Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A flowpath is defined between a rotating surface and a non-rotating surface. The flowpath is connected downstream of the heat exchanger and is configured to deliver air to at least one of the plurality of rotating components. At least a portion of the non-rotating surface and the rotating surface includes a base metal. An insulation material is disposed on a surface along the flowpath.

Abstract: Provided for a device and a method for controlling a gas turbine, a gas turbine control program, and the gas turbine are: a first opening setting unit that sets a first IGV opening command value for an inlet guide vane based on factors excluding a surge limit of a compressor; a second opening setting unit that sets a second IGV opening command value for the inlet guide vane based on the surge limit of the compressor, and corrects the second IGV opening command value based on a pressure in the compressor; an opening selecting unit that selects a maximum opening as an IGV opening command value for the inlet guide vane among the first IGV opening command value and the second IGV opening command value; and an opening control unit that adjusts the opening of the inlet guide vane based on the IGV opening command value S selected by the opening selecting unit.

Abstract: A swirler for creating a swirling fuel/air mix having vanes extending radially around a central axis of the swirler and positioned on an annular base. The swirler has mixing channels for mixing the fuel and the air. At least one mixing channel is defined by opposing walls of two adjacent vanes. At least one of the opposing walls includes a primary side injection opening for ejecting a stream of fuel into the mixing channel; and a secondary side injection opening which corresponds to the primary side injection opening and is for ejecting a jet of fuel into the mixing channel. The secondary side injection opening is positioned on the wall such that the jet from the secondary side injection opening creates turbulence, within the mixing channel, in the stream from the primary side injection opening. The turbulence increases mixing of fuel and air.

Abstract: A gas turbine fuel flow rate setting device includes a flow rate ratio computing unit and a flow rate computing unit. The flow rate ratio computing unit is configured to accept a parameter having a correlation with a gas turbine output, and use a predetermined flow rate ratio relationship between the parameter and a fuel flow rate ratio between a plurality of premixing nozzle groups to determine a fuel flow rate ratio corresponding to the parameter which has been accepted. The flow rate computing unit is configured to use the fuel flow rate ratio determined by the flow rate ratio computing unit to determine a flow rate of a fuel to be supplied to each of the plurality of premixing nozzle groups.

Abstract: A gas turbine engine component assembly is provided. The gas turbine engine component assembly, comprising: a first component having a first surface, a second surface opposite the first surface, and a cooling hole extending from the second surface to the first surface through the first component; a second component having a first surface and a second surface, the first surface of the first component and the second surface of the second component defining a cooling channel therebetween in fluid communication with the cooling hole for cooling the second surface of the second component; and a particulate capture device attached to at least one of the first component and the second component, the particulate capture device configured to aerodynamically separate the airflow from the particulate.

Abstract: An apparatus, system and method for generating ionosonic lift is provided. First and second control electrodes are attached in a spaced-apart relation proximal to a top surface of a wing. An alternating potential applied to the first and second control electrodes alternately attracts and repels ions of a fluid in which the wing is immersed, causing a reciprocating flow of the ions of the fluid and of neutral molecules between the first and second control electrodes. The reciprocating flow of the ions and neutral molecules causes a reduction in pressure at the top surface of the wing, resulting in net lift applied to the wing under the Bernoulli relation.

Abstract: A spacecraft propulsion system comprises two thrusters, each operating in accordance to a corresponding propulsion technique. A controller is configured to direct collected solar energy to heat a propellant for consumption in one of the two thrusters, or to generate electric energy for the other one of the two thrusters.

Abstract: A mobile fracking system and methods may include a gas turbine housed at least partially inside a trailer and an exhaust attenuation system configured to receive exhaust gas from the gas turbine. The exhaust attenuation system may include a lower elongated plenum configured to receive exhaust gas from the gas turbine and an upper noise attenuation system that is movably connected relative to a distal end of the lower elongated plenum.

Abstract: A system and method for operating a combustion system comprising a fuel nozzle defining at least one main fuel circuit and at least one pilot fuel circuit is generally provided. The method includes determining an overall flow of fuel, the overall flow of fuel defining a sum total fuel through the main fuel circuit and the pilot fuel circuit; determining a plurality of ranges of ratios of main fuel flow through the main fuel circuit versus pilot fuel flow through the pilot circuit from the overall flow of fuel, wherein each range of ratios is based on a combustion criterion different from one another; determining a resultant range of ratios of main fuel flow versus pilot fuel flow based on a hierarchy of combustion criteria, wherein the hierarchy of combustion criteria provides a priority ranking of the combustion criterion; and flowing the overall flow of fuel to the main fuel circuit and the pilot fuel circuit based on the resultant range of ratios of main fuel flow versus pilot fuel flow.

Abstract: A propulsion system coupled to a vehicle. The system includes a diffusing structure and a conduit portion configured to introduce to the diffusing structure through a passage a primary fluid produced by the vehicle. The passage is defined by a wall, and the diffusing structure comprises a terminal end configured to provide egress from the system for the introduced primary fluid. A constricting element is disposed adjacent the wall. An actuating apparatus is coupled to the constricting element and is configured to urge the constricting element toward the wall, thereby reducing the cross-sectional area of the passage.

Abstract: A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

Abstract: A gas turbine engine component assembly comprising: a first component having a first surface and a second surface opposite the first surface, wherein the first component includes a cooling hole extending from the second surface to the first surface; a second component having a first surface and a second surface, the first surface of the first component and the second surface of the second component defining a cooling channel therebetween; and a lateral flow injection feature integrally formed in the first component and fluidly connecting a flow path located proximate to the second surface of first component to the cooling channel, the lateral flow injection feature being configured to direct airflow from the airflow path through a passageway and into the cooling channel at least partially in a lateral direction parallel to the second surface of the second component such that a cross flow is generated in the cooling channel.

Abstract: An assembly is provided for a gas turbine engine. This assembly includes a combustor wall, a case and an inlet. The combustor wall is configured with a quench aperture. The case is displaced from the combustor wall such that a plenum is formed by and extends between the combustor wall and the case. The case includes a case wall, a deflector and a conduit. The deflector projects out from the case wall into the plenum towards the combustor wall. The deflector is arranged upstream of the quench aperture. The inlet to the conduit is arranged next to and downstream of the deflector.

Abstract: An integrated combustor nozzle includes a combustion liner that extends radially between an inner liner segment and an outer liner segment. The combustion liner includes a forward end portion, an aft end portion, a first side wall, and a second side wall. The aft end portion of the combustion liner defines a turbine nozzle. The integrated combustor nozzle further includes an impingement panel having an impingement plate disposed along an exterior surface of one of the inner liner segment or the outer liner segment. The impingement plate defines a plurality of impingement holes that direct coolant in discrete jets towards the exterior surface of the inner liner segment or the outer liner segment. The impingement panel is radially spaced from the exterior surface to form a cooling flow gap therebetween. The impingement panel includes a collection duct that extends from the impingement panel and defines a collection passage.

Abstract: A turbomachine includes a compressor and a turbine with a burner and a combustor between the compressor and the turbine. The burner is downstream of the compressor and upstream of the turbine. The burner is connected to the combustor at a front panel of the burner. The front panel includes a frame, a rim extending around a central aperture within the frame, and a seal segment. The frame, the rim, and the seal segment are all integrally formed as a single unitary body.

Abstract: A pilot burner assembly for a combustion volume in a gas turbine engine includes a pilot burner, a pilot-fuel supply line, and a pilot-air supply line. The pilot burner has a burner face that includes a plurality of pilot-fuel injection holes. The pilot-fuel injection holes provide a pilot-fuel to the combustion volume. The pilot-fuel supply line is fluidly connected to the pilot-fuel injection holes for supplying the pilot-fuel to the pilot-fuel injection holes. The pilot-air supply line provides a pilot-air to the pilot burner. The pilot-air is supplied to the combustion volume through the burner face. Pilot-air injection holes are located on the burner face and fluidly connected to the pilot-air supply line. The pilot-air injection holes inject the pilot-air into the combustion volume. A gas turbine has the pilot burner assembly.

Abstract: Electrical power generation system for an aircraft, including: a gas turbine having a high-pressure shaft and outputting a gas flow to actuate a free turbine, a polyphase permanent magnet generator coupled to the gas turbine and able to output a first alternating voltage for supplying through first conversion means a primary power supply network of the aircraft, the free turbine and the permanent magnet generator being mounted on the same drive shaft concentric with the high-pressure shaft of the gas turbine, a starter mounted on the high-pressure shaft able to ensure the starting of the gas turbine, the starter being a permanent magnet starter/generator able to output a second alternating voltage for supplying a secondary power supply network of the aircraft and auxiliary equipment of the gas turbine and the first conversion means include two first AC/DC converters controlled from a management unit.

Abstract: A plasma propulsion system with no internal electrodes is described. Gas is flowed into an insulated axisymmetric plasma liner. A radio frequency antenna generates an inductive or helicon plasma discharge within the liner. The plasma is accelerated through a converging/diverging magnetic field out of the liner, generating thrust.

Abstract: The present invention relates to an ion thruster for propulsion of spacecrafts, including: a reservoir for a propellant, an emitter for emitting ions of the propellant, the emitter having one or more projections of porous material and a base with a first side supporting said projections and a second side connected to the reservoir, and an extractor facing the emitter for extracting and accelerating the ions from the emitter, wherein the base is impermeable to the propellant at least on said first side and has pores or channels for providing flow of propellant from the reservoir to said projections.

Abstract: The present invention relates to electrospray thrusters, processes of making electrospray thrusters, and methods of using such electrospray thrusters. Applicant's thruster incorporates a unique geometry for the emitters and extractor grid that effectively eliminates ion interception on the grid, which is the primary failure mechanism of current devices, yet maintains the electric field conditions necessary for ion emission to occur. Without grid impingement, the thrust produced by the thruster is increased and thruster operational lifetime is increased substantially. Additionally, this non-traditional geometry also allows for higher electric fields at the emitter tip for a given applied voltage, thus enabling lower operational voltage of the thruster as compared to conventional designs.