Patent Application
20250198292
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 fan section or other sections of the engine may be formed of composite materials. Composite materials are generally lighter in weight, but can challenge durability.
A blade according to an example of the present disclosure includes an airfoil section spanning along a radial direction between first and second radial ends, and a dovetail at the second end of the airfoil section. The airfoil section and the dovetail are comprised of a fiber-reinforced composite including inter-section fiber plies each extending from the airfoil section into the dovetail and wedges interleaved with the inter-section fiber plies in the dovetail. The dovetail is asymmetric about the radial direction and has first and second circumferential sides that are sloped to face radially outwardly and form non-zero angles with the radial direction, one of the non-zero angles being 10° or less and a sum of the non-zero angles being from 35° to 45°.
In a further embodiment of any of the foregoing embodiments, the dovetail has a forward face, an aft face, and a dovetail body that extends from the forward face to the aft face. The aft face is circumferentially offset from the forward face such that the dovetail body curves from the forward face to the aft face.
In a further embodiment of any of the foregoing embodiments, the dovetail has a radially-inwardly facing side that is parallel to an axial direction that is perpendicular to the radial direction.
In a further embodiment of any of the foregoing embodiments, the one of the non-zero angles that is 10° or less is from 5° to 10°.
In a further embodiment of any of the foregoing embodiments, the first and second circumferential sides are flat.
In a further embodiment of any of the foregoing embodiments, the airfoil section defines a pressure side and a suction side. The first circumferential side has the non-zero angle that is less than 10°, and the first circumferential side is on the pressure side.
A gas turbine engine according to an example of the present disclosure includes a fan section disposed about an engine axis, a hub including a plurality of axial doveslots, and a plurality of fan blades. Each fan blade includes an airfoil section spanning along a radial direction between first and second radial ends, and a dovetail at the second end of the airfoil section. The dovetail is disposed in one of the doveslots to attach the fan blade with the hub. The airfoil section and the dovetail are comprised of a fiber-reinforced composite having inter-section fiber plies each extending from the airfoil section into the dovetail and wedges interleaved with the inter-section fiber plies in the dovetail. The dovetail defines first and second circumferential sides that are sloped to face radially outwardly and form non-zero angles with the radial direction, one of the non-zero angles being 10° or less and a sum of the non-zero angles is from 35° to 45°.
In a further embodiment of any of the foregoing embodiments, the dovetail has a forward face, an aft face, and a dovetail body that extends from the forward face to the aft face. The aft face is circumferentially offset from the forward face such that the dovetail body curves from the forward face to the aft face.
In a further embodiment of any of the foregoing embodiments, the dovetail has a radially-inwardly facing side that is parallel to an axial direction that is perpendicular to the radial direction.
In a further embodiment of any of the foregoing embodiments, the one of the non-zero angles that is 10° or less is from 5° to 10°.
In a further embodiment of any of the foregoing embodiments, the first and second circumferential sides are flat.
In a further embodiment of any of the foregoing embodiments, the airfoil section defines a pressure side and a suction side, the first circumferential side has the non-zero angle that is less than 10°, and the first circumferential side is on the pressure side.
In a further embodiment of any of the foregoing embodiments, the airfoil section is connected by a neck to the dovetail, the neck being a location at which the airfoil section flares to form the dovetail, and the neck is radially outboard of the doveslot.
A gas turbine engine according to an example of the present disclosure includes a fan hub having an axial doveslot, and a fan blade including an airfoil section defining a pressure side and a suction side. The airfoil section spans along a radial direction between first and second radial ends. A dovetail at the second end of the airfoil section is disposed in the doveslot to attach the fan blade with the fan hub. The airfoil section and the dovetail are comprised of a fiber-reinforced composite including inter-section fiber plies each extending from the airfoil section into the dovetail and wedges interleaved with the inter-section fiber plies in the dovetail. The dovetail defines a first circumferential side on the suction side and second circumferential side on the pressure side upon an impact event on the pressure side of the airfoil section. A bending is induced in which the first circumferential side presses against a side of the doveslot, creating stress at a neck at which the airfoil section flares to form the dovetail, and the first and second circumferential side is sloped to face radially outwardly and form non-zero angles with the radial direction. The non-zero angle of the first circumferential side being 10° or less in order to reduce concentration of the stress at the neck.
In a further embodiment of any of the foregoing embodiments, the one of the non-zero angles that is 10° or less is from 5° to 10°.
In a further embodiment of any of the foregoing embodiments, the first and second circumferential sides are flat.
In a further embodiment of any of the foregoing embodiments, the dovetail has a forward face, an aft face, and a dovetail body that extends from the forward face to the aft face. The aft face is circumferentially offset from the forward face such that the dovetail body curves from the forward face to the aft face.
In a further embodiment of any of the foregoing embodiments, the dovetail has a radially-inwardly facing side that is parallel to an axial direction that is perpendicular to the radial direction.
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).
Turning again to the fan section 22, the fan 42 includes fan blades 60 that are attached with a fan hub 62.
As shown in
The fan hub 62 includes a plurality of axial doveslots 68 that extend from the forward face to the aft face of the hub 62. The doveslots 68 are also curved and are generally of complementary geometry to the geometry of the dovetails 66 such that the dovetails 66 interlock with the doveslots 68 to radially retain the blades 60 on the hub 62. Although not shown, a wear pad, wear coating, or the like may be provided between the dovetail 66 and the sides of the doveslot 68 to reduce wear on the dovetail 66, the doveslot 68, or both.
Referring to
The composite material 70 includes inter-section fiber plies 74. A “ply” refers to a layer of fiber tows. For example, the layer includes a woven fabric, a tape with multiple fiber tows, or multiple fabrics and/or tapes that together make-up a distinct layer. The prefix “inter-” means that each ply 74 extends in both the airfoil section 64 and the dovetail 66. In the dovetail 66, there are wedges 76 interleaved with the plies 74 and that serve to increase the dovetail width and form the flared shape of the dovetail 66. The plies 74 and the wedges 76 extend the full axial length of the blade 62.
Engines may occasionally ingest foreign objects. The foreign objects can impact airfoils, particularly fan blades at the front of the engine 20. Such impacts can induce bending stresses on one or more blades in the fan. In that regard, fan blades that are formed of fiber-reinforced composites that have fiber plies may experience fiber ply delamination along the boundary between plies when subjected to bending stresses from foreign object impacts. In particular, at the neck where the blade flares to form the dovetail there is a curvature that can act as a stress riser, concentrating induced bending stress in a relatively small region, which may then cause delamination. Extensive delamination and blade damage could in some instances lead to blade detachment. As will be discussed below, the dovetails 66 of the disclosed blades 60 have an asymmetric geometry that is configured to reduce stress concentration from impacts on the blade, thus enhancing blade durability.
The sides 66d/66e each form a different non-zero angle, designated at X1 and X2 in
Additionally, the sum of the angles X1 and X2 (from 35° to 45°)is thought to represent a useful balance between the shallowness of the angle X1 to reduce stress concentration versus the need to have enough of a radial component to the slopes of the sides 66d/66e to ensure that the blade can remain retained in the doveslot 68 under the centrifugal force on the blade 60 during rotation. Accordingly, the blade 60 is configured to mitigate impact damage by managing the induced bending stresses to avoid high stress concentrations that could otherwise lead to substantial delamination.
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.
