Abstract: In gas turbines, compressed air is supplied via an air duct to combustion chambers and is heated there. Pressure losses in the air duct should be minimized in order to ensure good overall efficiency. This is achieved by the compressed air flowing with approximately constant velocity in the air duct from the compressor to the inlet into the combustion chamber. This is supported by the effective cross section of the air duct being almost constant over this distance.

Abstract: 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.

Abstract: An outer crossfire tube for attachment between adjacent combustors in a land-based gas turbine includes a substantially cylindrical bellows portion and a pair of uniform diameter free ends on opposite sides of the bellows portion, the uniform diameter free ends adapted to seat within apertures provided in the adjacent combustors. The outer crossfire tube is surrounded by a telescoping sleeve assembly, with remote ends of a pair of telescoping sleeve members fixed to respective combustor flanges.

Abstract: A three-stage lean burn combustion chamber (28) comprises a primary combustion zone (36), a secondary combustion zone (40) and a tertiary combustion zone (44). Each of the combustion zones (36,40,44) is supplied with premixed fuel and air by respective fuel and air mixing ducts (54,70,92). The fuel and air mixing ducts (54,70,92) have a plurality of air injections apertures (62,64,76,98) spaced apart in the direction of flow through the fuel and air mixing ducts (54,70,92). The apertures (62,64,76,98) reduce the magnitude of the fluctuations in the fuel to air ratio of the fuel and air mixture supplied into the at least one combustion zone (36,40,44). This reduces the generation of harmful vibrations in the combustion chamber (28).

Abstract: A cyclone combustor of the present invention uses a novel pre-mixture injection scheme to optimize performance. The cyclone combustor includes a cylindrical combustor can and three fuel/air premixing tubes entering the combustor can radially, with a tangential offset. The tangential offset is designed to provide an optimized circulation in the combustor can for improvement of liner life span, flame stability and engine turn-down. The ignition and pilot fuel systems are placed to take advantage of the premixing tube entry locations and the tangential direction of the mixture flow momentum in the combustor can. The special combination of the parallel axes of the combustor can and the mixing tubes provides a right angle between an outlet section and the major tube section of each premixing tube. The cyclone combustor of the present invention can meet the requirements for low NOx and CO emissions.

Abstract: A burner configuration includes a combustion chamber. A multiplicity of burners are disposed in the combustion chamber. Each of the burners has an outlet opening into the combustion chamber. Flow-guidance elements each at least partly form a respective one of the outlets of at least some of the burners. The flow-guidance elements project into the combustion chamber for guiding a fuel-gas flow discharging from the burners into the combustion chamber. As a result, combustion oscillations are suppressed.

Abstract: In a combustion device (10), particularly for driving gas turbines, comprising a plurality of burners (12, . . . , 15) of identical thermal power output, which work parallel to an axis (28) into a common combustion chamber (11), an effective suppression of thermoacoustic combustion instabilities is achieved in a simple way in that the burners (12, . . . , 15) are designed differently from one another in such a way that the flames (24, . . . , 27) or flame fronts generated by them are positioned so as to be distributed along the axis (28).

Abstract: A combustion chamber arrangement for gas turbines includes a multiplicity of individual combustion chambers which open out in a common annular combustion chamber and are arranged in the shape of a circle. The arrangement is such that it can be cooled highly efficiently using convective cooling, while also as far as possible dispensing with the use of impingement cooling. In such a combustion chamber arrangement, the annular combustion chamber, in the transition region to the individual combustion chambers, is of a height which fluctuates periodically and which is minimal in the region between adjacent individual combustion chambers.

Abstract: In gas turbines, compressed air is supplied via an air duct to combustion chambers and is heated there. Pressure losses in the air duct should be minimized in order to ensure good overall efficiency. This is achieved by the compressed air flowing with approximately constant velocity in the air duct from the compressor to the inlet into the combustion chamber. This is supported by the effective cross section of the air duct being almost constant over this distance.

Abstract: A combustor for a gas turbine engine includes outer and inner liners defining a combustion chamber and an igniter mounted to the outer liner. A dome plate is disposed between the outer and inner liners and has a plurality of circumferentially spaced openings formed therein. A fuel-air mixer is disposed in each one of the openings; each fuel-air mixer includes a swirler mounted in the corresponding opening and a fuel nozzle received in the corresponding swirler. Two adjacent ones of the swirlers are low air flow swirlers in general circumferential alignment with the igniter and the rest of the swirlers are higher air flow swirlers. Each one of the fuel nozzles provides a similar amount of fuel so that the two fuel-air mixers having low air flow swirlers produce a local region of increased fuel-to-air ratio in the vicinity of the igniter.

Abstract: A combustion chamber according to the invention has a number b0 of annularly arranged burners, of which a number k of modulatable burners have means for modulating a fuel mass flow, k being k<b0, and the modulatable burners being arranged in such a way that between every pair of adjacent modulatable burners are arranged in each case a1, a2, . . . ak nonmodulatable burners, and that the values a1+1, a2+1, . . . , ak+1 are not integral divisors of b0.

Abstract: A gas turbine is provided which has a combustor comprising a plurality of can-type combustor elements and an annular element. Isolated combustion takes place in the can-type combustor thereby generating individual component hot gas streams. The component hot gas streams are merged in the annular part. Combustion oscillations are substantially avoided by decoupling combustion flames. Accordingly, temperature and pressure homogenization as well as combustion stabilization by cross-ignition is effected.

Abstract: An integral aft seal for sealing the interface between a gas turbine transition piece and first stage nozzle exhibits sufficient variable thickness to adjust for relative component movement. The seal comprises a contiguous material having opposing planar faces and stiffeners attached at each of the faces. The contiguous material has corrugations to provide variable thickness. The stiffeners are attached to the planar faces of the contiguous material and generally have the same planar shape as the seal. Flexibility in the thickness dimension due to the corrugations biases the stiffeners away from one another and into contact with the turbine components. The aft seal is affixed to the transition piece by an attachment structure that eliminates relative movement between the seal and the transition piece to thereby eliminate wear.

Abstract: A fuel nozzle guide and method of assembly/disassembly is disclosed with the nozzle guide having a frusto-conical hub section, an annular base, a pair of retaining tabs for securing the guide to the outer wall of a combustor bulkhead for limited movement relative to the bulkhead, and a radial inflow swirler. The method of assembly/disassembly comprises assembling/disassembling the nozzle guide in the bulkhead from the outer wall (cold) side of the bulkhead and mechanically connecting/disconnecting the nozzle guide to/from the outer wall of the bulkhead with mechanical fasteners.

Abstract: A gas turbine combustion device comprises a plurality of combustion devices. Each combustion device has an inner cylinder and a tail cylinder in a vehicle chamber. An acoustic sleeve is provided between each inner cylinder and a corresponding outer cylinder in the vehicle chamber. An acoustic coupling effect with a vehicle chamber can be reduced.

Abstract: A turbine combustion system comprises at least one turbine combustion component that is provided with a protective coating on the outer surface otherwise known as the non-flame surface of at least one turbine combustion component. The at least one turbine combustion component comprises an inner surface that defines a hot flame path area and an outer surface. The coating on the outer surface of the combustion components provides protection for the components from at least one of oxidation, nitridation, and hot corrosion.

Abstract: A gas turbine engine for generating electricity having low NOx emissions that includes a turbine system linearly and axially connected by a power shaft to a compressor section and having a combustion section mounted vertically relative to the turbine section and the compressor section. The combustion system includes a plurality of individual combustors mounted in a circular or annular array around the upper cap or dome of the combustion system, each of said combustors being a dual mode, two-stage, emitting low levels of NOx. Each combustor exhausts its combustion gases into a common central plenum chamber that is vertically oriented relative to the turbine engine centerline. The plenum provides the hot gases to the turbine blades through an annular chamber 360 degrees around the shaft. The gas turbine engine vertical combustion system provides for a highly efficient, low nitric oxide emissions while allowing for uniform mixing of the combusting gases powering the turbine system.

Abstract: A computer-implemented model of fan section of a gas turbine engine accounts for the turbulence in the gas flow emanating from the rotor assembly and impinging upon an inlet to the stator vane cascade. The model allows for user-input variations in the sweep and/or lean angles for the stator vanes. The model determines the resulting acoustic response of the fan section as a function of the turbulence and the lean and/or sweep angles of the vanes. The model may be embodied in software that is rapidly executed in a computer. This way, an optimum arrangement in terms of fan noise reduction is quickly determined for the stator vane lean and sweep physical positioning in the fan section of a gas turbine engine.

Abstract: A sealing device for resisting the flow of gas through a labyrinth seal between a pair of adjacent combustor transitions of a gas turbine engine includes an elongated frame formed with a seat and a seal member disposed within the seat. The seal member is formed out of a metallic felt material and is progressively compressed within the labyrinth seal during initial engine startup, and provides resistance to leakage during steady state operation of the engine. The frame includes a first connection tab extending outwardly from the seat, the first connection tab being fixedly mounted on a ridge of the labyrinth seal. A second embodiment of the present invention is disclosed in which the frame includes a seat and both a first connection tab and a second connection tab extending outwardly from the seat in opposite directions.

Abstract: A crossfire tube for connecting adjacent combustors in a gas turbine, the crossfire tube includes a hollow tubular body having opposite end portions adapted to be secured to the adjacent combustors; an annular chamber surrounding a mid-section of the hollow tubular member; a first plurality of purge air feed holes in an outer wall of the chamber, and a second plurality of purge air feed holes in an inner wall of the chamber opening into the crossfire tube.

Abstract: A combustor for a gas turbine engine uses compressed air to cool a combustor liner and uses at least a portion of the same compressed air for combustion air. A flow diverting mechanism regulates compressed air flow entering a combustion air plenum feeding combustion air to a plurality of fuel nozzles. The flow diverting mechanism adjusts combustion air according to engine loading.

Abstract: In a central portion of inner tube 28 of combustor 20, pilot fuel nozzle 22 and pilot cone 33 are arranged and main fuel nozzles 21 and main swirlers 32 therearound. Air intake portion (X-1) is provided with rectifier tube 11 for making air intake uniform. In air intake portion (X-2), air holes of appropriate number of pieces are provided in circumferential wall of the inner tube 28. In main swirler portion (X-3) and pilot cone portion (X-4), bolt joint of the main swirlers 32 is employed and optimized welded structure having less influence of thermal stress of the pilot swirler 33 is employed, respectively. Tail tube cooling portion (X-5) is provided with cooling structure having less influence of thermal stress to cool flange 71 portion of tail tube 24 uniformly. By the improvements in the portions (X-1) to (X-5), obstacles in attaining higher temperature in the combustor 20 is dissolved and combustor performance is enhanced.

Abstract: A gas-turbine engine combustion system comprises a plurality of combustors interconnected by crossfire tube assemblies adapted to pass an ignition flame from an ignited combustor to another combustor on start-up of the engine. Each crossfire tube assembly comprises a cooling air inlet for introducing air into the assembly to film-cool its inner, ignition flame-facing surface, and a cooling sleeve surrounding the crossfire tube assembly in the region in which it opens into the combustor. The cooling sleeve directs the cooling air so as to cool the outer surface of the assembly, the inside of the sleeve and the inside surface of the combustor wall adjacent the cooling sleeve.

Abstract: A fuel control system for controlling the ratio of ring temperature adjustment in rings of a combustor in a gas turbine. The fuel control system consists of employing a computer to perform the steps of; defining an operational boundary inner ring temperature adjustments versus outer ring temperature adjustments that defines a safe operating region for the annular combustor of the gas turbine; calculating a point of operation of the inner ring temperature adjustment versus the outer ring temperature adjustment within a safety margin of the operational boundary, wherein nitrous oxide (NOx) emission levels of the gas turbine is reduced; and regulating the ring temperatures to maintain a near global minimum point of operation, while maintaining normal operation of the gas turbine.