Patent Publication Number: US-2016237903-A1

Title: High Pressure Compressor Rotor Thermal Conditioning Using Conditioned Compressor Air

Description:
BACKGROUND OF THE INVENTION 
     This application relates to extracting compressed air for thermal conditioning of a high pressure compressor rotor. 
     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. 
     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. 
     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. 
     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. Thus, there are severe stresses due to the thermal gradient between the disk bore and the outer rim region. 
     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 
     In a featured embodiment, a gas turbine engine comprises a compressor including a disk and a blade. A turbine rotor has a disk and a blade. Turbine conditioning air is supplied to the turbine rotor. The turbine conditioning air passes across the disk of the compressor to condition the disk. 
     In another embodiment according to the previous embodiment, a tie shaft connects the compressor disk and the turbine rotor. The turbine conditioning air passes along an axial length of the tie shaft. 
     In another embodiment according to any of the previous embodiments, a drive shaft extends radially inwardly of the tie shaft and an inner cooling air passes radially between an outer peripheral surface of the drive shaft and an inner peripheral surface of the tie shaft. 
     In another embodiment according to any of the previous embodiments, the turbine conditioning air is received from a compressor air supply. 
     In another embodiment according to any of the previous embodiments, the compressor air supply passes across a tangential on board injector before passing radially inwardly. 
     In another embodiment according to any of the previous embodiments, the air leaving the tangential onboard injector also has a portion passing radially outwardly to cool the disk of the turbine rotor. 
     In another embodiment according to any of the previous embodiments, the turbine conditioning air is received from a compressor air supply. 
     In another embodiment according to any of the previous embodiments, a fan rotor is positioned upstream of the compressor. The fan rotor delivers air into a bypass duct, and toward the compressor as core air. The fan rotor is driven by a second turbine rotor through a gear reduction. 
     In another embodiment according to any of the previous embodiments, the turbine conditioning air is received from a compressor air supply. 
     In another embodiment according to any of the previous embodiments, the air compressor supply passes across a tangential on board injector before passing radially inwardly. 
     In another featured embodiment, a method of operating a gas turbine engine comprises the steps of supplying a turbine conditioning air to a turbine rotor. The turbine conditioning air passes across a disk of a compressor to condition the disk of the compressor. 
     In another embodiment according to the previous embodiment, a tie shaft connects the disk of the compressor to the turbine rotor. The turbine conditioning air passes ng along the tie shaft before passing across the disk of the compressor. 
     In another embodiment according to any of the previous embodiments, an inner cooling air is supplied between an outer peripheral surface of a drive shaft which is radially inwardly of the tie shaft, and an inner peripheral surface of the tie shaft, such that the turbine conditioning air is cooled by the inner cooling air as the turbine conditioning air passes along the axial length of the tie shaft. 
     In another embodiment according to any of the previous embodiments, the turbine conditioning air is received from a compressor air supply. 
     In another embodiment according to any of the previous embodiments, the air compressor supply passes across a tangential on board injector before passing radially inwardly. 
     In another embodiment according to any of the previous embodiments, the air leaving the tangential onboard injector also has a portion passing radially outwardly to cool the disk of the turbine rotor. 
     In another embodiment according to any of the previous embodiments, the turbine conditioning air is received from a compressor air supply. 
     In another embodiment according to any of the previous embodiments, a fan rotor is positioned upstream of the compressor, and delivers air into a bypass duct, and toward the compressor as core air. The fan rotor being driven by a second turbine rotor through a gear reduction. 
     In another embodiment according to any of the previous embodiments, the turbine conditioning air is received from a compressor air supply. 
     In another embodiment according to any of the previous embodiments, the air compressor supply passes across a tangential on board injector before passing radially inwardly. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a gas turbine engine. 
         FIG. 2  shows a prior art arrangement for cooling a compressor disk. 
         FIG. 3  shows a disclosed cooling arrangement. 
     
    
    
     DETAILED DESCRIPTION 
       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. 
     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. 
     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. 
     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 . 
     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. 
     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). 
     A prior art engine  79  is illustrated in  FIG. 2 . A shaft  80  defines an outer surface for receiving a cooling air  81  headed to downstream turbine sections. 
     Compressor rotor blades  82  (only one of which is illustrated) are associated with a compressor disk or rim  84 . A tie shaft  85  connects to first turbine blades  86  (only one of which is illustrated) The first turbine blades  86  are associated with a disk  88 . As known, a combustor  89  sits between disks  84  and  88 . 
     As known, the temperature defined downstream of the blades  82  would be desirably increased. However, there are challenges to doing so. In particular, the temperature of the compressed air being moved by the blades  82  heats the outer peripheral portions, such as the outer rim  84 , much more rapidly than bores within the compressor rotor disk  84 . This can cause challenges as mentioned above in the Background of the Invention section. 
     It is known that air such as shown schematically at  90 , passes across a seal  92  and across the disk  84 . This air passes, as shown at  94 , along an outer surface of the tie shaft  85 . 
     Turbine conditioning air  96  is received from an air supply  97 . Air supply  97  taps the air from a compressor discharge section and passes it through a tangential on board injector  98 , which causes a swirl in the air. This air is then driven inwardly, as shown at  99 , and the airflow  99  mixes with the airflow  94 , and exits through an exit  100  to vane 100V, shown schematically. A further portion of the air at  96  moves radially outwardly and across a seal  101 , as shown at  102 , to cool disk  88 . 
     The airflow  90  is bled directly from the gas path. The gas path temperature reacts very quickly to engine power changes. Thus the airflow at  90  also responds quickly, and subsequently the disk  84  responds quickly. This is not desirable because of thermal gradients. 
       FIG. 3  shows an engine  180 . The shaft  80  again defines an outer surface for a cooling airflow  181 . The last compressor blades  182  rotate with disk  184 . The tie shaft  185  connects the turbine rotor blades  186  and disk  188  to drive the compressor  182 / 184 . 
     The outer flow  102  is generally as shown in  FIG. 2 . However, the air supply  97  now passes as airflow  196  leaving tangential on board injector  198 , which might have a somewhat large cross-sectional area than the  FIG. 2  prior art engine. The air  196  deflects off a surface  188 D of disk  188 . The airflow  196  then passes radially inwardly at  199  scrubbing the turbine disk web. Flow  199  also passes along an axial length of the tie shaft  185  scrubbing its surface. Some airflow  200  from a location  202  downstream of blade  182  passes across seal  203 , as shown at  204 , to mix with the airflow  199 . The mixed air  206  passes across a gap  208  and conditions the disk  184 . 
     When the engine  180  moves from lower power to higher power (as an example from idle to take-off), the air  199  will be hotter than a number of features it will pass along. As an example, the turbine disk surface  188 D and tie shaft  185  act as heat sinks. The heat sinks are heated by air  199 , which cools the air. Thus, when the air reaches the compressor disk  184 , it will be cooler then gaspath temperature, which would slow the thermal response of the compressor disk  184  rim compared to prior art. This will delay the transient behavior of the compressor disk  184 , or condition the disk. This makes the compressor disk thermal response slower, and will better match the response within the bore  209  of the disk. 
     Conversely, when the engine  180  cycles from high power to low power, the heat sinks will act to heat the air flow at  199 , such that the compressor disk  184  will be heated by the air  206 , to better match the response of the bore  209 . As such, the air flow  206  will condition the compressor disk  184  at both conditions. 
     For purposes of the claims, the air  199  is referred to as turbine conditioning air. 
     In contrast to the  FIG. 2  engine, the air  199  passing radially inwardly and downstream of the turbine disk  188  is now utilized to condition the compressor disk  184 . The air  206  has passed along the entire axial length of the tie shaft  185  and has been dramatically cooled/heated. Thus, the airflow  206  is able to better condition the disk  184  and the thermal gradients and stresses, as described above, are less pronounced. 
     While the turbine disk  188  and tie shaft  185  are disclosed as the heat sinks, the air may be routed from the turbine disk  188  to the compressor disk  184 , along or through other components. 
     The disclosed gas turbine engine  180  could be said to have a compressor including a disk  184  and a blade  182 . A turbine rotor has a disk  188  and a blade  186 . Turbine air is supplied to condition the turbine rotor and passes radially inwardly  199 . A tie shaft  185  connects the compressor and the turbine rotor. The air passes along an axial length of the tie shaft  185 . The turbine conditioning air is cooled as it passes along an axial length of the tie shaft  185 . The turbine conditioning air then passes across the disk  184  of the compressor to cool the disk. 
     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.