General Electric Company

United States of America As Represented by the Secretary of the Air Force

John W Vdoviak

Robert Alexander Nicoll

Barry Weinstein

Peter J Rock

Alexander C Bryans

Roy E Moyer

Donald Patrick McHugh

Chester J Lamando, Jr

Dennis C Evans

Neil Roger Brookes

Russell E Motsinger

Joseph K Bussichella

The present invention provides coolable exhaust nozzle flap and seal assemblies having liners that are spaced inward and apart from their respective flap and seal cool walls, to form a cooling passage therebetween that converges in the axial downstream flow direction. One embodiment provides single wall multi-hole film cooled nozzle flap and seal liners particularly for use individually or in combination on convergent and divergent hollow flaps and seals for variable exhaust nozzles. Flap and seal coolable assemblies may be unitary structures that are integrally fabricated or formed such as by casting.

Cooling apparatus for aircraft gas turbine engine exhaust nozzles

An afterburning, gas turbine aircraft engine with an annular duct for carrying bypass airflow is provided with a sliding valve type of rear Variable Area Bypass Injector (VABI) system. This rear VABI varies the direction and volume of the flow of bypass air into separate regions at an afterburning section of the engine. The system includes a sliding valve mounted in the bypass duct that is translated axially to open and close slots leading to core engine exhaust flow in the afterburning section. The unique structure of the injector system enables the axial translation of this sliding valve to simultaneously control bypass airflow into: first, a mixer just upstream of the afterburning section; second, a plurality of slots leading to the afterburning section; and third, an exhaust nozzle cooling liner that is cooled with the remaining portion of the bypass air. The system is actuated with a mechanical design that uses both rotation and translation to simplify the actuation motion so that one actuator can translate the valve. Because the system employs a valve that translates over a plurality of slots, it is commonly referred to as a variable slot bypass injector.

Variable slot bypass injector system

A turbine nozzle includes a circumferentially continuous annular outer band, and a plurality of circumferentially spaced apart vanes extending radially inwardly therefrom. A circumferentially segmented inner band is joined to the vanes and is disposed coaxially with the outer band. The inner band includes a plurality of circumferentially adjoining arcuate segments, adjacent ones of which have circumferentially facing end faces spaced apart to define a gap therebetween. In an exemplary embodiment, strip seals are inserted in corresponding slots formed in the end faces to seal leakage therebetween.

Low leakage turbine nozzle

The present invention provides a cooling apparatus with two spaced apart walls, one of which is a coolable wall, having a cooling flowpath between them wherein the cooling flowpath is converging in the direction of the cooling flow. In one embodiment the coolable wall is an afterburner cooling liner that includes a film cooling means on the hot side of the liner. Another embodiment provides a corrugated or wavy wall afterburner cooling liner wherein each wave includes a forward facing surface, from crest to trough of the wave in the direction of the hot gas flow, and an aft facing surface, from trough to crest. A high density pattern of multi-hole film cooling holes is disposed on the hotter of the two surfaces and a lower density pattern on the cooler one.

Multi-hole film cooled afterburner combustor liner

Multi-hole film cooled afterburner cumbustor liner

This disclosure describes a single shell, double dome combustor design configuration with the intermediate dividing wall between the two annuli removed, thereby reducing weight and cost, and improving durability. An annular daisy type mixer is positioned between the domes to improve the mixing of air and combustor gases prior to the exiting of the gases from the combustor to the turbine vanes. The daisy mixer positioned downstream of the domes to provides effective, low pressure loss mixing at a a controlled axial rate, by creating effective cross-flow barriers to avoid uncontrolled fuel/air ratio transfer between the inner and outer annuli, and thereby allows a concentric and controlled axial air admission in conjunction with the individual swirlers.

Double dome, single anular combustor with daisy mixer

A gas turbine engine of the type having generally cylindrical casings defining concentric bypass and turbine discharge passages including a support structure in the turbine discharge passage. The support structure is cooled with low pressure compressor discharge air through a scoop located in the bypass gas passage. The support structure includes a plurality of radial fairings, each of which includes a discharge chute which is approximately perpendicular to its respective fairing. The discharge chutes each communicate with the compressor discharge cool air flow in its respective fairing so as to provide a discharge flow in the form of a stratified annulus to cool downstream members which are aligned with the discharge chutes. The cooling air further reduces the temperature of the downstream gases so as to reduce the potential for flashback in the engine. The aerodynamic form and placement of the discharge chutes provide enhancements to the operation of the diffuser, of which the support structure is a part.

Aircraft engine frame construction

A turbine assembly includes an annular fairing and strut subassembly which provides mechanical strength to the turbine assembly and support for the rotor aft bearing. A readily demountable annular outlet guide vane assembly is mounted to the aft of the fairing and strut assembly for deswirling the exhaust gasses which exit the turbine assembly. The aft mounting of the outlet guide vanes allows for an axially shorter and lighter turbine engine design which removes greater inlet swirl than prior designs, with equal or less pressure loss.

Turbine engine assembly with aft mounted outlet guide vanes

An annular acoustic liner for gas turbine engine ducts comprising Helmholtz resonator chambers wherein the resonator's turnable neck length and chamber depth lie on a surface which is concentric to the liner centerline. In another embodiment of the invention the liner is generally rectangular and the resonator's turnable neck length and chamber depth lie on a surface which is parallel to the duct flow. A more particular embodiment of the invention the duct is an afterburner duct which may have a rectangular or circular cross section.

Closed cavity noise suppressor

The present invention comprises a gas turbine engine augmentor including a new and improved fuel spraybar therein. The spraybar includes a base having a manifold for receiving fuel, and a plurality of fuel tubes extending from the base and in flow communication with the manifold. The invention further includes means for cooling the fuel tubes for allowing independent thermal movement between the cooling means and the fuel tubes. In one embodiment of the invention, the cooling means comprises an elongate shield disposed upstream of the fuel tubes to block direct impingement of combustion gases against the fuel tubes while channeling a cooling fluid such as air over the fuel tubes. In accordance with another embodiment, the invention includes means for securing the shield to the fuel tubes for allowing unrestrained longitudinal thermal movement while restraining lateral movement therebetween beyond a predetermined amount. In yet another embodiment, the invention includes means for laterally dispersing fuel from discharge ports disposed in the fuel tubes. The means may include a delta wing.

Fuel spraybar

A new and improved gas turbine engine augmentor including a liner having a film-coolable first portion joined to a convection-coolable second portion is disclosed. A cooling gas, such as fan bypass air, is channeled to the liner for film cooling the first portion and providing a residual cooling film to the second portion. The second portion is also convection coolable by air passing over an outer surface thereof. In a particular embodiment of the invention, the augmentor liner also includes a screech portion disposed at an upstream end of the first portion and downstream of a flameholder. The screech portion includes a substantially imperforate outer wall and an inner wall spaced therefrom which define a manifold therebetween. The inner wall includes a plurality of apertures in flow communication with the combustion zone. The manifold and apertures are effective for suppressing screech while substantially preventing the leakage of cooling air therethrough.

Gas turbine engine augmentor

Means are provided for controlling and modulating the temperature of the hot gas exiting the combustor associated with a gas turbine engine. The means establish a preselected temperature gradient in the hot gas so as to provide a flow of hot gas at a relatively higher temperature through gaps between adjacent turbine vanes and a flow of hot gas at a relatively lower temperature upon the vanes.

Combustion selective temperature dilution

Airflow exiting a centrifugal compressor is diffused in multiple shaped passages tangential to the compressor impeller dumped into a plenum to reduce its Mach number and then directed through deswirl vanes to the entrance of a combustor. Deswirl at reduced Mach number significantly reduces pressure losses resulting in reduced complexity and manufacturing cost for the diffuser deswirl system.

Centrifugal compressor advanced dump diffuser

A flameholder for the afterburner of an aircraft gas turbine engine is provided which comprises a pair of concentric rings forming an annular flow passage therebetween which has a continuous annular inlet for receiving the gas flow from the turbine exhaust. A plurality of discrete pilot fuel jets eject pilot fuel into the passage. A plurality of flow vanes are positioned within and around the circumference of the annular flow passage in order to impart a circumferential swirl to the fuel-air mixture and thereby improve the mixing and vaporizing thereof. An igniter of the spark type is positioned downstream of the swirl vanes. The outer annular wall of the flow passage extends downstream of and diverges conically outward from the inner wall of the flow passage. An annular V-shaped gutter is secured to the downstream end of the inner wall of the annular flow passage in order to provide a stable flameholding surface. The pre-mixed, vaporized fuel recirculates into the V-gutter to provide combustion stability. A plurality of radially extending V-shaped vanes may also be secured to the downstream end of the inner annular wall of the annular flow passage in order to further spread the flame.

Growing community of inventors

Marblehead, MA, United States of America

John W Vdoviak