Abstract: A supersonic inlet employs relaxed isentropic compression to improve net propulsive force by shaping the compression surface of the inlet. Relaxed isentropic compression shaping of the inlet compression surface functions to reduce cowl lip surface angles, thereby improving inlet drag characteristics and interference drag characteristics. Supersonic inlets in accordance with the invention also demonstrate reductions in peak sonic boom overpressure while maintaining performance.

Abstract: A turbine nozzle includes a row of vanes extending radially in span between inner and outer bands. The vanes include opposite pressure and suction sidewalls and opposite leading and trailing edges. Each vane includes an inner pattern of inner cooling holes and an outer pattern of outer cooling holes distributed along the leading edge. The inner and outer holes diverge toward the corresponding inner and outer bands to preferentially discharge cooling air.

Abstract: A variable fan nozzle for use in a gas turbine engine includes a nozzle section, such as an inflatable bladder, associated with a fan bypass passage for conveying a bypass airflow. The nozzle section has an internal fluid pressure that is selectively variable to influence the bypass airflow.

Abstract: An arrangement (10) for conveying combustion gas from a plurality of can annular combustors to a turbine first stage blade section of a gas turbine engine, the arrangement (10) including a plurality of interconnected integrated exit piece (IEP) sections (16) defining an annular chamber (18) oriented concentric to a gas turbine engine longitudinal axis (20) upstream of the turbine first stage blade section. Each respective IEP (16) includes a first flow path section (40) receiving and fully bounding a first flow from a respective can annular combustor along a respective common axis (22) there between, and delivering a partially bounded first flow to a downstream adjacent IEP section (42). Each respective IEP further includes a second flow path section (112) receiving a partially bounded second flow from an upstream adjacent IEP (66) and delivering at least part of the second flow to the turbine first stage blade section.

Abstract: A fuel injector for a gas turbine engine comprises a housing stem and a nozzle, the nozzle including an internal wall in heat transfer relation with fuel flowing through the nozzle, and an external wall in heat transfer relation with ambient air. The internal and external walls have downstream tip ends that are relatively moveable at an interface due to relative thermal growth during operation of the engine. An internal insulating gap is disposed between the internal and external walls to provide a heat shield for the internal wall, and a bellows internal to the injector has an upstream end sealingly attached to an upstream portion of one of the internal and external walls, and a downstream end sealingly attached to a downstream portion of the other wall to fluidly separate the insulating gap from any fuel entering into the nozzle through the interface.

Abstract: An arrangement for delivering gasses from can combustors of a can annular gas turbine combustion engine to a turbine first stage section including a first row of turbine blades, the arrangement including a flow-directing structure for each combustor, wherein each flow-directing structure includes a straight path and an annular chamber end, wherein the annular chamber ends together define an annular chamber for delivering the gas flow to the turbine first stage section, wherein gasses flow from respective combustors, through respective straight paths, and into the annular chamber as respective straight gas flows, and wherein the annular chamber is configured to unite the respective straight gas flows along respective shear planes to form a singular annular gas flow, and wherein the annular chamber is configured to impart circumferential motion to the singular annular gas flow before the singular annular gas flow exits the annular chamber to the first row of blades.

Abstract: A turbomachine including an annular combustion chamber fitted with fuel injectors and nozzle vanes arranged at the outlet from the chamber, the number of nozzle vanes being an integer multiple of the number of fuel injectors, and the head of each injector being situated angularly half-way between the leading edges of two consecutive nozzle vanes, these leading edges being in alignment with primary air holes and/or with dilution air holes.

Abstract: A duplex turbine nozzle includes a row of different first and second vanes alternating circumferentially between radially outer and inner bands in vane doublets having axial splitlines therebetween. The vanes have opposite pressure and suction sides spaced apart in each doublet to define an inboard flow passage therebetween, and corresponding outboard flow passages between doublets. The vanes have different patterns of film cooling holes with larger cooling flow density along the outboard passages than along the inboard passages.

Abstract: Disclosed is a turbomachine including at least one exhaust pathway along which exhaust is directed and released to ambient and at least one exhaust processor capable of removing regulated substances from the exhaust. One or more ambient air inlets are located at the at least one exhaust pathway upstream of the at least one exhaust processor. The at least one exhaust pathway is configured such that ambient air is capable of being urged into the at least one exhaust pathway through the one or more ambient air inlets by an acceleration of the exhaust along the at least one exhaust pathway. The ambient air urged into the at least one exhaust pathway reduces a temperature of the exhaust to increase effectiveness of the at least one exhaust processor. Further disclosed is a method for releasing turbomachine exhaust to ambient.

Abstract: A gas turbine engine is piloted with a pilot flow of fuel delivered to a combustor as a liquid. A first additional flow of the fuel is also delivered to the combustor as a liquid. A second additional flow of the fuel is vaporized and delivered to the combustor as a vapor. A fuel injector may have passageways associated with each of the three flows.

Abstract: The invention relates to a thermodynamic circuit with a working medium comprising at least two substances with non-isothermal evaporation and condensation whereby the working medium can decompose above a given temperature. According to the invention, the heat from heat sources at temperatures above the decomposition temperature of the working medium may be made useful with little complexity and with high operational security, whereby the heat from the heat source is transferred in a first step to a hot liquid circuit and, in a second step, from the hot liquid circuit to the circuit with the working medium comprising at least two substances with non-isothermal evaporation and condensation. The heat introduced to the circuit with the working medium comprising at least two substances with non-isothermal evaporation and condensation can be reduced by means of the intermediate hot liquid circuit, such that a decomposition of the working medium can be avoided.

Abstract: A mounting system supporting and positioning an internal fuel manifold of a gas turbine engine includes at least a heat shield surrounding a fuel inlet of the fuel manifold, the heat shield bearing at least a portion of a load to support the fuel manifold.

Abstract: A gas turbine system including a source of gas coupled to a source of fuel wherein the gas and the fuel are combined to form a mixture of gas and fuel prior to the mixture being introduced to a fuel nozzle of the gas turbine system.

Abstract: An assembly including a stack structure made up of the following elements assembled together by brazing: the titanium-based metal piece; a first intermediate piece suitable for deforming to accommodate differential expansion between the metal piece and a piece made of ceramic material based on silicon carbide and/or carbon; a second intermediate piece that is rigid, having a coefficient of expansion close to that of said ceramic material piece and made of aluminum nitride (AlN) or of tungsten (W); and the ceramic material piece is disclosed.

Abstract: The invention relates to a control method for opening or closing a turbojet engine thrust reverser (1) by using at least one mobile cowl (2) displaceable by means of at least one electric motor (7) consisting in analyzing at least one parameter representative for a pressure in the turbojet engine jet and in carrying out an operating sequence in which the operating parameters of the electric motor (7) are adjusted to a situation.

Abstract: The injector device for injecting a liquid mono-propellant with a large degree of flow rate modulation and an injection speed that is stable, and that is closable for extinction and re-ignition purposes, is disposed at an upstream end of the wall of a combustion chamber of a rocket engine. The device includes at least one feed channel for feeding mono-propellant from a tank, and first and second concentric annular speed-up channels connected to the feed channels and having outlets opening out respectively via first and second annular injection sections situated in a plane that is substantially perpendicular to the axis of the chamber.

Abstract: A method for assembling a fuel supply system for use with a power generation system. The method includes coupling a fuel heater to a fuel source for heating a fuel. A first heating assembly is coupled to the fuel heater for heating a first flow of water channeled to the fuel heater. A heat recovery steam generator assembly is coupled to the fuel heater for channeling a second flow of heated water to the fuel heater. A valve assembly is coupled between the first heating assembly, the heat recovery steam generator assembly, and the fuel heater to enable a flow of heated water from the first heating assembly and the heat recovery steam generator assembly to be selectively channeled to the fuel heater.

Abstract: A high voltage propellant isolator includes at least two different types of isolator rings or segments, in alternating lateral arrangement, that direct the flow of propellant, such as xenon gas, in a tortuous path through the isolator.

Abstract: A method of operating a gas turbine engine having a turbine and a compressor connected via a shaft, a main fuel supply line for supplying fuel to a combustor that is positioned to release expanding hot gases to the turbine, the engine further including a starter/generator connected to the shaft via a gearbox assembly, the method is characterised the step of during engine start up fuel is circulated in a re-circulating fuel circuit positioned on the main fuel supply line and which has a first fuel/oil heat exchanger, for cooling the oil, and a fuel accumulator.

Abstract: The air exit hole of the vent line (6) connected to a venting apparatus for the lubricating oil system of a jet engine is arranged behind the nozzle throat (2) on the periphery of the exiting engine jet (4). The exit hole is tangentially arranged on the periphery of the engine jet, or slightly enters the rim area of the engine jet. The vent line extends under the protection of an aerodynamically shaped fairing (5, 11). This arrangement of the air exit, while being simply designed, cost-effective and weight-saving, provides for clean, invisible discharge of air from the oil venting apparatus.