United Technologies Corporation

Raytheon Technologies Corporation

Carrier Corporation

Ramons Andris Reba

Wesley K Lord

Jonathan Gilson

Bruce L Morin

David A Topol

Oliver V Atassi

Duane C Mccormick

David M Rockwell

Stephen A Morford

Satish Narayanan

Andre M Hall

Mark W Wilson

Daniel L Gysling

JinQuan Xu

Jose R Paulino

Brian Ernest Wake

Carson A Roy Thill

Gregory A Kohlenberg

Marios C Soteriou

Arthur Blanc

Thierry Pascal Maeder

Robert Hans Schlinker

Benjamin J Blechman

John C Simonich

Burce L Morin

Jonthan Gilson

Thierry Maeder

A compressor has an inlet port and a discharge port. The discharge port communicates into a resonator chamber. The resonator chamber includes a first stage resonator array and a second stage resonator array downstream of the first stage resonator array with a connecting passage intermediate the first and second resonator array. Each of the resonator arrays includes a pair of spaced resonator arrays sub-portions, with each of the sub-portions including a plurality of cells extending into a housing member, and have a bottom wall and an open outer wall communicating with the flow passage, with a plurality of orifices extending into each of the cells. The orifices have a smaller diameter than a hydraulic diameter of the cells.

Multi-stage resonator for compressor

A fan section for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan rotor having fan blades, and a plurality of fan exit guide vanes positioned downstream of the fan rotor. The fan rotor is configured to be driven through a gear reduction. A ratio of a number of fan exit guide vanes to a number of fan blades is defined. The fan exit guide vanes are provided with optimized sweep and optimized lean.

Geared gas turbine engine with reduced fan noise

A section for a gas turbine engine includes a rotating structure, a stationary structure, and a flow guide assembly arranged generally between the rotating structure and the stationary structure. A flow path is defined between the flow guide assembly and one of the rotating structure and the stationary structure. The flow guide assembly includes a plurality of apertures configured to disrupt acoustic waves of air in the flow path. A seal is configured to establish a sealing relationship between the rotating structure and the stationary structure, and wherein an inlet to the flow path is adjacent the seal. A gas turbine engine and a method of disrupting acoustic waves in a flow path of a gas turbine engine are also disclosed.

Gas turbine engine component for acoustic attenuation

Acoustic treatments for components of gas turbine engines are described. The acoustic treatments include an acoustic resonator having a backing chamber defining a respective volume and a neck arranged relative to the backing chamber and defining an opening, wherein the neck has a length and a cross-sectional area. The acoustic resonator cell satisfies the following relationships: (1) l/L=0.2-0.8, where l is a length of the neck and L is a depth of the backing chamber and (2) a/A=0.02-0.20, where a is a cross-sectional area of the neck and A is a cross-sectional area of the backing chamber.

Airfoil-mounted resonator

A compressor has a housing assembly having a plurality of ports including a suction port and a discharge A male rotor is mounted for rotation about an axis. A female rotor is enmeshed with the male rotor and mounted in the housing for rotation about an axis for drawing a flow from the suction port, compressing the flow, and discharging the compressed flow through the discharge port. A cavity group is between the discharge port and the male rotor and female rotor. The cavity group has a first member separating a plurality of cells and a foraminate cover member atop the first member.

Screw compressor resonator arrays

A nacelle assembly according to an exemplary aspect of the present disclosure includes, among other things, a core nacelle defined at least partially about a core engine, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path, and a variable area fan nozzle in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section downstream of the first fan nacelle section. The first fan nacelle section and the second fan nacelle section are axially movable relative to one another to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The auxiliary port is defined between the first fan nacelle section and the second fan nacelle section. The first fan nacelle section comprises a first acoustic system which provides an acoustic impedance configured to attenuate a noise characterized by a leading edge of the second fan nacelle section. The first acoustic system is defined at least in part within a trailing edge region of the first fan nacelle section. A method of reducing a total effective perceived noise level of a gas turbine engine with a variable area fan nozzle is also disclosed.

Variable area fan nozzle for gas turbine engine

A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

Variable area fan nozzle with acoustic system for a gas turbine engine

A noise control cassette for a gas turbine engine includes a perforated face sheet configured for exposure to an airflow, a non-perforated backing sheet, a core arranged between the face sheet and the backing sheet and defining a cavity between the face sheet and the backing sheet having an effective length tuned so as to provide acoustic reactance control, and an attachment face for attaching the cassette to an airfoil-shaped structure.

Airfoil acoustic impedance control

A nozzle for a gas turbine engine is provided with undulations on a core and/or fan exhaust flow nozzle. An annular wall defines a fluid flow passage and includes a base portion and an adjoining exit portion. The base portion is typically generally frustoconical in shape and includes an arcuate contour in an axial direction. The exit portion includes undulations in a generally radial direction that provide lobes and troughs each respectively including trailing edges. One of the lobe and trough trailing edges are recessed from the other of the lobe and trough trailing edges in a generally axial direction. The other of the lobe and trough trailing edges form apexes with the apexes provided on tabs. The troughs extend radially inward in the axial direction towards the trough trailing edges.

Serrated nozzle trailing edge for exhaust noise suppression

A system for reducing jet noise emission from an internally mixed gas turbine engine exhaust, comprising a fan/core flow mixer having a plurality of mixer lobes and a common flow nozzle having an equal number of tabs located along a circumferential edge of an aft end of the nozzle. There is a predetermined clocking relationship between the plurality of mixer lobes and the plurality of nozzle tabs that results in reduced exhaust noise emission, most evident in the lower frequency range. A method for reducing jet noise emission from an internally mixed gas turbine engine exhaust comprises selectively aligning a circumferential distribution of a mixed flow vorticity field produced by a fan/core mixer with a circumferentially distributed exhaust flow vorticity field produced by a modified common flow nozzle at an exit plane of the engine exhaust.

Jet exhaust noise reduction system and method

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South Windsor, CT, United States of America

Ramons Andris Reba