Patent ID: 8001793
Filing Date: 2011-08-23
Classification: B23P,F02C,F05D,F23M,F23R,Y02T,Y10T

Abstract:
1. A reverse flow combustor for an aircraft gas turbine engine, the combustor comprising: a generally axially-extending annular combustion chamber, the chamber provided between co-operating inner and outer liners, each liner being a single piece of sheet metal, the single piece of sheet metal extending from a radially extending flat dome portion through an axially extending cylindrical portion to a U-shaped exit duct portion, the liners mating to provide a liner assembly having a dome and an exit, the dome provided by overlapping flat dome portions of the inner and outer liners, the dome including a plurality of fuel nozzle openings, each sheet metal liner being louverless and thereby having a smooth and continuous inner surface along its length from dome to exit, the inner and outer liner each having an annular ring mounted to an outer side adjacent the liner assembly exit, the rings configured for connection of the liner assembly to the gas turbine engine; a dome heat shield assembly mounted inside the combustion chamber to the dome and spaced apart from the dome to provide a plenum between the heat shield and the dome, the combustor otherwise being heat shield-less, the dome heat shield assembly including a plurality of fuel nozzle openings aligning with the fuel nozzle openings of the dome; and a thermal barrier coating applied to at least a portion of the inner surface of each liner, wherein the liner assembly further includes a plurality of effusion patches, each patch having a plurality of effusion cooling holes provided at an acute angle through the liner, the effusion holes provided in a plurality of axially spaced rows extending around the liner circumference, each effusion patch extending around the circumference of its respective liner, each effusion patch bounded on both of its axial sides by an effusionless band, each effusionless band being a portion of the liner sheet metal which is free from effusion cooling holes, wherein a first said effusionless band is provided on each liner on the axially extending cylindrical portion immediately adjacent the radially extending flat dome portion, the first effusionless band being cooled by a film cooling jet fed from a first and a second source upstream of the of the band, the first source being at least one aperture through the heat shield feeding the plenum, the second source being a plurality of apertures provided through the radially extending flat dome portion feeding the plenum, the first and second sources sized to provide adequate an film cooling flow over a complete length of the first effusionless band, the heat shield and liner co-operating to direct cooling air from the plenum top provide the film cooling jet over the first effusionless band, and wherein a second said effusionless band provided on each liner adjacent its exit end, the band provided in an at least partially overlapping relationship with a respective one of said mounting rings, the rings having a plurality of impingement holes therethrough to permit impingement cooling of the outer side of the liner at said effusionless band.