Abstract: A system for providing air to a compressor of an auxiliary gas turbine engine of an aircraft. A system turbine has an inlet adapted to receive compressed air from the aircraft. A system compressor is mechanically coupled to the system turbine and has an inlet adapted to receive atmospheric air. The outlets of the system turbine and the system compressor are fluidly connectable to the inlet of the compressor of the auxiliary gas turbine engine. A method includes providing compressed air from an aircraft to an inlet of a system turbine mechanically coupled to a system compressor. The method includes providing atmospheric air to an inlet of the system compressor. The method includes providing the inlet of the compressor of an auxiliary gas turbine engine with air from an outlet of the system turbine and with air from an outlet of the system compressor.

Abstract: A system of attaching an injection system to a turbojet combustion chamber base. It comprises a deflector welded onto the chamber base. The deflector comprises an annular portion having an edge forming a retaining shoulder directed toward the front of the turbojet and the injection system comprises a collar on which is formed a retaining shoulder directed toward the rear of the turbojet and pressing against the retaining shoulder of the deflector.

Abstract: A method for operating a turbofan engine assembly is provided. The turbofan engine assembly includes a core cowl which circumscribes the core gas turbine engine, a nacelle positioned radially outward from the core cowl, a fan nozzle duct having an area defined between the core cowl and a portion of the nacelle, a first cowl and a second cowl that define a portion of the nacelle wherein the second cowl is repositionable with respect to the first cowl, and a thrust reverser assembly wherein the second cowl surrounds the thrust reverser assembly. The method includes varying an operating speed of the fan assembly from a first operating speed to a second operating speed. The method further includes selectively positioning the second cowl between a first operational position and a second operational position to vary the area of the fan nozzle duct to facilitate improving engine efficiency at the second operating speed.

Abstract: A combustor assembly for a gas turbine engine comprises a casing arrangement defining an aperture and a seal element mounted on the casing arrangement through which a component of the combustor can extend. The seal element comprises a relatively flexible force absorbing portion and a relatively rigid sealing portion. The force absorbing portion is provided between the sealing portion and the casing arrangement. The force absorbing portion is resiliently deformable to absorb force during movement of the casing arrangement. Such resilient deformation restricts transmission of the forces to the sealing portion. The seal element may be configured to carry a seal member.

Abstract: A combustor liner includes a liner having an upstream end and a downstream end having a longitudinal axis extending therethrough, and a plurality of cooling holes formed in the liner, the cooling holes are arranged along the longitudinal axis into a plurality of circumferentially extending rows that are variably spaced apart along the longitudinal axis.

Abstract: A power generation system includes a combustion turbine assembly (11) having a main compressor (12) receiving ambient inlet air, a main expansion turbine (14), a main combustor (16), and an electric generator (15) for generating electric power. An air storage (18?) has a volume for storing compressed air. A source of humidity (23, 34, 42) humidifies the compressed air that exits the air storage and thus provides humidified compressed air. A heat exchanger (24) receives a source of heat and the humidified compressed air so as to heat the humidified compressed air. A air expander (30) expands the heated, humidified compressed air to exhausted atmospheric pressure for producing additional power, and permits a portion of airflow expanded by the air expander to be injected into the combustion turbine assembly for power augmentation. An electric generator (31) is associated with the air expander for producing additional electrical power.

Abstract: An inner rotor shaft for use in a small twin spool gas turbine engine, the inner rotor shaft having a hollow middle section formed of a smaller diameter hollow section on a compressor end and a larger diameter hollow section on the turbine end of the shaft. Solid shaft end extend from the hollow section to form a forward solid shaft end to secure the fan rotor disk and an aft solid shaft end to secure the turbine rotor disk. A parabolic shaped transition section joins the forward shaft end to the smaller diameter hollow section, and a conical shaped transition section joins the aft shaft end to the larger diameter hollow section. A conical shaped transition piece joins the two hollow sections together to form an inner rotor shaft that can fit within a minimal space between the compressor rotor disk and the annular combustor assembly of the engine.

Abstract: A method facilitates operating a gas turbine engine. The method comprises supplying steam to a nozzle, supplying primary fuel to the nozzle, discharging the steam into a combustor from a plurality of circumferentially-spaced steam outlets defined in a tip of the nozzle, and discharging the primary fuel into the combustor from at least one outlet that is spaced circumferentially between the steam outlets.

Abstract: Apparatus including an exhaust nozzle assembly having a translatable structure operable to open and close a flow diverting port in an exhaust duct. The translatable structure includes a forward region partly defining a converging nozzle region. The translatable structure includes a rearward region partly defining a diverging nozzle region. The forward region and the rearward region join at an inner edge that partly defines a throat constriction. The translatable structure is translatable between a plurality of operational positions each associated with a longitudinal position of the throat constriction. In a first operational position, the flow diverting port is fully closed. In a second operational position, the flow diverting port is fully opened. In a first intermediate operational position, the flow diverting port is fully closed. In a second intermediate operational position, the flow diverting port is at least partly opened.

Abstract: To provide a gas turbine engine which is simple in structure but can effectively remove foreign matters from the combustion air. The intake passage 21 includes an inlet portion 29, a curved portion 30 and a reduced diameter portion 31. The bypass duct 24 is curved away from the central axial line CL in a region corresponding to the curved portion 30 and reduced diameter portion 31 of the intake passage 29. Between the intake passage 29 and bypass duct 24 is defined an annular space 32. In the reduced diameter portion 31, the outer liner 20 is formed with a large number of foreign matter introduction openings 33 in a circumferential arrangement for communication between the intake passage 21 and annular space 32. The inner casing 4 is formed with a plurality of foreign matter ejection holes 34 in a circumferential arrangement in a part thereof that curves outward for communication between the annular space 32 and bypass duct 24.

Abstract: A heat exchange system for use in lubricating systems for aircraft turbofan engine equipment in which a lubricant is provided under pressure to spaces bounded at least in part by surfaces moving relative to one another, the heat exchange system for providing air and lubricant heat exchanges to cool the lubricant at selectively variable rates in the engine fan airstreams. A heat exchanger core is provided in a controlled air flow duct system opening at its plural entrances to the engine fan airstreams and having its outlet end opening about at the end of the fan duct nozzle.

Abstract: A static structure is provided such as that used as an engine structure 30 within a gas turbine engine. A structure 30 comprises a hub or core 33 with a concentric ring or casing 32. The hub 33 and casing 32 are secured together through spoke or vane elements 31 which are sloped axially and tangentially relative to a principal axis X-X. In such circumstances any roll torque reaction caused by differentially rotating hub and ring motions is retained by the tangential inclination of the spokes or vanes whilst axial sloping of these spokes or vanes 31 and axially sloping of the struts 34 creates bracing in the direction of the principal axis X-X. In such circumstances essentially robust triangles are formed radially and longitudinally respectively each comprising spokes or vanes 31 and struts 34 secured at axially spaced locations on the hub or core 33.

Abstract: It is known that power is normally extracted within a gas turbine engine from a shaft in order to drive auxiliary devices such as electrical power generators. With high extraction rates from such shafts there is mismatching in rotational speeds between the shafts which must be adjusted through leakage taken from valves 8, 9. By measuring parameters and in particular pressure at a number of positions or parts within an engine arrangement 1 and comparing those parameters either through ratios or directly with reference values a controller can be configured in order to choose the correct sequence and scheduling of the valves 8, 9 opening to balance shaft speed. In such circumstances more efficient operation of the engine arrangement can be achieved over a wider range of engine speeds including idling.

Abstract: A method for altering the shape of an exhaust mixing structure extending from a downstream end of a jet engine nacelle nozzle. The method may involve attaching a first end of at least one actuator to the nacelle nozzle, and attaching a second end of the actuator to an inner skin of the exhaust mixing structure. At least one shape memory alloy (SMA) tendon connected to the actuator first and second ends may be activated such that the SMA tendon longitudinally constricts to move the inner skin of the exhaust mixing structure slidably along a longitudinal path. This causes a portion of the exhaust mixing structure to be flexed into a path of flow exiting the nacelle nozzle.

Abstract: There is provided an injector assembly having two or more oxidizer manifolds and/or two or more fuel manifolds for delivery of liquid propellants to a combustion chamber such that combustion instability is reduced or eliminated during throttling. Delivery of the oxidizer to the oxidizer manifolds is controlled by an oxidizer valve, which may comprise an integral valve. The oxidizer passes from the oxidizer manifolds into the oxidizer element and then into the combustion chamber. The multiple oxidizer manifolds allow the oxidizer to be provided through selective openings of the oxidizer element thus reducing the change in pressure drop across the oxidizer element to thereby reduce or eliminate combustion instability and other problems. Additionally, the injector assembly may also include a lift-off seal or a filler fluid source to fill any temporarily unused oxidizer manifolds with an oxidizer or filler fluid.

Abstract: Device for guiding an element in an orifice in a wall of a turbomachine combustion chamber, including a ring and a bush, the ring being traversed axially by the element and including an external annular rim guided transversely in an internal annular channel of the bush, the ring and the bush delimiting around the element a cylindrical annular passage opening into the chamber, air circulation being provided in the ring and/or in the bush and distributed about their axis so as to establish an air circulation in the cylindrical passage from the outside of the combustion chamber toward the inside.

Abstract: An apparatus and method converts a power generation combined cycle (CC) power plant to a load management compressed air energy storage (CAES) power plant. The CC power plant includes at least one combustion turbine, a heat recovery steam generator (HRSG) to receive exhaust heat from an associated combustion turbine, a steam turbine associated with the HRSG, and an electric generator associated with the steam turbine. An air storage stores compressed air. At least one compressor supplies the air storage with compressed air so that off peak energy can be converted to compressed air energy stored in the air storage. Compressed air from the storage is received by the HRSG and the HRSG provides heat to compressed air received from the air storage. The steam turbine receives heated compressed air from the HRSG and expands the heated compressed air to produce power.

Abstract: Unbalanced mass effects can occur at the low pressure compressor (5) of an aircraft engine. Stiffeners (20) are proposed to increase the stiffness of the assembly and therefore reduce possible cracks due to accidental unbalanced masses, one end (22) of the stiffener can be fixed to the engine casing (7) at the structural arm (10), and the other end (32) can be fixed to the compressor (5), preferably at the upstream face of the third guide vane (RD3). Advantageously, the stiffeners (20) are made of two parts fixed to each other; their manufacturing is optimized to limit the additional weight due to their presence.

Abstract: The present invention relates to a thrust orienting nozzle, shaped in such a way as to divide a principal propulsion gas flow coming from at least one gas generator into a first flow and a second flow for an ejection in a first half-nozzle and in a second half-nozzle and comprising at least one of following two piloting means: a means of dividing the principal flow into each of the two half-nozzles, and a means of orienting the thrust vector produced by each of the two half-nozzles. The invention applies in particular to the yaw control of an aircraft without a vertical tail unit.

Abstract: Disclosed is an abrupt opening apparatus of a pressurized chamber in an intended procedure to accomplish thrust termination for a stage separation of the solid rocket motor. It can open trust termination ports with simple mechanical structure to thereby terminate the normal thrust of a nozzle reliably, if it is intended to terminate the normal thrust produced from the nozzle of the rocket motor to thereby impede the forward movement of the rocket motor by producing the reverse thrust. The thrust termination device for the rocket motor comprises a closure 210 for closing the thrust termination ports in an air-tight structure; a snap ring 220 for supporting the closure 210; a wedge 230 arranged between both free ends of the snap ring; a wedge fixing plate 240 for fixing the wedge; and, a pulling wire 250 linked to the wedge at one end thereof for extracting the wedge by being drawn when the thrust termination is performed.