38 results about "Gas turbines" patented technology

There is disclosed a gas turbine shroud structure comprising: a shroud support component 12 attached to an inner surface of a gas turbine casing 3; a shroud segment 14 divided in a peripheral direction and supported by the inner surface of the shroud support component; and a heat-resistant restricting spring 18 held between the shroud segment and shroud support component to urge the shroud segment inwards in a radial direction. The shroud segment 14 is formed of a ceramic composite material, and includes a coating layer 15 having heat insulation and impact absorption effect on an inner surface of the segment.

Noise reduction conduit for non rotary components of aircraft engines, subjected to a characteristic range of temperatures of a gas turbine engine, constituted by an annular structure composed of an aerodynamic wet wall (10a, 10b), perforated and resistant mechanically and thermally; of a dry wall (12a, 12b), not resistant and of light weight; and of some intermediate elements to which both walls are mechanically attached and that define a jump or difference of temperature between the wet and dry walls; between which wet (10a, 10b) and dry walls (12a, 12b) there are partitions that define cavities (16a, 16b) isolated with regard to each other.

A gas turbine airfoil (10) includes a pressure sidewall (12) and a suction sidewall (14) joined along respective leading (16) and trailing edges (18) and extends radially outward from a root (20) to a tip (22). The airfoil also includes a trailing edge corner (24) comprising a metering hole (36) receiving a cooling fluid flow (26) from an interior fluid flow channel (e.g., 32) and discharging a metered flow (40). A dispersion cavity (42) receives the metered flow and discharges a dispersed flow (44). The dispersion cavity includes a cross sectional area (46) greater than a cross sectional area (48) of the metering hole. An open flow channel (52) receives the dispersed flow and conducts the dispersed flow to a periphery (54) of the airfoil, the open flow channel controlling mixing of the cooling fluid flow with a process gas (e.g., 28) flowing around an exterior (34) of the airfoil.

The invention relates to gas turbine engine seal systems having a rotating member with an abrasive tip surface disposed in rub relationship to a stationary seal member with an abradable surface. The abrasive tip surface is coated with a metallic alloy matrix having ceramic abrasive particles embedded in and projecting from the matrix. The abradable seal surface is coated with a ceramic coating. The rub relationship affords a tight operating clearance between the rotating member and the stationary seal member, thereby improving engine efficiency, reducing fuel consumption and minimizing overhaul downtime.

A system, including, a turbine combustor, including, a first wall disposed about a flow path of hot combustion gases, a second wall disposed about the first wall, and a damping system disposed between the first and second walls, wherein the damping system is configured to dampen vibration, and the damping system is tuned to dynamic drivers in the turbine combustor.

A method for assembling a gas turbine engine is provided. The method comprises coupling a first turbine nozzle within the engine, coupling a second turbine nozzle circumferentially adjacent the first turbine nozzle such that a gap is defined between the first and second turbine nozzles and providing at least one spline seal including a substantially planar body. The method also comprises forming at least one catch to extend outward from the body portion of the at least one spline seal, and inserting the at least one spline seal into a slot defined in at least one of the first and second turbine nozzles to facilitate reducing leakage through said gap, such that a portion of the at least one spline seal is received within a recess defined within the turbine nozzle slot to facilitate retaining the at least one spline seal within the turbine nozzle slot.

A head part of an annular combustion chamber for a gas turbine, the head part comprising: an end wall with a passage opening for accommodating a burner, the end wall comprising a back side which faces towards the combustion chamber; a burner collar adapted to fit within the passage opening and receive the burner, the burner collar comprising a protruding portion radially protruding from an outer surface of the burner collar; and a ring member for receiving the protruding portion of the burner collar; wherein the head part of the annular combustion chamber is configured such that in an installed configuration a radially outer portion of the ring member engages with the end wall and a radially inner portion of the ring member receives the protruding portion of the burner collar.

A gas turbine engine system includes a gas turbine engine using vaporized liquefied gas as fuel. The gas turbine engine is configured to ingest intake air from an intake port and direct exhaust gases to an exhaust port. The gas turbine engine system also includes a heat exchanger with a heat transfer fluid therein. The heat exchanger is configured to create at least a part of the fuel by heating the liquefied gas. The gas turbine engine system also includes an intake air temperature control circuit. The intake air temperature control circuit includes conduits configured to direct the heat transfer fluid to the intake port and the exhaust port. The gas turbine engine system also includes a control system configured to selectively direct the heat transfer fluid to one of the intake port and the exhaust port.

A gas turbine engine including a component having a first and a second component that are attached or integrally formed with each other. The first and the second component define an enclosed void. An insert is positioned within the enclosed void which is formed of a material having a relatively low Young's modulus. The insert may prevent resin from filling the enclosed void during assembly/manufacture of the component, while also reducing a weight of the component and an amount of stress on the component. The component may be a fan platform positioned between a plurality of fan blades, an outlet guide vane, or another component of the gas turbine engine.

A fuel cooled cooling air heat exchanger includes a fuel injector and an airflow body. The fuel injector has a fuel flow passage formed therein that includes a fuel inlet port and a fuel outlet port. The airflow body is coupled to and surrounds at least a portion of the fuel injector. The airflow body has an inner surface that is spaced apart from the fuel injector to define an airflow passage between the airflow body and the fuel injector, and the airflow passage includes an air inlet port and an air outlet port.

At least one of the main injectors of the engine, forming a starter main injector, is fed directly by the pressurized feed pipe, while a head loss is imposed between the pressurized feed pipe and the other main injectors. Ignition is caused to take place at the starter main injector, and after ignition, said head loss imposed between the feed pipe and the other main injectors is eliminated so that all of the main injectors are fed with fuel at substantially the same pressure, without any imposed head loss.

A gas turbine engine has a fan drive turbine for driving a gear reduction. The gear reduction drives a fan rotor. A lubrication system supplies oil to the gear reduction, and includes a lubricant pump to supply an air/oil mixture to an inlet of a deaerator. The deaerator includes a separator for separating oil and air, delivering separated air to an air outlet, and delivering separated oil back into an oil tank. The separated oil is first delivered into a pipe outwardly of the oil tank, and then into a location beneath a minimum oil level in the tank. Air within the oil tank moves outwardly through an air exit into the deaerator. A method of designing a gas turbine engine is also disclosed.

A system and method are provided for effectively removing ice that may have formed on gas turbine engine compressor inlet guide vanes and/or preventing, or at least inhibiting, reformation of ice on gas turbine engine compressor inlet guide vanes after the ice has been removed. A determination is made as to whether actual inlet guide vane position differs from the commanded inlet guide vane position by a predetermined amount. If so, then the inlet guide vanes are repeatedly commanded to move in at least two predetermined directions to remove ice that may have formed on the inlet guide vanes.

A combustion apparatus includes a combustion chamber, a pre-chamber and a first device for mixing a fuel with an oxidant, the first device being located upstream of the pre-chamber. The combustion chamber and the pre-chamber share a common first and second wall spaced to each other building a cavity. The first wall exhibits at least one first opening in the area of the combustion chamber for introducing a coolant into the cavity. Further, the combustion apparatus has a channel for feeding the oxidant to the first device, whereby the channel is formed by a housing element of the combustion apparatus and the common first wall of the combustion chamber and the pre-chamber. The at least one first opening is located in the area of the combustion chamber. The first wall exhibits at least one second opening, located in the area of the pre-chamber, for introducing the coolant into the cavity.

Apparatus and methods for detecting a threshold condition associated with the operation of a gas turbine engine are disclosed. In one embodiment, the method comprises: generating one or more sensor signals associated with an operating parameter of the gas turbine engine and providing the one or more sensor signals to a controller of the gas turbine engine; disturbing the one or more sensor signals provided to the controller in response to the threshold condition being met; and detecting the disturbance in the one or more sensor signals provided to the controller and generating one or more output signals indicative of the detected disturbance.

In one example, a gas turbine engine includes an exhaust nozzle. The exhaust nozzle includes a flap supported relative to a static structure by a strut. The flap includes a backbone providing a slot. A slider interconnects a strut end to the backbone. In one example, the slider includes a body that is slidingly received within the slot. A boss extends from the body and provides a first feature. The strut end includes a second feature that cooperates with the first feature to prevent relative rotation between the slider and a portion of the strut end.

The combustor has a laser ignitor mounted to the casing, remotely from the liner of the combustion chamber. The laser ignitor has an igniter beam path for igniting the fuel and air mixture in the combustion chamber, the igniter beam path extending at least partially across the air plenum surrounding the liner and into the combustion chamber through a corresponding beam path aperture provided in the liner.

A rotor blade of a gas turbine engine includes a blade root defining a cooling airflow entry cavity therein, in fluid communication with internal cooling air passages through the blade. The cavity includes opposed side walls and at least one divider wall extending therebetween. At least one end portion of the divider walls adjoins one of the side walls in an angled direction relative to a perpendicular direction of the side walls.

A process for forming a thermal barrier coating system on a substrate is disclosed including preparing a slurry including a donor powder, an activator powder, and a binder. The donor powder includes a metallic aluminum alloy having a melting temperature higher than aluminum, and the binder includes at least one organic polymer gel. The process further includes applying the slurry to the substrate, heating the slurry to form an aluminide bond coating including an additive aluminide layer and an aluminide interdiffusion zone disposed between the substrate and the additive aluminide layer, and applying a thermal barrier coating to the aluminide bond coating. The thermal barrier coating may be a dense vertically-cracked thermal barrier coating, and the substrate may be a gas turbine component. Thermal barrier coating systems formed by the process are also disclosed.

The invention designs a water inlet and ice prevention detection system of a gas turbine. The water inlet and ice prevention detection system is used for improving safety and efficiency of a gas turbine device running under a low-temperature environment. A frozen state of an air inlet system is accurately monitored and judged by the gas turbine device by the proposed water inlet and ice preventiondetection system of the gas turbine, and the purposes of energy saving and efficiency improvement are achieved on the premise that an air inlet condition is enabled to conform to safety running of the gas turbine by controlling an air inlet heating device to start and close within reasonable time. The water inlet and ice prevention detection system has the characteristics of high frozen detectionaccuracy, rapid response, simple structure and arrangement flexibility, and the detection function demand of the air inlet and ice prevention detection system of a land gas turbine and a ship gas turbine can be satisfied.

Operation temperature of combustion turbine heat components, which is made of costly high-temperature alloy, reaches 1100-1400 DEG C, if having no protecting layer, the components are fast damaged by high-temperature oxidation and high-temperature corrosion. The invention relates to a novel high-temperature alloy coating with lasting oxidization resistant ability for prolonging lifetime of the components in operation environment of high-temperature oxidation and high-temperature corrosion and avoiding cracking problem. The other use of the high-temperature alloy is preventing ceramic heat insulation layer coated on surface of the combustion turbine heat components from desquamating with the alloy used as high-temperature oxidation and corrosion resistant bottom layer of the ceramic heat insulation layer.

The invention discloses a novel gas air film cooling structure for an aircraft engine and a turbine guide blade of a gas turbine. The cooling structure comprises a blade basal body, an air film gap, a connecting rib and a step surface, and is characterized in that a step-shaped gap air film flow outlet structure is formed by an inner blade and an outer blade on the blade pressure surface basal body and the connecting rib, so that cooling air flows out in the tangential direction of the blade surface. The cooling flow outlet structure of a novel blade can form a uniform and consistent cooling air film on the blade surface; the air outlet air film is higher in stability; the vortex effect of cold air flowing out is weakened; the cooling air film is large in covering area and high in cooling efficiency; and meanwhile, the cooling air in the tangential direction is mixed with a high-temperature main flow, so that the pneumatic loss of the blade is reduced, and the blade has both high cooling efficiency and high-efficiency pneumatic performance. The structure also can enable the blade to be produced by a piece connecting structure, and can simplify the blade processing and maintenance process.

The invention belongs to the technical field of combustion gas turbine equipment and particularly relates to a long-service-life composite bearing for a combustion gas turbine. The long-service-life composite bearing comprises a ball bearing and a thrust bearing. The thrust bearing comprises a thrust bearing bush and a thrust bearing disc, wherein the thrust bearing bush and the thrust bearing disc are mounted on a thrust bearing seat. The mean thickness of an oil film of the thrust bearing is adjusted through a first adjustment ring. The ball bearing is mounted on a ball bearing seat. The ball bearing is matched with a main elastic ring and a subsidiary elastic ring. The matching predeformation amount of the main elastic ring and the subsidiary elastic ring is adjusted and fixed through asecond adjustment ring and a check ring. The safety clearance between the thrust bearing disc and the ball bearing seat is set through a third adjustment ring. The ball bearing and the axial thrust bearing are fixed to the end face of a same flange. Compared with a common ball bearing, the maximum bearing capacity of the ball bearing can be set by designing the main elastic ring and the subsidiary elastic ring; and when the bearing load exceeds a set value, the remaining load can be borne by the thrust bearing, so that the service life of the ball bearing is greatly prolonged.

The present invention relates to a balancing body (1) for fastening to a ring (4, 5) of a continuous blade arrangement of a compressor or turbine stage of a gas turbine, wherein the balancing body has a first stop (11) for the form-fitting attachment of the balancing body in one peripheral direction (R) to a first axial shoulder (6; 7) of the ring.