Abstract: There is disclosed a double skin combustor for a gas turbine engine. The combustor has a radially inner liner and a radially outer liner defining therebetween an annular combustion chamber. The radially inner liner and the radially outer liners both have a hot skin and a cold skin defining a cooling cavity therebetween. The cold skin has a plurality of segments joined to one another by sliding joints. The cooling cavity between the hot skin and the cold skin may be compartmentalized into individual cavities between the sliding joints.

Abstract: A burner for a combustor includes (a) a swirl generator enclosing a burner interior on an inlet side and including a tangential air inlet relative to a longitudinal center axis; (b) a mixing chamber enclosing the burner interior on an outlet side and defining a burner outlet fluidly connecting the burner interior with a combustion chamber; and (c) a lance arranged coaxially with the longitudinal center axis. The lance introduces fuel through a nozzle tip at or near the burner outlet into the combustion chamber. The nozzle tip includes a cartridge defining a center fuel passage; fuel swirl vanes within the center fuel passage at an outlet end of the nozzle tip; a first tube surrounding the center fuel passage and defining a first fluid passage; a second tube surrounding the first tube and defining a second fluid passage; and air swirl vanes in the second fluid passage.

Abstract: Methods and apparatus for sealing variable area fan nozzles of jet engines are disclosed. An apparatus in accordance with the teachings of this disclosure includes a frame and a seal to be coupled to the frame. The seal to enclose petals of a variable area fan nozzle to substantially prevent airflow between the petals.

Abstract: A portable jig and fixture having at least one base plate, a column assembly removably coupled to the at least one base plate, a slide assembly removably coupled to the column assembly, a power feed head removably coupled to the slide assembly, and a spindle unit removably coupled with the slide assembly having a motor removably engaged it. At least one cutter assembly is removably engaged with the spindle unit and at least one drill jig is also removably engaged with the spindle unit. Also, a turbomachine having complementary structure on the compressor discharge casing can engage with the at least one base plate alignment mounting structure on the portable jig and fixture.

Abstract: A system may include a transition piece, coupled to a support of the system, to receive combustion products from a combustion chamber. The system may also include a turbine nozzle, coupled to a support of the system, to receive combustion products from the transition piece. Additionally, a web plate disposed at a radial position about a turbine axis between the transition piece and the turbine nozzle may form a first seal between the web plate and the transition piece and form a second seal between the web plate and the turbine nozzle. In addition, the web plate may extend in a circumferential direction about the turbine axis and couple to a support of the system. Further, the web plate may include an inner surface, an outer surface, and an arm extending in a radial direction between the inner surface and the outer surface.

Abstract: A combustion chamber structure, particularly for a rocket engine, which comprises a liner surrounding a combustion chamber with an outer surface facing away from the combustion chamber on which coolant channels extending in a longitudinal direction of the liner are formed. The liner forms, at one longitudinal end, a coolant chamber extending in the circumferential direction of the liner for collecting and/or distributing a coolant. The combustion chamber structure further comprises a connecting structure that fluidly connects the coolant chamber to the coolant channels.

Abstract: A gas turbine engine includes a fan, a compressor, a combustor, a turbine, a bypass duct, and a bearing compartment assembly. The bearing compartment assembly includes a fluid pump, a compartment, a fluid line between the fluid pump and the compartment, and a damper check valve located in the fluid line. The damper check valve is a unitary, monolithic component that is configured to restrict a reverse flow from the compartment to the fluid pump substantially more than the damper check valve restricts a standard flow from the fluid pump to the compartment.

Abstract: A disclosed micro gas turbine system includes a micro gas turbine apparatus and an extracting cycle apparatus. The micro gas turbine apparatus includes a first compressor, a burner, and a first turbine. The first turbine expands a combustion gas generated by the burner. The extracting cycle apparatus includes a second compressor and a second turbine. The second compressor receives a flow of extracted air that is generated by extracting a part of a working fluid discharged from the first compressor. The second turbine expands the working fluid discharged from the second compressor. The working fluid discharged from the second turbine cools down the first turbine.

Abstract: A bypass valve for a combustor, including: a housing; an opening in the housing; a pneumatically actuated component within the housing; a passage coupling the opening in the housing and a duct of the combustor; a chamber formed by the housing and the pneumatically actuated component, the chamber containing a volume of air; and a control valve for controlling a pressure within the chamber to selectively displace the pneumatically actuated component to open and close the passage.

Abstract: A flow modulator for a fuel nozzle, including: a fixed outer tube, the fixed outer tube including a set of apertures; a movable inner tube concentrically positioned within the fixed outer tube, the movable inner tube including a set of apertures; and a pneumatically actuated component for displacing the movable inner tube within the fixed outer tube to selectively align the set of apertures of the fixed outer tube with the set of apertures of the movable inner tube.

Abstract: A gas turbine engine comprises a compressor section, a combustor, and a turbine section. The turbine section includes a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap taps air having been compressed by the compressor, the tapped air being passed through a heat exchanger. A vane section has vanes downstream of the combustor, but upstream of the first stage blade, and the air downstream of the heat exchanger passes radially inwardly of the combustor, along an axial length of the combustor, and then radially outwardly through a hollow chamber in the vanes, and then across the blade outer air seal, to cool the blade outer air seal.

Abstract: A gas turbine engine defines a radial direction and an axial centerline. The gas turbine engine includes a core turbine engine that defines a core inlet. The core inlet is oriented with respect to the axial centerline and positioned along the radial direction such that the area available to capture foreign object debris is minimized. In one aspect, the gas turbine engine defines a capture ratio less than about 35%, wherein the capture ratio is a ratio of an area between a splitter radius and a tangency radius to an area encompassed by the splitter radius. The splitter radius is defined as a radial distance between the axial centerline and an outer lip of a splitter of the core turbine engine. The tangency radius is defined as a radial distance between the axial centerline and a tangency point, which can be defined at an inner lip of the core inlet.

Abstract: An apparatus and a method for detecting a threshold vibration condition associated with the operation of a gas turbine engine are disclosed. The apparatus includes a vibration-sensitive trigger mounted to the gas turbine engine for monitoring engine vibration and configured to disturb a sensor signal when a threshold vibration condition is met. The method includes 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 vibration 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.

Abstract: A combustor adapted for use in a gas turbine engine is disclosed. The combustor includes a metallic case forming a cavity and a ceramic liner arranged in the cavity of the metallic case. The ceramic liner defines a combustion chamber in which fuel is burned during operation of a gas turbine engine. The ceramic liner is located in the metallic case using a plurality of cross key connectors.

Abstract: An aspect includes a system for pre-start motoring synchronization for multiple engines of an aircraft. The system includes a first engine starting system of a first engine and a controller. The controller is operable to synchronize a motoring time of the first engine starting system with one or more other engine starting systems of one or more other engines of the aircraft by extending the motoring time of the first engine starting system to match, within a synchronization tolerance, the motoring time of the one or more other engine starting systems in a pre-start motoring sequence.

Abstract: A method of cooling a gas turbine engine case assembly includes moving a fan air valve that is operatively connected to a precooler having a bypass inlet that is configured to receive bypass air that bypasses a gas turbine engine core to facilitate a provision of bypass air through a fan air valve inlet to the bypass inlet to a first open position, in response to a core compartment temperature being greater than a target core compartment temperature. The method further includes bleeding the bypass air through a bypass outlet of the precooler into a core compartment.

Abstract: Gas turbine combustion systems and fuel cartridge assemblies are provided. An exemplary combustion system may comprise a combustor including one or more components, such as a cylindrical combustion liner, a flow sleeve, a main mixer, a radial inflow swirler, a combustor dome, and a fuel cartridge assembly. An exemplary fuel cartridge assembly may comprise first and second fuel manifolds which are connected to respective fuel circuits which supply fuel, such as liquid fuel, through a plurality of fuel passages within the fuel cartridge assembly or to other locations within an associated combustor. The fuel cartridge assembly may further include a plurality of fuel injector tips located at a tip plate of the fuel cartridge assembly through which fuel may be supplied to an associated combustor.

Abstract: An exhaust section of a gas turbine engine includes an exhaust plug defining a plurality of exhaust plug apertures circumferentially spaced from each other. Also included is an exhaust nozzle radially offset from the exhaust plug defining an exhaust pathway between the exhaust plug and the exhaust nozzle. Further included is an exhaust plug liner having a non-uniform outer surface axially aligned with the exhaust plug apertures. The exhaust plug liner is rotatable relative to the exhaust plug between a first position and a second position to selectively change a cross-sectional area of the exhaust pathway during thrust reversal operation to reduce an amount of reverse thrust necessary.

Abstract: Combustor panels including panel bodies with first and second sides, a pin array extending from the first side, wherein each pin extends a first height, has a pin diameter, and is separated from adjacent pins by a pin array separation distance. A structural protrusion extends from the first side. No pins of the pin array are located within a flashing distance that is equal to a protrusion separation distance plus half of the pin diameter, wherein a location of the pin is measured from a center point of the pin to a closest point on the exterior surface of the structural protrusion. At least one pin array extension is integrally formed with the structural protrusion, the pin array extension extending along the first side to a position that replaces a pin of the pin array that would be within the flashing distance.

Abstract: A gas turbine engine fuel system including a combustor with at least one fuel injector, a fuel manifold that supplies fuel to each injector, a drain tank in fluid communication with the fuel manifold, and a vacuum generator that selectively provides below atmospheric pressure to the drain tank and includes first and second chambers separated by a flexible diaphragm. The fuel system also includes a first check valve that allows one-way flow from the fuel manifold into the fuel injector, and a second check valve that allows one-way flow from the fuel injector to the fuel manifold. The vacuum generator reduces a pressure within the second chamber causing the diaphragm to increase a volume of the first chamber and reduce a volume of the second chamber, such that expanding the first chamber causes below atmospheric pressure in the drain tank and fuel manifold, and the second check valve opens.