Abstract: The present invention provides a combustor for an aerospace gas turbine engine comprising two stages wherein each stage defines an inlet and an exit. The second stage inlet is in fluid communication with the first stage exit such that a first flowpath is defined and it passes substantially through the second stage. A plurality of flow channel tubes is positioned within the second stage and each flow channel tube passes sealingly through a header plate positioned upstream of the second stage inlet thereby defining a second flowpath that also passes substantially through the second stage. The first flowpath exit and the second flowpath exit are positioned adjacent and proximate to one another to provide for the generation of microflames or microflame jets exiting the second stage from between and around the flow channel tube exits. The first stage of the combustor provides a gasifier and a reformer.

Abstract: A method for operating a burner including a burner outlet opening with at least two sectors, each sector is assigned at least one fuel nozzle, is provided. The method is characterized in that fuel is supplied separately to the fuel nozzles of different sectors. Also described is a burner which includes at least two sectors wherein each sector is assigned at least one fuel nozzle. The burner includes at least two separate fuel supply lines and a device for adjusting the fuel mass flow which flows through the respective fuel supply line. The fuel supply lines supply fuel to the fuel nozzles of different sectors. Also described is a gas turbine which is fitted with at least one burner.

Abstract: A turbomachine including a combustor in which fuel is combustible to produce a working fluid, a turbine section, which is receptive of the working fluid for power generation operations, a transition piece in which additional fuel is combustible, the transition piece being disposed to transport the working fluid from the combustor to the turbine section and a staged combustion system coupled to the combustor and the transition piece. The staged combustion system is configured to blend components of the fuel and the additional fuel in multiple modes.

Abstract: An embodiment of the present invention provides a method and system of operating a combustion system that has Lean direct injection (LDI) functionality. The method and system provides a passive cooling system for an injector of the LDI system. The method and system may also provide a means to direct the flow of the fluid exiting the injector of the LDI system.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a flow sleeve that circumferentially surrounds the combustion chamber. A plurality of fuel injectors are circumferentially arranged around the flow sleeve to provide fluid communication through the flow sleeve to the combustion chamber. A distribution manifold circumferentially surrounds the plurality of fuel injectors, and a fluid passage through the flow sleeve and into the distribution manifold provides fluid communication through the flow sleeve to the plurality of fuel injectors.

Abstract: This invention provides a gas turbine combustor including: a main burner at a head portion of a combustor cylinder; and a pre-mixing type supplemental burner at a downstream portion of the combustor cylinder and extending through a circumferential wall thereof. The supplemental burner includes: an introducing passage configured to deflect a part of the compressed air radially inward, the compressed air flowing from an air passage between the circumferential wall of the combustor cylinder and a housing surrounding the circumferential wall toward the head portion of the combustor cylinder, and introduce the compressed air into the combustor cylinder; and a fuel nozzle configured to supply the fuel from fuel injection holes to the compressed air introduced into the introducing passage to produce a pre-mixed gas in the introducing passage.

Abstract: A method is provided for fuel injection in a sequential combustion system comprising a first combustion chamber and downstream thereof a second combustion chamber, in between which at least one vortex generator is located, as well as a premixing chamber having a longitudinal axis downstream of the vortex generator, and a fuel lance having a vertical portion and a horizontal portion, being located within said premixing chamber. The fuel injected is an MBtu-fuel. In said premixing chamber the fuel and a gas contained in an oxidizing stream coming from the first combustion chamber are premixed to a combustible mixture. The fuel is injected in such a way that the residence time of the fuel in the premixing chamber is reduced in comparison with a radial injection of the fuel from the horizontal portion of the fuel lance.

Abstract: Methods and systems are provided for premixing combustion fuel and air within gas turbines. In one embodiment, a combustor includes an upstream mixing panel configured to direct compressed air and combustion fuel through a premixing zone to form a fuel-air mixture. The combustor also includes a downstream mixing panel configured to mix additional combustion fuel with the fuel-air mixture to form a combustion mixture.

Abstract: An object of the present invention is to provide a gas turbine combustor that supports hydrogen-containing gas having a high burning velocity and is capable of performing low NOx combustion without reducing reliability of a burner. A first fuel nozzle is provided upstream of a combustion chamber and supplies fuel for activation and hydrogen-containing gas. The combustor has a primary combustion zone, a reduction zone and a secondary combustion zone. In the primary combustion zone, the fuel supplied from the first fuel nozzle is combusted under a fuel rich condition to form a burned gas containing a low concentration of oxygen. In the reduction zone, a hydrogen-containing gas is injected into the combustion chamber through a second fuel injection hole from a second fuel nozzle so that NOx generated in the primary combustion zone is reduced by an oxygen reaction of the hydrogen.

Abstract: A combustor section of a gas turbine includes a combustor liner, a sleeve and a fuel-air mixing tube. The combustor liner defines a combustion chamber. The sleeve surrounds the combustor liner. The combustor liner and the sleeve define an annular flow space. The fuel-air mixing tube is configured to channel a mixture of fuel and air and includes an inlet and an outlet. The inlet is in fluid communication with an exterior of the sleeve, and the outlet is in fluid communication with the combustion chamber. The combustor casing encloses the combustor section upstream relative to the inlet of the mixing tube and extends downstream. The sleeve and the combustor casing define a discharge air space. The discharge space is in fluid communication with the mixing tube. The fuel supplying device is located exteriorly of the combustor casing and is configured to inject fuel into the mixing tube.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a flow sleeve that circumferentially surrounds the combustion chamber. A plurality of fuel injectors are circumferentially arranged around the flow sleeve to provide fluid communication through the flow sleeve to the combustion chamber. A distribution manifold circumferentially surrounds the plurality of fuel injectors, and a fluid passage through the flow sleeve and into the distribution manifold provides fluid communication through the flow sleeve to the plurality of fuel injectors.

Abstract: A system for supplying fuel to a combustor includes a combustion chamber, a liner that circumferentially surrounds at least a portion of the combustion chamber, and a flow sleeve that circumferentially surrounds at least a portion of the liner. A tube provides fluid communication for a working fluid to flow through the flow sleeve and the liner and into the combustion chamber, and the tube includes a tube wall. A plurality of fuel ports through the tube wall provide fluid communication for fuel to flow through the tube wall and into the tube, and the fuel ports are non-circular.

Abstract: A system for supplying fuel to a combustor includes a combustion chamber and a fuel nozzle that provides fluid communication into the combustion chamber. A plurality of passages circumferentially arranged around the combustion chamber provide fluid communication into the combustion chamber. A liquid fuel plenum provides fluid communication to the plurality of passages. A baffle circumferentially surrounds at least a portion of the liquid fuel plenum inside the plurality of passages and forms a plurality of lobes around the liquid fuel plenum.

Abstract: A combustion apparatus in a gas turbine engine comprises a combustor shell for receiving air, a fuel injection system associated with the combustor shell, a first fuel supply structure, and a shield structure. The fuel supply structure is in fluid communication with a source of fuel for delivering fuel from the source of fuel to the fuel injection system and comprises a first fuel supply elements including a first section extending along a first path having a component in an axial direction and a second section extending from the first section along a second path having a component in a circumferential direction. The shield structure is associated with at least a portion of the second section of the first fuel supply element.

Abstract: In one embodiment, a method is provided fuel staging for a trapped vortex (TVC) combustion apparatus comprising an inlet premixer, for injecting fuel-air mixture into the inlet of the combustion apparatus and a vortex premixer, for injecting fuel-air mixture into the recirculating vortex. The combustion apparatus may be part of an engine, such as a gas turbine engine. The method comprises varying the relative proportion of mixture introduced through the inlet and vortex premixers as a function of load.

Abstract: A system for conditioning a working fluid in a combustor includes a primary combustion chamber, a liner circumferentially surrounding the primary combustion chamber, and a primary nozzle in fluid communication with the primary combustion chamber. A secondary combustion chamber located outside of the primary combustion chamber includes a shroud that defines a fluid passage, a secondary nozzle, and means for igniting fuel in the secondary combustion chamber. A method for conditioning a working fluid in a combustor includes flowing the working fluid through a primary combustion chamber and flowing at least a portion of the working fluid through a secondary combustion chamber located outside of the primary combustion chamber. The method further includes flowing a fuel through the secondary combustion chamber, combusting the fuel, and flowing the combustion gases from the secondary combustion chamber into the primary combustion chamber.

Abstract: In an SEV combustor and a method for reducing emissions in an SEV combustor of a sequential combustion gas turbine, an air/fuel mixture is combusted in a first burner and the hot gases are subsequently introduced into the SEV combustor (1) for further combustion. The SEV combustor (1) includes a chamber having a chamber wall (5) defining a mixing portion (8), for mixing the hot gases with a fuel, and a combustion region (9), at least one inlet (2) for introducing the hot gases into the mixing region (8), at least one inlet (12) for introducing a fuel into the mixing region (8) and at least one inlet (10, 13) for introducing steam into the mixing region.

Abstract: An apparatus and process is provided for combining fuel and combustion air to produce a mixture. The mixture is burned in a combustion chamber to produce a flame.

Abstract: A system for supplying a working fluid to a combustor includes a fuel nozzle, a combustion chamber downstream from the fuel nozzle, and a flow sleeve that circumferentially surrounds the combustion chamber. A plurality of fuel injectors are circumferentially arranged around the flow sleeve to provide fluid communication through the flow sleeve and into the combustion chamber. A separate cap for each fuel injector defines a separate volume around a different fuel injector outside of the flow sleeve, and a fluid passage through the flow sleeve and into each separate volume provides fluid communication through the flow sleeve and into each separate volume.

Abstract: A flow divider assembly includes a dividing valve and a balancing valve. The dividing valve includes a dividing cavity configured to receive a flow of fuel and a piston device arranged within the dividing cavity and selectively positioned to divide the flow of fuel in the dividing cavity into the first dividing valve port and the second dividing valve port. The balancing valve is fluidly coupled to the dividing valve and configured to receive the first portion and the second portion of the flow. The balancing valve includes a balancing cavity and a balancing device arranged within the balancing cavity and selectively positioned to direct the first portion of the flow into the first balancing cavity port at a third pressure and the second portion of the flow into the second balancing cavity port at a fourth pressure such that the first pressure is approximately equal to the second pressure.

Abstract: A combustor for a gas turbine capable of performing stable combustion even by using high temperature air, comprising a first burner (5) jetting fuel and air into a combustion chamber (2) and a second burner (8) causing the circulating jet of the fuel and air installed at a position corresponding to the tip part of a flame caused by the first burner (5).

Abstract: A combustor (28) for a gas turbine engine is provided comprising a primary combustion chamber (30) for combusting a first fuel to form a combustion flow stream (50) and a transition piece (32) located downstream from the primary combustion chamber (30). The transition piece (32) comprises a plurality of injectors (66) located around a circumference of the transition piece (32) for injecting a second fuel into the combustion flow stream (50). The injectors (66) are effective to create a radial temperature profile (74) at an exit (58) of the transition piece (32) having a reduced coefficient of variation relative to a radial temperature profile (64) at an inlet (54) of the transition piece (32). Methods for controlling the temperature profile of a secondary injection are also provided.

Abstract: A burner for stabilizing the combustion of a gas turbine is provided. The burner includes a combustion chamber and a plurality of nozzles opening out into the combustion chamber wherein fluid is introduced into the combustion chamber by the nozzles in the form of a fluid jet, wherein the fluid is burned in the combustion chamber to fatal a hot gas, wherein an annular gap is provided in the case of at least one nozzle for feeding the fluid to the nozzle at a nozzle inlet, wherein a preheater is provided for heating the fluid before entering the nozzle, wherein the preheater is a pre-burner with or without a combustion chamber, and wherein the pre-burner with or without a combustion chamber is arranged in the annular gap.

Abstract: The disclosure relates to a burner for a combustion chamber of a gas turbine, with an injection device for the introduction of at least one gaseous and/or liquid fuel into the burner, wherein the injection device has at least one body which is arranged in the burner with at least one nozzle for introducing the at least one fuel into the burner, the at least one body being configured as a streamlined body which has a streamlined cross-sectional profile and which extends with a longitudinal direction perpendicularly or at an inclination to a main flow direction prevailing in the burner. The at least one nozzle has its outlet orifice at or in a trailing edge of the streamlined body, and with reference to a central plane of the streamlined body, the trailing edge is provided with at least two lobes extending in opposite transverse directions.

Abstract: A fuel injector in a combustor apparatus of a gas turbine engine. An outer wall of the injector defines an interior volume in which an intermediate wall is disposed. A first gap is formed between the outer wall and the intermediate wall. The intermediate wall defines an internal volume in which an inner wall is disposed. A second gap is formed between the intermediate wall and the inner wall. The second gap receives cooling fluid that cools the injector. The cooling fluid provides convective cooling to the intermediate wall as it flows within the second gap. The cooling fluid also flows through apertures in the intermediate wall into the first gap where it provides impingement cooling to the outer wall and provides convective cooling to the outer wall. The inner wall defines a passageway that delivers fuel into a liner downstream from a main combustion zone.

Abstract: A method of combustion stability control for a gas turbine engine is provided, and includes the steps of receiving by a stability controller, information regarding environmental and operating conditions, and comparing the environmental and operating conditions to pre-programmed information to determine if a likelihood of combustion instability exists. The method further includes the steps of determining optimal fuel modulation frequency and amplitude for the environmental condition to reduce combustion instability, if a likelihood of combustion instability exists, and actuating at least one fuel modulation valve to, at the optimal fuel modulation frequency and amplitude, reduce combustion instability, if a likelihood of combustion instability exists. Systems for modulating fuel flow are also provided.

Abstract: A combustor (1) for a gas turbine engine, particularly for a gas turbine having sequential combustion, includes a combustor wall (4) defining a mixing region (5) and a combustion region (6). The mixing region (5) has at least one first inlet (2) for introducing combustion air into the mixing region (5) and at least one second inlet for introducing fuel into the mixing region (5), the combustion region (6) extending downstream of the mixing region. The mixing region (5) crosses over to the combustion region (6) in a transition region (14). A baffle (9) extends from the transition region (14) generally in the downstream direction (15), forming at least one space (10) between the combustor wall (4) and the baffle (9).

Abstract: A fuel lance (7) for introducing fuel into a gas flow in a combustor (1) of a gas turbine engine includes a region of the lance (7) through which the fuel is introduced into the gas flow having a generally helical formation (12).

Abstract: A catalytic reactor for a gas turbine engine comprising an air inlet, a premixing zone, a reacting zone comprising a reactive portion and a nonreactive portion, a post reaction mixing zone, a first fuel injection system for introducing fuel into the reactive portion, and a second fuel injection system for introducing fuel into the nonreactive portion.

Abstract: The invention relates to a method for starting a burner for combusting synthesis gases, wherein said burner comprises first and second fuel passages, the first fuel passage encompasses the second fuel passage in a substantially concentric manner and the gas transferred to the burner is mixed with combusting air and is combusted. According to said invention, in order to start the burner, the second fuel passage is first loaded with a synthesis gas to a predefined burner power at a first starting phase and the first fuel passage is loaded with the synthesis gas at a second starting phase.

Abstract: The low emission combustor includes a combustor housing defining a combustion chamber having a plurality of combustion zones. A liner sleeve is disposed in the combustion housing with a gap formed between the liner sleeve and the combustor housing. A secondary nozzle is disposed along a centerline of the combustion chamber and configured to inject a first fluid comprising air, at least one diluent, fuel, or combinations thereof to a downstream side of a first combustion zone among the plurality of combustion zones. A plurality of primary fuel nozzles is disposed proximate to an upstream side of the combustion chamber and located around the secondary nozzle and configured to inject a second fluid comprising air and fuel to an upstream side of the first combustion zone. The combustor also includes a plurality of tertiary coanda nozzles. Each tertiary coanda nozzle is coupled to a respective dilution hole.

Abstract: A combustion chamber and method for operating a combustion chamber are disclosed. The combustion chamber having at least a mixing device connected to a combustion device. The mixing device has a first fuel feeding stage, to inject fuel into the mixing device and mix the fuel with an oxidizer to then burn the fuel in the combustion device. The combustion device has a second fuel feeding stage to inject fuel into the combustion device. The fuel of the second stage is mixed with only an inert fluid to form a mixture that is then injected into the combustion chamber.

Abstract: A consumption amount of high-calorific gas such as coke oven gas (COG) during operation of a gas turbine is reduced, halt of the gas turbine due to clogging of a pilot system, a malfunction of a compressor which compresses high-calorific gas is prevented, and reliability of the gas turbine is improved. When operation of the gas turbine (11) starts, with use of both a first fuel supply system (31) which supplies a high-calorific fuel for a first nozzle constituting a combustor (17), and a second fuel supply system (32) which supplies a low-calorific fuel for a second nozzle constituting the combustor (17), the high-calorific fuel and the low-calorific fuel are supplied to the combustor (17), and at a time when the gas turbine (11) reaches output power which enables continuous operation with only the low-calorific fuel, supply of the high-calorific fuel to the combustor (17) is shut off, and only the low-calorific fuel is supplied to the combustor (17).

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 gas turbine combustor includes a combustion chamber defined by a combustor liner, the combustor liner having an upstream end cover supporting one or more nozzles arranged to supply fuel to the combustion chamber where the fuel mixes with air supplied from a compressor. A transition duct is connected between an aft end of the combustion chamber liner and a first stage turbine casing, the transition duct supplying gaseous products of combustion to the first stage turbine nozzle. One or more additional fuel injection nozzles are arranged at an aft end of the transition duct for introducing additional fuel into the transition duct upstream of the first stage turbine nozzle.

Abstract: 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.

Abstract: A method for assembling a gas turbine combustor system is provided. The method includes providing a combustion liner including a center axis, an outer wall, a first end, and a second end. The outer wall is orientated substantially parallel to the center axis. The method also includes coupling a transition piece to the liner second end. The transition piece includes an outer wall. The method further includes coupling a plurality of lean-direct injectors along at least one of the liner outer wall and the transition piece outer wall such that the injectors are spaced axially apart along the wall.

Abstract: A gas turbine combustion system is provided comprising a combustion chamber (16) having a central axis (44), a primary combustion stage (28) located at a front end (32) of the combustion chamber (16) for injecting fuel, air, or mixtures thereof substantially along the central axis (44), a plurality of secondary combustion stages (30A-D) spaced apart in flow series along a length of the combustion chamber (16), wherein each of the plurality of secondary combustion stages (30A-D) comprises a plurality of circumferentially-spaced secondary injectors (48) for injecting fuel, air, or mixtures thereof, toward the central axis (44), and wherein an internal diameter of the combustion chamber (16) decreases from at least a first one of the plurality of secondary combustion stages (30AD) to at least a second one of the plurality of secondary combustion stages (30A-D).

Abstract: A method and apparatus for measuring a start fuel flow to a pilot nozzle of a fuel system of a gas turbine engine using pressure differential between fuel passages leading to fuel nozzles.

Abstract: A combustion apparatus and duct structure for a gas turbine engine are provided. The combustion apparatus comprises a combustion system to receive fuel and air, ignite at least a portion of the fuel and air and output a stream of first combustion products and any remaining fuel and air. The combustion apparatus further comprises structure positioned adjacent to the combustion system for receiving and accelerating the first combustion products and any remaining fuel and air from the combustion system. The duct structure receives the first combustion products and any remaining fuel and air from the combustion apparatus, allows any remaining fuel and air to combust to generate second combustion products, accelerates the first and second combustion products and outputs the first and second combustion products to a first row blade assembly.

Abstract: A catalytic reactor for a gas turbine engine comprising an air inlet, a premixing zone, a reacting zone comprising a reactive portion and a nonreactive portion, a post reaction mixing zone, a first fuel injection system for introducing fuel into the reactive portion, and a second fuel injection system for introducing fuel into the nonreactive portion.

Abstract: A gas turbine engine air valve assembly has first and second valving elements. The second element is rotatable about a first axis relative to the first element. The rotation controls a flow of air through the first and second elements. An actuator is coupled by a linkage to the second element. A plurality of follower member assemblies each have an insertion portion extending through a corresponding first aperture in the first valving element and a corresponding second aperture in the second element. The insertion portion is circumferentially fixed relative to one of the first and second valving elements and circumferentially moveable relative to the other of the first and second valving elements.

Abstract: A method and apparatus for measuring a start fuel flow to a pilot nozzle of a fuel system of a gas turbine engine using pressure differential between fuel passages leading to fuel nozzles.

Abstract: A power generation system (10) and method of operating a power generation system (10). An exemplary power generation system (10) shown in FIG. 1 includes a gasification system (2) for generating a supply (33) of a first gas comprising hydrogen and a supply (39) of a second gas comprising carbon monoxide based on a chemical breakdown of hydrocarbons in a melt (17). A component system (3) produces a supply of steam (99) by reacting one of the streams of gas and a steam turbine (120) generates mechanical power by expanding the steam (99). In an associated method, a melt (17) is created in a gasifier (12) which then receives a hydrocarbon source (22) to generate a continuous supply (33) of a first gas comprising hydrogen while increasing concentration of dissolved carbon in the melt (17). Hydrogen present in the first gas is supplied to a first combustor (90). A continuous supply (39) of a second gas comprising carbon monoxide is generated with carbon present in the melt (17).

Abstract: A combustor assembly includes a combustor chamber having a primary and intermediate zone that provides for reduced flame temperatures. The combustor assembly includes first and second pluralities of injectors. The first plurality of injectors introduces fuel to a primary zone. A second plurality of injectors introduces fuel to an intermediate zone. During operation between initial start up and before the introduction of engine load, fuel is introduced into the primary zone only by the first plurality of injectors. Once engine load is applied to the engine, fuel is introduced into the intermediate zone by the second plurality of injectors. Introduction of additional volume of fuel allows the fuel-air ratio to remain constant regardless of engine operating conditions. The constant fuel-air ratio is maintained at a desired rate to lower flame temperatures and reduce nitrous oxide emissions.

Abstract: A combustor system for a miniature gas turbine engine includes fuel injection via orifices which direct fuel into fuel-air injection tubes which feed a fuel-rich mixture of fuel and air into a leading end of the combustor liner to form a primary burning region. Fuel system pressures are kept low and controlled by control of the fuel injection port size and number. Fuel breakup is via airblast and tube wall impingement. The fuel-air injection tubes are directed circumferentially, radially outward and toward the front end of the combustor. Air is fed into the combustor such that two distinct burning regions are created. Each region approximates a “well-stirred reactor” and the combination of the two regions results in an efficient use of combustion volume.

Abstract: A premixed air-fuel mixture supply device is joined to a combustor liner included in a combustor for a gas turbine or an aircraft engine. The premixed air-fuel mixture supply device has a pilot fuel injection unit having an inner wall, and a prevaporizing, premixing main fuel injection unit having an outer wall surrounding the inner wall. A combustion air passage is defined by the inner and the outer wall. An intermediate wall is disposed in the combustion air passage so as to divide an upstream part of the combustion air passage into a secondary combustion air passage surrounding the inner wall, and an outer combustion air passage surrounding the intermediate wall. Fuel inject holes are formed in the intermediate wall to inject fuel radially outward into the outer combustion air passage. A swirling device is disposed in the secondary combustion air passage to swirl combustion air flowing through the secondary combustion air passage. An atomization lip is formed in a tail part of the intermediate wall.

Abstract: An improved combination (110) of a premixing chamber (112) and a combustion chamber (111), with low emission of pollutants, for gas turbines running on liquid and/or gas fuel, in which the combustion chamber (111) has a truncated conical end (111a) and is surrounded by a cavity (115) for cooling air (120), the cavity (115) carrying combustion air (120a) to the premixing chamber (112), which has at its inlet apertures or ports (126) which can be constricted according to the quantity of fuel used, and at its outlet a circumferential set of burners (134) to create a corresponding set of additional flames; gas fuel supply means (125) fuel and liquid fuel injection devices (119) are provided in the premixing chamber (112), and a set of liquid fuel injectors (140) is present on the end (111a).

Abstract: A wear reduction means for a fuel nozzle interface of an industrial gas turbine combustor is disclosed. The wear reduction means includes a replaceable insert welded into the engagement location of the fuel nozzle tip. The fuel nozzle tip is coated with a material harder than the replaceable insert, such that any wear due to mechanical contact of the mating components is directed to the replaceable insert.

Abstract: A system for the destruction of volatile organic compounds while generating power. In a preferred embodiment the system comprises a combustor and a reaction chamber connected to an exit of the combustor. A primary inlet to the combustor supplies a primary fuel to the combustor. A secondary fuel, comprising air and an amount of one or more volatile organic compounds, is supplied to a compressor, which compresses the secondary fuel and directs the secondary fuel to the combustor and the reaction chamber. The system is suitably configured to enable the stoichiometric reaction of the two fuels in a manner sufficient to destroy the volatile organic compounds contained in the secondary fuel and power a turbine engine connected to an exit of the reaction chamber.