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 radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. 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 Patent Application No. 62/074,112, filed Nov. 3, 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, transient stresses 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. 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, from idle to take-off, cruise to decent, etc. 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 (P 3 , and temperature). 
       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 radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. 
         [0009]    In another embodiment according to the previous embodiment, the valve delivers the lower pressure tap to the bore of the disc at high power settings, including take-off. 
         [0010]    In another embodiment according to any of the previous embodiments, air is delivered through a strut downstream of the valve and into the bore of the disc. 
         [0011]    In another embodiment according to any of the previous embodiments, air is delivered from the strut radially inward of the disc and communicates with a downstream most portion of the disc. 
         [0012]    In another embodiment according to any of the previous embodiments, air is delivered through a strut downstream of the valve and into the bore of the disc. 
         [0013]    In another embodiment according to any of the previous embodiments, air is delivered from the strut radially inward of the disc and communicates with a downstream most portion of the disc. 
         [0014]    In another embodiment according to any of the previous embodiments, the air communicates with a downstream most portion of the disc. 
         [0015]    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 radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor section, and radially outwardly of the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when engine compressor is mounted on an aircraft. 
         [0016]    In another embodiment according to the previous embodiment, the valve delivers the lower pressure tap to the bore of the disc at high power settings, including take-off. 
         [0017]    In another embodiment according to any of the previous embodiments, air is delivered through a strut downstream of the valve and into the bore of the disc. 
         [0018]    In another embodiment according to any of the previous embodiments, air is delivered from the strut radially inward of the disc and communicates with a downstream most portion of the disc. 
         [0019]    In another embodiment according to any of the previous embodiments, air is delivered through a strut downstream of the valve and into the bore of the disc. 
         [0020]    In another embodiment according to any of the previous embodiments, air is delivered from the strut radially inward of the disc and communicates with a downstream most portion of the disc. 
         [0021]    In another embodiment according to any of the previous embodiments, the air communicates with a downstream most portion of the disc. 
         [0022]    In another featured embodiment, a method of operating a gas turbine engine includes the steps of tapping a lower pressure air tap and a higher pressure air tap from two distinct locations within a compressor and radially outwardly of an outer housing, and selectively delivering at least one of the lower pressure tap and the higher pressure tap to a bore of a compressor rotor disc. The higher pressure tap is delivered to the bore prior to an associated aircraft on which the gas turbine engine is mounted, moving to take-off and climb conditions. 
         [0023]    In another embodiment according to the previous embodiments, including the step of delivering the lower pressure tap to the bore of the disc at high power settings including take-off. 
         [0024]    In another embodiment according to any of the previous embodiments, including the step of delivering the tapped air through a strut and into the bore of the disc. 
         [0025]    In another embodiment according to any of the previous embodiments, including the step of delivering air from the strut radially inward of the disc and to a downstream most portion of the disc. 
         [0026]    In another embodiment according to any of the previous embodiments, including the step of delivering the tapped air through a strut and into the bore of the disc. 
         [0027]    In another embodiment according to any of the previous embodiments, including the step of delivering air to a downstream most portion of the disc. 
         [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. 2  shows details of a compressor section in a first condition. 
           [0031]      FIG. 3  shows the  FIG. 2  compressor in a second operational condition. 
       
    
    
     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 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. 
         [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 (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). 
         [0038]      FIG. 2  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 (including the blades) of the high pressure compressor  100  much more rapidly than bores  102 A of the disc  102  heat. This can cause challenges as mentioned above. 
         [0039]    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 . 
         [0040]    This disclosure taps compressed air from locations radially outwardly through an outer housing  107 . Thus, taps  108  and  110  extend through the outer housing  107  at two distinct locations in the high pressure compressor  100 . The tap  108  is shown to be at a lower pressure location than the tap  110 . Both taps pass through a valve  112  controlled by a controller  113 . Downstream of the valve  112 , the tapped air passes through a strut  114  and into a chamber  115  at a location upstream of the upstream most blade  116  of the high pressure compressor. This air passes into a chamber  115  and then radially inwardly along a path  117  radially inward of an innermost surface  118  of the high pressure compressor  100 . The air passing along path  117  is at a relatively high pressure still and, thus, provides good preheating to inner chambers  119  within the bores defined by the discs  102  and  103 . Thus, the challenges mentioned above in the prior art are reduced. 
         [0041]      FIG. 2  shows the valve  112  in a location to communicate the lower pressure tap  108  into the chamber  115 . This position is generally used at high power settings. 
         [0042]    At some lower power settings, such as at idle just before the aircraft is moving toward takeoff and climb, the valve  112  is moved to the position shown in  FIG. 3 . Controller  113  controls the movement of the valve  112 . In this position, higher temperature air from the tap  110  is delivered to preheat the inner chambers  119 . When the aircraft associated with the high pressure compressor  100  begins to move towards the most challenging times, such as takeoff and climb, the temperature gradient across the last compressor stages is reduced due to this preheating. 
         [0043]    As shown at  120 , after the air has preheated the disc bore  102 A, it passes downstream to provide cooling air for a turbine section (see  FIG. 1 ). Since the air being tapped to the inner chambers  119  is at a higher pressure, the air at  120  will also be at a higher pressure, which will provide a greater amount of cooling capacity, and increase the efficiency of the overall engine. 
         [0044]    During challenging times, and in particular, take-off and climb, the switch is moved to the  FIG. 2  position to provide cooling. After the more challenging condition has ended, the valve  112  may be returned to the  FIG. 3  position. 
         [0045]    It should be understood that some modulation between the two airflows, and mixing, could be provided at either position. However, in general, the hotter air from the  FIG. 3  position is the bulk of the air provided at a point in time just prior to take-off and climb. 
         [0046]    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.