Rolls-royce Corporation

Rolls-royce North American Technologies Inc.

Allison Advanced Development Company

Jack D Petty, Sr

Kevin Sauer

John D Holdcraft

Kenneth M Pesyna

James Carl Loebig

Michel Shawn Smallwood

Colin Faucett

Edward Claude Rice

Andrew J Eifert

Charles Bryce Graves

Todd S Taylor

Tab Michael Heffernan

Duane Smith

Mohan Razdan

Michael S Bell

Oran Watts

Andrew James Eifert

A system may be provided that includes a hot section component of a gas turbine engine. The hot section component includes a dual wall, which includes a first wall and a second wall. The first wall includes multiple impingement cooling holes extending through the first wall. The second wall is positioned adjacent the first wall. The first wall and the second wall together define a cooling passage between the first wall and the second wall. Multiple pockets are in a surface of the second wall. Each of the pockets is positioned opposite a respective one of the impingement cooling holes. Each of the pockets is configured to receive a cooling fluid from the respective one of the impingement cooling holes and direct the cooling fluid into the cooling passage. The cooling passage includes a single cooling passage into which the pockets are configured to direct the cooling fluid.

Impingement cooling dust pocket

A combustor liner assembly is provided including an inlet wall, an inner wall, an outer wall, and a plurality of fasteners. The inlet wall has an opening into a combustion chamber. The inner wall is coupled to the inlet wall at a first end of the combustor liner. The inner wall defines a radially inner end of the combustion chamber. The outer wall includes a plurality of segments. The plurality of segments define a radially outer end of the combustion chamber. A first portion of the plurality of fasteners couples each of the plurality of segments to another of the plurality of segments. A second portion of the plurality of fasteners couples at least one of the plurality of segments to the inlet wall at the first end of the combustor liner.

Combustor bolted segmented architecture

Devices, systems, and methods of a casing for a heat suppression system of a gas turbine engine exhaust include a heat shield and an insulation layer for arrangement between the heat shield and an outer skin.

Heat shield for signature suppression system

A fuel injection system may include a fuel spray nozzle including a nozzle head and a nozzle stem. The nozzle head may include an air channel and a swirler in fluid communication with the air channel. The nozzle stem may extend into a compressor discharge pressure cavity and convey fuel to the air channel. The air channel may combine air from the swirler with the fuel from the nozzle stem and convey a mixture of fuel and air to a combustor. The fuel injection system may further include a scoop. The scoop may be coupled to the fuel spray nozzle. The scoop may receive air that flows into the compressor discharge pressure cavity from a diffusor. An outer surface of the fuel spray nozzle and an inner surface of the scoop define a duct in fluid communication with the swirler. The swirler may receive air from the duct.

Ram pressure recovery fuel nozzle with a scoop

Devices, systems, and methods of a casing for a heat suppression system of a gas turbine engine exhaust include a floating heat shield.

A gas turbine engine includes a combustor having a dual-wall impingement convention effusion combustor tile assembly. The dual-wall tile assembly provides a cooling air flow channel and attachments for securing the tile to the cold skin liner of the combustor. Cooling is more efficient in part due to the dual wall construction and in part due to reduced parasitic leakage, and the design is less sensitive to attachment features which operate at lower temperatures.

Dual-wall impingement, convection, effusion combustor tile

A gas turbine engine includes a combustor having a combustor tile assembly with improved cooling air flow channels and enhanced cooling efficiency. A method of manufacturing same is provided which increases production capabilities and the geometric configurations of the exit ports which in turn improve the hot side operating temperature of the tiles in the combustion chamber.

Combustor tile

A gas turbine health monitoring fuel system including a fuel splitter, a first and second fuel nozzle, a first and second fuel pump, and a common drive coupled to the first and second fuel pumps. The first and second fuel pumps are downstream from the fuel splitter. The first fuel nozzle is fluidly coupled to the first fuel pump by a first fuel stream. The second fuel nozzle is fluidly coupled to the second fuel pump by a second fuel stream. The common drive is configured to drive the first and second fuel pumps at a same speed. The controller system is configured to determine a first fuel pressure in the first fuel stream, determine a second fuel pressure in the second fuel stream, and identify a difference between the first and second fuel pressures.

Fuel flow splitter and gas turbine fuel system health monitoring

An improved fastening system and method overcomes traditional thermal mismatch that occurs with standard fastener systems that are employed in extreme temperature environments. The fastening system could employ a metallic fastener that is combined with metallic spacer, each of which having a differing thermal growth coefficient. This arrangement may reduce thermal expansion by several orders of magnitude over the range of temperatures where the thermal fastener can be applied. Such an improved fastening system may be employed in gas turbine machines and in other locations where it is desirable to improve fastening of objects in extreme temperature environments.

Bi-metal fastener for thermal growth compensation

The present invention contemplates an apparatus for supporting an annular combustor liner in a gas turbine engine. In one embodiment, the apparatus includes a plurality of radial lugs extending outwardly from an outer surface of the combustor liner and being spaced circumferentially and uniformly about a periphery of the combustor liner. A band is disposed about the outer surface of the combustor liner and coupled to the radial lugs. The radial lugs each include a radially extending slot having opposing circumferentially facing radial surfaces. The band includes a plurality of pins, each pin being disposed within a respective one of the slots. The pin is configured to bear against at least one of the radial surfaces to provide circumferential support to the combustor liner by substantially preventing relative circumferential movement between the combustor liner and the band. The pin is slidable between the radial surfaces to allow relative radial movement between the combustor liner and the band.

Growing community of inventors

Indianapolis, IN, United States of America

Jack D Petty, Sr