Patent Application
20160298462
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.
Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow for the next set of blades.
Turbine vanes, blades, combustors, and other components include film cooling features to provide a boundary layer of cooling fluid along external surfaces, which protects the component from the hot combustion gases in the core flow path. The film cooling features include passages extending through a wall of the cooling component that may include a complex geometry that is difficult to manufacture. Therefore, there is a need to increase the efficiency of the film cooling features and the ease of manufacturing the film cooling features.
In one exemplary embodiment, a gas turbine engine component includes a wall which includes a first surface and a second surface opposing the first surface. A plurality of cooling passages extends between the first surface and the second surface each including a metering portion and a diffusion portion. The diffusion portion includes a cast or as-consolidated surface.
In a further embodiment of the above, the metering portion is located adjacent the first surface and the diffusion portion is located adjacent the second surface.
In a further embodiment of any of the above, a locating feature is adjacent the diffusion portion for locating a tool to machine the metering portion.
In a further embodiment of any of the above, the diffusion portion transitions between a cylindrical inlet and a non-cylindrical outlet.
In a further embodiment of any of the above, the diffusion portion includes an increasing cross-sectional area.
In a further embodiment of any of the above, the metering portion includes a machined surface.
In a further embodiment of any of the above, at least one of the metering portion and the diffusion portion extends transverse to the first surface of the second surface.
In a further embodiment of any of the above, a casting mold has a plurality of casting diffusion protrusions for forming the plurality of diffusion portions.
In another exemplary embodiment, a method of forming a cooling passage in a gas turbine component includes casting a plurality of diffusion portions in a component and forming a plurality of metering portions at least partially aligned with a corresponding one of the plurality of diffusion portions in the component.
In a further embodiment of any of the above, the plurality of metering portions is formed using a machining process.
In a further embodiment of any of the above, the component includes a wall having a first surface opposing a second surface.
In a further embodiment of any of the above, the plurality of metering portions is located adjacent the first surface and the plurality of diffusion portions is located adjacent the second surface.
In a further embodiment of any of the above, the first surface is located on an inner side of the component and the second surface is located on an outer side of the component.
In a further embodiment of any of the above, at least one diffusion portion includes an increasing cross-sectional area.
In a further embodiment of any of the above, the method includes locating a tool for forming the plurality of metering portions relative to the diffusion portion with a locating feature adjacent each of the plurality of diffusion portions.
In a further embodiment of any of the above, the method includes forming a mold having a plurality of diffusion protrusions for forming the plurality of diffusion portions.
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. 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).
The example gas turbine engine includes fan 42 that comprises in one non-limiting embodiment less than about twenty-six (26) fan blades. In another non-limiting embodiment, fan section 22 includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment low pressure turbine 46 includes no more than about six (6) turbine rotors schematically indicated at 34. In another non-limiting example embodiment low pressure turbine 46 includes about three (3) turbine rotors. A ratio between number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate fan section 22 and therefore the relationship between the number of turbine rotors 34 in low pressure turbine 46 and number of blades 42 in fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.
The pressure and suction sidewalls 66, 68 form the major opposing surfaces or walls of the blade 60, extending axially between the leading edge 62 and trailing edge 64, and radially outward from a root section 70, adjacent inner diameter (ID) platform 72, to a distal end 74 opposite the ID platform 72. In some designs, the distal end 74 may include a shroud forming an outer diameter (OD) platform.
Cooling passages 76 are located on one or more surfaces of the blade 60. In the illustrated example, the cooling passages 76 are located along the pressure sidewall 66 of the blade 60. In another example, the cooling passages 76 could be located adjacent the leading edge 62, the trailing edge 64, the pressure sidewall 66, the suction sidewall 68, or a combination thereof.
The cooling passages 76 are provided along one or more surfaces of the airfoil 80, for example the leading or trailing edge 82, 84, the pressure (concave) sidewall 86, the suction (convex) sidewall 88, or a combination thereof.
The blade 60 (
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The cooling passages 76 are oriented so that their inlets are positioned on the inner wall surface 96 and their outlets are positioned on the outer wall surface 98. During operation gas turbine engine 20, the outer wall surface 98 is in proximity to high temperature gases (e.g., combustion gases, hot air). Cooling airflow is delivered inside the pressure sidewall 66 where it exits an interior of the blade 60 through the cooling passages 76 and forms a cooling film on the outer wall surface 98.
Each of the cooling passages 76 include a metering portion 102 having a generally cylindrical cross section with a machined surface and a diffusion portion 100 having a cast surface. The diffusion portion 100 transitions between a cylindrical cross section at an inlet to a triangular cross section adjacent an outlet 104 to the diffusion portion 100. Although the outlet 104 to the diffusion portion 100 is triangular in the illustrated example, the outlet 104 to the diffusion portion 100 can be any shape as long as the outlet has a larger cross-sectional area than the inlet to the diffusion portion 100.
In the illustrated example, the metering portion 102 includes an inlet adjacent the inner wall surface 96 to allow cooling airflow to flow into the cooling passage 76 from internal passages 106 in the blade 60. The cooling air flows out of the internal passages 106 and flows through the metering section 102 the diffusion portion 100 to form a cooling film over the blade 60. The cooling passages 76 can be arranged in a linear row on pressure sidewall 66 and positioned axially so that the cooling air flows in substantially the same direction longitudinally as the high temperature gases flowing past the pressure sidewall 66.
Alternatively, the cooling passages 76 can also be located in a staggered formation or other formation on the pressure sidewall 66. Although the cooling holes 76 are described in relation to the blade 60, the cooling holes 76 in the vane 80 are similar to the cooling holes 76 in the blade 60 except where described below or shown in the Figures. The cooling passages 76 can be located on a variety of suitable components such as turbine vanes and blades, combustors, blade outer air seals, augmentors, and etc. The cooling passages 76 can be located on the pressure sidewalls 66, 86 or suction sidewalls 68, 88 of the blades 60 and the vanes 80, respectively. The cooling passages 76 can also be located on a tip or platforms of the blade 60 or vane 80.
The cooling passages 76 are formed during casting process and a machining process. The cooling passages 76 could also be formed during a consolidation process such as is produced from an additive manufacturing process such as powder bed laser sintering and a machining process.
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Although
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 the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.
