Abstract: A gas turbine which can detect ignition in a combustor regardless of startup conditions of the gas turbine, such as the hot startup or the cold startup. An ignition detecting method for the gas turbine comprises the steps of calculating a difference between the exhaust temperature detected at a particular time before outputting of an ignition command for a combustor and the exhaust temperature detected after the outputting of the ignition command, and determining that the combustor is ignited, when the calculated difference is not less than a predetermined value. As an alternative, the method includes a step of determining that the combustor is ignited, when a change amount or rate of the exhaust temperature exceeds a predetermined value in a predetermined period from the outputting time of the ignition command.

Abstract: A gas turbine combustor includes a substantially cylindrical combustor liner located substantially concentrically within a flow sleeve, the combustor liner composed through a ceramic matrix composite material, a forward end of the combustor liner provided with a plurality of circumferentially arranged bolt holes. An inner metal ring is located about an outside surface of the forward end of the combustor liner, the inner metal ring provided with a second plurality of circumferentially spaced bolt holes, with a plurality of bolts extending through the first and second pluralities of bolt holes and secured by self-locking nuts. An outer metal ring is spaced radially outwardly of the inner metal ring, with a plurality of circumferentially spaced struts extending between the inner and outer rings.

Abstract: A generator assembly for a gas turbine engine. The assembly includes an electrical generator connected via a first sprag clutch to an accessory gearbox driven by the engine. The generator is also connected by a second sprag clutch to an air turbine connected to engine bleed valves which are selectively open when the engine speed falls below a predetermined level. The clutches are arranged such that the generator is driven at any time by the fastest rotating of the gearbox and turbine.

Abstract: A plasma thruster having a magnetic circuit of a downstream bottom plate from which arms protrude. At least one of the arms includes a permanent magnet. The mass, overall dimensions, electricity consumption, and cost of the thruster are thereby reduced.

Abstract: A combustion liner for a gas turbine combustion system is provided. The combustion liner comprises a combustion zone between an inlet end and an exhaust end. The combustion liner comprises a one-piece casting construction. The combustion liner is formed from a nickel-based superalloy having strength characteristics.

Abstract: A bleed valve for a gas turbine engine having a chamber inward of the bypass flow path and in fluid flow communication with the bypass and primary flow paths, a valve member for alternatively closing and opening the fluid flow communication with the primary flow path remaining inward of the bypass flow path, and an actuator actuating the valve member across the bypass flow path.

Abstract: The emergency fuel shutoff system for a gas turbine engine comprises a pyrotechnic device which allows the fuel circuit to be closed upon triggering of the pyrotechnic device.

Abstract: Exhaust nozzle 14 for an engine of a flying craft, comprising a tubular body with two outlet ducts 16, 18 and boxes 26, 28 for injecting gas into the outlet ducts for the purpose of thrust vectoring, these boxes having windows 38 aligned with slots 39 formed in the outlet ducts, and controlled means for adjusting the outflow of gas injected into the outlet ducts.

Abstract: In a control system for a gas turbine engine, a degradation level of the gas turbine engine is estimated according to a deviation of an output of an internal temperature sensor when the engine is producing a target output at a target rotational speed from a prescribed reference temperature, and a correction value for correcting the engine operation map is computed according to the degradation level of the gas turbine engine. Thereby, even when the degradation of the engine has progressed to a certain extent typically owing to the aging of various component parts, the engine operation map can be corrected in such a manner that the level of fuel supply is optimally selected and the acceleration, efficiency and other items of performance of the engine can be improved.

Abstract: A pulsed plasma thruster (PPT) and a method of making the PPT are disclosed. The PPT includes no moving parts and can be made in a small size. The PPT can achieve long operating duration by using vapor as a fuel. Liquid used to form the vapor can be easily stored and can provide an ample supply of vapor. The PPT is also designed to facilitate easy and rapid manufacture. The process for making the PPT uses known techniques for making printed circuit board devices.

Abstract: Apparatus for regulating the flow rate of fuel to a turboshaft engine in acceleration or deceleration, the engine having a free turbine and a core engine, the apparatus including sensors transmitting information to regulator elements, the information relating: to a first speed of rotation NTL of the free turbine; to a second speed of rotation Ng of the engine's gas generator; to an internal temperature T4 of the gas at the inlet to the free turbine; to the external pressure; and to the external temperature. The apparatus further includes control elements activated by the regulator to actuate a fuel metering system of the engine. The regulator evaluates the flow rate of fuel to be supplied to the engine in acceleration or deceleration on the basis of at least one optimum regulation relationship, which determines a main modulated flow rate as a function of a modulated speed of rotation.

Abstract: A pulse detonation engine comprises a primary air inlet; a primary air plenum located in fluid communication with the primary air inlet; a secondary air inlet; a secondary air plenum located in fluid communication with the secondary air inlet, wherein the secondary air plenum is substantially isolated from the primary air plenum; a pulse detonation combustor comprising a pulse detonation chamber, wherein the pulse detonation chamber is located downstream of and in fluid communication with the primary air plenum; a coaxial liner surrounding the pulse detonation combustor defining a cooling plenum, wherein the cooling plenum is in fluid communication with the secondary air plenum; an axial turbine assembly located downstream of and in fluid communication with the pulse detonation combustor and the cooling plenum; and a housing encasing the primary air plenum, the secondary air plenum, the pulse detonation combustor, the coaxial liner, and the axial turbine assembly.

Abstract: A gas turbine engine comprising a core engine surrounded by a core casing, at least one C-duct having radially inner and outer walls and a thrust reverser unit; the engine is mounted to an aircraft pylon via at least one thrust member attached to the core engine; the thrust reverser unit is mounted within the outer wall of the C-duct; the C-duct is secured to the core casing via a vee-groove attachment comprising vee-groove and vee-blade parts, the attachment is arranged to transfer loads between the thrust reverser unit and the pylon; wherein one cooperating part of the attachment is mounted to a torsion box mounted on the inner wall of the C-duct.

Abstract: A variable-section turbomachine nozzle comprises a plurality of moving flaps, and a moving flap control system comprising control levers each associated with a respective moving flap, the control levers comprising both controlled levers that are actuated directly by actuators, and follower levers that are actuated by the controlled levers via transmission links. When the nozzle is at rest, the flaps associated with the follower levers present a radial offset relative to the flaps associated with the controlled levers.

Abstract: A multiple accessory support bracket that mounts a plurality of external accessories for an auxiliary power unit (APU) so that all the external accessories may be mounted to the APU in a single operation.

Abstract: Carbon dioxide recirculating apparatus (20, 120) is disclosed for use in an arrangement having combination means (115) and a path for the flow of a gas through the combustion means (115). The apparatus (20, 120) comprises extraction means (221) for extracting carbon dioxide from a first region of the path downstream of the combustion means (115). It further includes condensing means (26, 30) for condensing the extracted carbon dioxide, and feed means (36, 136) for feeding the condensed carbon dioxide to a second region of the path upstream of the combustion means.

Abstract: A combustor assembly for a gas powered turbine includes a premix section to mix a first selected volume of fuel with a selected oxidizer. The premix section includes an injector plate that includes a porosity according to selected characteristics, such as pore size, pore density, pore distribution, and other selected characteristics. Therefore, the fuel may be provided through the porous plate to the premix area in a selected uniform flux.

Abstract: Methods and systems for a dry low NOx gas turbine engine system include a gas turbine engine including at least one dry low NOx combustor. The combustor includes a plurality of injection points wherein at least some of the injection points are configured to inject a fuel into the combustor at a plurality of different locations. The system includes a water source coupled to the combustor and operable to inject water into others of the plurality of injection points. The system also includes a control system that includes a sensor configured to measure an exhaust gas concentration of the turbine, a processor programmed to receive a signal indicative of the turbine exhaust gas concentration, and to automatically control the water injection using the received exhaust gas concentration signal. Such systems in use together can mitigate visible emissions from the exhaust stack.

Abstract: The invention relates to a regulation of the temperature of a hot gas of a gas turbine, particularly of a stationary gas turbine used for generating electricity, which comprises an injecting device for injecting a liquid into an airflow, which can be drawn into a compressor and with the aid of which a fuel inside a combustion chamber, which is situated downstream, combusts while producing the hot gas that is subsequently expanded when flowing through the turbine part located downstream. A temperature measuring device is provided that measures the temperature of the airflow before the compressor, whereby the temperature of the hot gas is regulated by the quantity of fuel. The aim of the invention is to provide a regulation with which, during wet compression operation, the serviceable life of the components subjected to the action of hot gas is lengthened.

Abstract: Control logic, and a method implemented by the control logic, compensates for mismatched inter turbine temperatures between two gas turbine engines and displays a biased inter turbine temperature (ITT) reading to the pilot without jeopardizing the need to be properly informed of the health of all engines and the relationship engine temperature compared to the maximum allowed temperature. A temperature compensation value is added to the lower inter turbine temperature to thereby increase the ITT to a value that is substantially equal to the higher ITT value. As a result, any potential nuisance or distraction that the flight crew may experience from displaying the mismatch is eliminated.