Abstract: A nozzle system includes a multitude of circumferentially distributed divergent seals that circumscribe an engine centerline. Each divergent seal includes an interface between a forward seal bridge bracket and an aft seal bridge bracket with a forward bridge support and an aft bridge support which provides axial and radial support for the divergent seals between adjacent divergent flaps. A divergent-convergent seal joint structure includes a horn and a fork. By having the forward bridge bracket retain the divergent seal in the axial direction, there is no need for axial sliding of the divergent seal relative to the divergent flap. The joint structure provides circumferential support as the axial and radial support are provided by the bridge-bracket interface.

Abstract: The subject application is directed to burners for gas turbine engines that use a pilot flame to assist in sustaining and stabilizing the combustion process. An embodiment of the disclosed burners includes, inter alia, a burner housing, a pilot combustor and a quencher. The burner housing has axially opposed upstream and downstream end portions. Additionally, the housing has at least one main fuel inlet passage and at least one main air inlet passage which are adapted to supply fuel and air respectively to an internal chamber defined in the housing. The pilot combustor is disposed along the axis of the burner housing and has an inlet for receiving a rich fuel and air mixture, a combustion chamber within which the rich fuel and air mixture is combusted into combustion products, and an outlet for exhausting the combustion products from the combustion chamber. The quencher is disposed within the internal chamber of the burner housing along the central axis and positioned at the outlet of the pilot combustor.

Abstract: Methods and apparatus for fast starting and loading a combined cycle power system are described. In one example embodiment, the method includes loading the gas turbine at up to it's maximum rate, and loading the steam turbine at its maximum rate with excess steam bypassed to the condenser while maintaining the temperature of steam supplied to the steam turbine at a substantially constant temperature from initial steam admission into the steam turbine until all steam generated by the heat recovery steam generator is being admitted to the steam turbine while the gas turbine operates at up to maximum load.

Abstract: The present invention is concerned with an annular combustion chamber of a gas turbine engine with an external annular wall and an internal annular wall, including an upstream fairing for separating the gas stream at the inlet of the chamber into a combustion stream and a bypass stream which bypasses the inlet of the chamber, the fairing including an annular wall forming a cap, which includes a downstream portion for fastening to a wall of the chamber and an upstream portion forming an edge of the flow cross section for the combustion stream, wherein the upstream portion is continued into at least one additional downstream portion for fastening to the wall of the chamber.

Abstract: A CAES system (10) includes an air storage (18), a compressor (20) supplying compressed air to the air storage, a power generating structure (11, 102), a heat exchanger (24), an auxiliary combustor (27), an air expander (30), and an electric generator (32). The system operates in one of modes a) a main power production mode wherein the auxiliary combustor is inoperable and the power generating structure is operable, to produce power by the air expander, fed by the heated compressed air received from the air storage, in addition to power produced by the power generating structure, or b) a synchronous reserve power mode wherein the auxiliary combustor is operable and the power generating structure is inoperable, with compressed air withdrawn from the air storage being preheated by the auxiliary combustor that feeds the air expander, with the air expander expanding the heated air and the generator providing immediate start-up power.

Abstract: A lean blowout protection system and method is provided that facilitates improved lean blowout protection while providing effective control of turbine engine speed. The lean blowout protection system and method selectively and gradually biases the lean blowout (LBO) schedule based on current engine data. This facilitates improved lean blowout protection while providing effective control of turbine engine speed and temperature.

Abstract: The engine comprises a low pressure stage (1) and a high pressure stage (2) with two independent rotors (3,4) and associated fixed fluid preparation members (9,10,13,14), combining “compressor” means (15,16) and “expander” means (17,18) and at least one combustion chamber (19,20) in a central volume defined axially by the rotors (3,4) and laterally by three substantially cylindrical, coaxial walls (22,23,24), defining in pairs an outer duct (25) connecting the outlet of the LP compressor (15) to the inlet of the HP compressor (16) and an inner duct (26) connecting the outlet of the HP compressor (16) to the inlet of a first combustion chamber (19), whose outlet supplies the HP expander (18) with gases produced by combustion guided to the inlet of the LP expander (17), whose outlet is in communication with the outside via at least one exhaust opening (41) for the combustion gases.

Abstract: A combustion control device for a gas turbine is capable of determining respective fuel flow rate control valve position command values based on fuel ratios. The combustion control device for a gas turbine is configured to calculate fuel flow rate command values for respective types of fuel gas based on a total fuel flow rate command value and fuel gas ratios, to calculate fuel gas flow rates based on the fuel flow rate command values and on a function of the fuel flow rate command value and the fuel gas flow rate, to calculate Cv values of the respective fuel flow rate control valves in accordance with a Cv value calculation formula based on the fuel gas flow rates, a fuel gas temperature, and front and back pressures of the respective fuel flow rate control valves, and to calculate respective fuel flow rate control valve position command values based on the Cv values and on a function between the Cv value and apertures of the fuel flow rate control valves.

Abstract: A combustion control device for a gas turbine is capable of controlling fuel ratios such as a pilot ratio and controlling an aperture of a combustor bypass valve in response to a combustion gas temperature at an inlet of the gas turbine in accordance with the original concept by computing a combustion load command value which is proportional to the combustion gas temperature. The combustion control device is configured to calculate a 700° C. MW value and a 1500° C.

Abstract: A fuel delivery system (10) comprises a variable displacement pump (20) for pressurizing fuel. The variable displacement pump has a distinct first pump displacement setting for a first desired mass flow of fuel and a distinct second pump displacement setting for a second desired mass flow of fuel. The variable displacement pump operates in only one of the first and second pump displacement settings. A fuel control (22) with a metering valve downstream of the variable displacement pump selectively regulates fuel delivery.

Abstract: A method for reducing the nonsteady side loads acting on a nozzle of a rocket engine during a startup phase of said engine. The nozzle comprises a combustion chamber (1) where exhaust gases are generated, a divergent portion (3) in which a supersonic flow of said exhaust gases occurs, and a throat (2) connecting the combustion chamber to the divergent portion, which method comprises the positioning of a body of rounded shape (5) inside the divergent portion (3) along its axis corresponding to an axial position such that, during at least part of the startup phase, a shock wave (8), induced by the distrubance of the flow of the exhaust gases by the body of rounded shape (5) is incident to the wall of the divergent portion (3) at an axial incidence position where it produces a jet separation or a separation in the form of a toroidal separation bulb.

Abstract: A heat engine includes a turbine (10), compressor (20), heat pump (30) and combustion source (40) and uses heated and compressed air as a motive fluid in an open Brayton cycle. In the process of the invention, the compressor pressurizes air (21) from the environment. A closed-cycle heat pump (30), having a high-temperature condensation side (33) and a low-temperature evaporation side (35), increases the energy in the compressed air through heat exchange with the high-temperature condensation side (33) of the heat pump. Upon heating in a constant pressure process, the motive fluid is expanded through the turbine (10). The combustion source (40), which is isolated from the motive fluid, further heats the low-temperature evaporation side (35) of the heat pump.

Abstract: An LNG (liquefied natural gas) power plant has a heating/cooling device. The heating/cooling device is controlled by a command generated by a temperature control device. The temperature control device receives a heating value and a temperature of the fuel gas. The heating value of the fuel gas introduced to a gas turbine power generation unit from an LNG vaporization facility is detected by a heating value detector. The temperature of the fuel gas is detected by a temperature detector. A target temperature calculator installed in the temperature control device calculates a target temperature based on the heating value. A command generator installed in the temperature control device generates the command by comparing the target temperature and the fuel temperature.

Abstract: A gas turbine load control device includes load setting means, first bias setting means, second bias setting means, target output setting means, and so forth. The target output setting means sets up a target output by adding a positive side bias value to a load set value when the load setting means gradually increases the load set value in response to an increase in a requested load set value inputted from requested load setting means. The target output setting means sets up the target output by subtracting a negative side bias value from the load set value when the load setting means gradually decreases the load set value in response to a decrease in the requested load set value.

Abstract: A method of detecting onset of a gas turbine condition, such as compressor stall, includes receiving data indicative of an operating parameter of a compressor of the gas turbine. The method also includes performing a wavelet transformation on the data to generate wavelet transformed data. The wavelet transformation is configured to affect a processing characteristic regarding a performance of the wavelet transformation. Features indicative of onset of the gas turbine condition in the wavelet transformed data are then identified to provide an indication for controlling the gas turbine to prevent compressor stall from occurring. A system for detecting onset of compressor stall in a gas turbine includes a sensor for providing data indicative of an operating parameter of the compressor and a processor for performing a wavelet transform on the data to identify features of the optimized wavelet transformed data indicative of onset of stall.

Abstract: Many variables in processes such as those using turbocompressors and turbines must be limited or constrained. Limit control loops are provided for the purpose of limiting these variables. By using a combination of closed loop and open loop limit control schemes, excursions into unfavorable operation can be more effectively avoided. Transition between open loop and closed loop may be enhanced by testing the direction and magnitude of the rate at which the limit variable is changing. If the rate of change indicates recovery is imminent, control is passed back to the closed loop limit control function.

Abstract: A gas turbine which can be easily employed in an area where it is hard to obtain a sufficient amount of water, such as an isolated island. Heated and pressurized heavy oil and water in a supercritical state are mixed with each other in a modifying unit to produce fuel-purpose modified oil. The fuel-purpose modified oil is depressurized by a depressurizing valve. Due to a temperature fall caused by adiabatic expansion with the depressurization, the fuel-purpose modified oil is brought into a two-phase state where moisture is in a gas phase (steam) and modified oil is in a liquid phase. The fuel-purpose modified oil is separated into the steam and the modified oil by a gas-liquid separator. The separated steam is condensed to water in a condenser and returned to a water supply line. The modified oil in the liquid phase is supplied to a combustor, thereby driving a gas turbine.

Abstract: A fuel control system for a gas turbine engine includes a metering valve for metering a flow of fuel, a throttling valve for maintaining a pressure drop across the metering valve, the throttling valve being shiftable between an open state and a shutoff state blocking an outlet of the metering valve, a primary control pressure supply for supplying a primary control pressure to the metering valve for controlling the position of the metering valve, a backup control pressure supply, a transfer valve shiftable between a first position connecting the primary control pressure to supply to the metering valve and a second position connecting the backup control pressure to the metering valve, and an electrohydraulic servovalve (EHSV) operably connected to the throttling valve and the transfer valve and controlling the state of the transfer valve and the position of the throttling valve. Also a method for controlling a transfer valve and a throttling valve.

Abstract: A fuel system for a gas turbine engine having at least one augmentor zone which fuel system includes a fuel pump (12) drawing fuel from a fuel tank (10) and pumping the fuel in a downstream direction, at least one fuel line (14) connecting the fuel pump (12) to the at least one augmentor zone, at least one augmentor metering valve (20) in the at least one augmentor fuel line (14) for metering fuel to the at least one augmentor zone, and a fuel return path (32) returning fuel from a point in the fuel line (14) downstream of the fuel pump (12) to the fuel tank (10), wherein the point is downstream of the at least one augmentor metering valve (20). A selector valve (30) is provided to allow thermal recirculation of fuel using an augmentor metering valve. Also a method of controlling fuel recirculation in a gas turbine engine.

Abstract: A method for operating a gas turbine set comprises the compression of an air mass flow in a compressor, feeding the compressed air mass flow into a first combustion chamber, burning the first mass fuel flow in the first combustion chamber in the compressed mass air flow, expanding the resulting hot gas in the first turbine, feeding the partially expanded hot gas into a second combustion chamber, and burning a second mass fuel flow in the second combustion chamber in the partially expanded hot gas. The method furthermore comprises supplying to the gas turbine set the first mass fuel flow and the second mass fuel flow preferably in a common supply line with a fuel supply pressure, and stipulating division of the total mass flow into a first and second mass fuel flow according to a normal operation concept. When the boundary value of the fuel supply pressure for example in the preferably common supply line is not reached, there is a deviation from the division according to the normal operating concept.