Patent Application
20190078469
Exemplary embodiments pertain to the art of gas turbine engines and, more particularly, to a fan exit stator assembly retention system.
In a gas turbine engine used for propulsion, a fan case and a smaller diameter compressor case cooperate to radially bound an annular fan duct. Fan exit guide vanes, or stators, span across the fan duct to de-swirl working medium fluid flowing therethrough. Typically, fan exit stators do not radially retain an outer diameter shroud of the stator, but rigid bolting at the inner diameter shroud may be present. Upon impact with an object during operation, such as a bird or ice, the outer diameter shroud experiences excessive radial deflection, and a radial load passes through the vane and into a joint between the vane and the inner diameter shroud. No mechanical retention is present at the vane to the inner diameter shroud joint, apart from a thin layer of silicone adhesive. A lack of a robust retention system may result in shroud damage and/or vane withdrawal.
Disclosed is a retention system for a stator vane assembly including a stator vane having a radially inner end and a radially outer end. Also included is an outer diameter shroud coupled to the radially outer end of the stator vane. Further included is an inner diameter shroud coupled to the radially inner end of the stator vane. Yet further included is a flange of the outer diameter shroud coupled to a frame member with a mechanical fastener. Also included is an inner shroud flange extending radially inwardly and defining a radial recess, the radial recess allowing radial movement of the radially inner end of the stator vane.
In addition to one or more of the features described above, or as an alternative, further embodiments may include a slot defined by the stator vane proximate the radially inner end of the stator vane. Also included is a retainer bar insertable in the slot, the retainer bar located on a radially inner side of the inner diameter shroud to prevent withdrawal of the stator vane from the inner diameter shroud.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the retainer bar has a primarily rectangular cross-section.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the radially inner end of the stator vane is a base portion that includes a width that is greater than a width of the remainder of the stator vane.
In addition to one or more of the features described above, or as an alternative, further embodiments may include a slot defined by the base of the stator vane. Also included is a retainer bar insertable in the slot, the retainer bar located on a radially inner side of the inner diameter shroud to prevent withdrawal of the stator vane from the inner diameter shroud.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the retainer bar has a primarily rectangular cross-section.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the stator vane is a fan exit stator located proximate an inlet of a low pressure compressor of a gas turbine engine.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the frame member is a forward center body frame of the gas turbine engine.
Also disclosed is a gas turbine engine that includes a compressor section, a combustion section, and a turbine section. Also included is a retention system for a fan exit stator located proximate an inlet of the compressor section. The retention system includes an outer diameter shroud coupled to a radially outer end of the fan exit stator. The retention system also includes an inner diameter shroud coupled to a radially inner end of the fan exit stator. The retention system further includes a flange of the outer diameter shroud coupled to a forward center body frame with a mechanical fastener. The retention system yet further includes an inner shroud flange extending radially inwardly and defining a radial recess, the radial recess allowing radial movement of the radially inner end of the fan exit stator.
In addition to one or more of the features described above, or as an alternative, further embodiments may include a slot defined by the fan exit stator proximate the radially inner end. Also included is a retainer bar insertable in the slot, the retainer bar located on a radially inner side of the inner diameter shroud to prevent withdrawal of the fan exit stator from the inner diameter shroud.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the retainer bar has a primarily rectangular cross-section.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the radially inner end of the fan exit stator is a base portion that includes a width that is greater than a width of the remainder of the fan exit stator.
In addition to one or more of the features described above, or as an alternative, further embodiments may include a slot defined by the base of the fan exit stator. Also included is a retainer bar insertable in the slot, the retainer bar located on a radially inner side of the inner diameter shroud to prevent withdrawal of the fan exit stator from the inner diameter shroud.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the retainer bar has a rectangular cross-section.
Further disclosed is a method of retaining a fan exit stator of a gas turbine engine. The method includes coupling an outer diameter shroud to a forward center body frame with a mechanical fastener. The method also includes inserting a radially inner end of the fan exit stator through an opening of an inner diameter shroud. The method further includes operatively coupling the inner diameter shroud to a frame member of the gas turbine engine at an inner shroud flange, the inner shroud flange defining a radial recess to allow radial movement of the fan exit stator.
In addition to one or more of the features described above, or as an alternative, further embodiments may include inserting a retainer bar through a slot defined by the fan exit stator proximate the radially inner end of the fan exit stator subsequent to insertion of the radially inner end of the fan exit stator through the opening of the inner diameter shroud to radially retain the fan exit stator relative to the inner diameter shroud.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the frame member is the forward center body frame.
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 feet (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 to
For purposes of description and clarity, one of the fan exit stators 62 is shown and described herein. The fan exit stator functions as an airfoil to remove a substantial circumferential flow component from air exiting the fan section 22. The core air flow C air passes over the fan exit stator 62. A pressure side of an aft section of the fan exit stator 62 guides the entering air so that upon complete passage of the fan exit stator 62, the air flow is in an axial direction. Air exiting the fan section 22 flows to the low pressure compressor 44. The air entering the low pressure compressor 44 first flows past the fan exit stator 62 and then through a front center body duct 64. The air with reduced swirl then flows through inlet guide vanes 66 and first rotors 68 of the low pressure compressor 44.
The fan exit stator 62 is radially bound by an inner diameter shroud 80 proximate a radially inner end 84 of the fan exit stator 62 and by an outer diameter shroud 86 proximate a radially outer end 87 of the fan exit stator 62. However, prior fan exit stators do not include radial retention of the outer diameter shroud 86. The embodiments described herein, provide such outer diameter retention, as well as a more structurally reliable inner diameter retention assembly.
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The features of the retention system described above, provide a rigid outer diameter shroud connection to the supporting frame, a radial spline connecting the inner diameter shroud to the frame, an extended vane inner diameter, and a retainer inserted into the vane extension. The system substantially reduces outer diameter shroud deflection and radial load on the fan exit stator 62 while the retainer and vane extension prevent stressing the adhesive joint and vane withdrawal. By rigidly constraining the outer diameter shroud, damage to the shroud due to excessive deflection is avoided. A rigid outer diameter, combined with an inner diameter shroud radial spline reduces radial load on the vane to inner diameter shroud joint, and combined with a retainer protects the adhesive joint and avoids vane pull out from the inner diameter shroud.
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. For example, “about” can include a range of ±8% or 5%, or 2% of a given value.
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.
