Abstract: A gas turbine engine includes a fan duct including a fan duct inner structure that surrounds a core engine, a fan case that surrounds a fan, a core engine frame, and at least one mechanism configured to secure a portion of the fan duct inner structure to a portion of the core engine frame. The at least one mechanism includes a castellated arcuate portion mounted to one of the fan duct inner structure and the core engine frame and an inwardly projecting retaining feature mounted to the other of the fan duct inner structure and the core engine frame. The castellated arcuate portion is rotatable about an engine central longitudinal axis to position a feature of the castellated arcuate portion proximate to a portion of the inwardly projecting retaining feature to latch the fan duct inner structure.

Abstract: The present disclosure relates to systems and methods that provide a low pressure liquid CO2 stream. In particular, the present disclosure provides systems and methods wherein a high pressure CO2 stream, such as a recycle CO2 stream from a power production process using predominately CO2 as a working fluid, can be divided such that a portion thereof can be expanded and used as a cooling stream in a heat exchanger to cool the remaining portion of the high pressure CO2 stream, which can then be expanded to form a low pressure CO2 stream, which may be in a mixed form with CO2 vapor. The systems and methods can be utilized to provide net CO2 from combustion in a liquid form that is easily transportable.

Abstract: A fuel delivery system is provided for a turbine engine. This fuel delivery system includes a variable orifice jet configured to spray fuel received from a spray bar fuel conduit. The variable orifice jet includes a flexible body and a pintle that extends through a sidewall of the flexible body. An area of an orifice between the sidewall and the pintle is variable.

Abstract: The present disclosure describes a micro gas turbine flameless heater, in which the heat is generated by burning fuel in a gas turbine engine, and the heater output air mixture is generated by transferring the heat in the gas turbine exhaust to the cold air drawn from the ambient environment. The present disclosure also describes component geometries and system layout for a gas turbine power generation unit that is designed for simple assembly, disassembly, and component replacement. The present disclosure also allows for quick removal of the rotating components of the gas turbine engine in order to reduce assembly and maintenance time. Furthermore, the present disclosure describes features that help to maintain safe operating temperatures for the bearings and structures of the gas turbine engine power turbine.

Abstract: A variable area fan nozzle actuation (“VAFN”) system is disclosed. The VAFN system may include an electrohydrostatic actuator (“EHA”) arranged to translate a VAFN panel relative to a translating sleeve. An electrical coupling may extend between a translating sleeve associated with the VAFN system and the fixed structure. The electrical coupling may be movable so that as the translating sleeve and fixed structure move relative to each other, power may be provided to the EHA by a wiring harness extending across the space between the translating sleeve and the fixed structure and connecting the EHA to an EHA power source.

Abstract: A combustor assembly (17) including guide vanes (44) located between an inner cylinder (24) and a flow sleeve (25). Each guide vane (44) includes a circumferentially angled flow directing portion (60) adjacent to a leading edge (46). The leading edge (46) of at least one guide vane (44) can be located radially inward along the longitudinal axis (54) relative to the leading edge (46) of at least one other of the guide vanes (44). The length of the guide vanes (44) may vary, and the circumferential spacing between a first pair of the guide vanes (44) can be different from a spacing between a second pair of the guide vanes (44).

Abstract: Burners having a fuel plenum in a base are disclosed. One disclosed example apparatus includes a base of a burner, the base comprising a fuel plenum and coupled to fuel nozzles, where at least one of the fuel nozzles is in fluid communication with the fuel plenum. The disclosed example apparatus also includes a burner head of the burner comprising nozzle passages in fluid communication with an airflow path, where the burner head defines a pilot combustion space that opens towards a flame tube of the burner, and is in fluid communication with the airflow path, and where each nozzle passage is to receive a fuel nozzle to provide fuel to entrain with air from the airflow path.

Abstract: A combustor assembly for a gas turbine engine is provided. The combustor assembly generally includes an annular dome and a liner. The liner at least partially defines a combustion chamber and includes the forward end received within a slot defined by the annular dome. Additionally, a heat shield is provided. The heat shield includes an end also received within the slot defined by the annular dome. A mounting assembly attaches the forward end of the liner and the end of the heat shield to the annular dome, such that the forward end of the liner and the end of the heat shield are co-mounted within the slot defined by the annular dome.

Abstract: An environmental control system includes a low pressure tap at a location on a first compressor section of a main compressor section. The low pressure tap communicates airflow to a first passage leading to a downstream outlet. A compressor is driven by an electric motor. A combined outlet intermixes airflow from the first passage and from the compressor driven by the electric motor and passes the airflow downstream to be delivered to an aircraft use.

Abstract: A system includes a control system. The control system includes a coherence derivation system configured to derive a coherence between respective outputs of each of a plurality of combustors coupled to a gas turbine system, and a phase derivation system configured to derive a phase difference between the respective outputs of each of the plurality of combustors coupled to the gas turbine system. The control system is configured to derive an indication of an impending lean blowout (LBO) or an actual LBO of at least one of the plurality of combustors based at least in part on the coherence derivation, the phase derivation, or a combination thereof.

Abstract: The present disclosure is directed to a gas turbine engine assembly having a compressor configured to increase pressure of incoming air, a combustion chamber, at least one turbine coupled to a generator, a torsional damper, and a controller. The combustion chamber is configured to receive a pressurized air stream from the compressor. Further, fuel is injected into the pressurized air in the combustion chamber and ignited so as to raise a temperature and energy level of the pressurized air. The turbine is operatively coupled to the combustion chamber so as to receive combustion products that flow from the combustion chamber. The generator is coupled to the turbine via a shaft. Thus, the torsional damper is configured to dampen torsional oscillations of the generator. Moreover, the controller is configured to provide additional damping control to the generator.

Abstract: A combustor includes: a pilot burner; a plurality of main burners, each of which is provided radially and outwardly along a circumferential direction about the pilot burner, and including a main nozzle disposed in a main burner cylinder; and an extension tube provided to extend toward a downstream side from the main burner cylinder of each of the main burners. The extension tube has a circular inlet communicating with the main burner cylinder, and an outlet at the downstream side comprising two radial edges parallel to the radial direction and two circumferential edges formed along the circumferential direction so as to connect both ends of the radial edges. The combustor further includes an air passage formed outside of the main burner cylinder; and an inner communication hole (formed at the side of the inlet of the extension tube and at a position corresponding to the radially and inwardly circumferential edge for allowing communication between the air passage and the inside of the extension tube.

Abstract: A fluid plenum including a body defining an internal cavity having an inlet and an outlet. The fluid plenum further includes at least one wall positioned in the internal cavity that divides the internal cavity into first and second passageways, and which also divides the inlet into first and second inlet portions, and divides the outlet into first and second outlet portions. The first passageway receives fluid through the first inlet portion and directs fluid to the first outlet portion, and the second passageway receives fluid through the second inlet portion and directs fluid to the second outlet portion. The first and second passageways extend first and second lengths that are different from one another, and also generate a substantially common back-pressure at the first and second inlet portions during flow of a fluid stream through the inlet, including a first sub-stream of the fluid stream through the first passageway and a second sub-stream of the fluid stream through the second passageway.

Abstract: A combined liner and panel for use in a gas turbine engine comprise a ceramic matrix composite panel having a plurality of extending members extending away from a face of ceramic matrix composite panel which will face hot products of combustion. A liner includes a plurality of spring members that apply a bias force biasing the extending members. The spring members bias the extending members on the panel in a direction away from the face of the panel which will face the hot products of combustion. A gas turbine engine, and a method of attaching a ceramic matrix composite panel to a liner are also disclosed.

Abstract: The present invention discloses a novel apparatus and methods for augmenting the power of a gas turbine engine, improving gas turbine engine operation, and reducing the response time necessary to meet changing demands of a power plant. Improvements in power augmentation and engine operation include additional heated compressed air injection from a power augmentation system and a motor-generator in selective operation with the power augmentation system.

Abstract: A gas turbine engine has a compressor section received within an inner housing. An is an outer housing is spaced radially outwardly of the inner core housing. A nacelle has an anti-icing system which taps compressed air from the compressor section through an anti-ice valve and to the nacelle. The anti-ice valve is opened at startup of the gas turbine engine to assist compressor stability.

Abstract: A fuel manifold apparatus includes an annular array of fuel nozzles interconnected by a plurality of manifold tubes, wherein each manifold tube includes at least two fuel passages integrally formed therein, the fuel passages being configured for conduction heat transfer therebetween.

Abstract: A multi stream aircraft fixed geometry nozzle includes an inner nozzle, an outer nozzle, and a supersonic ejector. The outer nozzle is configured to channel a third stream from an aft end of a third stream duct surrounding a bypass duct of a multi stream aircraft engine to the supersonic ejector to merge the third stream with the primary stream. The fixed geometry nozzle is configured to operate between an SFC mode and a thrust mode such that, when the inner nozzle accelerates the primary stream supersonically to the supersonic ejector, at which the primary stream is merged with the third stream, in the SFC mode the total pressure of the primary stream is substantially the same as the total pressure of the third stream, and in the thrust mode the total pressure of the primary stream is substantially greater than the total pressure of the third stream.

Abstract: An airflow control system for a gas turbine system according to an embodiment includes: a compressor component of a gas turbine system; a mixing area for receiving an exhaust gas stream produced by the gas turbine system; an air extraction system for extracting a supply of bypass air from an excess flow of air generated by the compressor component of the gas turbine system; an enclosure surrounding the gas turbine system and forming an air passage, the bypass air flowing through the air passage and around the gas turbine system into the mixing area to reduce a temperature of the exhaust gas stream; and an exhaust processing system for processing the reduced temperature exhaust gas stream.

Abstract: Systems and methods are disclosed herein for passively managing cooling air in a gas turbine engine. A cooling air supply line may supply cooling air to a component in the gas turbine engine. A metering coupon may have a negative coefficient of thermal expansion. The metering coupon may allow more airflow through the metering coupon and through the component in response to an increase in temperature.