Patent Application
20190176228
This disclosure relates to a gas turbine engine component, such as an airfoil. More particularly, the disclosure relates to a component cooling passage having an obstruction and a core for making the component cooling passage.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. 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 are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
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 and prepare it for the next set of blades.
Many blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages. In some applications, such as vanes, a baffle is provided in the cooling passage to reduce the cross-sectional area of the cooling passage and increase cooling fluid velocity. Typically, the baffles are constructed from sheet metal inserted into the vane after the vane has been cast.
In one exemplary embodiment, a gas turbine engine includes a structure that has walls that provide a cooling passage and a cooling surface. A non-ferrous obstruction is relative to the walls. The obstruction includes a portion spaced from the cooling surface to provide a gap which is configured to receive a cooling fluid.
In a further embodiment of any of the above, the structure is an airfoil.
In a further embodiment of any of the above, the structure is a blade outer air seal.
In a further embodiment of any of the above, the structure is a platform that supports an airfoil.
In a further embodiment of any of the above, the structure is a combustor line
In a further embodiment of any of the above, the structure is an exhaust liner.
In a further embodiment of any of the above, the obstruction is arranged to block a flow of the cooling fluid through the cooling passage.
In a further embodiment of any of the above, the obstruction includes a refractory metal.
In a further embodiment of any of the above, the obstruction provides multiple gaps adjacent to the cooling surface.
In a further embodiment of any of the above, the structure includes a nickel alloy, and the obstruction is provided by a material that is different than the nickel alloy.
In another exemplary embodiment, a core structure for providing a gas turbine engine a cooling passage includes a first material provided on a second material and together provides a perimeter. The first material is provided along multiple portions of the perimeter with the second material separating the multiple portions.
In a further embodiment of any of the above, the first and second materials are different than one another.
In a further embodiment of any of the above, the first material is ceramic.
In a further embodiment of any of the above, the second material includes a refractory metal.
The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, 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; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
The example engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis X 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 connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46. The inner shaft 40 drives the fan 42 through a speed change device, such 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 high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis X.
A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
The example low pressure turbine 46 has a pressure ratio that is greater than about five (5). The pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
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 as well as setting airflow entering the low pressure turbine 46.
The core airflow C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes vanes 59, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 59 of the mid-turbine frame 57 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 57. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
The disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
In one disclosed embodiment, the gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
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 (′TSFC)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (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 1.50. In another non-limiting embodiment the low fan pressure ratio 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.
The disclosed cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade 64 is described. It should be understood that the cooling passage may also be used in vanes, blade outer air seals, and turbine platforms, for example.
Referring to
The airfoil 78 of
The airfoil 78 includes a cooling passage 90 provided between the pressure and suction walls 86, 88. The exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage 90. Example cooling passages 90 illustrated in
A cooling source 92 is in fluid communication with the cooling passage 90. In the example, the cooling passage 90 is illustrated as having an obstruction 94 arranged within the cooling passage to restrict flow from the cooling source 92, which increases the velocity of the cooling fluid within the cooling passage 90.
An example cross-sectioned through the airfoil 78 is illustrated in
In the example, the obstruction 94 is non-ferrous and engages the cooling surface 98 to locate the obstruction 94 relative to the walls 96. The obstruction 94 includes portions spaced from the cooling surface 98 to provide one or more gaps 102. The gaps 102 receive the cooling fluid from the cooling source 92 (
A multi-material core is used to produce the obstruction 94. Referring to
Generally, to provide the obstruction 94 within the cooling passage 90c, the core structure 104 is arranged within a wax mold, as is known in conventional investment casting processes, to produce a wax structure in the shape of the component to be cast. A perimeter 105 of the core structure 104 provides the cooling surface 98 within a space 100 of the cast component, as shown in
Referring to the method 110 shown in
Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For example, the obstruction may be provided in passages other than cooling passages. The passages may be formed in other structures, such as blade outer air seals, platforms, combustor liner and exhaust liners. For that and other reasons, the following claims should be studied to determine their true scope and content.
This disclosure is a divisional of U.S. patent application Ser. No. 14/766,061 filed Aug. 5, 2015, which is a U.S. National Phase of PCT/US2014/011833 filed Jan. 16, 2014, which claims priority to Provisional Application No. 61/763,661 filed Feb. 12, 2013.
| Number | Date | Country | |
|---|---|---|---|
| 61763661 | Feb 2013 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | 14766061 | Aug 2015 | US |
| Child | 16275468 | US |
