1. Field of the Invention
This invention relates to film cooled combustor liners for use in gas turbine engines, and more particularly, to aircraft gas turbine engine combustor liners having cooling holes that are angled in the circumferential direction.
2. Description of Related Art
Combustor liners are generally used in the combustion section of a gas turbine engine which is located between the compressor and turbine sections of the engine. Combustor liners are also used in the exhaust section of aircraft engines that have afterburners. Combustors generally include an exterior casing and an interior combustor wherein fuel is burned producing a hot gas usually at an intensely high temperature such as 3,000.degree. F. or even higher. To prevent this intense heat from damaging the combustor before it exits to a turbine, a heat shield or combustor liner is provided in the interior of the combustor. This combustor liner thus prevents the intense combustion heat from damaging the combustor or surrounding engine.
Prior methods for film cooling combustion liners provided circumferentially disposed rows of film cooling slots such as those depicted in U.S. Pat. No. 4,566,280 by Burr and U.S..Pat. No. 4,733,538 by Vdoviak et al. which are typified by complex structures that have non-uniform liner thicknesses which give rise to thermal gradients which cause low cycle fatigue in the liner and therefore shorten their potential life expectancy and reduce their durability. The complex shapes and machining required to produce these liners negatively effects their cost and weight.
A more detailed discussion of the related art may be found in a related U.S. patent application Ser. No. 07/614,418 entitled "GAS TURBINE ENGINE MULTI-HOLE FILM COOLED COMBUSTOR LINER AND METHOD OF MANUFACTURE", invented by Wakeman et al., filed Nov. 15, 1990, assigned to the same assignee, and incorporated herein by reference.
Engine designers have long sought to incorporate low weight single wall combustor liners capable of withstanding the temperatures and pressure differentials found in combustors. To that end the invention described in the Wakeman reference provides a single wall, preferably sheet metal, annular combustor liner having multi-hole film cooling holes which are disposed through the wall of the liner at sharp downstream angles. The multi-hole film cooling holes are spaced closely together to form at least one continuous pattern designed to provide film cooling over the length of the liner. The present invention provides multi-hole film cooling holes which have a diameter of about 20 mils with a nominal tolerance of about .+-.2 mils, are spaced closely together about 61/2 to 71/2 hole diameters apart, have a downstream angle of 20 degrees with a nominal tolerance of about .+-.1 degree, and a circumferential angle with respect to the engine center-line of between 30 and 65 degrees. Axially adjacent holes are circumferentially offset by half the angle between circumferentially adjacent holes to further enhance the evenness of the cooling film injection points. The Wakeman invention further provides an embodiment wherein the combustor liner may be corrugated so as to form a way wall which is designed to prevent buckling and is particularly useful for outer burner liners in the combustion section of gas turbine engines and exhaust duct burner liners in aircraft gas turbine engines having afterburners.
A phenomena which occurs both in the main combustion section and in the afterburner combustion section is swirl, wherein swirled patterns of higher thermal degredation areas are formed on the liner. The patterns generally coincide with the swirl of the combustor flow induced by swirlers in the fuel nozzles to promote better combustion and in the exhaust section by turbine nozzles.