Abstract:
An integrally bladed rotor has an outer rim with a plurality of blades extending radially outwardly of the outer rim. A plurality of channels are formed radially inwardly of the outer rim. A discontinuity formed at a radially outer surface of the outer rim includes a first thin slot at a radially outer face of the outer rim with an enlarged seal holding area. A second thin slot is positioned radially inwardly of the seal holding. The first and second thin slots are thinner circumferentially than the enlarged seal holding area. A seal is inserted into the seal holding area. The seal does not extend into the first and second thin slots, nor into the channels.

Description:
RELATED APPLICATIONS 
     This application is a continuation-in-part of U.S. Ser. No. 11/965,883 filed Dec. 28, 2007 now U.S. Pat. No. 9,133,720. 
    
    
     BACKGROUND OF THE INVENTION 
     This application relates to an integrally bladed rotor, such as utilized in gas turbine engines, wherein an outer rim has a discontinuity. 
     Gas turbine engines typically include a plurality of sections mounted in series. A fan section may deliver air to a compressor section. The compressor section may include high and low compression stages, and delivers compressed air to a combustion section. The air is mixed with fuel in the combustion section and burned. Products of this combustion are passed downstream over turbine rotors. 
     The compressor section includes a plurality of rotors having a plurality of circumferentially spaced blades. Recently, these rotors and blades have been formed as an integral component, called an “integrally bladed rotor.” 
     In one known integrally bladed rotor, blades extend from an outer rim. The outer rim in integrally bladed rotors is subject to a number of stresses, and in particular, hoop stresses. The hoop stresses can cause the life of the integrally bladed rotor to be reduced due to thermal fatigue. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, an integrally bladed rotor has an outer rim with a plurality of blades extending radially outwardly of the outer rim. A plurality of channels are formed radially inwardly of the outer rim. A discontinuity is formed at a radially outer surface of the outer rim, including a first thin slot at a radially outer face of the outer rim with an enlarged seal holding area. A second thin slot is positioned radially inwardly of the seal holding. The first and second thin slots are thinner circumferentially than the enlarged seal holding area. A seal is inserted into the seal holding area, but does not extend into the first and second thin slots nor into the channels. 
     In another embodiment according to the previous embodiment, there are a plurality of discontinuities formed between the blades. 
     In another embodiment according to any of the previous embodiments, the discontinuity is formed non-parallel to a blade stacking line of the blades. 
     In another embodiment according to any of the previous embodiments, the discontinuity is formed non-parallel to a central axis of the integrally bladed rotor. 
     In another embodiment according to any of the previous embodiments, the discontinuity is formed such that it shields a leading edge of one of the blades and a trailing edge of the adjacent blade. 
     In another embodiment according to any of the previous embodiments, the channels are formed non-parallel to a blade stacking line of the blades. 
     In another embodiment according to any of the previous embodiments, the channels are formed non-parallel to a central axis of the integrally bladed rotor. 
     In another embodiment according to any of the previous embodiments, the channels are formed radially inward of the blades. 
     In another embodiment according to any of the previous embodiments, the channels include a first essentially flat portion on a radially inward side of the channel, and a second essentially flat portion on a radially outward side of the channel. 
     In another embodiment according to any of the previous embodiments, the leading edges and trailing edges of the blades are softened. 
     In another featured embodiment, a gas turbine engine has a compressor section including at least one rotor with a plurality of blades. The blades are non-parallel to a central axis of said integrally bladed rotor, with the at least one rotor being an integrally bladed rotor. The compressor delivers compressed air downstream into a combustion section. The combustion section delivers products of combustion downstream across a turbine rotor. 
     In another embodiment according to the previous embodiment, the integrally bladed rotor of the compression section has an outer rim with a plurality of blades extending radially outwardly of the outer rim. A discontinuity is formed at a radially outer surface of the outer rim, including a first thin slot at a radially outer face of the outer rim with an enlarged seal holding area. A second thin slot is positioned radially inwardly of the seal holding. The first and second thin slots are thinner circumferentially than the enlarged seal holding area. A seal is inserted into the seal holding area but does not extend into the first and second thin slots nor into the channels. A plurality of channels is formed radially inwardly of the outer rim. 
     In another embodiment according to any of the previous embodiments, there are a plurality of discontinuities formed between the blades. 
     In another embodiment according to any of the previous embodiments, the discontinuity is formed non-parallel to a blade stacking line of the blades. 
     In another embodiment according to any of the previous embodiments, the discontinuity is formed non-parallel to a central axis of the integrally bladed rotor. 
     In another embodiment according to any of the previous embodiments, the discontinuity is formed such that it shields a leading edge of one of the blades and a trailing edge of the adjacent blade. 
     In another embodiment according to any of the previous embodiments, the channels are formed non-parallel to a blade stacking line of the blades. 
     In another embodiment according to any of the previous embodiments, the channels are formed non-parallel to a central axis of the integrally bladed rotor. 
     In another embodiment according to any of the previous embodiments, the channels are formed radially inward of a leading edge of the blades and a trailing edge of the blades. 
     In another embodiment according to any of the previous embodiments, the channels include a first essentially flat portion on a radially inward side of the channel, and a second essentially flat portion on a radially outward side of the channel. 
     In another featured embodiment, an integrally bladed rotor has an outer rim with a plurality of blades extending radially outwardly of the outer rim. A plurality of channels is formed radially inwardly of the outer rim. A plurality of discontinuities is formed at a radially outer surface of the outer rim and between the blades, each including a first thin slot at a radially outer face of the outer rim with an enlarged seal holding area. A second thin slot is positioned radially inwardly of the seal holding area. The first and second thin slots are thinner circumferentially than the seal holding area. A seal is inserted into the seal holding area but does not extend into the first and second thin slots nor into the channels. The discontinuities are formed non-parallel to a blade stacking line of the blades. The discontinuities are formed non-parallel to a central axis of the integrally bladed rotor The channels are formed non-parallel to a blade stacking line of the blades. The channels are formed non-parallel to a central axis of the integrally bladed rotor. The channels include a first essentially flat portion on a radially inward side of the channel, and a second essentially flat portion on a radially outward side of the channel. 
     These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows a gas turbine engine. 
         FIG. 2  shows an integrally bladed rotor according to an embodiment of the present invention. 
         FIG. 3  shows a perspective of the  FIG. 2  inventive integrally bladed rotor. 
         FIG. 4A  is a detail view of the  FIG. 2  integrally bladed rotor. 
         FIG. 4B  shows a further detail of  FIG. 4A . 
         FIG. 5  shows a perspective of the  FIG. 4  integrally bladed rotor. 
     
    
    
     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 flowpath B while the compressor section  24  drives air along a core flowpath C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . 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 including three-spool architectures. 
     The 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. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low pressure compressor  44  and a low pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through 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 compressor  52  and high pressure turbine  54 . A combustor  56  is arranged 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. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. 
     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 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 5. 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 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 2.5:1. 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. 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 (&#39;TSFC&#39;)”—is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf 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 [(Tambient deg 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. 
       FIG. 2  shows an integrally bladed rotor  80 , such as may be utilized for the high stage compression section. The integrally bladed rotor  80  includes an outer rim  82 , a plurality of circumferentially distributed blades  84 , a central hub  48 , and a plurality of channels  86 . The channels  86  extend through the axial width of the rotor  80 . The radial dimension of lugs  125  between channels  86  is small in order to reduce dead weight outside of the outer rim  82 . 
     Channels  86  and discontinuities  88 ,  90  and  92  (see  FIGS. 3 through 5 ) address the hoop stresses discussed earlier. 
     As is clear from  FIG. 3 , the discontinuity  88  and channel  86  lies on plane  125 . The aerodynamic stacking line of the blades  84  and associated airfoils lie on plane  130 . Planes  125  and  130  are not parallel to one another or the central axis  140  of the integrally bladed rotor  80 . This arrangement serves to relieve stress on the leading  300  and trailing  301  edges of the blades  84  while still providing support for the center of the integrally bladed rotor  80 . The angle of the discontinuities  88  on plane  125  relative to the aerodynamic stacking line of the blades  84  on plane  130  can allow the discontinuity to shield the leading edge  300  of one blade from stress and the trailing edge  301  of the adjacent blade from stress. 
       FIG. 4A  shows integrally bladed rotor  80 . In integrally bladed rotor  80 , a discontinuity  88 ,  90 ,  92  is formed through a radial extent of the outer rim  82 . 
     As shown, a central enlarged, seal holding portion  90  is formed between two smaller slots  88  and  92 . As can be appreciated from  FIGS. 4A-B , a radially inner slot  92  extends to the channel  86 . As is clear from  FIGS. 4A-B , the slots  88  and  92  extend for a thinner circumferential extent than does a seal holding portion  90 . 
     As is shown in  FIG. 3  and  FIGS. 4A-B , the cavities  86  are formed radially inward from the blades  84 . The edges of the blades are softened in order to reduce thermally driven stresses at the edges and reduce the thickness of the outer rim  82 . 
       FIG. 4B  shows that the channels  86  have a modified oval shape. The channel has essentially flat portions  115  and  120  where it meets the outer rim  82  on the radially outer side. The channel  86  also has an essentially flat portion  110  on the radially inner side. Curved edges  130  connect the essentially flat portions. These essentially flat portions serve to reduce the high stresses due to centrifugal force at these areas. Other modified shapes with essentially flat portions could be used as well. 
     As shown in  FIG. 5 , the outer slot  88  extends across the axial width of the rotor  80 . Seals  96  may be inserted in the enlarged portion  90  of the discontinuity. The seal  96  is shown as a wire seal, however, other seals, such as brush seals, W seals or feather seals, may be utilized. The seals prevent recirculation of gases from the radially outer face of the outer rim  82  into the channels  86 . As is clear from  FIG. 5 , the seal material  96  is inserted into the seal holding portion  90 , and not into the slots  88  and  92 . In addition, the channel  86  does not receive the seal material. 
     Although embodiments of this invention have 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.