Patent Application
20180328207
This disclosure relates to a gas turbine engine, and more particularly to a component having at least one tip vortex creation feature.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
Each of the compressor section and the turbine section may include multiple stages. Each stage typically includes alternating rows of rotating structures called rotor blades followed by stationary structures called stators. The rotor blades create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine. The stators direct the core airflow to the blades to either add or extract energy.
Some gas turbine engines incorporate cantilevered stator designs. Cantilevered stators include a stationary structure that is affixed at a radially outer portion and unsupported at a radially inner portion. A portion of a rotating structure surrounds a tip of each cantilevered stator. A clearance may extend between the tip and the rotating structure. Gas turbine engine efficiency may depend on minimizing this clearance.
A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a static structure that extends between a radially outer portion and a radially inner portion and at least one vortex creation feature formed on the static structure.
In a further non-limiting embodiment of the foregoing component for a gas turbine engine, the component is a cantilevered stator.
In a further non-limiting embodiment of either of the foregoing components for a gas turbine engine, the cantilevered stator is a compressor cantilevered stator.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the cantilevered stator is a turbine cantilevered stator.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one vortex creation feature is formed on a tip of the cantilevered stator.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one vortex creation feature includes a plurality of serrations.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one vortex creation feature includes a plurality of teeth.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one vortex creation feature includes a plurality of grooves.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one vortex creation feature includes a combination of at least one serration, tooth and groove.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one vortex creation feature establishes a tortuous flow path between a tip of the static structure and a rotating structure surrounding the tip.
A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a first stage of cantilevered stators and a second stage of cantilevered stators disposed downstream from the first stage of cantilevered stators. The first stage of cantilevered stators includes first vortex creation features and the second stage of cantilevered stators includes second vortex creation features.
In a further non-limiting embodiment of the foregoing gas turbine engine, the first vortex creation features and the second vortex creation features include one of a serration, a tooth and a groove.
In a further non-limiting embodiment of either of the foregoing gas turbine engines, each of the first stage of cantilevered stators and the second stage of cantilevered stators include a plurality of static structures that extend between a radially outer portion and a radially inner portion.
In a further non-limiting embodiment of any of the forgoing gas turbine engines, a tip is located at the radially inner portion of each of the plurality of static structures. The first vortex creation features and the second vortex creation features are formed on the tips.
In a further non-limiting embodiment of any of the forgoing gas turbine engines, the first vortex creation features are different from the second vortex creation features.
A method of operating a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, forcing an airflow to bypass a tip clearance between a static structure and a rotating structure of the gas turbine engine by generating flow vortices within the tip clearance.
In a further non-limiting embodiment of the foregoing method, the method includes the step of providing at least one vortex creation feature on a tip of the static structure.
In a further non-limiting embodiment of any of the foregoing methods, the step of forcing airflow includes generating a tortuous flow path at a tip of the static structure.
In a further non-limiting embodiment of any of the foregoing methods, the step of generating the tortuous flow path includes providing at least one vortex creation feature on the tip of the static structure.
In a further non-limiting embodiment of any of the foregoing methods, the step of forcing airflow includes forcing the airflow across a portion of the static structure that is radially outward from a tip of the static structure.
The various features and advantages of this 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 gas turbine engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine centerline longitudinal axis A. The low speed spool 30 and the high speed spool 32 may be mounted relative to an engine static structure 33 via several bearing systems 31. It should be understood that other bearing systems 31 may alternatively or additionally be provided.
The low speed spool 30 generally includes an inner shaft 34 that interconnects a fan 36, a low pressure compressor 38 and a low pressure turbine 39. The inner shaft 34 can be connected to the fan 36 through a geared architecture 45 to drive the fan 36 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 35 that interconnects a high pressure compressor 37 and a high pressure turbine 40. In this embodiment, the inner shaft 34 and the outer shaft 35 are supported at various axial locations by bearing systems 31 positioned within the engine static structure 33.
A combustor 42 is arranged between the high pressure compressor 37 and the high pressure turbine 40. A mid-turbine frame 44 may be arranged generally between the high pressure turbine 40 and the low pressure turbine 39. The mid-turbine frame 44 can support one or more bearing systems 31 of the turbine section 28. The mid-turbine frame 44 may include one or more airfoils 46 that extend within the core flow path C.
The inner shaft 34 and the outer shaft 35 are concentric and rotate via the bearing systems 31 about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low pressure compressor 38 and the high pressure compressor 37, is mixed with fuel and burned in the combustor 42, and is then expanded over the high pressure turbine 40 and the low pressure turbine 39. The high pressure turbine 40 and the low pressure turbine 39 rotationally drive the respective high speed spool 32 and the low speed spool 30 in response to the expansion.
The pressure ratio of the low pressure turbine 39 can be pressure measured prior to the inlet of the low pressure turbine 39 as related to the pressure at the outlet of the low pressure turbine 39 and prior to an exhaust nozzle of the gas turbine engine 20. In one non-limiting embodiment, the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 38, and the low pressure turbine 39 has a pressure ratio that is greater than about five (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 disclosure is applicable to other gas turbine engines, including direct drive turbofans.
In this embodiment of the exemplary gas turbine engine 20, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.
Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of [(Tram° R)/(518.7° R)]0.5, where T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
Each of the compressor section 24 and the turbine section 28 may include alternating rows of rotor assemblies and stator assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades 25, while each vane assembly can carry a plurality of stators 27 that extend into the core flow path C. The blades 25 create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine 20 along the core flow path C. The stators 27 direct the core airflow to the blades 25 to either add or extract energy.
This disclosure relates to tip vortex creation features that may be incorporated into one or more components of the gas turbine engine 20. Among other benefits, the exemplary tip vortex creation features can increase gas turbine engine efficiency.
The portion 100 includes multiple stages of alternating rows of rotating rotor blades 25 and stationary stators 27. Each row of rotor blades 25 and stators 27 is circumferentially disposed about the engine centerline longitudinal axis A. Although four stages are depicted, it should be understood that the portion 100 could include a greater or fewer number of stages. The rotor blades 25 are attached to rotating structures 50, such as disks, that rotate about the engine centerline longitudinal axis A to move the rotor blades 25. Each rotating structure 50 includes a rim 52 that supports one or more rotor blades 25. The rotating structure 50 may additionally include a sealing structure 54, such as a rotor seal land or other rotating structure, which extends between the rims 52 of adjacent rotor blades 25.
In this exemplary embodiment, the stators 27 are cantilevered stators. That is, the stators 27 include a static structure 58 that extends into the core flow path C. In one embodiment, the static structure 58 is an airfoil. Each static structure 58 may be affixed to an engine casing 56 at a radially outer portion 60 and is unsupported at a radially inner portion 62. A tip 64 of the radially inner portion 62 of the static structure 58 is disposed adjacent to the rotating structure 50. In one embodiment, the sealing structure 54 surrounds the tips 64. A clearance X extends across the open space between the tip 64 and the rotating structure 50.
One or more of the static structures 58 may include tips 64 having at least one vortex creation feature 66 that is formed on the tips 64 of the stators 27. At least one of the stages of the stators 27 may exclude any vortex creation features 66. For example, in this embodiment, a fourth stage of stators 27-4 is formed without vortex creation features 66 at the tips 64.
Referring to
The flow vortices 74 that are generated within the clearance X force the airflow AF to bypass the clearance X between the tip 64 and the rotating structure 50. The airflow AF is instead forced across a portion of the static structure 58 that is radially outward from the tip 64. In other words, incorporating the vortex creation features 66 into the tip 64 results in the creation of pockets of local turbulent flow vortices 74 that force more airflow AF to pass over the static structure 58 residing in the core flow path C, thereby improving gas turbine engine efficiency through effective tip clearance control. Efficiency benefits may occur based on a higher percentage of flow path airflow being forced onto the static structure 58 to be guided and directed towards the next stage to minimize flow path turbulence.
Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.
It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure.
The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claims should be studied to determine the true scope and content of this disclosure.
This application is a divisional of U.S. patent application Ser. No. 14/764,760, filed Jul. 30, 2015, which is a national stage entry of International Application No. PCT/US2014/013979, filed Jan. 31, 2014, which claims the benefit of U.S. Provisional Application No. 61/760,817, filed Feb. 5, 2013.
| Number | Date | Country | |
|---|---|---|---|
| 61760817 | Feb 2013 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | 14764760 | Jul 2015 | US |
| Child | 16044611 | US |
