of the compressor. 
         [0008]    In a further non-limiting embodiment of any of the foregoing methods of cooling a compressor section, the first direction is an axial direction and the second direction is an axial direction opposite from the first axial direction. 
         [0009]    In a further non-limiting embodiment of any of the foregoing methods of cooling a compressor section, the method further comprises the step of moving a portion of the flow from the heat exchanger to a compressor hub in the first axial direction. 
         [0010]    In a further non-limiting embodiment of any of the foregoing methods of cooling a compressor section, the flow is diverted from a midpoint of a core airflow through the compressor. 
         [0011]    A spacer for a compressor of a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a first side portion,
       a second side portion spaced apart from the first side portion, and a middle web arranged between the first and second side portions. At least one of the first and second side portions and the middle web include at least one orifice to communicate flow in a direction that is different from a core flowpath flow direction.       
 
         [0013]    In a further non-limiting embodiment of the foregoing spacer, the middle web includes at least one orifice to communicate flow in a direction that is opposite from the core flowpath flow direction. 
         [0014]    In a further non-limiting embodiment of any of the foregoing spacers, one of the first and second side portions includes at least one orifice in a direction that is perpendicular to the core flowpath direction. 
         [0015]    In a further non-limiting embodiment of any of the foregoing spacers, the at least one orifice include a valve, the valve configured to vary a flowrate of the flow through the at least one orifice. 
         [0016]    In a further non-limiting embodiment of any of the foregoing spacers, the flow is radially inside a core flowpath of the compressor. 
         [0017]    In a further non-limiting embodiment of any of the foregoing spacers, the first side portion is parallel to the second side portion. 
         [0018]    A compressor for a gas turbine engine according to an exemplary aspect of the present invention includes, among other things, a first compressor stage, a second compressor stage, and a spacer arranged between the first and second compressor stages. The spacer including a first side portion, a second side portion spaced apart from the first side portion, and a middle web arranged between the first and second side portions, wherein at least one of the first and second side portions and the middle web includes at least one orifice. 
         [0019]    In a further non-limiting embodiment of the foregoing compressor, one of the first and second compressor stages is the aftmost compressor stage of a high pressure compressor. 
         [0020]    In a further non-limiting embodiment of the foregoing compressor, the spacer is received between first and second rims of the first and second compressor stages, respectively. 
         [0021]    In a further non-limiting embodiment of any of the foregoing compressors, the at least one orifice includes a valve, the valve configured to vary a flowrate of the flow through the at least one orifice. 
         [0022]    In a further non-limiting embodiment of any of the foregoing compressors, the first side portion is arranged radially outward from the second side portion. 
         [0023]    In a further non-limiting embodiment of any of the foregoing compressors, the second side portion and the middle web include first and second orifices, respectively. 
         [0024]    In a further non-limiting embodiment of any of the foregoing compressors, the first and second side portions and the middle web include first and second sets of orifices, respectively. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0025]      FIG. 1  schematically illustrates an example gas turbine engine. 
           [0026]      FIG. 2  illustrates a section and partial schematic view of a portion of a high pressure compressor of the engine of  FIG. 1 . 
           [0027]      FIG. 3  schematically illustrates a close-up view of a portion of the high pressure compressor of  FIG. 2 . 
           [0028]      FIG. 4 a    illustrates a spacer for the high pressure compressor of  FIGS. 2-4 . 
           [0029]      FIG. 4 b    illustrates a cutaway view of a portion of the spacer of  FIG. 5   a.    
           [0030]      FIG. 4 c    illustrates a close-up cutaway view of a portion of the spacer of  FIG. 5   b.    
           [0031]      FIG. 5  schematically illustrates a close-up view of a portion of the high pressure compressor blades of  FIG. 3 . 
       
    
    
     DETAILED DESCRIPTION 
       [0032]      FIG. 1  schematically illustrates an example gas turbine engine  20 . The example gas turbine engine  20  of  FIG. 1  is a two-spool turbofan that generally incorporates 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 flowpath while the compressor section  24  drives a core airflow C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . The compressor section  24  may include a low pressure compressor  44  and a high pressure compressor  52 . In this example, the gas turbine engine  20  is a geared gas turbine engine wherein the fan section  22  rotates at a different speed than the turbine section  28 . However, the examples in this disclosure are not limited to implementation in the geared gas turbine architecture described, and may be used in other architectures such as a direct drive two-spool gas turbine engine, a three-spool gas turbine engine, or a single spool turbojet. 
         [0033]    There are various types of gas turbine engines, and other turbomachines, that can benefit from the examples disclosed herein. Also, although depicted as a 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 turbofans as the teachings may be applied to other types of turbine engines. 
         [0034]    Referring to  FIGS. 2-4   c  with continuing reference to  FIG. 1 , a high pressure compressor  52  of the compressor section  24  includes several stages  60 ,  62 ,  64 . In the example shown, stage  64  is the aftmost stage. The stages  60 ,  62 ,  64  are connected to one another by way of a tie rod  66  assembly. In another example, the stages  60 ,  62 ,  64  may be interconnected by bolted assemblies, welded assemblies, or by other fastening means. While a high pressure compressor  52  is shown, it should be understood that the examples in this disclosure may be used in any other type of compressor, such as the low pressure compressor  44 , or an intermediate pressure compressor (for the three-spool gas turbine engine). 
         [0035]    Each of the stages  60 ,  62 ,  64  includes a disc  68  with a rim  70  at the disc  68  periphery. A blade  72  is attached to the rim  70 . Between each of the discs  68  are air spaces known as bores  74 . Between each of the rims  70  are spacers  76 . The spacers  76  may support stators  77  ( FIG. 3 ). In another example, cantilevered stators interface with the spacers  76 . 
         [0036]    During operation, the core airflow C flows past the blades  72  and is compressed. Core airflow C exits the compressor  52  from the aftmost stage  64 . A portion of the core airflow C may be drawn off into a cooling stream D. As is shown in  FIG. 2 , in one example, the cooling stream D is drawn from the midpoint of the core airflow C flow path. This allows that the highest pressure and lowest temperature air from the core airflow C is provided to the cooling stream D. Cooling stream D may also be drawn from any radial point of the core airflow C flow path (i.e. any point other than the midpoint). In another example, cooling stream D may be drawn off from an upstream (i.e. axially forward) compressor stage  60 ,  62 . The cooling stream D may be less than 3% of the mass flow of the core airflow C exiting the compressor  52 . 
         [0037]    The cooling stream D may be used to provide initial cooling to the aftmost stage  64  of the compressor  52 . However, as the cooling stream D heats up due to heat exchange from the hot compressor  52 , additional cooling air may be routed from bores  74  radially outward to supplement the cooling stream D. In one example, additional cooling air may be radially provided from the bores  74  to each stage  60 ,  62 ,  64 . This additional cooling air serves to make up any losses due to leakage within the compressor  52  as well as provide the coolest air to the forward-most stages of the compressor  52 . 
         [0038]    In the example shown in  FIG. 2 , the cooling stream D passes through a heat exchanger (HEX)  79  to remove thermal energy from the cooling stream D. The heat exchanger  79  may be any type of heat exchanger, for example, an air-air cooler, an oil-air cooler, etc. The amount of thermal energy removed from the cooling stream D by the heat exchanger  79  may be selectively variable, allowing for optimal conditioning of the cooling stream D. For example, in some engine  20  operating modes, the heat exchanger  79  may be turned off so effectively no thermal energy is removed from the cooling stream D. In other modes, the heat exchanger  79  may provide substantial cooling of the cooling stream D by removing a substantial amount of thermal energy. Once cooled, cooling stream D is used to reduce temperature gradients through components of the compressor  52  to improve component lifetimes. 
         [0039]    Conditioned cooling stream D is supplied to the rim  70  of the compressor stage  64 . The conditioned cooling stream D may pass through the spacers  76  and rims  70  and down into the bores  74  between stages  60 ,  62 ,  64 . The conditioned cooling stream D flows progressively in a direction opposite the direction of the core airflow C through the spacers  76  and rims  70  to provide cooling to the rims  70  and to the bores  74 . That is, core airflow C defines a downstream flow direction, while cooling stream D flows in an opposite upstream direction. 
         [0040]    A portion E of the cooling stream D may be diverted to flow down a compressor hub  78 , arranged aft of the last compressor stage  64 . After flowing through the rims  70  and bores  74  or along the hub  78 , the cooling air D and E may be expelled from the compressor  52  and used to cool another part of the engine  20 , such as the turbine section  28 . 
         [0041]      FIG. 3  shows a close up view of a portion of the compressor  52 , and  FIGS. 4 a - c    show the spacer  76 . The spacer  76  is a ring with an “H”-shaped cross section. That is, the spacer  76  includes first and second sections  80 ,  82  with a middle web  84  arranged between the first and second sections  80 ,  82 . In the example of  FIG. 3 , the first and second sections  80 ,  82  are generally parallel to one another, and the web  84  is generally perpendicular to the first and second section  80 ,  82 . However, in another example, the first and second sections  80 ,  82  may not be parallel to one another. In the example shown, the first section  80  is arranged radially inward from the second section  82 . 
         [0042]    The spacer  76  includes axial flow orifices  86  in the middle web  84 , which allows the cooling stream D to flow axially through the compressor  52  to the next of the stages  60 ,  62 ,  64 . The rims  70  include axial orifices  87  as well. The spacer also includes radial flow orifices  88 , which allows the cooling stream D to flow radially through the compressor  52  and down into the bores  74 . In the example shown, the radially inner second parallel section  82  of the spacer  76  includes the radial orifice  88 . The orifices  86 ,  87 ,  88  allow air to pass through the spacer  76  while the air is rotating at or near the speed of the disc  68  rotation. As is shown in  FIGS. 4 a   - c,  there may be more than one orifice  86 ,  88  in the spacer  76  at each compressor stage  60 ,  62 ,  64 . 
         [0043]    In one example, the orifices  86 ,  87 ,  88  may include a variable valve  100  ( FIG. 3 ) in order to provide optimal cooling to the compressor  52 . For example, during certain engine  20  modes, the rims  70  may become particularly hot and all of the cooling stream D may diverted through the axial orifices  86 ,  87  by closing the radial orifices  88 . In turn, the orifices  86 ,  87 ,  88  may be regulated open at idle engine  20  conditions when the compressor  52  cooling is turned off to achieve a more uniform temperature distribution from the blade rims  70  to the bores  74 . The orifices  86 ,  87 ,  88  may include any type of valve, for example, thermostatic or inertia-activated valves. 
         [0044]      FIG. 5  shows a close-up view of a portion of the blade  72 . The blade  72  may extend over the spacer  76 . A seal  90  may be arranged on a radially inward side of the blade  72 , between the blade  72  and the rim  70  and the spacer  76 . The seal  90  prevents the cooling stream D from mixing with the core airflow C to maintain efficiency of the compressor  52 . 
         [0045]    While  FIGS. 2-3 and 5  depict axially-installed compressor blades  72 , it should be understood that the present disclosure can be applied to other types of compressor discs  68 , such as integrally bladed rotors (IBRs). In the case of IBRs, the discs  68  may include holes or slots under the blades  72  to allow the cooling stream D to pass through. 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 invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.