Patent Application
20200157959
A gas turbine engine typically includes a fan section, 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 the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction.
A gas turbine engine according to an example of the present disclosure includes a combustor that has a combustor wall and a combustion chamber. The combustor wall has a lip at an exit region of the combustion chamber. There is a circumferential row of vanes adjacent the exit region. Each vane includes a platform and an airfoil section extending from the platform. The platform defines forward and trailing edges and first and second circumferential side edges joining the forward and trailing edges. The forward edge is adjacent the lip of the combustor wall. The lip defines a first annular slot and the forward edges collectively defining a second annular slot. The first and second annular slots together define an annular seal slot. An annular feather seal is entrapped in the annular seal slot between the combustor wall and the platform.
In a further embodiment of any of the foregoing embodiments, the platform is radially inwards of the airfoil section.
In a further embodiment of any of the foregoing embodiments, the lip of the combustor abuts the leading edge of the platform.
In a further embodiment of any of the foregoing embodiments, the annular feather seal is a split ring.
In a further embodiment of any of the foregoing embodiments, the split ring has overlapping ends.
In a further embodiment of any of the foregoing embodiments, the first annular slot is defined by radially inner and outer walls of the combustor wall and the second annular slot is defined by radially inner and outer walls of the platform.
In a further embodiment of any of the foregoing embodiments, the radially outer wall of the first annular slot and the radially outer wall of the second annular slot abut, and the radially inner wall of the first annular slot and the radially inner wall of the second slot are axially spaced apart such that there is a gap there between.
In a further embodiment of any of the foregoing embodiments, at least one of the radially inner wall of the first annular slot or the radially inner wall of the second annular slot is scalloped.
In a further embodiment of any of the foregoing embodiments, the seal slot is radially thicker than the annular feather seal.
A gas turbine engine according to an example of the present disclosure includes a combustor that has a combustor wall and a combustion chamber. The combustor wall has a lip at an exit region of the combustion chamber. There is a vane adjacent the exit region that has a platform and an airfoil section extending along a radial axis from the platform. The platform defines forward and trailing edges and first and second circumferential side edges joining the forward and trailing edges. The vane has rotational play about the radial axis under aerodynamic loads such that the vane moves relative to the combustor between a seated state in which the forward edge abuts the lip of the combustor wall and an unseated state in which there is a divergent gap between the forward edge and the lip of the combustor. The lip defines a first slot that extends circumferentially and the forward edge defines a second slot that extends circumferentially. The first and second slots together define a seal slot. An annular feather seal extends in the seal slot. The annular feather seal is wider than the divergent gap to maintain sealing when in the vane is in the unseated state.
In a further embodiment of any of the foregoing embodiments, the first side of the airfoil section is a suction side, the second side of the airfoil section is a pressure side, the second circumferential side of the platform is located to the second side of the airfoil section, and the divergent gap diverges toward the second circumferential side.
In a further embodiment of any of the foregoing embodiments, the annular feather seal is a split ring.
In a further embodiment of any of the foregoing embodiments, the split ring has overlapping ends.
In a further embodiment of any of the foregoing embodiments, the first slot is defined by first radially inner and outer walls of the combustor wall and the second slot is defined by radially inner and outer walls of the platform.
In a further embodiment of any of the foregoing embodiments, at least one of the radially inner wall of the first annular slot or the radially inner wall of the second annular slot is scalloped.
In a further embodiment of any of the foregoing embodiments, the seal slot is radially thicker than the annular feather seal.
An airfoil according to an example of the present disclosure includes a vane that has a platform and an airfoil section that extends from the platform. The platform defines forward and trailing edges and first and second circumferential side edges that join the forward and trailing edges. The forward edge includes a bearing surface for abutting a lip of a combustor wall. The forward edge defines a slot arc segment for receiving a portion of an annular feather seal.
In a further embodiment of any of the foregoing embodiments, the platform is radially inwards of the airfoil section.
In a further embodiment of any of the foregoing embodiments, the slot arc segment is defined by radially inner and outer walls, and the radially inner wall is scalloped.
The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.
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 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 a 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 may be 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 the low pressure compressor, or aft of the combustor section 26 or even aft of turbine section 28, and fan 42 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 and less than about 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 (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)]{circumflex over ( )}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).
The high pressure turbine 54 includes a circumferential row of turbine vanes 72 adjacent the exit region 70. Each vane 72 includes an inner or first platform 74, an outer or second platform 76, and an airfoil section 78 that spans in a radial direction between the first and second platforms 74/76. A radial view of a portion of the row of turbine vanes 72 is also shown in
The airfoil section 78 includes an airfoil outer wall 80 that delimits the profile of the airfoil section 78. The outer wall 80 defines a leading end 80a, a trailing end 80b, and first and second sides 80c/80d that join the leading and trailing ends 80a/80b. The first and second sides 80c/80d span in the radial direction between first and second ends 80e/80f that are attached, respectively, to the first and second platforms 74/76. In this example, the first side 80c is a suction side and the second side 80d is a pressure side.
The first platform 74 (
The lip 65 defines a first annular slot 82 and the forward edges 74a of the platforms 74 of the vanes 72 collectively define a second annular slot 84. The first and second annular slots 82/84 together define an annular seal slot 86. An annular feather seal 88 is entrapped in the annular seal slot 86 between the combustor wall 62 and the platform 74. As will be described further below, the platform 74 can move axially away from the lip 65, thereby opening a gap through which combustion gases can escape. In this regard, the annular feather seal 88 serves as a secondary seal across the interface between the platform 74 and the combustor wall 62 to prevent the escape of combustion gases from the core gaspath.
As shown in
In particular, the divergent gap 92 presents an unusual sealing challenge because the vanes 72 may dynamically move between the seated and unseated positions during engine operation and the forward edges 74a of the platforms 74 and the lip 65 are non-parallel when in the unseated position. Thus, seals that cannot accommodate dynamic movement or seals that rely on parallel sides may not provide a desired level of sealing. In this regard, the annular feather seal 88 is able to address both the dynamic movement and the non-parallel nature of the divergent gap 92.
For example, the configuration of the feather seal 88 and the seal slot 86 facilitate dynamic sealing of the divergent gap 92. In one example, the seal slot 86 defines a slot radial thickness t1 and the feather seal 88 defines a seal radial thickness t2, where t2 is less than t1. The seal slot 86 also defines a slot axial width w1 and the feather seal 88 defines a seal axial width w2, where w2 is less than w1. That is, the feather seal 88 is smaller in cross-section than the seal slot 86. This permits the feather seal 88 to shift dynamically within the seal slot 86 to accommodate shifts in the position of the vanes 72. Additionally, the seal axial width w2 is larger (i.e., wider) than the divergent gap 92, to maintain sealing when the vane 72 is in the unseated state. For instance, the play in the vanes 72 may be determined or estimated during engine design to determine or estimate the maximum size of the divergent gap 92. The seal axial width w2 is then selected to be larger than the maximum size in order to ensure sealing entirely along the divergent gap 92.
The first and second annular slots 82/84 that define the annular seal slot 86 are also configured to bias the feather seal 88 to a sealed position. For example, the first slot 82 is defined by radially inner and outer walls 82a/82b of the combustor wall 62, and the second slot 84 is defined by radially inner and outer walls 84a/84b of the platform 74. The outer wall 82b and the outer wall 84b abut (at bearing surface 75 and lip 65). The inner wall 82a and the inner wall 84a are axially spaced apart such that there is a gap 94 there between. There is a high pressure region “P” (
To further permit communication of the high pressure in addition to the gap 94, and also reduce weight, at least one of the inner wall 82a or the inner wall 84a may be scalloped.
The feather seal 88 is also configured to dynamically adapt in diametric size to maintain sealing.
Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.
The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.
