Abstract: A fuel system for a turbine engine for improving efficiency in a fuel system where a major stage and secondary stage can be combined and held to a relatively constant fuel ratio while maintaining acceptable engine dynamics and NOx emissions is disclosed. The fuel system may be formed from a first premix injector assembly stage positioned upstream from a combustor basket, whereby the first premix injector assembly stage is a secondary injector system. The fuel system may be formed from a first primary injector assembly stage, which is a main injector system, positioned downstream from the first premix injector assembly stage. The first premix injector assembly stage and the first primary injector assembly stage may be coupled together such that the fuel system is capable of emitting fuel into a combustor of the turbine engine via the first premix injector assembly stage and the first primary injector assembly stage simultaneously.

Abstract: The invention relates to a fuel injector (101) for a turbine engine such as an airplane turboprop or turbojet, the injector comprising a body (102) including admission means (104) for admitting fuel under pressure, a stop valve (107) mounted in the body (102) downstream from the admission means (104) and designed to open at a first determined fuel pressure and to remain open beyond that first pressure in order to feed a primary fuel circuit (136), and a metering valve (115) mounted in the body (102) downstream from the stop valve (107) and designed to open above a second determined fuel pressure, greater than the first pressure, and to remain open above the second pressure in order to feed a secondary fuel circuit (117). The stop valve (107) and the metering valve (115) form a common movable assembly (133).

Abstract: The present application and the resultant patent provide a dual fuel combustor for a gas turbine engine. The combustor may include a primary premixer positioned within a head end plenum of the combustor, and a dual fuel, injection system positioned within the head end plenum and upstream of the premixer. The injection system may be configured to inject a gas fuel about an inlet end of the premixer when the combustor operates on the gas fuel. The injection system also may be configured to vaporize and inject a liquid fuel about the inlet end of the premixer when the combustor operates on the liquid fuel. The present application and the resultant patent also provide a related method of operating a dual fuel combustor.

Abstract: The invention refers to a sequential combustor arrangement including a first burner, a first combustion chamber, a dilution burner for admixing a dilution gas and a second fuel via a dilution-gas-fuel-admixer to the first combustor combustion products. The dilution-gas-fuel-admixer has at least one streamlined body which is arranged in the dilution burner for introducing the at least one second fuel into the dilution burner through at least one fuel nozzle, and 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 dilution burner. The streamlined body includes a dilution gas opening for admixing dilution gas into the first combustor combustion products upstream of the at least one fuel nozzle. The disclosure further refers to a method for operating a gas turbine with such a sequential combustor arrangement.

Abstract: A fuel injector for a gas turbine engine is provided. The fuel injector includes a central body, an air inlet duct, a mixing duct, a swirler, and a flow conditioner. The air inlet duct and the mixing duct are positioned around the central body to define an air flow passage. The swirler is positioned between the air inlet duct and the mixing duct. The flow conditioner is disposed in the air flow passage upstream with respect to the swirler. The flow conditioner has a perforated plate configured to uniformly distribute air circumferentially within the air inlet duct.

Abstract: An internal combustion engine in which the power output is controlled by modulating at least one of the compression ratio, expansion ratio, ratio of expansion rate to compression rate, air to fuel ratio, and steam to air ratio. Continuous isobaric catalytic combustion followed by isothermal expansion and the use of separate compressor and expander devices are used. Control dynamically maximizes fuel efficiency for the given power demand conditions. Power output is controlled by modulating flame temperature and/or pressure instead of by throttling. Lean combustion, high compression ratio, exhaust heat recuperation, and high power density and fuel economy are provided. External cooling is minimized or eliminated. Insulation of the engine effectively reduces energy losses to friction. Interchangeable use of gasoline, hydrogen and ammonia at high fuel efficiency is made possible for transitional periods of fuel availabilities. An injector suitable for isothermal expansion is provided.

Abstract: A fuel injector includes pilot and main fuel injectors. The pilot fuel injector includes at least one pilot air swirler and the main fuel injector includes a main air blast fuel injector located between inner main and outer main air swirlers. A first air splitter is located between the pilot and inner main air swirlers and a second air splitter is located between the pilot and inner main air swirlers. The first air splitter has a downstream portion converging to a downstream end. The second air splitter has a downstream portion diverging to a downstream end. The second air splitter downstream end is downstream of the first air splitter downstream end and the ratio of the distance from the first air splitter downstream end to the second air splitter downstream end to the diameter of the second air splitter downstream end is in the range of 0.22 to 0.30.

Abstract: A combustion chamber including a body with a fuel supply duct for supplying a fuel into the body and a carrier air supply duct for supplying air into the body. An adjusting system adjusts the carrier air mass flow supplied into the body according to the features of the fuel.

Abstract: A method is provided for the measurement of parameters of a gas present in a gas turbine combustion chamber. The method includes tuning a laser to a range containing the absorption lines of species to be analyzed in the gas, and directing the laser light through the combustion chamber and detecting laser light reflected off boundary walls of the combustion chamber. In order to analyze the absorption spectrum measured at high temperatures and pressures, a signature recognition algorithm is applied to the spectrum. The measured absorption spectrum is cross-correlated with a calibration absorption model spectrum for the absorption lines at several temperatures, pressures, and concentrations generated prior to the measurement. Values for pressure, temperature, and concentrations of selected species in the gas are determined simultaneously allowing direct control of the combustion chamber process. An apparatus for carrying out the method is also provided.

Abstract: A fuel supply system for a gas turbine includes a combustion section, a transition duct downstream from the combustion section, a turbine section downstream from the transition duct, and a first stage of stationary vanes circumferentially arranged inside the turbine section. A hot gas path is between the transition duct and the stationary vanes, and a fuel injector provides fluid communication into the hot gas path. A method of supplying fuel to a gas turbine includes combusting a first fuel in a combustion chamber to produce combustion gases, flowing the combustion gases through a transition duct to a hot gas path, and flowing the combustion gases through the hot gas path to a first stage of stationary vanes in a turbine section. The method further includes injecting a second fuel into the hot gas path downstream from the transition duct and upstream from the first stage of stationary vanes.

Abstract: A gas turbo set including a first turbine, a second turbine, and a combustion chamber connected between the first and second turbines and operated by auto-ignition is provided. The turbines and the combustion chamber are located on a common shaft that may be rotated about an axis. To increase the efficiency of the gas turbo set, the outer periphery of the second turbine is at a greater distance from the axis than that of the first turbine, leading to a reduction in the size and/or the number of turbine blades.

Abstract: A turbine arrangement is provided for use in a gas turbine engine that includes a combustion chamber and a nozzle. The turbine arrangement includes a source of liquid and a turbine blade assembly that is rotatable about a central shaft. The blade assembly further includes a plurality of turbine blades, the blade having a forward edge that faces the combustion chamber and an opposite rear edge. A hollow interior of at least one blade is in fluid communication with the source of liquid. The blade (e.g. a rear edge thereof) including a plurality of openings in communication with the hollow interior and sized to produce liquid droplets for discharge downstream of the turbine blade to generate a gas (e.g., steam) due to contact with hot gases generated by the combustion chamber.

Abstract: A gas turbine includes downstream side members that face a passage and have guide faces arranged in the passage side as it goes toward the upstream side. In a cross section including an axis center of a rotor shaft extending in a flow direction of a gas main stream, extension sections of the guide faces continuing to the upstream end portion of the guide faces are arranged downstream from the cavity inner wall portion of the upstream side member facing the cavities.

Abstract: A fuel injection device for an annular combustion chamber of a turbine engine, the device including a pilot circuit feeding an injector, a multipoint circuit feeding injection orifices formed in a front face of an annular chamber, and an annular ring mounted in the annular chamber and including fuel-passing orifices opening out into the injection orifices, each fuel-passing orifice formed in the annular ring including a zone of small section that is extended at least downstream or upstream by an orifice portion of increasing section.

Abstract: Provided is a fuel flow control method of a gas turbine combustor provided in a humid air gas turbine, by which method NOx generation in the gas turbine combustor is restricted before and after the starting of humidification and combustion stability is made excellent.

Abstract: A gas turbine combustor 1 of the present invention comprises a fuel injector 13 for injecting a fuel F toward a combustion chamber 11; a swirler 14 which takes-in compressed air CA generated in a compressor and swirl the compressed air CA, in the vicinity of the fuel injector 13; a tubular guide member 34 for guiding the compressed air CA taken-in from the swirler 14, to the combustion chamber 11; and a heat shield 23 having a cylindrical portion 23b located outward relative to the guide member 34; wherein the cylindrical portion 23b has a purge hole 40; and air is introduced through the purge hole 40 and is supplied to a space 39 formed between the guide member 34 and the cylindrical portion 23b.

Abstract: A power generation system is provided and includes a compressor to increase a pressure of inlet ambient air and a turbine downstream from the compressor having a first stage and additional stages downstream from the first stage, the first stage structurally incorporating a combustor in which the compressed ambient air is mixed with fuel and in which the mixture is combusted to generate high temperature exhaust gas from which mechanical energy is derived in the first and additional stages.

Abstract: A constant volume combustor device includes detonative combustion. In one form the wave rotor of the constant volume combustor is supported by magnetic bearings.

Abstract: A combined cycle power plant (10) that is optimized to reduce total life cycle cost rather than only thermal efficiency. A steam bottoming cycle (16) is fed with exhaust gas (14) from a gas turbine topping cycle (12). The topping cycle incorporates multiple stages of combustion in order to maintain a desired high temperature of the exhaust gas being delivered to the bottoming cycle while at the same time avoiding peak combustion temperatures that would produce undesirable levels of NOx emissions. Accordingly, no post-combustion exhaust gas treatment is needed. In one embodiment, an aspirated compressor (36) provides compressed air to a torus-configured combustion chamber (20) that is integrated with a first stage of vanes (22), and a plurality of stages of in-situ reheat combustion (26, 28) are utilized. In other embodiments, torus-configured combustion chambers may be utilized for more than one of the stages of combustion.

Abstract: A gas turbine output learning circuit is configured to compute a current combustion gas temperature TIT at an inlet of a gas turbine by linear interpolation by use of two characteristic curves A and B respectively representing relations between a pressure ratio and an exhaust gas temperature in the cases of the combustion gas temperature at the inlet of the gas turbine at 1400° C. and 1500° C., then to compute ideal MW corresponding to this combustion gas temperature TIT at the inlet of the gas turbine by linear interpolation according to 1400° C.MW and 1500° C.MW (temperature controlled MW), and then to correct the 1400° C.MW and the 1500° C.MW so as to match the ideal MW with a measured gas turbine output (a power generator output).

Abstract: A modular fuel nozzle configuration is defined which permits lower-cost manufacturing operations such as injection moulding to be employed. Also described is a method of making such a component.

Abstract: A gas turbine engine comprising a compressor, a combustion chamber, and at least two turbines mounted oppositely to the combustion chamber, such that the gas turbine engine is capable of generating multidirectional thrust.

Abstract: The present invention relates to a sealing body for sealing an outer component with respect to an inner component, in particular for a combustion chamber of a gas turbine. The outer component has a passage opening, through which the inner component projects out of the outer component. In the installed state, the sealing body is arranged in the region of the passage opening, surrounds the inner component in a circumferentially continuous manner, bears in a circumferentially continuous manner, by means of an inner sealing zone, against an outer sealing surface of the inner component and bears in a circumferentially continuous manner, by means of an outer sealing zone, against an inner sealing surface of the outer component. The sealing body is produced by deformation from at least one metal sheet that is in strip form and is circumferentially continuous. The inner sealing zone and the outer sealing zone are formed at a profiled section of the sealing body along the entire circumference.

Abstract: A secondary combustion system for a stage one turbine nozzle. The secondary combustion system may include a supply tube extending into the stage one nozzle, a number of injectors extending from the supply tube to an outer surface of the stage one nozzle, and an air gap surrounding each of the number of injectors.

Abstract: A gas turbine engine includes a combustor having a fuel nozzle assembly for delivering fuel to a combustion chamber. A sensor is at least partially housed within the fuel nozzle assembly for sensing a parameter associated with the delivery of fuel to the combustion chamber. In one example, the sensor is an optical sensor coupled to an optical decoder for ascertaining the parameter. A trim valve is in communication with the fuel nozzle assembly. A controller is in communication with the sensor and is programmed to actuate the trim valve to adjust the amount of fuel delivered through the fuel nozzle assembly in response to the parameter. In one example, an end of the optical sensor is arranged near a fuel exit of the fuel nozzle assembly. The controller actuates the trim valve for the fuel nozzle assembly to achieve a desired air/fuel ratio and/or balance the fuel delivery of the fuel nozzle assembly relative to another fuel nozzle assembly within the gas turbine engine's combustor system.

Abstract: A noise control cassette for a gas turbine engine includes a perforated face sheet configured for exposure to an airflow, a non-perforated backing sheet, a core arranged between the face sheet and the backing sheet and defining a cavity between the face sheet and the backing sheet having an effective length tuned so as to provide acoustic reactance control, and an attachment face for attaching the cassette to an airfoil-shaped structure.

Abstract: A fuel cell system comprising a plurality of fuel cells combined together to form a fuel cell stack, with the fuel cell stack having an anode side with an inlet for a fuel and an outlet for non-consumed fuel and exhaust gases arising at the anode side as well as a cathode side with an inlet for a gaseous oxidising agent such as air and an outlet for exhaust gases which arise at the cathode side, and a compressor having a compressor inlet and a compressor outlet, the compressor outlet of which is connected to the oxidising agent inlet of the fuel cell stack provided at the cathode side, is characterised in that a gas dynamic pressure wave machine having a driven rotor is provided, with an inlet for exhaust gases arising at the cathode side, an inlet for fresh oxidizing agent, an outlet for oxidising agent compressed by the cathode side exhaust gases and also an outlet for the cathode side exhaust gases being associated with the rotor and with the outlet for the oxidising agent compressed by the cathode side ex

Abstract: A gas turbine combuster ia provided with a dome heat shield having a fuel nozzle opening, the opening receiving a floating collar assembly for permitting relative movement between nozzle and heat shield. The floating collar is provided with a louver to provide film cooling to the face of the combustor heat shield and, thus, improve cooling thereof.

Abstract: A turbine nozzle (2) that, among components constructing a turbine, reaches particularly high temperature is efficiently cooled with a relatively simple structure. A double casing structure in which a turbine casing (7) is provided outside a turbine shell (5) is formed. The turbine casing (7) functions as a flow path (24) for compressed air (20?21) before combustion. The turbine shell (5) covers a turbine nozzle (2) and a radial turbine impeller (3) and forms flow paths (15, 16) for combustion gas (10?11?12?13). The compressed air (21) before combustion flowing in the flow path (24), for compressed air, having air-tightness between itself and the outside air is blown to the turbine nozzle (2) through a through-hole (51) penetrating both wall surfaces of the turbine shell (5). By this, the turbine nozzle (2) is cooled and the compressed air used to cool the turbine nozzle is made to flow toward the turbine impeller (3).

Abstract: The invention relates to a gas turbine (3) with reheat in the turbine section (11), whereby a fuel injection device (37) is provided for the injection of fuel (33) in to a combustion air flow (61) in the direction of flow before a cooling air flow (65) branching. By means of the early mixing of fuel (63) in the combustion air flow (61) and concomitant additional use of the cooling air flow (65) as combustion air, the specific output and the efficiency of the gas turbine (3) are improved.

Abstract: A dual fuel premix nozzle and method of operation for use in a gas turbine combustor is disclosed. The dual fuel premix nozzle utilizes a fin assembly comprising a plurality of radially extending fins for injection of gas fuel and compressed air in order to provide a more uniform injection pattern and homogeneous mixture. The premix fuel nozzle includes a plurality of coaxial passages, which provide gaseous fuel and compressed air to the fin assembly. When in liquid fuel operation, the gas circuits are purged with compressed air and liquid fuel and water pass through coaxial passages to the tip of the dual fuel premix fuel nozzle, where they inject liquid fuel and water into the secondary combustion chamber. An alternate embodiment includes an additional gas fuel injection region located along a conically tapered portion of the premixed fuel nozzle, downstream of the fin assembly.

Abstract: A combustion turbine with a secondary combustor is provided. The secondary combustor is disposed within the turbine assembly of the combustion turbine. The secondary combustor assembly is structured to maintain the working gas at a working temperature within the turbine assembly. The secondary combustor assembly re-heats the working gas by injecting a combustible gas into the working gas. The secondary combustor assembly includes a plurality of openings disposed among the vanes and/or blades in the turbine assembly which are coupled to a combustible gas source. As the combustible gas is injected into the working gas, the combustible gas will auto-ignite. That is, the combustible gas will combust without the need for an igniter or pre-existing flame. Combustion of the combustible gas in the turbine assembly re-heats the working gas.

Abstract: A gas turbo-machine and method of designing and constructing such machine includes preselecting specific operating conditions for the gas turbo-machine, and constructing a master stage as a model to have a given design and geometric shape which results in substantially the optimum efficiency during operation of the master stage at the preselected operating conditions.

Abstract: A gas turbine system in which substantially all of the compressed air from the compressor portion of the gas turbine is used to fluidize a bed of solid fuel, such as char, in a pressurized fluidized bed gasifier so as to produce a hot gas. The hot gas flows through a heat recovery steam generator so as to produce steam, which is expanded in a steam turbine so as to produce shaft power. A first portion of the hot gas from the heat recovery steam generator is expanded in the turbine section of the gas turbine so as to produce additional shaft power. A second portion of the hot gas from the heat recovery steam generator is then used to cool the turbine section of the gas turbine, after further cooling and filtering. Since none of the turbine cooling fluid is obtained by bleeding compressed air directly from the compressor, the mass flow of the hot gas flowing through the gasifier, and, therefore, the hot gas it produces is maximized, thereby maximizing steam generation in the heat recovery steam generator.

Abstract: A low-cost, lightweight mobile gas turbine capable of rotating in excess of ne hundred thousand RPM is coupled to a centrifugal air compressor which supplies high volume airflow to a diesel fueled combustion chamber for supplying hot gas to the turbine. Three energy outputs are obtained in the form of pressurized airflow, hot gas flow and large volumes of directed ambient air. The hardware components are essentially state-of-the-art, but their combination for use in multi-purpose military combat operations has not heretofore been known. Such uses include smoke generation, dissemination of infre-red and radar absorbant clouds, decontamination of large masses, hot water supply and air pressure source for mechanics tool operation.

Abstract: The process of increasing the efficiency and output of a gas-turbine having blades in a combustion chamber which includes cooling the gas-turbine blades with a liquid-fuel and mixing the fuel used for cooling with air in the combustion chamber and burning the mixture and driving the turbine by the exhaust from the combustion chamber. Preferably the fuel is a liquid-gas such as hydrogen. The turbine may be cooled by injecting liquid-fuel onto its disk to provide a film coating which is centrifuged around the blades whereby the liquid will cool both the interior and exterior of the blades and the absorbed heat is introduced into the combustion chamber. The combustion chamber includes a plurality of air inlets. The turbine drives a compressor and the compressor and the air inlet may be cooled with the liquid-fuel, which may be cryogenically cooled. An improved jet engine which is convertible from a turbo to a ramjet is provided.

Abstract: An expansion turbine is used to develop power, wherein combustion of fuel is carried out within the turbine. Increments of fuel are added to two or more points in the presence of a catalyst whereby temperature of the gas or equipment is controlled at substantially isothermal conditions.