Patent Application
20190063452
Exemplary embodiments of the present disclosure are direct to fan hubs of gas turbine engines and more particularly a conical fan hub that reduces a blade off load.
Gas turbine engines, such as turbofan gas turbine engines, typically include a core engine having a fan section, a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and mixed with fuel in the combustor section for generating hot combustion gases. The hot combustion gases flow through the turbine section which extracts energy from the hot combustion gases to power the compressor section and drive the fan section.
The core engine includes an engine casing structure that includes a fan containment case (FCC) and a fan case downstream from the FCC. The FCC and the fan case surround the fan section of the gas turbine engine and contain the fan section components in the event of a fan blade out event. A fan blade out event occurs where a fan blade of the fan section becomes dislodged from the fan section and strikes the FCC.
Accordingly, it is desirable to limit the mass of the blade in the event of a fan blade out event.
In one embodiment, a conical hub for a fan of a gas turbine engine is provided. The conical hub having: a plurality of attachment features located on an outer circumferential surface of the conical hub, wherein at least some of the plurality attachment features are axially aligned with each other and at least some of the plurality of attachment features are off set from each other, and wherein each of the plurality of attachment features have an opening configured to receive a portion of a pin; and wherein the outer circumferential surface of the conical hub increases in diameter with respect to an axis of the conical hub in a forward to aft direction of the conical hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, as least some of the plurality of attachment features are proximate to a forward leading edge of the conical hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein as least some of the plurality of attachment features are arranged in a plurality of rows on the outer circumferential surface of the conical hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, at least some of the plurality of attachment features are arranged in a plurality of rows on the outer circumferential surface of the conical hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub has at least two different Gaussian curvatures.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub has at least two different Gaussian curvatures.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub has at least two different Gaussian curvatures.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub undulates.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein at least some of the plurality of attachment features are located proximate to a leading edge of the conical hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein at least some of the plurality of attachment features are arranged in a plurality of rows on the outer circumferential surface of the conical hub.
In yet another embodiment, a gas turbine engine is provided. The gas turbine engine having: a conical fan hub; and a plurality of blades secured to the conical fan hub via a plurality of attachment features located on an outer circumferential surface of the conical hub, wherein at least some of the plurality attachment features are axially aligned with each other and at least some of the plurality of attachment features are off set from each other, and wherein each of the plurality of attachment features have an opening configured to receive a portion of a pin for securing the plurality of blades to the conical fan hub; and wherein the outer circumferential surface of the conical fan hub increases in diameter with respect to an axis of the conical hub in a forward to aft direction of the conical fan hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein at least some of the plurality of attachment features are located proximate to a forward leading edge of the conical hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein at least some of the plurality of attachment features are arranged in a plurality of rows on the outer circumferential surface of the conical hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the outer circumferential surface of the conical hub has at least two different Gaussian curvatures.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub has at least two different Gaussian curvatures.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub has at least two different Gaussian curvatures.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub undulates.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein at least some of the plurality of attachment features are located proximate to a leading edge of the conical hub.
In yet another embodiment, a method of reducing blade off loads during a blade out event in a gas turbine engine is provided. The method including the steps of: securing a plurality of blades to a conical fan hub of the engine via a plurality of attachment features located on an outer circumferential surface of the conical hub, wherein at least some of the plurality attachment features are axially aligned with each other and at least some of the plurality of attachment features are off set from each other, and wherein each of the plurality of attachment features have an opening configured to receive a portion of a pin for securing the plurality of blades to the conical fan hub; and wherein the outer circumferential surface of the conical fan hub increases in diameter with respect to an axis of the conical hub in a forward to aft direction of the conical fan hub.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, wherein the outer circumferential surface of the conical hub has at least two different Gaussian curvatures.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
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 low pressure compressor 44 and a 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 high pressure compressor 52 and 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. An engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 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 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 disclosure 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 (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 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 m/sec).
Referring now to
Also illustrated in
For a blade containment test under 14 CFR 33.94, the fan blade 70 is cut at the blade to dovetail interface represented by the dashed line 88. This releases at least portion 86 of the fan blade 70 into the illustrated flow paths B and C.
Referring now to
By providing a coned hub with a radially extending midline 91 and/or coned hub or rotor 90 as illustrated herein, the cut line 88 for use in a blade containment test under 14 CFR 33.94, allows portion 86 of the fan blade 70 to be significantly smaller, which benefits rotating imbalances as well as reducing the impact energy of a released blade into the fan containment case (FCC).
In addition and as also illustrated, the cone angle φ of the hub or rotor 90 allows reconfiguration of the static structure 80, the shaft 82 and thus the bearing 84 closest the hub 90 may be relocated to an area that results in improved rotor or hub dynamics.
In order to secure the fan blade 70 to the coned or conical hub 90, a plurality of attachment features 94 extend from a surface 93 and the fan blade is secured thereto by a plurality of ligaments or connecting members 96 which are secured to the fan blade 70 at one end and extend to the connecting member or members 96 at the other end.
In one embodiment, the ligaments or connecting members 96 are secured to the attachment features 94 by a pin or pins 98. In one embodiment, pins 98 may be press fit into its corresponding opening in order to secure the ligaments or connecting members 96 to the hub 90. Of course, alternative methods of securement are considered to be within the scope of the present disclosure. Still further and as illustrated in at least
Referring now to
By varying the Gaussian curvature of the hub, the related blade design may also vary. As such, the hub and the blade securement thereto below the core flow path C can vary. This allows the blade attachment to be configured in order to account for centripetal forces or stresses encountered by the blade and/or areas of its securement to the hub.
Referring now to
In one non-limiting embodiment, the walls or attachment features 94 may be formed in the hub 90 via a lathing process. As such, the hub may be placed on a turning machine or lathe and a cutting tool is used to remove surface material in order to form the walls or attachment features 94.
As mentioned above and in one embodiment, the walls or attachment features 94 may extend continuously about the hub 90. In yet another embodiment, the walls or attachment features may extend partially about hub 90 (e.g., not completely around) or some of the walls or attachment features 94 may extend completely around and some may not.
As illustrated, the ligaments or connecting members 96 have an opening 130 for receipt of pin or member 98 therein. In addition, the ligaments or connecting members 96 may also have a slot or opening 132 in order to receive a portion of wall 94 therein. Accordingly and as the ligaments or connecting members 96 are placed on a portion of wall 94 a portion of the wall or feature 94 is received in slot or opening 132. Once opening 130 is aligned with opening 100 a pin or member 98 is inserted therein in order to secure the ligaments or connecting members 96 to the hub 90. In this embodiment, the ligaments or connecting members 96 will have a portion on either side of wall or feature 94.
In an alternative embodiment and as illustrated in at least
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
