Abstract: An aerodynamically stabilized premixing burner includes a swirl generator for the production of a rotating combustion air flow, and a device for the introduction of at least one fuel into this combustion air flow. The burner is advantageously provided with a device for the introduction of an axial air flow into the center of the generated rotational flow. This axial air flow is controllable in order to affect the position and intensity of the flame-stabilizing recirculation zone at the burner mouth.

Abstract: A combustion control system controls a turbine engine that includes a fuel manifold and a plurality of fuel injectors. The control system includes a fuel pulsator and a controller. The fuel pulsator is coupled in flow communication with the plurality of injectors and the fuel manifold. The controller is coupled to the fuel pulsator such that the pulsator is between the controller and the fuel manifold. The controller is variably selectable and configured to facilitate promoting stable combustion.

Abstract: Respective combustion gas streams are generated in a can combustor. The streams are channeled downstream into an annular turbine nozzle. And, dynamic interaction of circumferentially adjacent combustion gas streams is suppressed axially between the cans and the nozzle.

Abstract: In a method of minimizing thermoacoustic vibrations in gas-turbine combustion chambers, a modulated spraying of liquid or gaseous premix fuel into a premix burner is carried out. The premix fuel is fed by means of two fuel lines (13, 14), and the modulation of the quantity of fed fuel is effected in each case by one fuel valve per fuel line.

Abstract: A gas turbine combustor includes a side wall, for defining a combustion volume, having upstream and downstream ends, a pilot nozzle, disposed adjacent the upstream end of the side wall, for discharging a pilot fuel to form a diffusion flame in the combustion volume, and a plurality of main nozzles, provided around the pilot nozzles, for discharging a fuel-air mixture to form premixed flames in the combustion volume. Film air is supplied into the combustion volume downstream of the main nozzles along the inner surface of the side wall to reduce the fuel-air ratio in a region adjacent the inner surface of the side wall and to restrain a combustion-driven oscillation in the combustion volume.

Abstract: A gas turbine combustor includes a side wall, for defining a combustion volume, having upstream and downstream ends, a pilot nozzle, disposed adjacent the upstream end of the side wall, for discharging a pilot fuel to form a diffusion flame in the combustion volume, and a plurality of main nozzles, provided around the pilot nozzles, for discharging a fuel-air mixture to form premixed flames in the combustion volume. Film air is supplied into the combustion volume downstream of the main nozzles along the inner surface of the side wall to reduce the fuel-air ratio in a region adjacent the inner surface of the side wall and to restrain a combustion-driven oscillation in the combustion volume.

Abstract: A fuel nozzle device suitable for use in a gas turbine engine or the like is provided. The fuel nozzle device includes a fuel line and a plurality of gas orifices disposed at a downstream end of the fuel line, the plurality of gas orifices operable for injecting fuel into an air stream. The acoustic resistance of each of the plurality of gas orifices is chosen to match the acoustic impedance of the fuel line such that the maximum acoustic energy may be transferred between the fuel nozzle device and the combustor, thus enhancing the ability of the fuel nozzle device to control the combustion dynamics of the gas turbine engine system. A fuel injection resonator assembly suitable for use in a gas turbine engine or the like is also provided. The fuel injection resonator assembly includes a plurality of orifices separated by a variable length tube.

Abstract: A combustion chamber suitable for a gas turbine engine is provided with at least one Helmholtz resonator having a resonator cavity and a damping tube in flow communication with the chamber interior. The damping tube is provided with at least one cooling hole extending through its wall for improved damping and cooling.

Abstract: A gas turbine engine combustor defining a combustion zone therein and being adapted for receiving compressed air from a compressor. The combustor comprises inner and outer walls spaced-apart by a predetermined spacing distance and defines a cavity therebetween. The outer wall has a first area defining at least an impingement aperture therein, the impingement aperture permits fluid flow communication between the compressor and the cavity. The inner wall has a second area corresponding to the first area and defines a plurality of effusion apertures therein, the effusion apertures permitting fluid flow communication between the cavity and the combustion zone, and each of the effusion aperture defines a cross-sectional area smaller than that of the impingement aperture. The effusion apertures are disposed in groups having a predetermined geometric arrangement relative to the impingement aperture and define a ratio of number of effusion apertures to the impingement aperture of between about 2:1 and about 4:1.

Abstract: A gas turbine engine having an exhaust casing that includes a plurality of fairings that shield struts that are affixed to and extend between concentric rings. Each fairing has a moveable tail section that is adjusted during turbine operation to change the angle of attack of the exhaust gases to minimize vortex shedding and the corresponding vibrations that result.

Abstract: Combustion driven pressure pulsations may limit the range of operating conditions where a modern gas turbine can operate with low emission and high efficiency performance. The control of acoustic vibrations has been consequently growing as an essential issue in gas turbine design, development and maintenance. The basic idea of the present invention is to use cooling air leakage through the gaps (7) of combustor liner segments (3;4) to get acoustic damping of combustion pulsations. The main component of the design is the sealing device (8) which covers the gap (7) between the two liner segments (3) and (4). The sealing device (8) collects in the plenum (9) some of the cooling air (6) flowing between the outer casing (1) and the inner liner structure (2). For this purpose the sealing device (8) is provided with openings (12) along the side walls (10) which allow the cooling air (6) to enter the plenum (9).

Abstract: A combustion control system controls a turbine engine that includes a fuel manifold and a plurality of fuel injectors. The control system includes a fuel pulsator and a controller. The fuel pulsator is coupled in flow communication with the plurality of injectors and the fuel manifold. The controller is coupled to the fuel pulsator such that the pulsator is between the controller and the fuel manifold. The controller is variably selectable and configured to facilitate promoting stable combustion.

Abstract: A method of controlling a gas turbine engine to provide protection against damaging pressure transients in the combustion process and to ensure compliance with emission requirements. Pressure fluctuations are monitored in a plurality of frequency ranges, and unacceptable pressure transients in different frequency ranges trigger different corrective actions. Unacceptable pressure transients in low and intermediate frequency ranges trigger a change in the pilot fuel fraction of a dual-mode combustor, while unacceptable pressure transients in a high frequency range trigger immediate power reduction in the engine. A control system for a gas turbine engine includes a plurality of timers for defining consecutive time periods for alternate monitoring of pressure transients and not monitoring pressure transients. Corrective action is taken only if unacceptable pressure transients are detected in each of the monitored time periods.

Abstract: A fuel nozzle device suitable for use in a gas turbine engine or the like is provided. The fuel nozzle device includes a fuel line and a plurality of gas orifices disposed at a downstream end of the fuel line, the plurality of gas orifices operable for injecting fuel into an air stream. The acoustic resistance of each of the plurality of gas orifices is chosen to match the acoustic impedance of the fuel line such that the maximum acoustic energy may be transferred between the fuel nozzle device and the combustor, thus enhancing the ability of the fuel nozzle device to control the combustion dynamics of the gas turbine engine system. A fuel injection resonator assembly suitable for use in a gas turbine engine or the like is also provided. The fuel injection resonator assembly includes a plurality of orifices separated by a variable length tube.

Abstract: Respective combustion gas streams are generated in a can combustor. The streams are channeled downstream into an annular turbine nozzle. And, dynamic interaction of circumferentially adjacent combustion gas streams is suppressed axially between the cans and the nozzle.

Abstract: A gas turbine combustor includes a side wall, for defining a combustion volume, having upstream and downstream ends, a pilot nozzle, disposed adjacent the upstream end of the side wall, for discharging a pilot fuel to form a diffusion flame in the combustion volume, and a plurality of main nozzles, provided around the pilot nozzles, for discharging a fuel-air mixture to form premixed flames in the combustion volume. Film air is supplied into the combustion volume downstream of the main nozzles along the inner surface of the side wall to reduce the fuel-air ratio in a region adjacent the inner surface of the side wall and to restrain a combustion-driven oscillation in the combustion volume.

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: Described are a method of and an appliance for suppressing flow eddies within a turbomachine, having a burner in which a fuel/air mixture is caused to ignite and in which hot gases are formed which leave the burner at the burner outlet and discharge into a combustion chamber, which follows the burner in the flow direction of the hot gases. The invention is characterized in that a mass flow is mixed into the hot gases directly at the location of the burner outlet.

Abstract: A gas turbine includes an annular combustion chamber and an outer wall of an annular combustion chamber. A straining ring is arranged on the outer wall of the annular combustion chamber and enables oscillations of the outer wall to be damped via friction. The effects of combustion oscillations produced by damaging vibrations of the annular combustion chamber are thus reduced. A method is further for damping an oscillation of an outer wall of an annular combustion chamber.

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: In a gas turbine control apparatus, a frequency analyzing section frequency-analyzes at least one of pressure oscillation in combustors of a gas turbine and acceleration oscillation of each of the combustors and outputs a first frequency analysis result as the result of frequency analysis for a plurality of predetermined frequency bands. A control unit controls at least one of a first fuel flow rate of fuel and a first air flow rate of air based on the first frequency analysis result for the plurality of frequency bands. The fuel and the air are supplied to the gas turbine.

Abstract: In a method for operating a premix burner a water quantity is introduced into at least one of the burner and the reaction zone of the burner depending on at least one command variable formed from at least one measured value. The method can also be used if a burner is operated dry, i.e., without water injection for nitrogen oxide injection, and the injected water quantity is less than 20% of the fuel quantity. The method can be used advantageously especially if a characteristic value derived from combustion pulsations or material temperatures is used as a command variable. The method enables measured parameters to be kept below a permissible upper limit.

Abstract: A combustion control system controls a turbine engine that includes a fuel manifold and a plurality of fuel injectors. The control system includes a fuel pulsator and a controller. The fuel pulsator is coupled in flow communication with the plurality of injectors and the fuel manifold. The controller is coupled to the fuel pulsator such that the pulsator is between the controller and the fuel manifold. The controller is variably selectable and configured to facilitate promoting stable combustion.

Abstract: For the purpose of reduced NOx gas emission, a gas turbine engine comprises a cylinder having a combustion region inside of the cylinder; a resonator having a cavity and provided around the surface of the cylinder and sound absorption holes formed on the cylinder and having opening ends on the cylinder.

Abstract: For the purpose of reduced NOx gas emission, a gas turbine engine comprises a cylinder having a combustion region inside of the cylinder; a resonator having a cavity and provided around the surface of the cylinder and sound absorption holes formed on the cylinder and having opening ends on the cylinder.

Abstract: A method and a device are described for the controlled damping of combustion-driven oscillations in a turbomachine with a burner system providing at least one burner, into which is introduced, via at least one burner nozzle arranged centrally in the burner, fuel which is intermixed with combustion inflow air flowing into the burner, to form a fuel/air mixture which is ignited in a combustion chamber following the burner system.

Abstract: A combustor can simultaneously reduce the amount of NOx exhaust and combustion oscillation. The combustor has an internal cylinder which accommodates a premixing nozzle and an external cylinder which accommodates the internal cylinder, and includes an air flow passage which supplies air from a compressor to the premixing nozzle. The air flow passage is provided with a punched metal plate near the maximum velocity fluctuation position whereat the velocity fluctuation of the air flow increases to a maximum.

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. In a preferred embodiment of the invention, a highest value of LCM(b0, a1+1), LCM(b0, a2+1), . . . LCM(b0, ak+1) is maximum, LCM designating the lowest common multiple. It thereby becomes possible to damp a maximum number of azimuthal vibration modes of the combustion chamber by means of a minimum number of modulatable burners. Each pair of modulatable burners gives rise to at least one undesirable vibration or instability which, however, is damped by the other modulatable burner or burners arranged according to the invention.

Abstract: A combustor structure of a gas turbine, in which a sheet-like vibration damper, which resonates with the air vibration in the intake chamber and absorbs the energy of the air vibration, is attached to the inner wall of the casing by an attaching member via a space.

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: In a combustion chamber arrangement, especially an annular combustion chamber arrangement for a gas turbine, one or more burners has, at its mouth, a deflecting device by which a combustion chamber arrangement is deflected. This achieves the effect of acoustic detuning, whereby the formation of a combustion oscillation is suppressed.

Abstract: An apparatus for use in a system having a gas fuel supply and a combustor includes a fuel actuator, a controller; and a premixer. The controller generates a command signal indicative of actuation sought from the fuel actuator. The actuator receives gas fuel from the liquid fuel supply and provides liquid fuel at a modulated rate of flow in response to the command signal. The premixer receives the fuel from the actuator, mixes the fuel with air, and provides the mixture to the combustor.

Abstract: This invention relates to air breathing engines, such as ramjets, scramjets and internal combustion, and more particularly to an active combustion control device for a combustor. In more particularity, the present invention relates to a method and apparatus that applies active combustion control technology to advanced propulsion devices and closed-loop fuel injection at sub-harmonic frequencies of the instability frequency of the combustor. The problem of limited actuator frequency response is addressed by injecting fuel pulses at sub-harmonic frequencies of the instability. The fuel may be liquid, solid or gas. To achieve this desired result, a closed loop controller is designed to determine sub-harmonic frequencies using a divider to divide the instability frequency of a combustor, yielding a fraction of the harmonic frequency.

Abstract: The invention relates to a damper arrangement for reducing combustion-chamber pulsations arising inside a gas turbine (1), having a combustion-chamber housing (8) which upstream comprises a front plate (2) with a plurality of individual burners (6) and damping elements (7, 7a, 7b) projecting through the front plate (2) and downstream is connected to a turbine stage (9) and is surrounded by a turbine housing (3) which comprises first openings (5a) which are adapted to the burners (6) and through which the burners (6) project upstream. The invention is characterized in that closable second openings (5b), through which it is possible to insert and tune the damping elements (7, 7a, 7b), are provided inside the turbine housing (9) adjacent to the first openings (5a) adapted to the burners (6).

Abstract: A combustion control system controls a turbine engine that includes a fuel manifold and a plurality of fuel injectors. The control system includes a fuel pulsator and a controller. The fuel pulsator is coupled in flow communication with the plurality of injectors and the fuel manifold. The controller is coupled to the fuel pulsator such that the pulsator is between the controller and the fuel manifold. The controller is variably selectable and configured to facilitate promoting stable combustion.

Abstract: A dampening device for suppressing vibrations of a tube assembly in a catalytic combustor which includes, a plurality of closely oriented, parallel tubes with each tube having at least one expanded region and at least one narrow region. The expanded regions being structured to contact at least one adjacent tube, thus providing support and minimizing degradation of the joint connecting the tubes to the tube sheet, and degradation of the tubes themselves. Such degradation can result from vibration due to flow of cooling air inside of the tubes, flow of the fuel/air mixture passing over the tubes transverse and longitudinal to the tube bundle, and/or other system/engine vibrations.

Abstract: The systems and methods of the invention offer a method for providing a corrective modulation signal to suppress an acoustic pressure wave in an operational system. The method includes sampling the acoustic pressure wave generated in the operational system and sampling a previously generated corrective modulation signal, the previously generated corrective modulation signal having parameters. The method further includes performing fast Fourier transform processing on the sampled acoustic pressure wave. A pair of single frequency discrete Fourier transform processing is performed on the sampled acoustic pressure wave. Also, the method includes determining the frequency, phase and magnitude of a dominate pressure wave in the acoustic pressure wave based on the fast Fourier transform processing and the discrete Fourier transform processing.

Abstract: The objective is to provide a gas turbine and the combustor thereof in which super high frequency combustion oscillation and the generation of NOx are reduced.

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 resonator module for a combustion turbine power plant, where the combustion turbine power plant defines a flow path. The resonator module includes a first member, an a second member. The first member has a size substantially smaller than the diameter of said flow path and a plurality of openings therethrough. The openings are in fluid communication with the flow path. The second member has a size generally equal to said first member. The second member is in a generally spaced relation to the first member and encloses a volume of gas between the first and second members.

Abstract: This invention relates to air breathing engines, such as ramjets, scramjets and internal combustion, and more particularly to an active combustion control device for a combustor. In more particularity, the present invention relates to a method and apparatus that applies active combustion control technology to advanced propulsion devices and closed-loop fuel injection at sub-harmonic frequencies of the instability frequency of the combustor. The problem of limited actuator frequency response is addressed by injecting fuel pulses at sub-harmonic frequencies of the instability. The fuel may be liquid, solid or gas. To achieve this desired result, a closed loop controller is designed to determine sub-harmonic frequencies using a divider to divide the instability frequency of a combustor, yielding a fraction of the harmonic frequency.

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 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 (76, 78, 80, 92). Secondary fuel injectors (106) and two secondary fuel manifolds (105A, 105B) supply fuel into different circumferential sectors, halves, of the secondary fuel and air mixing duct (80). The secondary fuel manifolds (105A, 105B) have secondary fuel valves (107A, 107B) which supply a greater proportion of fuel to the secondary fuel manifold (105A) than the secondary fuel manifold (105B) so that there is circumferential biasing of fuel in the secondary combustion zone (40). This circumferential biasing of fuel in the secondary combustion zone (40) reduces the generation of harmful pressure oscillations in the combustion chamber (28).

Abstract: In a combustion apparatus (10), in particular for driving gas turbines, in which combustion apparatus (10) a gaseous fuel in a burner (11) is sprayed through a plurality of separate fuel-injection devices (15, 16) into a gas flow containing combustion air, and the resulting mixture flows into a combustion chamber (12) for combustion and burns there, the acoustic impedance or stiffness of the fuel-injection devices (15, 16) is selected to be different in order to avoid thermoacoustic combustion instabilities.

Abstract: This invention relates to air breathing engines, such as ramjets, scramjets and internal combustion, and more particularly to an active combustion control device for a combustor. In more particularity, the present invention relates to a method and apparatus that applies active combustion control technology to advanced propulsion devices and closed-loop fuel injection at sub-harmonic frequencies of the instability frequency of the combustor. The problem of limited actuator frequency response is addressed by injecting fuel pulses at sub-harmonic frequencies of the instability. The fuel may be liquid, solid or gas. To achieve this desired result, a closed loop controller is designed to determine sub-harmonic frequencies using a divider to divide the instability frequency of a combustor, yielding a fraction of the harmonic frequency.

Abstract: In a method of controlling thermoacoustic vibrations in a combustion system having a combustion chamber and a burner, the fluid shear layer forming in the region of the burner is acoustically excited. An apparatus for controlling thermoacoustic vibrations in a combustion system having a combustion chamber and a burner is distinguished by the fact that device(s) for the acoustic excitation of the working gas are arranged in the region of the burner.

Abstract: In a method of suppressing or controlling thermoacoustic vibrations which develop in a combustion system having a burner working in a combustion chamber due to the formation of coherent or vortex structures and a periodic heat release associated therewith, in which method the vibrations are detected in a closed control loop and acoustic vibrations of a certain amplitude and phase are generated as a function of the detected vibrations and induced in the combustion system, improved suppression is achieved in that, within the control loop, the amplitude of the generated acoustic vibrations is selected to be proportional to the amplitude of the detected vibrations.

Abstract: In a method for minimizing thermoacoustic oscillations in gas turbine combustion chambers, an inert gas (N2, CO2, H2O) is additionally mixed with the stream of fuel.

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: The invention refers to a combustion chamber device including a combustion chamber (2) and a set of burners (6) with an outlet opening (7) in the combustion chamber. Each burner includes members (10) for the supply of a flow of oxygen-containing gas, members (8) for the supply of fuel, and a space (9), which extends to the combustion chamber (2) for mixing of the fuel and the oxygen-containing gas. Each burner (6) is arranged to convey a substantially equally large flow of oxygen-containing gas through the burner to the combustion chamber. Each burner (6) is defined by at least one parameter including a characteristic length (a) of the space (9) along the longitudinal direction (x). At least the characteristic length (a) of at least one of the burners is selected in such a way that it deviates from the corresponding length (a) of another burner and in such a way that the influence of the oscillations which arise during operation of the combustion chamber device are reduced.