United Technologies Corporation

Raytheon Technologies Corporation

Carrier Corporation

William M Proscia

Christopher D Jones

Scott A Liljenberg

Justin M Locke

Michael G McCaffrey

Timothy S Snyder

Conway Chuong

James B Hoke

Dennis M Moura

Jayant Sabnis

Jeffrey Mark Cohen

Duane C Mccormick

Gregory S Hagen

Arthur S Kesten

John E Paul

William P Patrick

William J Roy

Christopher W Robak

Theodore R Strickland

Prashant G Mehta

Andrzej Banaszuk

Caroline Karanian

Kristine Marie Carnavos

Christopher T Chipman

Joseph D Evetts

Torger J Anderson

Duane McCormick

Jagannath Pranatharthikaran

Alexander Vranos

Purity Dele-Oni

Jenny N Baarck

Daren Krayenvenger

A method of reducing noise from a combustor of a gas turbine engine includes the steps of establishing a maximum noise limit that may be for a particular frequency range. A primary fuel flow percentage, which may be emitted from a fuel nozzle arrangement having various groupings of simplex and duplex nozzles, is then established. An immersion depth measured between an aft rim of a swirler and a distal tip of the fuel nozzles may then be reduced thereby reducing the noise amplitude.

Reducing noise from a combustor of a gas turbine engine

An assembly is provided for a turbine engine. This assembly includes a swirler and a fuel nozzle. The swirler is configured with an outer wall, an inner wall, an outer passage and an inner passage. The outer wall circumscribes the inner wall and extends axially along an axis to a distal outer wall end. The inner wall extends axially along the axis to a distal inner wall end that is axially recessed within the swirler from the distal outer wall end. The outer passage is formed by and radially between the inner wall and the outer wall. The inner passage is formed by and radially within the inner wall. The fuel nozzle projects into the inner passage. The fuel nozzle is configured with a plurality of orifices axially aligned with the inner wall and arranged circumferentially about the axis.

High shear swirler with recessed fuel filmer for a gas turbine engine

An injector of a radial fuel injection system for a combustor of a gas turbine engine includes a swirler; and a fuel nozzle located within the swirler, the fuel nozzle located within the swirler, the fuel nozzle operable to provide a biased radial fuel distribution within the swirler. A method of controlling a fuel flow to a combustor of a gas turbine engine includes selectively controlling a biased radial fuel distribution within a swirler of a radial fuel injection system.

Radial fuel shifting and biasing in an axial staged combustor for a gas turbine engine

A rotational equipment assembly includes a plurality of seal shoes, a seal base, a plurality of spring elements and a secondary seal assembly. The seal shoes are arranged circumferentially around an axial centerline and include a first seal shoe. The first seal shoe includes a first seal shoe base and a first seal shoe rib that projects axially out from the first seal shoe base to an axial distal end of the first seal shoe rib. The seal base extends circumferentially around the axial centerline. The spring elements include a first spring element that connects and extends between the first seal shoe and the seal base. The secondary seal assembly is configured to seal a gap between the seal base and the seal shoes. The secondary seal assembly includes a free floating seal plate that axially contacts and is configured to slide radially along the first seal shoe rib.

Controlled contact surface for a secondary seal in a non-contact seal assembly

A combustor a gas turbine engine includes an axial fuel injection system in communication with a combustion chamber, the axial fuel injection system operable to supply a first percentage of fuel and a radial fuel injection system that communicates with the combustion chamber downstream of the axial fuel injection system, the radial fuel injection system operable to supply a second percentage of fuel, the first percentage and the second percentage of fuel scheduled in response to an engine operating parameter.

Pilot/main fuel shifting in an axial staged combustor for a gas turbine engine

An injector of a radial fuel injection system for a combustor of a gas turbine engine includes a swirler; and a fuel nozzle located within the swirler, the fuel nozzle located within the swirler, the fuel nozzle operable to provide a biased circumferential fuel distribution within the swirler. A method of controlling a fuel flow to a combustor of a gas turbine engine includes selectively controlling a biased circumferential fuel distribution within a swirler of a radial fuel injection system.

Circumferential fuel shifting and biasing in an axial staged combustor for a gas turbine engine

Pilot fuel injectors for combustors of gas turbine engines are provided. The pilot fuel injectors include a swirler having an exit into a combustion chamber of the combustor and a nozzle located within the swirler. The nozzle includes a primary fuel circuit configured to atomize fuel dispensed therethrough and a secondary fuel circuit having a first sub circuit having at least one injection aperture a first injection depth from the exit of the swirler and a second subcircuit having at least one second injection aperture located at a second injection depth from the exit of the swirler, wherein the first injection depth is greater than the second injection depth.

Pilot injector fuel shifting in an axial staged combustor for a gas turbine engine

A method of noise control from a combustor of a gas turbine engine includes selectively forming a plurality of local circumferential zones with different fuel-air ratios within the combustor.

Fuel-air mixture distribution for gas turbine engine combustors

An improved induced draft gas burner assembly for low NOresidential and light commercial gas furnaces is disclosed. The improved burner includes an upstream assembly coupled to a fuel inlet and an air inlet. The air inlet is coupled to a neck extending between the air inlet and the outlet of the neck, both of which may be disposed within an optional chamber. If used, such a chamber also includes an outlet coupled to a burner. In practice, the chamber and neck behave as a Helmholtz resonator that can be tuned to provide an upstream impedance Zthat exceeds the downstream impedance Zof the components downstream of the burner assembly.

[object Object]

A method for controlling a temperature distribution within a combustor having a plurality of chamber sections comprising controlling a fuel-to-air ratio in the chamber sections. At least two chamber sections have different fuel-to-air ratios to create a non-uniform temperature distribution within the combustor to reduce thermoacoustic instabilities.

Method for control of thermoacoustic instabilities in a combustor

A turbine engine resonator comprising a panel and a vessel. The panel is along a flow path within the engine and has an aperture. The vessel has a mounting flange secured to the panel. The flange has an aperture mated to the aperture of the panel. The vessel has a body secured to the mounting flange with an interior of the body in communication with the flange aperture. A first portion of the panel extends into the mounting flange aperture and may redirect a leakage flow, if any, to limit an effect of leakage upon resonator performance.

Turbine engine tailcone resonator

An acoustic linerincludes a remote panel, a perforated proximate paneltransversely spaced from the remote panel and one or more partitionsspanning across the space between the panels. At least one bafflecooperates with the partitions to define a resonator chamberand an associated neckfor establishing communication between the resonator chamber and a fluid stream G flowing past the proximate panel.

Architecture for an acoustic liner

A cooled acoustic liner useful in a fluid handling duct includes a resonator chamberwith a neck, a face sheet, and a coolant plenumresiding between the face sheet and the chamber. Coolant bypasses the resonator chamber, rather than flowing through it, resulting in better acoustic admittance than in liners in which coolant flows through the resonator chamber and neck. In one embodiment, the liner also includes a graze shield. Openingspenetrate both the face sheet and the shield to establish a relatively low face sheet porosity and a relatively high shield porosity. The shielded embodiment of the invention helps prevent a loss of acoustic admittance due to fluid grazing past the liner. Another embodiment that is not necessarily cooled, includes the resonator chamber, low porosity face sheet and high porosity shield, but no coolant plenum for bypassing coolant around the resonator chamber. An associated method of retrofitting an acoustic treatment into a fluid handling module includes installing openings in the module and mounting a resonator boxon the module so that the inlets to the resonator necks register with the installed openings.

High admittance acoustic liner

An acoustic linerincludes a remote panel, a proximate paneltransversely spaced from the remote panel and a resonator chamberresiding between the panels. Perforationspenetrate the proximate panel in registration with the resonator chamber. A neckwith an inletrecessed from the proximate panel establishes communication between the chamber and a fluid stream G flowing past the proximate panel. A bypass coolant passageguides coolant through the perforations without guiding it through the resonator chamber.

Growing community of inventors

Marlborough, CT, United States of America

William M Proscia