Patent Application
20160377288
This application relates to components which are to be attached in a hot gas path in a gas turbine engine.
Gas turbine engines are known, typically include compressor compressing air and delivering it into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. Eventually the products of combustion leave through an exhaust nozzle. In some engine types, an after burner may be provided adjacent the exhaust nozzle.
The combustor, and everything downstream of the combustor, could be in the path of hot products of combustion. Components utilized in this hot flow path are subject to challenges due to the high temperatures. Thus, liners are utilized at many of these locations. As an example, the combustor is often provided with combustor liners, as are the exhaust nozzle, and the after burner. Historically, these liners have been attached to an outside housing, and the liners have a web facing the products of combustion, and end legs bent back toward the housing at a sharp angle. The sharp angle creates a corner.
The corner provides a location for initiation of cutting or burning of the metal. In addition, while it is desirable to provide coatings on such panels, it is difficult to apply a coating to a sharp corner.
In a featured embodiment, a gas turbine engine section has a housing and a plurality of panels attached to the housing. The panels face toward a flow path of hot products of combustion. The panels include a central web and extending legs. A bend between the central web and the extending legs is formed at a radius.
In another embodiment according to the previous embodiment, the central web extends along a direction having at least a component parallel to an axis of an engine which is to receive the section.
In another embodiment according to any of the previous embodiments, the housing is part of a combustor in a gas turbine engine.
In another embodiment according to any of the previous embodiments, a bend from the central web at a first end extends into one of the legs and leads to a foot which extends axially away from the web. A bend into one of the legs at a second end of the panel has the leg positioned on the foot of an adjacent one of the panels.
In another embodiment according to any of the previous embodiments, a coating is provided on the panel.
In another embodiment according to any of the previous embodiments, a bend from the central web at a first end extends into one of the legs and leads to a foot which extends axially away from the web. A bend into one of the legs at a second end of the panel has the leg positioned on the foot of an adjacent one of the panels.
In another embodiment according to any of the previous embodiments, a coating is provided on the panel.
In another embodiment according to any of the previous embodiments, the housing is part of an exhaust nozzle.
In another embodiment according to any of the previous embodiments, the housing is part of a turbine section.
In another featured embodiment, a gas turbine engine has a combustor section, a turbine section and an exhaust nozzle, with one of the combustor, the turbine section, and the exhaust nozzle being formed with a plurality of panels attached to a housing. The plurality of panels faces toward a flow path of hot products of combustion. The panels include a central web and extending legs, with a bend between the central web and the extending legs formed at a radius.
In another embodiment according to the previous embodiment, the central web extends along a direction having at least a component parallel to an axis of rotation of the engine.
In another embodiment according to any of the previous embodiments, a bend from the central web at a first end extends into one of the legs and leads to a foot which extends axially away from the web. A bend into one of the legs at a second end of the panel has the leg positioned on the foot of an adjacent one of the panels.
In another embodiment according to any of the previous embodiments, a coating is provided on the panel.
In another embodiment according to any of the previous embodiments, the housing is part of a combustor in a gas turbine engine.
In another embodiment according to any of the previous embodiments, a bend from the central web at a first end extends into one of the legs and leads to a foot which extends axially away from the web. A bend into one of the legs at a second end of the panel has the leg positioned on the foot of an adjacent one of the panels.
In another embodiment according to any of the previous embodiments, a coating is provided on the panel.
In another embodiment according to any of the previous embodiments, the housing is part of an exhaust nozzle.
In another embodiment according to any of the previous embodiments, the housing is part of a turbine section.
These and other features of this application may be best understood from the following specification and drawings, the following which is a brief description.
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. A mid-turbine frame 57 of the engine static structure 36 is 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 over 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 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 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. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (“TSFC”)”—is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf 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.
An exhaust nozzle 19 is shown. An afterburner may be included in the nozzle.
The liners 104 can be seen to have a web or face 108 which extends generally along the axis of rotation A of the engine. It could be said that the web 108 extends along a direction having at least a component parallel to the axis of rotation A. Of course, the web can deviate from being directly parallel. The web 108 face radially inwardly, and face a hot gas flow H. Legs 110 are formed at ends of the web 108, and extend generally radially outwardly from the web 108. This creates a sharp corner 112. As mentioned above, sharp corners 112 provide a location for initiation of burning or cutting of the metal, and further complicate the application of a coating.
Forming the bend 156 at a radius eliminates the sharp corner 112 of the prior art. The other end of the panels 162 bends into a leg 160 which extends radially outwardly and, in this embodiment, contacts the foot 158. Of course, there may not be contact in other embodiments. Here again, the bend 162 is formed on a radius.
By eliminating the sharp corner, the localized spot for initiation of cutting or burning is also eliminated.
Further, as shown in
While the panels are illustrated in a combustor section, such as combustor section 56, the panels could also be utilized in the turbine section, or in the exhaust nozzle of the
Although an embodiment of this invention has been disclosed, a worker of ordinary skill in the art would recognize that certain modifications would come within a scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
This application claims priority to U.S. Provisional Application No. 61/846,649, filed Jul. 16, 2013.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US14/44232 | 6/26/2014 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| 61846649 | Jul 2013 | US |
