Abstract: A combustor for a gas turbine engine is provided which includes a plurality of liner segments and a support shell. Each liner segment includes a panel, a forward wall, a trailing wall, a pair of side walls, and a plurality of mounting studs. The panel includes a face surface and a back surface. The forward wall is positioned along a forward edge of the panel and the trailing wall is positioned along a trailing edge of the panel. The side walls connect the forward and trailing walls. The forward, trailing, and side walls extend out from said back surface a particular distance. The plurality of mounting studs extend out from the back surface, and each includes liner segment attachment apparatus.

Abstract: A high performance annular combustor for a gas turbine engine powering aircraft consisting of a primary combustion zone with the fuel nozzles and air swirlers in the dome and having typical combustion/dilution air holes in the liner and a secondary zone downstream of the primary zone having fuel nozzles and air swirlers operational solely when the primary zone is at a predetermined stoichiometric condition and issuing fuel which is at parity with the fuel/air ratio of the primary zone.

Abstract: An annular combustor for a gas turbine engine is constructed to include a sector that has a reduced airflow distribution than the remainder of the circumference by supplying a reduced airflow in a judiciously located portion of the fuel nozzles and the single stage of radially ingested combustion airflow. This single stage circumferential zoned annular combustor serves to accommodate higher fuel/air ratios.

Abstract: An annular combustor for a gas turbine engine includes a sector in the primary combustion zone that operates at a reduced level of airflow than the remaining portion of the combustor by providing fixed air openings in the dome or fuel nozzles and in the combustion air admission holes. This allows the combustor to operate at higher fuel/air ratios with a minimum of complexity in the hardware.

Abstract: An air blast fuel nozzle for a gas turbine engine having a variable valve for controlling the flow of fuel to the combustor of the engine and valve bypass means for the flow of a low quantity of fuel to said combustor without passing through said variable valve to provide transient operability during stall recovery and spool down starting.

Abstract: The fuel nozzles in the combustor of a gas turbine engine are concentrically disposed in two tiers about the circumference of the dome and are arranged in a triangular pattern with the air swirl orientation providing an increased intensity of combustion so as to enable a reduction in combustor size and weight while increasing temperature rise and the overall thrust-to-weight ratio of the engine.

Abstract: A gas turbine engine combustor has a plurality of circumferentially fuel injectors. Two lower nozzles are pilot nozzles and are of the fuel pressure atomized type, and the majority of nozzles are of the air atomized type. A pressurizing valve maintains high fuel pressure to the pilot nozzles at all operating conditions. The fuel/air ratio for the pilot nozzles is optimized facilitating start up and increased stability during transient conditions.

Abstract: A method for distributing a flow of fuel over the gas flow area of a gas turbine engine includes the steps of dividing the gas flow area into a plurality of sector shaped sub-areas (48), (50), (52), (54) disposed within a surrounding, annular sub-area (46) which is in turn disposed adjacent the augmentor wall (12). During operation, the method according to the present invention distributes the fuel over one or more of the sub-areas (46-54) in response to the current rate of fuel flow. At the lowest and intermediate fuel flow rates, fuel is distributed only over one or more of the sector shaped sub-areas (48-54). Only at the very highest rates of augmentor fuel flow is fuel distributed finally in the outer annular sub-area (46).

Abstract: An augmentor in a turbofan jet engine utilizes a crossover chute assembly direct portions of the bypass fan stream inwardly and portions of the hot core stream outwardly into alternating radial zones of fan and core stream flow; and also annularizes a portion of the core stream around those zones. A plurality of angularly positionable radial vanes causes rotational swirl of the radial zones in one direction to produce buoyant "g" forces that effect rapid inward propagation of a flame front from the annularized zone.

Abstract: This invention relates to the counter-parallel cooling flow for a combustor liner with means for loosely mounting one of the two spaced liner walls relative to the other to support it so that it "floats". The counter-paralled flow provides both convenctive and film cooling permitting the floating wall to be in axially spaced segments relative to the flow of combustor air. The floating wall is segmented in the circumferential direction to permit relatively minimal stress during thermal expansion.

Abstract: An improved cooling arrangement is provided for cooling a portion of a combustor dome assembly of a gas turbine engine. The improvement is comprised of an annular dome plate having a generally V-shaped cross section and including one or more circumferentially spaced apart, generally radial cooling slots within a surface which is exposed to the combustion zone.

Abstract: In a gaseous fuel delivery system for delivering gaseous fuel to the combustor of a gas turbine engine wherein the delivery system includes a fuel injector having a nozzle orifice, a gaseous fuel supply passage providing fluid communication between the nozzle orifice and a source of gaseous fuel under pressure, an improvement is presented comprising means for preventing pressure perturbations occurring within said combustor from compressing gaseous fuel in at least a portion of the fuel delivery system.

Abstract: A combustion liner cooling slot is provided with a plurality of circumferentially spaced stabilizing dimples disposed in an annular lip adapted to facilitate attachment of a cooling fluid to the combustion liner in a protective film barrier. The stabilizing dimples are characterized by a depressed portion having a dimple ceiling of constant radial height and further having a decreasing circumferential width portion.

Abstract: A combustion chamber is provided with a fuel source and an igniter for initiating combustion. The igniter projects into the combustion chamber and into proximity with the fuel source. A heated tip portion of the igniter initiates burning and is cooled by means of cooling air passing through a plurality of air passages defined by a plurality of channels near the external surface of the igniter housing and further defined by a surrounding ferrule. Certain of the channels extend axially of the housing while at least one channel extends circumferentially thereabout and intersects the former channels. An opening in the ferrule provides communication between preselected of the channels and the interior of the combustion chamber.

Abstract: An apparatus for passing a portion of a cooling fluid, flowing in a plenum in a predetermined direction, in a protective film upon a liner defining a hot gas passage. The apparatus is comprised of a first liner portion or segment partially defining the plenum and a second liner portion or segment partially defining the hot gas passage and cooperating with the first liner segment to form a pocket. A first lip is located at the downstream end of the second liner segment and is radially adjacent a second lip located at the upstream end of the first liner segment. The first lip is spaced from the first liner segment to form an exit from the pocket for directing cooling air along the first liner segment in the predetermined direction. The apparatus further comprises a plurality of apertures in the first liner to pass cooling fluid in a plurality of streams into the pocket and along the second lip in a direction substantially opposite from said predetermined direction.