Abstract: An afterburner for use with a gas turbine engine includes a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from a pilot can be injected. The exit apertures can be protrusions or slots in some forms. In some embodiments, cooling passages are arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

Abstract: An afterburner is disclosed for use with a gas turbine engine and, in one form, is structured to receive a bypass air. The afterburner can be situated in a bypass duct and can be a toroidal combustor or a can combustor. In one embodiment, the afterburner includes a combustor arranged to receive bypass air and a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from the afterburner combustor can be injected. The exit apertures can be protrusions or slots in some forms. In one embodiment, a cooling passage is arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

Abstract: An afterburner for use with a gas turbine engine includes a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from a pilot can be injected. The exit apertures can be protrusions or slots in some forms. In some embodiments, cooling passages are arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

Abstract: An afterburner is disclosed for use with a gas turbine engine and, in one form, is structured to receive a bypass air. The afterburner can be situated in a bypass duct and can be a toroidal combustor or a can combustor. In one embodiment, the afterburner includes a combustor arranged to receive bypass air and a plurality of vanes distributed downstream of a turbine of the gas turbine engine. The vanes can include one or more exit apertures through which hot combustion flow from the afterburner combustor can be injected. The exit apertures can be protrusions or slots in some forms. In one embodiment, a cooling passage is arranged around the exit apertures. An upstream vane portion can be positioned to inject fuel to be combusted via interaction with hot flow that is discharged through the exit apertures.

Abstract: A combustion system including a plurality of axially staged tubular premixers to control emissions and minimize combustion noise. The combustion system includes a radial inflow premixer that delivers the combustion mixture across a contoured dome into the combustion chamber. The axially staged premixers having a twist mixing apparatus to rotate the fluid flow and cause improved mixing without causing flow recirculation that could lead to pre-ignition or flashback.

Abstract: A combustion system including a plurality of axially staged tubular premixers to control emissions and minimize combustion noise. The combustion system includes a radial inflow premixer that delivers the combustion mixture across a contoured dome into the combustion chamber. The axially staged premixers having a twist mixing apparatus to rotate the fluid flow and cause improved mixing without causing flow recirculation that could lead to pre-ignition or flashback.

Abstract: A combustor diffuser assembly for a gas turbine engine includes a prediffuser for receiving compressed air from a gasifier impeller of a gas turbine engine and for directing it into the inlet of a vortex controlled downstream diffuser and wherein the prediffuser and vortex controlled diffuser are spaced to form gaps for air bleed to form vortices to minimize bleed and main diffuser flow pressure drop and to minimize distortion in compressed air flow to plural passages leading to primary and secondary air openings in the inner and outer walls of an annular combustor assembly for the gas turbine engine.

Abstract: A combustion liner for a gas turbine engine operating at high temperatures has portions of the combustion liner double-walled, with a narrow clearance between the inner and outer walls through which cooling air flows to cool the liner. The surfaces of the walls which define these cooling air inlets are roughened to increase heat transfer from the air to the wall. The walls are joined and held apart by spacing strips extending axially of the liner and bonded to the overlapping wall sections. Large combustion air inlet holes extend through the overlapping wall sections at some points. Auxiliary air inlets for liner cooling are provided downstream of these combustion air inlet holes.

Abstract: A gas turbine engine combustor assembly of unique configuration has an outer wall made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween by a reinforcing and heat dissipation ring and a unique weld configuration to prevent thermal erosion of the ends of the porous metal panels at the butt joints; the combustor further including a unique inner wall made up of a plurality of like axially extending multi-layered porous metal panels joined at butt joints by a reinforcing and heat dissipation ring on the inner surface of the inner wall panels and an improved butt weld joint that prevents thermal erosion of the ends of the porous metal inner wall panels.

Abstract: A gas turbine engine combustor assembly of annular configuration has outer and inner walls made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween and each outer and inner wall including a transition panel of porous metal defining a combustor assembly outlet supported by a combustor mount assembly including a stiffener ring having a side undercut thereon fit over a transition panel end face; and wherein an annular weld joins the ring to the end face to transmit exhaust heat from the end face to the stiffener ring for dissipation from the combustor; a combustor pilot member is located in axially spaced, surrounding relationship to the end face and connector means support the stiffener ring in free floating relationship with the pilot member to compensate for both radial and axial thermal expansion of the transition panel; and said connector means includes a radial gap for maintaining a controlled flow of coolant from outside of the transition pa

Abstract: A combustion liner for a gas turbine combustion apparatus or the like has a head wall and a side wall extending downstream to an outlet for combustion products. A reaction zone is contained within the upstream end of the side wall and a dilution zone downstream. A barrier ring extends inwardly from the side wall to provide a constricted path between the zones. Combustion air is admitted from a cluster of nozzles in the head wall and a ring of nozzles extending forwardly from the side wall. Each nozzle in the head wall is axially aligned with one of the nozzles in the side wall so that air jets entering into the combustion zone through each such pair of nozzles impinge to create a high degree of turbulence. Fuel is jetted against a spatter plate in each nozzle in the head wall so that the fuel is atomized and entrained by the air flowing through the head wall air nozzle.

Abstract: A combustion apparatus for gas turbine engines particularly adapted to reduce emissions to meet automotive requirements. The fuel is laid on the wall of a cylindrical prechamber and evaporated from the wall by combustion air which is introduced through a swirler at the upstream end of the prechamber. The inner surface of the prechamber is artificially roughened by a grid of grooves to improve fuel evaporation. The fuel is laid on the wall from an annular manifold extending around the upstream end of the prechamber through tangential orifices leading from the manifold into the interior of the prechamber. The fuel manifold is cooled and shielded from heat by an air jacket. More air enters through entrance ports distributed around the prechamber toward its downstream end. The resulting lean fuel-air mixture is delivered past an annular flow dam at the outlet of the prechamber into a combustion or reaction zone which is abruptly enlarged from the prechamber.