Abstract:
A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve for selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve is operable to selectively block flow of either of the lower pressure and higher pressure tapped paths to the bore of the disc, with the disc including holes to allow air from compressor chambers to communicate with the bore of the disc. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

Description:
CROSS REFERENCE TO RELATED APPLICATION 
       [0001]    This application claims priority to U.S. Provisional Application No. 62/075,281 which was filed on Nov. 5, 2014. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    This application relates to extracting compressed air for thermal conditioning of a high pressure compressor rotor. 
         [0003]    Gas turbine engines used on aircraft typically include a fan delivering air into a bypass duct and into a compressor section. Air from the compressor is passed downstream into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. 
         [0004]    Turbine rotors drive compressor and fan rotors. Historically, the fan rotor was driven at the same speed as a turbine rotor. More recently, it has been proposed to include a gear reduction between the fan rotor and a fan drive turbine. With this change, the diameter of the fan has increased dramatically and a bypass ratio or volume of air delivered into the bypass duct compared to a volume delivered into the compressor has increased. With this increase in bypass ratio, it becomes more important to efficiently utilize the air that is delivered into the compressor. 
         [0005]    One factor that increases the efficiency of the use of this air is to have a higher pressure at the exit of a high pressure compressor. This high pressure results in a high temperature increase. The temperature at the exit of the high pressure compressor is known as T 3  in the art. 
         [0006]    There is a stress challenge to increasing T 3  on a steady state basis due largely to material property limits called “allowable stress” at a given maximum T 3  level. At the maximum, a further increase in a design T 3  presents challenges to achieve a goal disk life. In particular, as the design T 3  is elevated, a transient stress in the disk increases. This is true since the radially outer portions of a high pressure compressor rotor (i.e., the blades and outermost surfaces of the disk or blisk), which are in the path of air, see an increased heat rapidly during a rapid power increase. Such an increase occurs, for example, when the pilot increases power during a take-off roll. However, a rotor disk bore does not see the increased heat as immediately. Similar high stresses occur with a change from high power back to low, when the outer rim cools more quickly than the rotor bore. Thus, there are severe stresses due to the thermal gradient between the disk bore and the outer rim region. 
         [0007]    Thermal gradient challenges are greatest during large changes in power setting. For instance, when an associated aircraft moves from idle to take-off, or cruise to decent. It is possible that the thermal stress in the disk is much greater than the stress due to the centrifugal force on the disk. The engine has typically been at low speed or idle as the aircraft waits on the ground and then, just before take-off, the speed of the engine is increased dramatically. Disk thermal gradient stresses may result in a compressor design that cannot achieve desired pressures. 
       SUMMARY OF THE INVENTION 
       [0008]    In a featured embodiment, a compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve for selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve is operable to selectively block flow of either of the lower pressure and higher pressure tapped paths to the bore of the disc, with the disc including holes to allow air from compressor chambers to communicate with the bore of the disc. 
         [0009]    In another embodiment according to the previous embodiment, the lower temperature tapped path passes through a heat exchanger before reaching the valve. 
         [0010]    In another embodiment according to any of the previous embodiments, the valve allows flow from the higher temperature tapped path when the associated engine is moving from a lower power operation to a higher power operation. 
         [0011]    In another embodiment according to any of the previous embodiments, the valve delivers the lower temperature tapped path to the bore of the disc when the engine is at higher power operation. 
         [0012]    In another embodiment according to any of the previous embodiments, the valve blocks flow of both the higher pressure and lower pressure tapped paths when the engine is operating at other lower power settings. 
         [0013]    In another embodiment according to any of the previous embodiments, air from the compressor chamber passes radially inwardly through the holes in the disc, and into the bore of the disc when the valve blocks flow of both the higher pressure and lower pressure paths. 
         [0014]    In another embodiment according to any of the previous embodiments, the valve communicates the higher temperature tap path radially inwardly of the blades, through the disc, and into the bore when the engine is operating at other lower power settings. 
         [0015]    In another embodiment according to any of the previous embodiments, a seal blocks flow of air at a radially inner portion of the compressor section from passing upstream. 
         [0016]    In another featured embodiment, a gas turbine engine comprises a compressor section, a combustor, and a turbine section. The compressor section includes a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve for selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve is operable to selectively block flow of either of the lower pressure and higher pressure tapped paths to the bore of the disc, with the disc including holes to allow air from compressor chambers to communicate with the bore of the disc. 
         [0017]    In another embodiment according to the previous embodiment, the lower temperature tapped path passes through a heat exchanger before reaching the valve. 
         [0018]    In another embodiment according to any of the previous embodiments, the valve allows flow from the higher temperature tapped path when the associated engine is moving from a lower power operation to a higher power operation. 
         [0019]    In another embodiment according to any of the previous embodiments, the valve delivers the lower temperature tapped path to the bore of the disc when the engine is at higher power operation. 
         [0020]    In another embodiment according to any of the previous embodiments, the valve blocks flow of both the higher pressure and lower pressure tapped paths when the engine is operating at other lower power settings. 
         [0021]    In another embodiment according to any of the previous embodiments, air from the compressor chamber passes radially inwardly through the holes in the disc, and into the bore of the disc when the valve blocks flow of both the higher pressure and lower pressure paths. 
         [0022]    In another embodiment according to any of the previous embodiments, the valve communicates the higher temperature tap path radially inwardly of the blades, through the disc, and into the bore when the engine is operating at other lower power settings. 
         [0023]    In another embodiment according to any of the previous embodiments, a seal blocks flow of air at a radially inner portion of the compressor section from passing upstream. 
         [0024]    In another featured embodiment, a method of operating a gas turbine engine includes the steps of tapping air from a compressor section exit, the compressor section having a rotor including blades and a disc, with a bore defined radially inwardly of the disc. The high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve selectively blocks flow of either, or both, of the lower pressure and higher pressure tapped paths to the bore of the disc. 
         [0025]    In another embodiment according to the previous embodiment, the lower temperature tapped path passes through a heat exchanger before reaching the valve. 
         [0026]    In another embodiment according to any of the previous embodiments, air from the compressor chamber passes radially inwardly through holes in the disc, and into the bore of the disc when the valve blocks flow of both the higher pressure and lower pressure paths. 
         [0027]    In another embodiment according to any of the previous embodiments, the valve communicates the higher temperature tap path radially inwardly of the blades, through the disc, and into the bore when the engine is operating at some lower power settings. 
         [0028]    These and other features may be best understood from the following drawings and specification. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0029]      FIG. 1  schematically shows a gas turbine engine. 
           [0030]      FIG. 2A  shows details of a compressor section in a first condition. 
           [0031]      FIG. 2B  shows the  FIG. 2  compressor section in a second operational condition. 
           [0032]      FIG. 2C  shows the  FIG. 2  compressor section in a third operational condition. 
           [0033]      FIG. 3  shows a second embodiment of the third operational condition 
       
    
    
     DETAILED DESCRIPTION 
       [0034]      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. 
         [0035]    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. 
         [0036]    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. 
         [0037]    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 . 
         [0038]    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. 
         [0039]    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 (10,668 meters). 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). 
         [0040]      FIG. 2A  shows a high pressure compressor section  100 . While a number of stages are illustrated, this disclosure focuses on the most downstream stages. Hubs or discs  102  and  103  are shown mounting a pair of blades  104  and  106 . As known, a temperature T 3  is defined downstream of an end blade  104 . As mentioned above, it is desirable to increase the T 3 , however, there are real world challenges in doing so. In particular, the temperatures of the compressed air being moved by the blades  104  and  106  heats the outer peripheral portions (such as the outer rim surface  102 A) of the high pressure compressor  100  much more rapidly than bores  102 B of the discs  102  and  103  heat. This can cause challenges as mentioned above. 
         [0041]    In the past, air has been tapped from the compressor stages radially inwardly through the disc at upstream locations and delivered to preheat the downstream areas, such as bores of the discs  102  and  103 . However, tapping the air radially inwardly through the hub decreases the pressure and, thus, the efficiency of the preheating. This becomes particularly acute as one moves to more downstream locations, such as the discs  102  and  103 . 
         [0042]    As shown in  FIG. 2A , the compressor  100  has two air taps  107  and  108  which are taken from the compressor exit  109 . Air tap  108  passes through a heat exchanger  110  where the air is cooled. Thus, air tap  107  is a non-cooled high pressure air that will be at a higher temperature than the air in tap  108  which is a cooler high pressure air. Both taps  107  and  108  pass to a valve  111 . Downstream of the valve  111 , the air may flow, as shown at  112 , radially inwardly of an inner end  113  of the hubs  102  and  103  and then through anti-vortex tubes  114 , and holes  115  in the hubs  102  and  103 . In this configuration, flowing radially outward, tubes  114  serve to increase the pressure of the flow like a centrifugal compressor. The air also flows, as shown at  117 , to cool the turbine section (see  FIG. 1 ). 
         [0043]    The vortex tubes  114  are positioned in chambers  119 . The chambers  119  are desirably preheated during certain conditions to address the stresses as mentioned above.  FIG. 2A  shows an initial transition from a low power setting as the engine anticipates moving to a high power setting, such as takeoff. A control  120  for the valve  111  moves the valve such that majority of the air, if not all of the air, delivered into the chambers  119  is the non-cooled higher temperature air from tap  107 . The air passing into the chambers  119  and through the anti-vortex tubes  114  preheats the hubs  102  and  103 . 
         [0044]    In fact, the valve  111  would likely have initially been in this position during idle and taxi, so that valve movement may not actually be required prior to the  FIG. 2A  time. The valve  111  may stay in the  FIG. 2A  position for a portion of time after the engine is transitioning to a higher power setting. 
         [0045]    A seal  121  blocks the flow of air to more upstream locations, such that it is directed to the chambers  119 . 
         [0046]    As can be appreciated, the flow paths shown in this figure are relatively schematic. An appropriate valve and communication structure would be well within the ability of a worker in the art, given the disclosure. 
         [0047]      FIG. 2B  shows the operation at higher temperature power settings, such as takeoff through climb, for example. In this position, the control  120  has moved the valve  117  such that the air delivered to  112  inwardly of radially inner end  113 , vortex tubes  114  and chambers  119  and holes  115  is now the cooled high pressure air from tap  108 . At the higher pressure settings, cooling is desirably provided. 
         [0048]    As mentioned above, the change from  FIG. 2A  to  FIG. 2B  may not occur simultaneously with the beginning of the higher power operation of the associated engine. Instead, there may be a short duration of time at high power before the valve moves to the  FIG. 2B  position. 
         [0049]      FIG. 2C  shows operation at stabilized low temperature power settings, such as cruise, In the  FIG. 2C  position, the control  120  has moved the valve  111  such that it blocks flow from both taps  107  and  108 . In this position, air may pass from the holes  115 , but now radially inwardly from compressor chambers  122  and into the chambers  119 . The air flows as shown at  124  through the vortex tubes  114 , along the radially inner end  113  of the compressor, and as shown at  126  and  128  passes downstream toward the turbine section. This provides more efficient operation in that the high pressure compressed air is not taken through the valve  111  in this position. The efficiency benefit is because the air is being extracted at a lower pressure compressor location. The flow is thus sourced from a less expensive stage/location. 
         [0050]    Thus, this embodiment reverses the flow of air such that under certain conditions and, typically, conditions leading up to high power and high power, a high pressure air source is passed through a valve  111  and then radially outward to activate chambers  119  and pre-condition the rotor disks  102 B. At lower power settings, the air passes radially inwardly through the compressor chambers  122  into the chambers  119 . 
         [0051]      FIG. 3  shows another embodiment  130 . Embodiment  130  operates in the  FIGS. 2A and 2B  conditions in a similar fashion. That is, the  FIG. 3  valve would communicate the hot air source  107  inwardly, and to preheat the chambers  119 , as in the  FIG. 2A  operation. Again, at some point, the valve  132  would then be moved to communicate the air source  108  to the chambers inwardly to the chambers  119 . As in the prior embodiment, the change may happen shortly after being moved to high power operation of the associated engine. 
         [0052]    However, in the condition shown in  FIG. 2C , the embodiment  130  operates differently from the embodiments of  FIGS. 2A-2C . Here, the control  134  directs at least the hotter air source  107  through a valve  132  that may be actually mounted in the compressor exit  133 . The control  134  may move the valve to tap the air as shown at  136  radially inward of the blades  104  and  106 . The air then passes through the ports  138  in the hubs  102  and  103 , and into anti-vortex tubes  140  before passing along the inner periphery at  142 , and to the turbine at  144 . 
         [0053]    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.