Illustrative embodiments pertain to the art of turbomachinery, and specifically to struts of gas turbine engines.
Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust. The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine extracts energy from the expanding combustion gas, and drives the compressor via a common shaft. Energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.
The individual compressor and turbine sections in each spool are subdivided into a number of stages, which are formed of alternating rows of rotor blade and stator vane airfoils. The airfoils are shaped to turn, accelerate and compress the working fluid flow, or to generate lift for conversion to rotational energy in the turbine.
The combustor section includes a combustor where combustion takes place. In a gas turbine engine, the combustor is fed high pressure air by the compressor section. The combustor then heats this air at constant pressure. After heating, air passes from the combustor section through the turbine section (vanes and rotating blades). A combustor must contain and maintain stable combustion despite very high air flow rates. To do so combustors are carefully designed to first mix and ignite the air and fuel, and then mix in more air to complete the combustion process. Combustors play a crucial role in determining many operating characteristics of a gas turbine engine, such as fuel efficiency, levels of emissions, and transient response (i.e., the response to changing conditions such as fuel flow and air speed).
In typical gas turbine engine arrangements, the combustor is supported by an on-board injector. Such support is typically accomplished by a tab which bolts a combustor inner aft support shell and the on-board injector with contact between the combustor and a first vane of a turbine section through conformal seals. The use of an aft combustor tab typically requires use of a seal between the combustor shell and the first vane/on-board injector to reduce air leakage at these interfaces. Although such design may provide weight efficiencies from a compact design perspective, there may be various drawbacks to such configurations. Accordingly, improved coupling and mounting of a combustor in gas turbine engines may be advantageous.
According to some embodiments, combustor mounting structures for gas turbine engines are provided. The combustor mounting structures include a shell portion defining an outer ring arranged at an outer radius relative to a central axis, a combustor connection element defining an inner ring arranged at an inner radius that is less than the outer radius relative to the central axis, and a plurality of struts extending radially between and connecting the shell portion to the combustor connection element, wherein one or more flow apertures are defined between the shell portion and the combustor connection element in a radial direction and between adjacent struts of the plurality of struts in a circumferential direction.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that the shell portion is part of a combustor shell.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that combustor mounting structure is configured to fixedly engage with at least one of a diffuser case and an on-board injector of the gas turbine engine, and the central axis is an engine central longitudinal axis.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that each strut of the plurality of struts has a geometry configured to provide flexibility or relative movement between the shell portion and the combustor connection element.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that each strut of the plurality of struts includes a geometry such that the shell portion and the combustor connection element are offset in an axial direction along the central axis passing through the center of the inner and outer rings.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that each strut of the plurality of struts includes at least one convolution.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that each strut of the plurality of struts includes at least one omega-shape geometry.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that each strut of the plurality of struts includes a portion that runs parallel to the central axis passing through the center of the inner and outer rings.
In addition to one or more of the features described above, or as an alternative, further embodiments of the combustor mounting structures may include that the combustor connection element comprises one or more mounting apertures configured to enable mounting of the combustor mounting structure within the gas turbine engine.
According to some embodiments, gas turbine engines are provided. The gas turbine engines include a combustor section having a combustor arranged within a diffuser case, a turbine section arranged aft of the combustor section along an engine central longitudinal axis, the turbine section having a first vane, and a combustor mounting structure for mounting the combustor within the gas turbine engine forward of the first vane. The combustor mounting structure includes a shell portion defining an outer ring at an outer radius, a combustor connection element defining an inner ring arranged at an inner radius that is less than the outer radius, and a plurality of struts extending radially between and connecting the shell portion to the combustor connection element, wherein one or more flow apertures are defined between the shell portion and the combustor connection element in a radial direction and between adjacent struts of the plurality of struts in a circumferential direction.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that the shell portion is part of a combustor shell of the combustor.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include an on-board injector arranged radially inward from the first vane.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that the combustor mounting structure is configured to fixedly engage with at least one of the diffuser case and the on-board injector.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include a fastener configured to fixedly connect the combustor connection element of the combustor mounting structure, the diffuser case, and the on-board injector.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that each strut of the plurality of struts has a geometry configured to provide flexibility or relative movement between the shell portion and the combustor connection element.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that each strut of the plurality of struts includes a geometry such that the shell portion and the combustor connection element are offset in an axial direction along the engine central longitudinal axis passing through the center of the inner and outer rings.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that each strut of the plurality of struts includes at least one convolution.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that each strut of the plurality of struts includes at least one omega-shape geometry.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that each strut of the plurality of struts includes a portion that runs parallel to the engine central longitudinal axis passing through the center of the inner and outer rings.
In addition to one or more of the features described above, or as an alternative, further embodiments of the gas turbine engines may include that the combustor connection element comprises one or more mounting apertures configured to enable mounting of the combustor mounting structure within the gas turbine engine.
The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: The subject matter is particularly pointed out and distinctly claimed at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which like elements may be numbered alike and:
Detailed descriptions of one or more embodiments of the disclosed apparatus and/or methods 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 Ax 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 can be 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 Ax 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)/(514.7° 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).
Although the gas turbine engine 20 is depicted as a turbofan, it should be understood that the concepts described herein are not limited to use with the described configuration, as the teachings may be applied to other types of engines such as, but not limited to, turbojets, turboshafts, and turbofans wherein an intermediate spool includes an intermediate pressure compressor (“IPC”) between a low pressure compressor (“LPC”) and a high pressure compressor (“HPC”), and an intermediate pressure turbine (“IPT”) between the high pressure turbine (“HPT”) and the low pressure turbine (“LPT”).
As discussed above, a combustor of the combustor section may be supported by an on-board injector (e.g., a tangential on-board injector or “TOBI”). For example, turning to
The combustor 210 includes a shell 214 that is mounted to, at least, a diffuser case 216 at a flange connection 218. The flange connection 218 fixedly connects the shell 214 of the combustor 210, an inner portion of the diffuser case 216, and an on-board injector 220. As shown, the flange connection 218 is located at a forward end (along the engine central longitudinal axis Ax) of the first vane 212 and slightly radially inward therefrom. The flange connection 218 bolts or otherwise fastens the combustor inner aft support shell (shell 214) to the on-board injector 220. Furthermore, there is contact between the combustor 210 and the first vane 212 through seals (e.g., conformal seals) along a gas path. For example, the use of an aft-combustor tab to join at the flange connection 218 typically requires use of a seal between the shell 214, the first vane 212, and the on-board injector 220 to reduce air leakage at the interface between the components.
Although there are benefits to this type of configuration (e.g., weight efficiency from a compact design), there may be drawbacks as well. For example, inclusion of the necessary seals results in additional components that are subject to the environment, and thus may wear, fatigue, and/or fail. Wear of the seal can cause loss of bolt preload which can cause both release of part of the seal and/or the bolt itself, which can result in Domestic Object Damage (DOD) in the engine. Further, wear of the seal can result in air loss from an ineffective seal.
Accordingly, embodiments of the present disclosure are directed to a configuration of gas turbine engine components that can eliminate the use of seals at a junction between a combustor section and a turbine section. In accordance with embodiments of the present disclosure, a ring-strut-ring design is provided to connect a combustor shell to a flange or other connection. For example, without limitation, twelve struts may be equally spaced about a circumference of the engine to connect the combustor shell to a flange that is removed or remove from the first vane of the turbine section. As such, a combustor support to the inner diffuser case and an on-board injector (“OBI”) flange may be moved radially inward, which adds weight but ensures a tight diffuser-combustor-OBI flange connection independent of the conformal seal wear.
Turning now to
As shown, the connection structure 316 is located inboard or radially inward (toward the engine central longitudinal axis Ax) relative to the combustor 306 and the first vane 310. The connection structure 316 comprises a diffuser case connection element 318, a combustor connection element 320, and an OBI connection element 322 that are joined or connected by a fastener 324. To enable the inboard or radially inward location of the connection structure 316, the combustor connection element 320 is arranged apart from a shell portion 326 of combustor shell 308 by a strut 328. The shell portion 326, the combustor connection element 320, and the strut 328 form a combustor mounting structure. The strut 328 extends radially inward from the shell portion 326 to the combustor connection element 320 when installed within the gas turbine engine 300. As shown, the shell portion 326 of the combustor shell 308 is arranged to contact or be positioned at a forward end or edge of the first vane 310, and defines, in part, a hot gaspath from the combustor section 302 to the turbine section 304 of the gas turbine engine 300. The shell portion 326, the strut 328, and the combustor connection element 320 are arranged as a ring-strut-ring design/configuration. That is, the shell portion 326 is a ring arranged at an outer diameter, the combustor connection element 320 is arranged as a ring at an inner diameter, and the strut 328 extends between and connects the combustor connection element 320 to the shell portion 326.
The strut 328 is one of a number of struts that are arranged about/between the ring-shapes of the shell portion 326 and the combustor connection element 320. The struts 328 can provide flexibility and allow for relative movement or adjustments between the combustor connection element 320 and the shell portion 326. Such relative movement may occur during operation of the gas turbine engine 300, e.g., at the connection structure 316 and/or between the elements thereof.
Turning now to
As noted, the shell portion 402 attached to the combustor connection element 404 by the struts 406. The struts 406 are structural elements that flexibly connect the shell portion 402 to the combustor connection element 404 to enable mounting of the combustor mounting structure 400 to a gas turbine engine, such as to an on-board injector and diffuser case (and proximate a first vane of a turbine section). The combustor connection element 404 includes one or more mounting apertures 408 for permitting installation of the combustor mounting structure 400 within a gas turbine engine using one or more fasteners. Further, the combustor mounting structure 400 defines one or more flow apertures 410 arranged circumferentially between adjacent struts 406. The flow apertures 410 allow for airflow through the combustor mounting structure 400, such as to enter and flow through an on-board injector located proximate the combustor mounting structure 400 when installed within a gas turbine engine.
The ring-strut-ring design of the combustor mounting structures described herein can be made to reduce thermal stress contributions and improve part life by reducing strut stiffness using a geometric strut design (e.g., a “wind-back” geometry that incorporates a bend, twist, omega-shape, convolution, etc. geometries). Such geometric designs can reduce radial and axial stiffness and minimize the ring-strut-ring thermal stress by allowing thermal flexibility. Various geometries may be employed for the strut design without departing from the scope of the present disclosure. The strut, and the geometry thereof, may be selected to provide flexibility and the ability of relative movement between components when installed within a gas turbine engine.
Turning now to
The various configurations of
Advantageously, embodiments described herein allow improved mounting and operation of gas turbine engines, particularly at a junction between a combustor section and a turbine sanction thereof. For example, advantageously, embodiments described herein can divorce the conformal seal wear from the combustor bolts, support or eliminate the need for inner conformal seal, and provide for improved durability life (e.g., low cycle fatigue and crack growth lives). Further, advantageously, the strut design of embodiments of the present disclosure does not block the combustor dilution holes or interfere with on-board injector hardware. Furthermore, advantageously, various geometries and/or embodiments shown and described herein may minimize the local peak stress concentration by use of a strut having a gradual thickness transition between the outer (shell portion) and inner (combustor connection element) rings.
As used herein, 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” may include a range of ±8%, or 5%, or 2% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein.
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. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” “radial,” “axial,” “circumferential,” and the like are with reference to normal operational attitude and should not be considered otherwise limiting.
While the present disclosure has been described with reference to an illustrative 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.
This application claims the benefit of U.S. Provisional Application Ser. No. 62/870,133 filed Jul. 3, 2019, the disclosure of which is incorporated herein by reference in its entirety.
This invention was made with Government support awarded by the United States. The Government has certain rights in this invention.
Number | Date | Country | |
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62870133 | Jul 2019 | US |