Abstract: A fuel and air injection handling system for a rotating detonation engine is provided. The system includes a compressor configured to compress air received via a compressor inlet and configured to output the air that is compressed as swirling, compressed air through a compressor outlet. The system also includes an annular rotating detonation combustor fluidly coupled with the compressor outlet. The annular rotating detonation combustor has a detonation cavity that extends around an annular axis, the annular rotating detonation combustor configured to combust the compressed air from the compressor in detonations that rotate within the detonation cavity around the annular axis of the annular rotating detonation combustor. The annular rotating detonation combustor is fluidly and directly coupled with the compressor outlet.

Abstract: An engine that uses detonation for generating energy includes a housing; an inlet configured to inject a fuel mixture into an ignition region of the housing, the inlet being attached to the housing; an ignitor located in the ignition region and configured to ignite the fuel mixture; a deflagration to detonation, DDT, region in fluid communication and downstream from the ignition region; a pair of electrodes located in the DDT region and configured to apply nanosecond repetitive voltage pulses to the DDT region; and a detonation region, within the housing, in fluid communication and downstream from the DDT region. The nanosecond repetitive voltage pulses generate a non-thermal plasma that transition a combustion front propagation through the housing from a deflagration mode to a detonation mode.

Abstract: A Brayton cycle engine including a longitudinal wall extended along a lengthwise direction. The longitudinal wall defines a gas flowpath of the engine. An inner wall assembly is extended from the longitudinal wall into the gas flowpath. The inner wall assembly defines a detonation combustion region in the gas flowpath upstream of the inner wall assembly.

Abstract: A Brayton cycle engine and method for operation. The engine includes an inner wall assembly and an upstream wall assembly each extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath between the inner wall assembly and the upstream wall assembly. The engine flows an oxidizer through the gas flowpath and the inner wall captures a portion of the oxidizer. The engine further adjusts the captured flow of oxidizer via the upstream wall and flows a first flow of fuel to the captured flow of oxidizer to produce rotating detonation gases. The engine flows the detonation gases downstream and to mix with the flow of oxidizer, and flows and burns a second flow of fuel to the detonation gases/oxidizer mixture to produce thrust.

Abstract: An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes an engine bleed tap coupled to a fan-air source or a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine and a motorized compressor in fluid communication with the engine bleed tap. The engine bleed control system also includes a controller operable to selectively drive the motorized compressor to boost a bleed air pressure as pressure augmented bleed air and control delivery of the pressure augmented bleed air to an aircraft use.

Abstract: A pulse combustor system for efficiently operating a pulse combustor. The pulse combustor system includes the pulse combustor and a duct. The pulse combustor has a combustion chamber defining an internal space, a conduit having a first end in fluid communication with the internal space and a second end in fluid communication with an environment outside of the pulse combustor system, and a fuel injector configured to inject fuel into the internal space of the combustion chamber. The duct has two openings, with one opening disposed adjacent to the second end of the conduit. The pulse combustor system has an average operating frequency, and the duct has a length that is about one quarter of a wavelength corresponding to the average operating frequency. The pulse combustor and the duct each has a central longitudinal axis, and the two axes are substantially aligned.

Abstract: An engine includes an inlet tube introducing air to a combustion process and a first plurality of fuel injectors disposed in the inlet tube and used for scram-jet engine operation. The engine includes a second plurality of fuel injectors used for ram-jet engine operation. The second plurality of fuel injectors is upstream from the first plurality of fuel injectors and is disposed in the inlet tube. The engine includes a combustor swirl zone downstream of and adjacent to the first plurality of fuel injectors.

Abstract: A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supporting the shaft relative to the inlet case. The second bearing is arranged radially outward from the shaft.

Abstract: A combustor assembly for a gas turbine engine includes a dome defining a slot. The combustor assembly also includes a liner at least partially defining a combustion chamber and extending between an aft end and a forward end. The forward end is received within the slot of the dome. In one exemplary aspect, the combustor assembly includes features that warm the forward end of the liner during transient operation of the engine. Furthermore, the combustor assembly includes features that reduce the thermal gradient between the forward end and the other portions of the liner. In this way, improved durability of the liner may be achieved.

Abstract: A combustion system includes an annular tube disposed between an inner wall and an outer wall, the annular tube extending from an inlet end to an outlet end; at least one annulus inlet disposed in the annular tube proximate the inlet end, the annulus inlet providing a conduit through which fluid flows into the annular tube; at least one outlet disposed in the annular tube proximate the outlet end; at least one inlet fluid plenum disposed upstream of the annulus inlet; and at least one fluid inlet disposed upstream of the inlet fluid plenum. The fluid inlet is linearly offset from the annulus inlet.

Abstract: A ramjet engine and system and method for operation is generally provided. The ramjet includes a longitudinal wall extended along a lengthwise direction. The longitudinal wall defines an inlet section, a combustion section, and an exhaust section. A fuel nozzle assembly is extended from the longitudinal wall. The fuel nozzle assembly defines a nozzle throat area. The fuel nozzle assembly is moveable along a radial direction to adjust the nozzle throat area based at least on a difference in pressure of a flow of fluid at an inlet of the inlet section and a pressure of the flow of fluid at the fuel nozzle assembly.

Abstract: A Brayton cycle engine including an inner wall assembly defining a detonation combustion region upstream thereof extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath. A method for operating the engine includes flowing an oxidizer through the gas flowpath; capturing a portion of the flow of oxidizer via the inner wall; flowing a first flow of fuel to the captured flow of oxidizer; producing a rotating detonation gases via a mixture of the first flow of fuel and the captured flow of oxidizer; flowing at least a portion of the detonation gases downstream to mix with the flow of oxidizer; flowing a second flow of fuel to the mixture of detonation gases and oxidizer; and burning the mixture of the second flow of fuel and the detonation gases/oxidizer mixture.

Abstract: In one embodiment, a power generation system includes a pulse detonation engine including a combustion chamber, a linear power generator including a working chamber, and a nozzle positioned between the combustion chamber and the working chamber that expands exhaust gas expelled from the combustion chamber, wherein the nozzle increases thermodynamic efficiency of the system.

Abstract: An axial-direction passage, a forced vortex passage, a first blade array passage, and a second blade array passage are formed in a rotor shaft of a turbine. Cooling air from an air extraction port of a compressor flows through the axial-direction passage which extends in an axial direction. The forced vortex passage is connected to the axial-direction passage and extends outwards in the radial direction relative to an axial line from a connecting portion between the forced vortex passage and the axial-direction passage. The first blade array passage is connected to an end portion on the outer side in the radial direction of the forced vortex passage and guides cooling air to a first blade row among a plurality of blade rows. The second blade array passage is connected to an end portion of the forced vortex passage and guides cooling air to a second blade row.

Abstract: A combustion system includes an annular tube disposed between an inner wall and an outer wall, the annular tube extending from an inlet end to an outlet end; at least one fluid inlet disposed in the annular tube proximate the inlet end, the fluid inlet providing a conduit through which fluid flows into the annular tube; at least one outlet disposed in the annular tube proximate the outlet end; at least one primary fuel injector, the primary fuel injector dispersing fuel into a fluid stream entering the annular tube via the fluid inlet; and at least one secondary fuel injector, the secondary fuel injector disposed in the annular tube.

Abstract: A scramjet engine has a first passage forming member and a second passage forming member. A passage is formed between a first surface and a second surface. The passage has an upstream zone, a combustion zone and a downstream zone. A cavity of a concave shape is provided on the first surface in the combustion zone. The first passage forming member has a convex section located in the upstream zone, a first fuel injection section configured to inject fuel into the passage from a first fuel nozzle provided for the convex section, and a second fuel injection section configured to inject fuel to the cavity. The second passage forming member has a third fuel injection section configured to inject fuel to a direction toward the first surface from the second surface in the passage through a second fuel nozzle provided in the downstream zone.

Abstract: Combustion equipment includes annular combustion chamber having an annular upstream wall structure. Annular outer casing surrounds and is spaced from annular combustion chamber and annular inner casing is within and spaced from annular combustion chamber. Ogee shaped annular cowl is positioned upstream of annular upstream wall structure and a stage of compressor outlet guide vanes is positioned upstream of annular cowl. Pre-diffuser extends in downstream direction from stage of compressor outlet guide vanes, pre-diffuser is positioned upstream of cowl. Stage of compressor outlet guide vanes is connected to outer and inner casing. Outer fairing extends radially inwardly from outer casing such that annular outer casing, outer fairing and pre-diffuser define annular outer cavity and outer fairing is spaced from cowl or inner fairing extends radially outwardly from inner casing such that annular inner casing, inner fairing and pre-diffuser define an annular inner cavity and inner fairing is spaced from cowl.

Abstract: A rotating detonation combustion system is generally provided. The rotating detonation combustion system includes an outer wall, an upstream wall, and a radial wall. The outer wall is defined circumferentially around a combustor centerline extended along a lengthwise direction. The outer wall defines a first radius portion generally upstream along the outer wall. A second radius portion is defined generally downstream along the outer wall and a transition portion is defined between the first and second radius portions. The first radius portion defines a first radius greater than a second radius at the second radius portion. The transition portion defines a generally decreasing radius from the first radius portion to the second radius portion. The upstream wall is defined circumferentially around the combustor centerline and is extended along the lengthwise direction and inward radially of the first radius portion of the outer wall. An oxidizer passage is defined within the upstream wall.

Abstract: A turbine engine assembly including an air flow guide assembly, including at least one guide vane and at least one structural arm, the vane and arm extending radially about an axis. The arm includes an upstream end portion having a guide vane profile and including a leading edge aligned with that of the vane; a downstream portion; and an intermediate portion including an upper surface extending between an upstream end point and a downstream end point. The upstream end point is separated from the leading edge of the arm by an axial distance of between 0.2c and 0.5c, c being the length of the axial chord of the vane, and the angle of the tangent to the upper surface at the upstream end point is equal to that at the downstream end point ±1 degree.

Abstract: The turbojet includes, linking two of its parts as a main nacelle portion and a rear nacelle portion, a device which moves them relative to one another, for example in order to open a reverse gas thrust slot. The device includes a sleeve and two rods which are coaxial connected together by two screw links and two locking devices which may be engaged independently, such that the deployment of one of the rods results in the opening of the slot and the movement of the other rod, controlled by the same device, only occurs under specific turbojet maintenance circumstances.