Annular one-piece corrugated liner for combustor of a gas turbine engine

#The present invention relates generally to a liner for the combustor of a gas turbine engine and, in particular, to an annular one-piece corrugated liner of substantially sinusoidal cross-section where the amplitude of the corrugations and/or the wavelength between adjacent corrugations is varied from an upstream end to a downstream end.

#Combustor liners are generally used in the combustion section of a gas turbine engine located between the compressor and turbine sections of the engine, although such liners may also be used in the exhaust sections of aircraft engines that employ afterburners. Combustors generally include an exterior casing and an interior combustor where fuel is burned to produce a hot gas at an intensely high temperature (e.g., 3000° F. or even higher). To prevent this intense heat from damaging the combustor case and the surrounding engine before it exits to a turbine, a heat shield or combustor liner is provided in the interior of the combustor.

#One type of liner design includes a number of annular sheet metal bands which are joined by brazing, where each band is subject to piercing operations after forming to incorporate nugget cooling holes and shaped dilution holes. Each band is then tack welded and brazed to the adjacent band, with stiffeners known as “belly bands” being tack welded and brazed to the sheet metal bands. The fabrication of this liner has been found to be labor intensive and difficult, principally due to the inefficiency of brazing steps applied to the stiffeners and sheet metal bands.

#FIG. 1 is a cross-sectional view of a gas turbine engine including a combustor liner in accordance with the present invention;

#FIG. 2 is an enlarged, cross-sectional view of the combustor depicted in FIG. 1;

#FIG. 3 is a partial perspective view of the outer liner for the combustor depicted in FIGS. 1 and 2 in accordance with the present invention;

#Referring now to the drawings in detail, wherein identical numerals indicate the same elements throughout the figures, FIG. 1 depicts an exemplary gas turbine engine 10 having in serial flow communication a low pressure compressor 12, a high pressure compressor 14, and a combustor 16. Combustor 16 conventionally generates combustion gases that are discharged therefrom through a high pressure turbine nozzle assembly 18, from which the combustion gases are channeled to a conventional high pressure turbine 20 and, in turn, to a conventional low pressure turbine 22. High pressure turbine 20 drives high pressure compressor 14 through a suitable shaft 24, while low pressure turbine 22 drives low pressure compressor 12 through another suitable shaft 26, all disposed coaxially about a longitudinal or axial centerline axis 28.

#As seen in FIG. 2, combustor 16 further includes a combustion chamber 30 defined by an outer liner 32, an inner liner 34, and a dome 36 located at an upstream end thereof. It will be seen that a fuel/air mixer 38 is located within dome 36 so as to introduce a mixture of fuel and air into combustion chamber 30, where it is ignited by an igniter (not shown) and combustion gases are formed which are utilized to drive high pressure turbine 20 and low pressure turbine 22, respectively.

#In accordance with the present invention, it will be noted from FIGS. 3 and 4 that outer liner 32 is annular in shape and preferably formed as a one-piece construction from a type of sheet metal. More specifically, outer liner 32 includes a first end 42 located adjacent to an upstream end of combustor 16, where first end 42 is connected to a cowl 44 and dome 36 by means of a rivet band 40 (which is in turn connected to cowl 44 and dome 36 via a mechanical connection such as bolt 46 and nut 48, a welded connection, or other similar form of attachment). Accordingly, it will be appreciated that outer liner 32 is preferably connected to rivet band 40 via rivets 41 and therefore eliminates the need for outer liner 32 to have a flange formed thereon at upstream end 42. Starter slots 55 and 57 are preferably provided in rivet band 40 and upstream outer liner end 42, respectively, to promote a cooling film along the hot side of outer liner 32. Outer liner 32 also includes a second end 50 located adjacent to a downstream end of combustor 16, where second end 50 is preferably connected to a seal assembly 52 by means of rivets 53. In this way, outer liner 32 is able to move axially in accordance with any thermal growth and/or pressure fluctuations experienced.