Abstract: A seal assembly can accommodate deflection between two components through a cartridge that is slidingly engaged with a slide plate.

Abstract: A seal assembly can accommodate deflection between two components through a cartridge that is slidingly engaged with a slide plate.

Abstract: A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated alloy structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

Abstract: A combustor for use in a gas turbine engine includes a chute and a combustion liner defining a combustion chamber and a chute-receiving aperture that extends through the combustion liner. The chute extends through the chute-receiving aperture of the combustion liner and defines a passageway sized to convey air from an environment outside the combustion chamber through the combustion liner into the combustion chamber.

Abstract: A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated allow structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

Abstract: A multi stream aircraft fixed geometry nozzle includes an inner nozzle, an outer nozzle, and a supersonic ejector. The outer nozzle is configured to channel a third stream from an aft end of a third stream duct surrounding a bypass duct of a multi stream aircraft engine to the supersonic ejector to merge the third stream with the primary stream. The fixed geometry nozzle is configured to operate between an SFC mode and a thrust mode such that, when the inner nozzle accelerates the primary stream supersonically to the supersonic ejector, at which the primary stream is merged with the third stream, in the SFC mode the total pressure of the primary stream is substantially the same as the total pressure of the third stream, and in the thrust mode the total pressure of the primary stream is substantially greater than the total pressure of the third stream.

Abstract: A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated alloy structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

Abstract: A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated alloy structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

Abstract: A combustor for use in a gas turbine engine includes a chute and a combustion liner defining a combustion chamber and a chute-receiving aperture that extends through the combustion liner. The chute extends through the chute-receiving aperture of the combustion liner and defines a passageway sized to convey air from an environment outside the combustion chamber through the combustion liner into the combustion chamber.

Abstract: A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated allow structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

Abstract: A multi stream aircraft fixed geometry nozzle includes an inner nozzle, an outer nozzle, and a supersonic ejector. The outer nozzle is configured to channel a third stream from an aft end of a third stream duct surrounding a bypass duct of a multi stream aircraft engine to the supersonic ejector to merge the third stream with the primary stream. The fixed geometry nozzle is configured to operate between an SFC mode and a thrust mode such that, when the inner nozzle accelerates the primary stream supersonically to the supersonic ejector, at which the primary stream is merged with the third stream, in the SFC mode the total pressure of the primary stream is substantially the same as the total pressure of the third stream, and in the thrust mode the total pressure of the primary stream is substantially greater than the total pressure of the third stream.

Abstract: A gas turbine engine variable porosity combustor liner has a laminated alloy structure. The laminated alloy structure has combustion chamber facing holes on one side and cooling plenum facing holes on a radially opposite side. The combustion chamber facing holes are in fluid communication with the cooling plenum facing holes via axially and circumferentially extending flow passages sandwiched between metal alloy sheets of the laminated alloy structure. Porous zones having respective different cooling flow amounts are formed in the laminated alloy structure based on at least one of an arrangement of the combustion chamber facing holes, an arrangement of the cooling plenum facing holes, and an arrangement of the flow passages.

Abstract: The present invention provides gas discharge technique useful on vehicle engines. The technique comprises an s-shaped duct having an ejector formed therein to mix hot exhaust gas with a relatively cooler airflow entrained by the ejector. The s-shaped duct may be comprised of a number of segments that when integrated form the s-shape and ejector. A variety of cross sections including circular and rectangular are provided for the s-shaped duct. The segments may be composed of smaller subsections and may further have cooling slots. A nacelle is provided to at least partially enclose the s-shaped conduit and may have an inlet that is in fluid communication with the ejector of the s-shaped conduit.

Abstract: An annular diffuser for a gas turbine engine includes inner and outer walls spaced apart to define a diffuser flow path, and a support member coupling the inner wall to the outer wall while allowing independent radial displacement therebetween. The support member includes a first end portion rigidly connected to the inner wall, and a second end portion extending outside of the diffuser flowpath and coupled to the outer wall by a radially extending pin. An aerodynamically-shaped shroud member surrounds the support member to thermally isolate the support member from the diffuser flowpath, thereby shielding the support member from transient thermal loads. The shroud member is pinned to the support member at a single axial location. A combustor dome panel is attached to the downstream end portion of the shroud member and is adapted to independently support inner and outer combustor liners in spaced relation to define a combustion chamber therebetween.

Abstract: A liner apparatus is provided which includes an inner wall, a plurality of fasteners extending out from the inner wall, an impingement sheet attached to the inner wall by the fasteners, apparatus for spacing the inner wall and the impingement sheet a distance apart, and apparatus for sealing between the inner wall and the impingement sheet along an edge of the liner.

Abstract: A thermal relief slot is formed by including a lateral slot in each of three overlying sheet metal sheets with the upper and lower slots being in overlying position and the middle slot being axially spaced therefrom and the three sheets being bonded into an integral sheet metal sheet assembly except for the area extending over the relief slots. The edge of the upper and/or lower outer edges of adjacent the relief slots may be chamfered to prevent fraying of the edges. A two sheet assembly includes the relief slots axially spaced and two sheet metal sheets are bonded except for the area including and extending between the two relief slots.

Abstract: A gap seal for convergent/divergent flap hinge ends to spaced disc and wall surfaces in a 2-D nozzle of a gas turbine engine is provided. The seal employs at least two cups resiliently mounted to the hinge ends, which cups independently translate axially and resiliently to contact the spaced surfaces at different or varying gaps to seal same. The translating cup seals of the invention thus maintain a constant seal in the gaps between the hinge ends and such surfaces at varying C/D hinge positions and varying operating conditions in the 2-D nozzle. That is, the cups independently maintain their seals across gaps as they expand or contract due to thermal changes as well as across gaps of different sizes.