Patent Application
20250116193
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-pressure and temperature core gas flow. The high-pressure and temperature core gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section may include low and high pressure compressors, and the turbine section may also include low and high pressure turbines.
Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature capability and lifetime. Ceramic matrix composite (“CMC”) materials are also being considered for airfoils. Among other attractive properties, CMCs have high temperature resistance. Despite this attribute, however, there are unique challenges to implementing CMCs in airfoils.
An airfoil according to an example of the present disclosure includes an airfoil section that defines pressure and suction sides, a leading edge, and a trailing edge region including a trailing edge. The airfoil section is formed of a ceramic matrix composite that includes fiber plies disposed in a ceramic matrix. The fiber plies have core fiber plies that define a radial tube that circumscribes an internal cavity, and skin fiber plies that define an exterior of the airfoil section and wrap around the core fiber plies from the pressure side of the trailing edge region, through the leading edge, and to the suction side of the trailing edge region. There is a filler element in the trailing edge region aft of the internal cavity that is sandwiched between the skin fiber plies on the pressure side and the skin fiber plies on the suction side and at least one cooling passage has at least one inlet orifice section that opens to the internal cavity and extends through the core fiber plies. A radially-elongated multi-dimensional plenum is connected with the at least one inlet orifice section. A plurality of outlet orifice sections extend through the trailing edge, and a plurality of intermediate passage sections are bound by at least one of the skin plies and the filler element and connect the radially-elongated multi-dimensional plenum with the plurality of outlet orifices.
In a further embodiment of any of the foregoing embodiments, the radially elongated multi-dimensional plenum has a serpentine shape.
In a further embodiment of any of the foregoing embodiments, the serpentine shape is provided by a first row of radially-spaced channel segments in one of the core fiber plies and a second row of radially-spaced channel segments in one of the skin fiber plies that is radially offset from the first row such that each one of the radially-spaced channel segments of the second row partially overlaps two of the radially-spaced channel segments of the first row.
In a further embodiment of any of the foregoing embodiments, each of the intermediate passage sections is of an equivalent constant cross-sectional area.
In a further embodiment of any of the foregoing embodiments, the intermediate passage sections are of non-equivalent constant cross-sectional areas.
In a further embodiment of any of the foregoing embodiments, each of the intermediate passage sections is divergent.
The airfoil as recited in claim 1, further including branch passage sections that interconnect the intermediate passage sections.
In a further embodiment of any of the foregoing embodiments, the intermediate passage sections are radially curved.
In a further embodiment of any of the foregoing embodiments, the radial tube has a radiused end in the trailing edge region, and the at least one inlet orifice section opens to the internal cavity at a location forward of the radiused end.
A gas turbine engine according to an example of the present disclosure includes a compressor section, a combustor in fluid communication with the compressor section, and a turbine section in fluid communication with the combustor. The turbine section has airfoils according to any of the foregoing embodiments.
A method of fabricating an airfoil according to any of the foregoing embodiments includes laying-up first core fiber plies to form a radial tube that circumscribes an internal cavity, partially densifying the first core fiber plies with ceramic matrix material, and laying-up an additional core fiber ply around the partially densified first core fiber plies to form a core preform. The additional core fiber ply include pre-formed slots. A filler element is arranged aft of the internal cavity of the core preform, a first skin fiber ply is laid-up around the core preform and the filer element to form a first intermediate preform, and the first intermediate preform is at least partially densified with additional ceramic matrix material. In the at least partially densified first skin fiber ply, intermediate passage section are formed, wherein the intermediate passage sections partially overlap with the preformed slots to provide a radially-elongated multi-dimensional plenum. At least one inlet orifice section is formed through the core preform to the internal cavity to provide a second intermediate preform, and one or more additional skin fiber plies are laid-up around the second intermediate preform to form a final preform. The final preform is then densified with additional ceramic matrix material, and a plurality of outlet orifice sections are formed that extend through the trailing edge. The at least one inlet orifice section, the a radially-elongated multi-dimensional plenum, the intermediate passage sections, and the plurality of outlet orifice sections are connected to provide at least one cooling passage.
The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof.
The various features and advantages of the present 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.
In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements. Terms such as “first” and “second” used herein are to differentiate that there are two architecturally distinct components or features. Furthermore, the terms “first” and “second” are interchangeable in that a first component or feature could alternatively be termed as the second component or feature, and vice versa.
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 the 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 through 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), and can be less than or equal to about 18.0, or more narrowly can be less than or equal to 16.0. 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. The gear reduction ratio may be less than or equal to 4.0. The low pressure turbine 46 has a pressure ratio that is greater than about five. The low pressure turbine pressure ratio can be less than or equal to 13.0, or more narrowly less than or equal to 12.0. 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 an 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. The engine parameters described above and those in this paragraph are measured at this condition unless otherwise specified. “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, or more narrowly greater than or equal to 1.25. “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 1150.0 ft/second (350.5 meters/second), and can be greater than or equal to 1000.0 ft/second (304.8 meters/second).
The airfoil section 60 is formed of a ceramic matrix composite (CMC) 64. Referring to the cutaway section in
The fiber plies 66 include core fiber plies 68 and skin fiber plies 70. The core fiber plies 68 define a radial tube 72 that circumscribes and immediately borders the internal cavity 61. The skin fiber plies 70 define the exterior profile of the airfoil section 60 and wrap around the core fiber plies 68 from the pressure side 60d in the trailing edge region 63, through the leading edge 60a, and to the suction side 60c in the trailing edge region 63. In the illustrated example, there are four core fiber plies 68 and three skin fiber plies 70, although it is to be understood that the numbers of plies 68/70 could be varied.
The radial tube 72 has a radiused end 74 in the trailing edge region 63. Except for the inner-most one of the core fiber plies 68, all of the core fiber plies 68 wrap completely around the internal cavity 61. The inner-most one of the core fiber plies 68 terminates short of the radiused end 74, as the small radius may exceed the capability of the ply to bend without fiber tow distress.
There is a filler element 76 in the trailing edge region 63, aft of the internal cavity 61. A filler element in general is often colloquially referred to as a “noodle” and serves as a non-structural space-filler, usually in an interstice where other fiber plies bend. Here, the filler element 76 fills in the region aft of the radiused end 74 and is sandwiched between the skin fiber plies 70 on the pressure side 60d and the skin fiber plies 70 on the suction side 60c. The filler element 76 may be formed of, but is not limited to, a bundle of densified ceramic fibers (a CMC) or a monolithic ceramic.
The sides 60c/60d and the trailing edge region 63 may require cooling. In that regard, as shown in
The inlet section 78a opens to the internal cavity 61 and serves for entry of the cooling air into the passage 78, and the outlet sections 78c extend through the trailing edge 60b, such as along a centerline of the trailing edge 60b, and serve to discharge spent cooling air to the exterior of the airfoil section 60 (into the core gaspath C). The plenum 78b serves to distribute the cooling air to the intermediate passage sections 78d, which connect the plenum 78b to the outlet orifice sections 78c. The plenum 78b extends along the outer-most one of the core plies 68 (designated 68a in
As indicated above, the plenum 78b is multi-dimensional. As used herein, the term “multi-dimensional” refers to the path (e.g., as represented by a central axis) of plenum 78b having portions that extend along the radial direction (relative to engine axis A) and also portions that extend along at least one direction that is orthogonal to the radial direction, such as an axial or chordal direction. In the example illustrated in
The radiused end 74 of the radial tube 72 is an area that may be under considerable stress in the airfoil section 60. For example, the core fiber plies 68 are under bending stresses from bending at the radius in the end 74, and dynamic stresses on the airfoil section 60 may concentrate at this area due to the bending. Such stresses on a CMC may tend to cause delamination between fiber plies. In this regard, the proposition of using an orifice in the radiused end is undesirable, as it would cause a discontinuity in the ceramic fibers and thereby potentially weaken the CMC in that area. Rather, in the example shown (see
In further embodiments, the location L is substantially further forward of location P by a multiple of the distance (the product of N and t). In such examples, the intermediate passage sections 78d are longer and thus run over a longer distance along the suction side 60c (or pressure side 60d), thereby providing additional cooling of the side 60c (or 60d).
In
A typical lay-up process might involve laying-up all of the fiber plies to form a fiber preform and then fully densifying the fiber preform. In the above method, however, the densifications are staged, to occur prior to all of the fiber plies being laid-up. For example, the ceramic matrix in the above densifications is formed by, but is not limited to, chemical vapor infiltration (CVI), melt infiltration (MI), a hybrid of CVI and MI, and/or polymer infiltration and pyrolysis (PIP). The disclosed methodology enables at least the intermediate passage sections 78d, inlet orifice section 78a, and outlet orifice section 78c to be formed after one or more densification steps. While machining and/or drilling is involved and may be somewhat time-consuming, forming these sections of the cooling passage 78 after densification permits precise control over the location and geometry of the passage sections.
Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.
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 this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.
