Honeywell International Inc.

Vericor Power Systems LLC

Ian L Critchley

David G Walhood

Nagaraja S Rudrapatna

Gregory O Woodcock

Bruce Anson

Jurgen C Schumacher

Philip Smalley

Jeffrey Spencer

Cristopher Frost

Frank Joseph Zupanc

Paul Yankowich

Stony Kujala

Robert B Sumegi

Michael W Scott

Matthew C Caspermeyer

Barry J Kingery

Michael T Todd Barton

Samir Rida

Kenneth R Tyrer

Mike W Scott

Lorenzo Hernandez

Stony W Kujala

A liner for a combustor includes a first wall and a second wall extending around at least a portion of the first wall to form a liner cavity with the first wall. The first wall defines a first wall orifice and the second wall defines a second wall orifice. The liner further includes a first insert mounted on the second wall within the second wall orifice and extending through the first wall orifice. The first insert is configured to direct a first air jet through the second wall and the first wall. The first insert is formed by a tubular body portion extending through the first wall orifice and the second wall orifice, and a shoulder extending around the body portion that abuts the first side of the second wall. The shoulder has a first diameter and the first wall orifice has a second diameter, greater than the first diameter.

Liner for a combustor with strain isolated inserts

A combustor for a turbine engine includes a first liner and a second liner forming a combustion chamber with the first liner. The combustion chamber is configured to receive an air-fuel mixture for combustion therein. The first liner is a first double wall liner with a first wall forming a portion of the combustion chamber and a second wall extending around at least a portion of the first wall to form a liner cavity with the first wall. The first wall defines a first wall orifice and the second wall defines a second wall orifice. The combustor further includes a first insert mounted on the second wall within the second wall orifice and extending through the first wall orifice. The first insert is configured to direct a first air jet through the second wall, through the first wall, and into the combustion chamber.

Double wall combustors with strain isolated inserts

An engine assembly includes a compressor section configured to supply compressed air and a combustion section coupled to the compressor section. The compressor section includes a diffuser with diffuser vanes to condition the compressed air, and the diffuser vanes have a vane count of a first value. The combustion section includes a combustion chamber to receive the compressed air and fuel injection assemblies configured to introduce fuel into the combustion chamber. The fuel injection assemblies have a fuel injection assembly count of a second value, and the first value is a whole number multiple of the second value.

Engine assemblies and methods with diffuser vane count and fuel injection assembly count relationships

An engine assembly includes a combustor having a combustion chamber in which an air and fuel mixture is combusted to produce combustion gases. The engine assembly further includes a transition scroll coupled to the combustor for receiving the combustion gases. The transition scroll includes an interior surface, an exterior surface, and effusion cooling holes for providing cooling air to the interior surface. The engine assembly further includes a turbine coupled to the transition scroll for receiving and extracting energy from the combustion gases.

Transition scrolls for use in turbine engine assemblies

An engine assembly includes a scroll assembly including an inlet, an outlet, an inner section, and an outer section. Each of the inner section and the outer section extend between the inlet and the outlet with the outer section being longer than the inner section. The engine assembly further includes a combustor having an entrance for receiving fuel to be mixed with air and combusted in the combustor, an exit coupled to the inlet of the scroll assembly for delivering the combusted air-fuel mixture, including a hot stream, thereto, and a wall coupled between the entrance and the exit and including a plurality of dilution holes therethrough for deflecting the hot stream toward the outer section of the scroll assembly.

Combustors for use in turbine engine assemblies

A combustion system includes a combustor having a forward annular liner having a first plurality of effusion holes, and an aft annular liner having a second plurality of effusion holes and forming a combustion chamber with the forward annular liner. The first plurality of effusion holes and the second plurality of effusion holes are adapted to receive compressed air from a compressor and contribute to a single toroidal recirculation air flow pattern in the combustion chamber. The combustion system further includes a rotary fuel slinger further adapted to receive a flow of fuel from a fuel source and to centrifuge the received fuel into the combustion chamber; and an igniter extending at least partially into the combustion chamber to ignite the fuel and compressed air in the combustion chamber, to thereby generate combusted gas.

Combustors and combustion systems for gas turbine engines

A combustion system includes a fuel manifold adapted to receive a flow of fuel from a fuel source, the fuel manifold including at least one orifice from which the received fuel is discharged. The combustion system further includes a rotary fuel slinger disposed adjacent to the fuel manifold. The rotary fuel slinger includes a coupler shaft and a slinger disc coupled to the coupler shaft. The slinger disc includes a first portion disposed generally perpendicular to the coupler shaft and a second portion disposed at an angle to the first portion. The fuel manifold is positioned relative to the rotary fuel slinger such that the discharged fuel impinges on the second portion. The rotary fuel slinger is configured to rotate such that the flow of fuel is centrifuged radially outwardly from the second portion onto the first portion and subsequently off the rotary fuel slinger to atomize the received fuel.

Combustion systems with rotary fuel slingers

The present invention provides an adaptive combustion controller and method for a turbine engine. The adaptive combustion controller and method modulates the fuel flow to the turbine engine combustor to reduce combustion instabilities. In particular, the adaptive combustion controller includes a fuel flow phase controller and a fuel flow magnitude controller. The adaptive combustion controller receives sensor data from the turbine engine. In response to the sensor data the fuel flow phase controller adjusts the phase of the modulated fuel flow to reduce instabilities in the combustor. Likewise, in response to the sensor data the fuel flow magnitude controller adjusts the magnitude of the modulated fuel flow to further reduce the instabilities in the combustor. By modulating the fuel flow to the combustor, and adaptively adjusting the phase and magnitude of the modulated fuel flow, the adaptive combustion controller is able to effectively reduce combustion instabilities.

System and method for turbine engine adaptive control for mitigation of instabilities

A variable penetration dilution jet array for a single can and scroll assembly includes a plurality of differentially sized dilution openings around the circumference of the combustor can. The alternating smaller and larger openings provide for circumferential and radial mixing uniformity with the smaller openings giving shallow penetration and the larger openings enabling deep core penetration. The larger openings provide dilution air to the hot gas flow core without the need for an increase in combustor pressure drop. The smaller openings provide a film cooling flow to the downstream scroll, reducing dedicated scroll cooling requirements.

Temperature variance reduction using variable penetration dilution jets

A gas turbine engine integrates the functions of an auxiliary power unit (APU) with one or more functions of an ECS, and that is capable of operating in both an unfired mode and a fired mode. The gas turbine engine includes a combustor system that is configured to allow the gas turbine to quickly transition from the unfired mode to the fired mode, by bypassing a portion of the air flowing to the combustor system around the combustor.

Dual mode power unit having a combustor bypass system

A system and method of combusting a hydrocarbon fuel is disclosed. The system combines the accuracy and controllability of an air staging system with the ultra-low emissions achieved by catalytic combustion systems without the need for a pre-heater. The result is a system and method that is mechanically simple and offers ultra-low emissions over a wide range of power levels, fuel properties and ambient operating conditions.

Hybrid lean premixing catalytic combustion system for gas turbines

A control and method for arresting flashback in a gas turbine engine combustion system having a venturi for delivering a mixture of air and fuel to a combustion chamber and a valve for controlling the amount of air flowing to the venturi. The control and method includes a flashback detection routine, a valve position trim routine, and a reference flame temperature adjustment routine. The flashback detection routine detects a flashback in the venturi by comparing the temperature at the venturi to the temperature of the air entering the valve. The valve position trim routine receives this flashback signal and in response causes the valve to open until the flashback is arrested. Finally, to prevent the flashback from recurring, the reference flame temperature adjustment routine adjusts a reference flame temperature in the control system so as to prevent the valve from returning to the position at which the flashback occurred.

Flashback control for a gas turbine engine combustor having an air bypass system

The present invention provides a premix fuel injector for use in gas turbine engines and in combustion systems having controllable pressure drops. The premix fuel injector comprises a premix chamber having an inlet for receiving a flow of pressurized air and having an exit. A venturi is coupled to the exit of the premix chamber and an inlet of a combustion chamber. Gaseous fuel is flowed into the premix chamber by a plurality of circumferentially disposed tubes extending into the premix chamber with each of said tubes having at least one hole for flowing a stream of the gaseous fuel.

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Phoenix, AZ, United States of America

Ian L Critchley