Patent Application
20210148245
The present disclosure relates generally to gas turbine engine flowpath construction, and more specifically to a contoured mateface configuration for utilization in a blade outer air seal.
Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section further drives the compressor section to rotate. In some examples, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine as well.
Each of the compressor, combustor and turbine section are fluidly connected via a primary flowpath, with the outer diameter of the primary flowpath being defined at least partially by a set of circumferentially arranged blade outer air seals.
In one example, a gas turbine engine includes one of a turbine section and a compressor section comprised of a plurality of stages, at least one of the stages in the plurality of stages defining an outer diameter comprised of a plurality of circumferentially arranged blade outer air seals, each blade outer air seal being spaced from each adjacent blade outer air seal in the plurality of circumferentially arranged blade outer air seals via a mateface gap, wherein the mateface gap is oblique to a radius of the gas turbine engine, such that air entering the mateface gap is directed to an inner diameter surface of at least one of the blade outer air seals in the plurality of blade outer air seals.
In another example of the above gas turbine engine, the mateface gap is defined by a first surface of a first blade outer air seal in the plurality of circumferentially arranged blade outer air seals and defined by a second surface of a second blade outer air seal in the plurality of circumferentially arranged blade outer air seals.
In another example of the any of the above gas turbine engines, the first surface and the second surface are complimentary angled planar surfaces relative to the gas turbine engine radius.
In another example of the any of the above gas turbine engines, the first surface and the second surface are planar surfaces and are oblique to each other.
In another example of the any of the above gas turbine engines, at least one of the first surface and the second surface includes a curvature.
In another example of the any of the above gas turbine engines, each of the first surface and the second surface includes a curvature.
In another example of the any of the above gas turbine engines, the curvature of the first surface and the curvature of the second surface is complimentary.
In another example of the any of the above gas turbine engines, at least one of the curvatures is a complex curvature.
In another example of the any of the above gas turbine engines, each of the curvatures is a complex curvature.
In another example of the any of the above gas turbine engines, at least one of the mateface gaps includes a nozzle fluid exit.
In another example of the any of the above gas turbine engines, each of the mateface gaps includes a nozzle fluid exit.
In another example of the any of the above gas turbine engines, an outer diameter portion of each mateface gap is sealed via an interstage seal.
In another example of the any of the above gas turbine engines, each interstage seal is a featherseal.
One exemplary method for film cooling an internal surface of a gas turbine engine component includes angling a mateface gap such that air exiting the mateface gap is directed along an inner diameter surface of one of an outer diameter component.
In another example of the above described method, the outer diameter component is a blade outer air seal.
Another example of the above described method includes diffusing air exiting the mateface gap using non-complimentary blade outer air seal surfaces.
Another example of the above described method includes accelerating air exiting the mateface gap.
An exemplary blade outer air seal for a gas turbine engine includes a platform defining an inner diameter surface, a first mateface surface extending from a first end of the inner diameter surface and a second mateface surface extending from the inner diameter surface, at least one of the first mateface surface and the second mateface surface being oblique to a radius of a gas turbine engine.
In another example of the above blade outer air seal, the first mateface surface and the second mateface surface are complimentary.
In another example of any of the above blade outer air seals, at least one of the first mateface surface and the second mateface surface is curved.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
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)]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).
Within the compressor section 24 and the turbine section 28 are multiple stages with each stage including a set of circumferentially arranged rotors paired with a set of circumferentially arranged stators. Radially outward of the rotors in a given stage is an outer diameter defined by circumferentially arranged blade outer air seals, with the blade outer air seals being connected via intersegment seals. In order to cool the blade outer air seals, air is flowed circumferentially through the blade outer air seal and expelled from a circumferential edge of the blade outer air seal, into a gap between the segments. The facing surfaces of adjacent blade outer air seals are referred to as matefaces. The air is ejected from between the matefaces into the core gaspath flow, and expelled along the core gaspath. Any excess cooling capacity in the ejected air is wasted and represents a loss of efficiency.
With continued reference to
Each of the blade outer air seals includes two circumferentially facing sides 126, 128 with at least one of the sides 126, 128 being at least partially skew relative to the radius of the engine 20. By skewing at least one of the sides 126, 128 the airflow passed into the intersegment gap 124 is directed along a radially inward facing surface 121 of the blade outer air seal 120, thereby enhancing the cooling effect provided to the blade outer air seal 120.
The directing and contouring of the sides 126, 128 serves a dual purpose of directing leakage or post-part-cooling flow such that the flow cools the facing surface 126, 128 and provides a film coverage of an adjacent blade outer air seal 120 before being disposed of in the primary flowpath. In addition to this, the contours of the facing surfaces 126, 128 are in some examples configured to reduce ingestion of gaspath air into the intersegment cavity 124, thereby reducing oxidization of the blade outer air seal 120.
In order to further enhance the film cooling effect, and prevent flowpath air from being drawn into the interstage gap 124, each interstage gap 124 is angled aligned with the rotation 102 direction of the rotors 110 such that the outlet of the interstage gap 124 is rotationally after the inlet (i.e. the outlet is counterclockwise of the inlet for a system with counterclockwise rotor rotation). In alternative examples, the rotation 102 can be reversed with the same alignment of the interstage gaps 124.
With continued reference to
With regards to the first example facing surface pairs 302, illustrated in
With regards to the second example facing surface pair 302,
With regards to the third example facing surface pair 302,
With regards to the fourth example facing surface pair 302, illustrated in
With regards to the fifth example facing surface pair 302, the facing surface pair 302 of
With continued reference to
While described above and illustrated in
It is further understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
