Patent Application
20150377029
This disclosure relates to a gas turbine engine, and more particularly to a gas turbine engine component that includes at least one curved turbulator.
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.
Due to exposure to hot combustion gases, numerous components of the gas turbine engine may include internal cooling passages that route cooling air through the part. A variety of interior treatments may be incorporated into the internal cooling passages to augment the heat transfer effect and improve cooling. For example, some cooling passages may include pedestals, air-jet impingement, or turbulator treatments.
A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a wall that forms a portion of an outer periphery of at least one cavity and at least one curved turbulator that extends from said wall.
In a further non-limiting embodiment of the foregoing component for a gas turbine engine, the component is one of a blade and a vane.
In a further non-limiting embodiment of either of the foregoing components for a gas turbine engine, the component is a blade outer air seal (BOAS).
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a plurality of curved turbulators are spaced along the wall.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator includes a contiguous body having at least one peak and at least one valley.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the contiguous body provides a smooth surface that excludes any sharp transition areas.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator is sinusoidal shaped.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a row of film cooling holes are spaced from the at least one curved turbulator.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the row of film cooling holes includes a first film cooling hole and a second film cooling hole staggered from said first film cooling hole.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a second turbulator extends from the wall and includes a different shape from the at least one curved turbulator.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator extends across a width of said wall.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, the at least one curved turbulator extends perpendicular to a direction of flow of cooling airflow communicated through the at least one cavity.
A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a first curved turbulator that protrudes into a cavity flow path and a second curved turbulator that protrudes into the cavity flow path at a position that is spaced from the first curved turbulator. A row of film cooling holes are disposed between the first curved turbulator and the second curved turbulator.
In a further non-limiting embodiment of the foregoing component for a gas turbine engine, the row of film cooling holes includes a first film cooling hole and a second film cooling hole that is staggered from the first film cooling hole.
In a further non-limiting embodiment of either of the foregoing components for a gas turbine engine, the first curved turbulator and the second curved turbulator are sinusoidal shaped.
In a further non-limiting embodiment of any of the foregoing components for a gas turbine engine, a pitch between the first curved turbulator and the second curved turbulator is continuously varied.
Nom A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor section in fluid communication with the compressor section and a turbine section in fluid communication with the combustor section. A component extends into a core flow path of at least one of the compressor section and the turbine section, The component includes a wall that forms a portion of an outer periphery of at least one cavity of the component. At least one curved turbulator extends from the wall.
In a further non-limiting embodiment of the foregoing gas turbine engine, the component is an airfoil of the turbine section.
In a further non-limiting embodiment of either of the foregoing gas turbine engines, the component is a blade outer air seal (BOAS).
In a further non-limiting embodiment of any of the foregoing gas turbine engines, the wall is part of a platform of the component.
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 vane 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 vanes 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 vanes 27 direct the core airflow to the blades 25 to either add or extract energy.
Various components of the gas turbine engine 20, including but not limited to the airfoils of the blades 25 and the vanes 27 of the compressor section 24 and the turbine section 28, may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section 28 is particularly subjected to relatively extreme operating conditions. Therefore, some components may require internal cooling circuits for cooling the parts during engine operation.
This disclosure relates to curved turbulators that can be incorporated into the walls of internal cooling cavities of gas turbine engine components. Among other benefits, the exemplary curved turbulators provide reduced stress concentrations and increased flexibility of film cooling hole placement as compared to prior art interior treatments.
In this embodiment, the body portion 52 is representative of an airfoil. For example, the body portion 52 could be an airfoil that extends between inner and outer platforms (not shown) where the component 50 is a vane, or could extend from platform and root portions (also not shown) where the component 50 is a blade. Alternatively, the component 50 could be a non-airfoil component, including but not limited to a blade outer air seal (BOAS), a combustor liner, a turbine exhaust case liner, or any other part that may require dedicated cooling.
A gas path 62 is communicated axially downstream through the gas turbine engine 20 along the core flow path C (see
One or more cavities 72 may be disposed inside of the body portion 52 as part of an internal cooling circuit for cooling portions of the component 50. The cavities 72 may extend radially, axially and/or circumferentially inside of the body portion 52 to establish cooling passages for receiving a cooling airflow 68 to cool the component 50. The cooling airflow 68 may be communicated into one or more of the cavities 72 from an airflow source 70 that is external to the component 50.
The cooling airflow 68 is generally of a lower temperature than the airflow of the gas path 62 that is communicated across the body portion 52. In one particular embodiment, the cooling airflow 68 is a bleed airflow that can be sourced from the compressor section 24 or any other portion of the gas turbine engine 20 that includes a lower temperature and higher pressure than the component 50. The cooling airflow 68 can be circulated through the cavities 72, such as along a serpentine path, to transfer thermal energy from the component 50 to the cooling airflow 68 thereby cooling the component 50. The cooling circuit can include any number of cavities 72. The cavities 72 may be in fluid communication with one another or could alternatively be isolated from one another.
One or more ribs 74 may extend between the first wall 58 and the second wall 60 of the body portion 52. The rib(s) 74 divide the cavities 72 from one another.
As discussed in greater detail below, at least one of the cavities 72 can include one or more curved turbulators 80 that protrude into a cavity flow path 82 of the cavity 72 to disrupt the thermal boundary layer of the cooling airflow 68 and increase the cooling effectiveness of the internal cooling circuit of the component 50. In one embodiment, the curved turbulators 80 are miniature walls protruding into the cavity flow path 82. The design, configuration and placement of the numerous curved turbulators 80 shown by
A curved turbulator 80 may extend from the wall 84. In this embodiment, the wall 84 of the cavity 72 includes a plurality of curved turbulators 80. The curved turbulators 80 can span a width W of the wall 84 and extend substantially perpendicular to the direction of flow of the cooling airflow 68 within a cavity flow path 82 of the cavity 72. Due to the continuous curvature of the curved turbulators 80, a pitch P (e.g., a spacing) between each adjacent curved turbulator 80 is continuously varied.
A row of film cooling holes 86 can be disposed between radially adjacent curved turbulators 80. In this embodiment, each row of film cooling holes 86 includes a first film cooling hole 86A and a second film cooling hole 86B that is radially staggered from the first film cooling hole 86A. Of course, additional film cooling holes than are shown in this embodiment could be disposed through the wall 84 in each row of film cooling holes 86. The film cooling holes 86A, 86B do not intersect through any curved turbulator 80 because of the wavy design of the curved turbulators 80. Other portions of the wall 84 may exclude film cooling holes 86 between adjacent curved turbulators 80.
The curved turbulators 80 are configurable in a variety of patterns. For example, as shown in
The curved turbulators of this disclosure may embody any curved or wavy geometry that provides a smooth transition surface that is capable of accommodating relatively large variations in the streamwise positioning of the turbulators relative to the cooling airflow that flows within the cavities. The exemplary curved turbulators also provide reduced stress concentrations as compared to treatments having more angular designs, such as V-shaped turbulators.
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.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US14/13981 | 1/31/2014 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| 61760795 | Feb 2013 | US |
