Abstract: An igniter arranged to ignite combustion in a primary flow including a fuel and air mixture, the igniter including one or more geometric features arranged to: induce a shockwave flow structure at least partially disposed in the primary flow; and ignite the fuel and air mixture by virtue of the shockwave flow structure. The present disclosure also relates to a method of igniting combustion in a primary flow including a fuel and air mixture, the method including: providing one or more geometric features arranged to induce a shockwave flow structure; inducing the shockwave flow structure at least partially in the primary flow; and igniting the fuel and air mixture by virtue of the shockwave flow structure.

Abstract: A fuel injection system for a gas turbine engine includes a fuel nozzle with a fuel injection aperture to inject a fuel jet and a multiple of airflow passages in the fuel nozzle to communicate a multiple of air streams to interact with the fuel jet.

Abstract: A system including a multi-tube fuel nozzle, including a plurality of tubes extending in an axial direction relative to a central axis of the multi-tube fuel nozzle, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet; and an inlet flow conditioner, including a plate extending in a radial direction relative to the central axis of the multi-tube fuel nozzle; an outer wall extending circumferentially about the plate, wherein the outer wall is coupled to the plate; and a plurality of air openings in the plate, the outer wall, or a combination thereof, wherein the plurality of air openings are disposed upstream from the air inlets in the plurality of tubes.

Abstract: A combustor (10) includes a cap (16), a liner (20), a transition piece (24), and a combustion chamber (22) located downstream from the cap (16) and defined by the cap and liner. A secondary nozzle (40) circumferentially arranged around the liner (20) or transition piece (24) includes a center body, a fluid passage through the center body, a shroud circumferentially surrounding the center body, and an annular passage between the center body and the shroud. A method for supplying fuel to a combustor (10) includes flowing fuel through a primary nozzle radially disposed in a breech end of the combustor and flowing fuel through a secondary nozzle (40) circumferentially arranged around and passing through at least one of a liner (20) or a transition piece. The secondary nozzle (40) includes a center body, a fluid passage through the center body, a shroud circumferentially surrounding at least a portion of the center body (44), and an annular passage between the center body and the shroud.

Abstract: A system including a multi-tube fuel nozzle, including a first plate having a first plurality of openings, a plurality of tubes extending through the first plurality of openings in the first plate, wherein each tube of the plurality of tubes includes an air inlet, a fuel inlet, and a fuel-air mixture outlet, and a resilient metallic seal disposed along the first plate about the plurality of tubes.

Abstract: Hybrid internal detonation-gas turbine engines incorporating detonation or pulse engine technology (such as an internal detonation engine), and methods of manufacturing and using the same are disclosed. The internal detonation engine includes a detonation chamber having a fuel igniter therein, a stator at one end of the detonation chamber having at least a first opening to receive fuel, a rotor adjacent to the stator, and an energy transfer mechanism configured to convert energy from igniting or detonating the fuel to mechanical energy. The detonation chamber and fuel igniter are configured to ignite or detonate a fuel in the detonation chamber. Either the stator or the detonation chamber has a second opening to exhaust detonation gas(es). The rotor has one or more third openings therein configured to overlap with at least the first opening as the rotor rotates.

Abstract: A system including a first multi-tube fuel nozzle including a plurality of first tubes extending in an axial direction, wherein each first tube of the plurality of first tubes includes a first air inlet, a first fuel inlet, and a first fuel-air mixture outlet, a second multi-tube fuel nozzle including a plurality of second tubes extending in an axial direction, wherein each second tube of the plurality of second tubes includes a second air inlet, a second fuel inlet, and a second fuel-air mixture outlet, and an aft plate including a plurality of first tube apertures and a plurality of second tube apertures, wherein the plurality of first tubes extend to the plurality of first tube apertures, and the plurality of second tubes extend to the plurality of second tube apertures.

Abstract: A gas engine assembly includes a compressor, a combustion system, a bypass line and a control system. The control system is configured to control gas supply parameters based on a transportation delay value. The transportation delay value corresponds to a delay between a time when a gas supply control mechanism is adjusted and a time that gas having a corresponding adjustment of a gas characteristic is received at a predetermined point downstream from the gas supply control mechanism.

Abstract: There is provided a method for improving the combustion efficiency of a combustor of a gas turbine engine powering an aircraft. The method comprises selectively using two distinct fuel injection units or a combination thereof for spraying fuel in a combustion chamber of the combustor of the gas turbine engine. A first one of the two distinct fuel injection units is selected and optimized for high power demands, whereas a second one of the two distinct fuel injection units is selected and optimized for low power level demands. In operation, the fuel flow ratio between the two distinct injection units is controlled as a function of the power level demand.

Abstract: An ignition/chemical reaction promotion/flame holding device and a high-performance speed-type internal combustion engine using this device are provided, whereby ignition and the spreading and holding of flames can be dramatically improved in a gas turbine, a ram machine, a rocket engine, or another speed-type internal combustion engine. An ignition/chemical reaction promotion/flame holding device of a speed-type internal combustion engine comprises a spark plug for preparing charged particles in a predetermined location in a combustor of the speed-type internal combustion engine, and a microwave oscillator and antenna for inducing plasma with a working fluid in the combustor as a starter material by irradiating the charged particles and their surrounding vicinity with microwave pulses; wherein a region in which sufficient conditions for performing combustion are met is formed in the combustor by supplying an active chemical species produced from the working fluid by the effect of the plasma.

Abstract: A turbine engine for an aircraft including a turbine engine shaft and a pumping module, including: a pump shaft, connected to the turbine engine shaft; a pump for supplying fuel to the turbine engine, mounted on the pump shaft, configured to deliver a flow of fuel as a function of a speed of rotation of the turbine engine shaft; and an electrical device mounted on the pump shaft and configured, according to a first mode of operation, to drive the pump shaft in rotation to actuate the supply pump and, according to a second mode of operation, to be driven in rotation by the pump shaft to supply electrical power to equipment of the turbine engine.

Abstract: A preparation of liquid fuel for combustion, particularly for ignition, in the burners of a gas turbine is proposed. A source of heated liquid fuel, particularly heated pressurized fuel, is arranged in a reservoir. The reservoir is arranged in parallel with a main fuel feed of a combustion system. The reservoir can be filled by the main fuel feed and/or evacuated into the main fuel feed. The reservoir contains a heating means, which heat the liquid fuel filled in the reservoir, particularly while circulating in the reservoir.

Abstract: A method and a device for controlling a lean fuel intake gas turbine engine, which can stably maintain operation of the gas turbine engine by preventing misfire and burnout of a catalytic combustor even when a catalyst in the combustor is deteriorated. A difference between measured temperatures of an inlet and an outlet of the combustor with respect to a methane concentration in an intake gas that is to be taken into the lean fuel intake gas turbine engine is compared with reference temperature difference data that is data indicating difference between temperatures of the inlet and the outlet of the combustor including a catalyst in its initial state to be a reference, and at least one of the inlet temperature and the outlet temperature of the combustor is controlled based on a difference between the measured temperature difference and the reference temperature difference data.

Abstract: A fuel system for an aircraft comprises a boost pump, a main fuel pump and a motive pump. The boost pump receives fuel from a storage unit. The main fuel pump receives fuel from the boost pump and delivers fuel to a distribution system. The motive fuel pump receives fuel from the boost pump, routes fuel through the storage unit, and delivers fuel to an actuator.

Abstract: A fuel nozzle system for enabling a gas turbine to start and operate on low-Btu fuel includes a primary tip having primary fuel orifices and a primary fuel passage in fluid communication with the primary fuel orifices, and a fuel circuit capable of controlling flow rates of a first and second low-Btu fuel gases flowing into the fuel nozzle. The system is capable of operating at an ignition status, in which at least the first low-Btu fuel gas is fed to the primary fuel orifices and ignited to start the gas turbine, and a baseload status, in which at least the second low-Btu fuel gas is fired at baseload. The low-Btu fuel gas ignited at the ignition status has a content of the first low-Btu fuel gas higher than that of the low-Btu fuel gas fired at the baseload status. Methods for using the system are also provided.

Abstract: A gas turbine engine combustion system including a mixing duct that separates into at least two branch passages for the delivery of a fuel and working fluid to distinct locations within a combustion chamber. The residence time for the fuel and working fluid within each of the two branch passages is distinct.

Abstract: A combustion chamber may include a first surface and a second surface interconnected by a wall forming a chamber having a central axis. The first surface may define an exhaust opening and the second surface defining a pilot opening, wherein the exhaust opening and the pilot opening align along the central axis. A plurality of inlet ports may be configured to deliver air to the chamber. A plurality of fuel ports may be arranged on an inside of the second surface to deliver fuel to the chamber. The air flow from the inlet ports and fuel from the fuel ports may oppose each other to create a vortex of product proximal to the second surface.

Abstract: A fuel system for a gas turbine engine includes an engine fuel manifold, a hydraulic actuator, and a drain piston assembly. The hydraulic actuator actuates in response to a change in pressures within the hydraulic actuator. The drain piston assembly is fluidically connected to both the hydraulic actuator and the engine fuel manifold. The drain piston assembly receives fuel from the engine fuel manifold and sends fuel to the hydraulic actuator during engine shut down.

Abstract: A system includes a multi-tube fuel nozzle of a turbine combustor. The multi-tube fuel nozzle includes a support structure defining an interior volume configured to receive an air flow; a plurality of mixing tubes disposed within the interior volume, wherein each of the plurality of mixing tubes comprises a respective fuel injector; and an outer annular wall configured to direct an air flow from an annulus between a liner and a flow sleeve of the turbine combustor at least partially radially inward into the interior volume through an air inlet and toward the plurality of mixing tubes, wherein the outer annular wall at least partially defines an air flow passage extending from the annulus to the interior volume.

Abstract: A slinger combustor has an annular combustor shell defining a combustion chamber having a radially inner fuel inlet for receiving a spray of fuel centrifuged by a fuel slinger. The combustion chamber has a fuel atomization zone extending radially outwardly from the fuel inlet and merging into a radially outwardly flaring expansion zone leading to a combustion zone. A plurality of nozzle air inlets are defined in the fuel atomization zone of the combustor shell. The nozzle air inlets have a nozzle axis intersecting the stream of fuel and a tangential component in a direction of rotation of the fuel slinger. A plurality of dilution holes are defined in the combustor shell and have a dilution axis intersecting the combustion zone. The dilution axis of at least some of the dilution holes has a tangential component opposite to the direction of rotation of the fuel slinger.

Abstract: A method for use in a gas turbine engine. The method includes the steps of: configuring a downstream injection system within the interior flowpath that includes two injection stages, a first stage and a second stage, wherein the first stage and the second stage are each axially spaced from the other; and circumferentially positioning the injectors of the first stage and the second stage based on: a) a characteristic of an anticipated combustion flow occurring just upstream of the first stage during a mode of operation; and b) the characteristic of an anticipated combustion flow just downstream of the second stage given an anticipated effect of the air and fuel injection from the first stage and the second stage.

Abstract: A method including providing a wet algae material having been subjected to a refinement process without a water separation phase; supplying the wet algae material including a water fraction and an algae-grown biofuel to a turbine engine; and operating the turbine engine with the wet algae material where a retained portion of the water fraction is retained in the wet biofuel during a manufacturing process not including a dehydration step and includes an amount sufficient to reduce generation of a quantity of nitrogen oxides, and where the turbine engine can further include a combustor capable of combusting the wet biofuel.

Abstract: The present application and the resultant patent provide a diffusion combustor fuel nozzle for a gas turbine engine. The fuel nozzle may include one or more gas fuel passages for one or more flows of gas fuel, a swirler surrounding the one or more gas fuel passages and positioned about a downstream face of the fuel nozzle, a number of swirler gas fuel ports defined in the swirler, and a number of downstream face gas fuel ports defined in the downstream face of the fuel nozzle. The swirler may include a number of swirl vanes and a number of air chambers defined between adjacent swirl vanes. The present application and the resultant patent further provide a method of operating a diffusion combustor fuel nozzle of a gas turbine engine.

Abstract: A combustor comprises an annular combustor chamber formed between the inner and outer liners, the annular combustor chamber having a central axis. Fuel nozzles are in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber. The fuel nozzles are oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber. Nozzle air inlets are in fluid communication with the annular combustor chamber to inject nozzle air generally radially in the annular combustor chamber. A plurality of dilution air holes are defined through the inner and outer liner downstream of the nozzle air inlets, the dilution holes configured for high pressure air to be injected from an exterior of the liners through the dilution air holes generally radially into the combustor chamber, a central axis of the dilution air holes having a tangential component relative to the central axis of the annular combustor chamber.

Abstract: Embodiments of the present disclosure are directed to systems and methods for premixing fuel and air prior to combustion within a combustion chamber. The system includes a plurality of fuel injectors and a plurality of mixing tubes, wherein each mixing tube has a first portion for receiving one of the plurality of fuel injectors and a second portion having a mixing chamber that is configured to mix fuel and air. The length of the mixing chamber varies among the plurality of mixing tubes to allow for different mixing times.

Abstract: A combustor comprises an annular combustor chamber formed between the inner and outer liners. Fuel nozzles each have an end in fluid communication with the annular combustor chamber to inject fuel in the annular combustor chamber, the fuel nozzles oriented to inject fuel in a fuel flow direction having an axial component relative to the central axis of the annular combustor chamber. A plurality of nozzle air holes are defined through the inner liner and the outer liner adjacent to and downstream of the fuel nozzles. The nozzle air holes are configured for high pressure air to be injected from an exterior of the liners through the nozzle air holes generally radially into the annular combustor chamber. A central axis of the nozzle air holes has a tangential component relative to the central axis of the annular combustor chamber.

Abstract: A fuel injector for a gas turbine engine may include a flow path for a fuel-air mixture extending longitudinally through the fuel injector, and a fuel gallery extending circumferentially around the flow path. The fuel gallery may be adapted to inject a liquid fuel into the flow path. The fuel injector may also include an annular casing positioned circumferentially around the fuel gallery to define an insulating chamber around the gallery. The fuel injector may also include an annular cover extending around the fuel injector to define a metering chamber. The fuel injector may further include one or more purge holes fluidly coupling the metering chamber to the insulating chamber, and one or more metering holes fluidly coupling the metering chamber to a volume exterior to the fuel injector.

Abstract: Systems and methods for controlling the composition of a combustion exhaust gas are provided.

Abstract: Systems, devices, and methods for controlling a fuel supply for a turbine or other engine using direct and/or indirect indications of power output and optionally one or more secondary control parameters.

Abstract: A method for purging fuel from a fuel system of a gas turbine engine on shutdown of the engine comprises, in one aspect, terminating a fuel supply to the fuel system and using the residual compressed air to create a reversed pressure differential in the fuel system relative to a forward pressure differential of the fuel system used to maintain fuel supply for engine operation, and under the reversed pressure differential substantially purging the fuel remaining in the system therefrom to a fuel source.

Abstract: A fuel nozzle includes an inner wall defining a central passage extending in an axial direction of the fuel nozzle, a hub wall surrounding the inner wall and defining a first annular passage, an outer wall surrounding the hub wall and defining a second annular passage, and a shroud surrounding the outer wall and defining a third annular passage. A swirler may receive air and direct the air into the first annular passage. The swirler includes at least one swirl vane extending from the shroud to the hub wall that has an air passage extending between the shroud and the hub wall. The air passage is coupled to the first annular passage and has a first width adjacent the shroud and a second width adjacent the hub wall. The second width is larger than the first width defining a diverging outlet into the first annular passage.

Abstract: One embodiment of the present disclosure is a unique method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft. Another embodiment of the present disclosure is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A protocol for assembling a fuel shut off pin valve assembly for a gas turbine engine. The protocol outlines a specific sequence of events in order to ensure failsafe incorporation of a fuel shut off valve assembly within the low pressure turbine area and specifically within the engine casing of the gas turbine.

Abstract: A valve control device is provided in a gas turbine having a combustor for generating combustion gas, a turbine driven by the combustion gas generated by the combustor, a flow rate regulating valve for regulating the flow rate of the fuel to be supplied to the combustor, and a pressure regulating valve disposed upstream of the flow rate regulating valve, for regulating the fuel pressure. The valve control device controls the opening degree of the valve. The valve control device includes a load decrease detection part which detects a load decrease of the gas turbine, and a pressure control part which controls the opening degree of the valve in accordance with the output of the gas turbine. The valve control device suppresses instability of the gas turbine output even when the load rapidly decreases.

Abstract: A method for preventing flashback in a burner having swirl generator with a central fuel distributor element and an outer wall enclosing the central fuel distributor element and bounding an axial flow channel for combustion air is provided. A separating wall encloses the central fuel distributor element radially within the outer wall, wherein the axial flow channel is divided into a radially inner channel segment and a radially outer segment by the separating wall. A tangential flow component in the radially outer channel segment is provided to the combustion air flowing through the axial flow channel, wherein, during combustion, the combustion air passes the radially inner channel segment without a tangential flow component.

Abstract: A fuel nozzle for a gas turbomachine includes an outer nozzle body including an inner surface defining a mixing zone, and an inner nozzle body arranged within the outer nozzle body. The inner nozzle body includes a fluid passage. At least one flow affector extends from the inner nozzle body to the outer nozzle body. The at least one flow affector includes an inner surface that defines an interior chamber having an inlet fluidly connected to the fluid passage and at least two openings fluidically linking the interior chamber and the mixing zone. One or more flow tuning elements are arranged at the interior chamber upstream of the at least two openings. The one or more flow affectors are configured and disposed to condition a fluid passing into the interior chamber to affect a substantially iso-kinetic distribution of the fluid within the interior chamber.

Abstract: A burner arrangement is disclosed with a conical swirler in the form of a double cone which is arranged concentric to a burner axis and which encloses a swirl chamber, and with a central fuel lance which lies in the burner axis and projects from the cone point of the swirler into the swirl chamber, wherein a first stage is provided for injecting premix fuel, in which the premix fuel is injected radially outwards into the swirl chamber through injection openings which are arranged on the fuel lance, and wherein a second stage is provided for injecting premix fuel, in which the premix fuel is injected into an air flow, which is guided in the double cone, through injection openings in the double cone. With such a burner arrangement, the gas pressure which is required in the first stage is reduced by the entire premix fuel being injected into the swirl chamber in the first stage through two oppositely-disposed injection openings with increased opening diameter.

Abstract: An object is to reduce a fluctuation in the gas-turbine output in a nozzle switching period. In the nozzle switching period during which a first nozzle group that has been used is switched to a second nozzle group that is going to be used, the amounts of fuel supplied through the first nozzle group and the second nozzle group are adjusted by using at least one adjustment parameter registered in advance, the adjustment parameter registered in advance is updated according to the operating condition of the gas turbine, and the updated adjustment parameter is registered as an adjustment parameter to be used next.

Abstract: The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles, a fuel manifold system in communication with the micro-mixer fuel nozzles to deliver a flow of fuel thereto, and a linear actuator to maneuver the micro-mixer fuel nozzles and the fuel manifold system.

Abstract: A method and device for igniting a turbomachine combustion chamber including alternately: an intake phase of a fluid into a chamber through an intake port, during which a piston compresses an elastic mechanism under pressure of the fluid such that the elastic mechanism applies onto a piezoelectric element a force sufficient for the piezoelectric element to induce between the electrodes an electric voltage enabling an electric arc to be generated, until the piston reaches a predetermined position for closing a valve for sealing the intake port; and an exhaust phase of the fluid, during which the elastic mechanism pushes back the piston to induce a fluid ejection out of the chamber through an exhaust port, and the valve is open.

Abstract: The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles, a pre-nozzle fuel injection system supporting the fuel nozzles, and a linear actuator to maneuver the fuel nozzles and the pre-nozzle fuel injection system.

Abstract: The present application provides a combustor for use with flow of fuel and a flow of air in a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles positioned within a liner and an air bypass system position about the liner. The air bypass system variably allows a bypass portion of the flow of air to bypass the micro-mixer fuel nozzles.

Abstract: The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a linear actuator so as to maneuver the micro-mixer fuel nozzles axially along the liner.

Abstract: In a method for the low-CO emissions part load operation of a gas turbine with sequential combustion, the air ratio (?) of the operative burners (9) of the second combustor (15) is kept below a maximum air ratio (?max) at part load In order to reduce the maximum air ratio (?), a series of modifications in the operating concept of the gas turbine are carried out individually or in combination. One modification is an opening of the row of variable compressor inlet guide vanes (14) before engaging the second combustor (15). For engaging the second combustor, the row of variable compressor inlet guide vanes (14) is quickly closed and fuel is introduced in a synchronized manner into the burner (9) of the second combustor (15). A further modification is the deactivating of individual burners (9) at part load.

Abstract: A method for operating a firing plant with at least one combustion chamber and at least one burner, especially a gas turbine, includes an operating characteristic for operating the combustion chamber close to the lean extinction limit defined as a burner group staging ratio (BGVRich). Pressure pulsations (PulsActual) measured in the combustion chamber are processed by a filter device (2) and converted into corresponding signals (PulsActual,Filter(t)). An exceeding/falling short of at least one pulsation limiting value (PulsLimit) is monitored by a monitoring device (3) and adapts a pulsation reference value (PulsRef) in dependence upon the monitoring. The processed pressure pulsations (PulsActual,Filter(t)) are then compared with the adapted pulsation reference value (PulsRef,adapt), and, from this, a correction value ?BGV is determined, by which the burner group staging ratio (BGVRich) is corrected, and as a result operation of the firing plant close to the lean extinction limit is realized.

Abstract: A mixture of air and fuel is received into a reaction chamber of a gas turbine system. The fuel is oxidized in the reaction chamber, and a maximum temperature of the mixture in the reaction chamber is controlled to be substantially at or below an inlet temperature of a turbine of the gas turbine system. The oxidation of the fuel is initiated by raising the temperature of the mixture to or above an auto-ignition temperature of the fuel. In some cases, the reaction chamber may be provided without a fuel oxidation catalyst material.

Abstract: Methods and systems for operating a gas turbine engine including a fuel delivery system and a plurality of combustor assemblies are provided. The fuel delivery system comprises a primary fuel circuit configured to continuously supply fuel to each of the plurality of combustor assemblies during a first mode of operation and a second mode of operation. At least one secondary fuel circuit of the fuel delivery system is configured to supply fuel to each of the plurality of combustor assemblies during the second mode of operation. The secondary fuel circuit includes at least one isolation valve coupled in flow communication with each of the plurality of combustor assemblies. The at least one isolation valve facilitates preventing fluid flow upstream into the secondary fuel circuit during the first mode of operation. The fuel delivery system, using the isolation valve, replaces a purging system in the gas turbine engine.

Abstract: An apparatus and method for lean/rich combustion in a gas turbine engine (10), which includes a combustor (12), a transition (14) and a combustor extender (16) that is positioned between the combustor (12) and the transition (14) to connect the combustor (12) to the transition (14). Openings (18) are formed along an outer surface (20) of the combustor extender (16). The gas turbine (10) also includes a fuel manifold (28) to extend along the outer surface (20) of the combustor extender (16), with fuel nozzles (30) to align with the respective openings (18). A method (200) for axial stage combustion in the gas turbine engine (10) is also presented.

Abstract: A method for running up a stationary gas turbine is provided. The turbine has a combustion chamber, the burners of which have a pilot burner and a main burner and by which various types of fuel are introduced into the combustion chamber for burning. The method, carried out while a rotor of the gas turbine is accelerating from a standstill to a nominal speed, includes feeding fuel of a first type of fuel to the pilot burners and feeding fuel of the first type of fuel to the main burners. In order to provide a method in which a comparatively low supply pressure is required in the fuel supply system and in which combustion vibrations that put the machine at risk are avoided during the running-up process, it is proposed that fuel of a second type of fuel is fed to the burner before the nominal speed is reached.

Abstract: A combustion method and combustor for use in a jet engine, the jet engine having a compressor portion and a turbine portion. The combustor includes an outer tube having a central axis that extends longitudinally intermediate the compressor portion and turbine portion and is positioned to receive air discharged by the compressor portion. An inner tube is positioned within the outer tube that includes an associated outer surface spaced from the inner surface of the outer tube thereby defining a combustion chamber. The outer tube and inner tube include fluid directing structure for communicating at least some of the air discharged by the compressor portion to the combustion chamber. The fluid directing structure directs air into the combustion chamber in a direction offset from the central axis, thereby causing rotation or swirling of the air about the central axis.