Abstract: The present disclosure relates to a thrust reverser for a turbojet engine nacelle, including cowls that move backwards to open the thrust reverser. Each cowl includes a primary axial guide rail that receives most of the loads of the cowl, and secondary rails guiding an outer portion of each cowl. The thrust reverser includes a rear end stop for each cowl is arranged on the primary axial guide rails.

Abstract: The invention relates to a ramjet including a detonation chamber and an aircraft comprising such a ramjet. According to the invention, the ramjet (S1) comprises an annular detonation chamber (2) having a continuous detonation wave and fuel injection means (6) for continuously injecting fuel (F2) directly into the chamber (2) just downstream of an air injection base (3). The fuel (F2) and the air (F1) are injected separately into the detonation chamber (2) in a permanent manner throughout the operation of the ramjet (S1).

Abstract: An apparatus comprises a leading edge of an inlet. The leading edge of the inlet is positioned relative to a direction of air flow such that a total pressure of air along the leading edge of the inlet is equalized within selected tolerances.

Abstract: An injector plate for a rocket engine assembly, for connection to the rocket engine assembly having a body with a fuel manifold that contains a fuel entry and a fuel passage ring encircling an axis, that fuels recirculation passages extending between portions of the fuel passage ring and fuel element passages, each fuel element passage extending to fuel outlets. The injector plate also has an oxidizer with an internal primary oxygen passage, secondary oxygen passages branching from the primary oxygen passage at an internal portion of the body to oxidizer outlets and tertiary oxygen passages branching from the secondary oxygen passages at an internal portion of the body to the oxidizer outlets. The fuel outlets and the oxidizer outlets are arranged to form outer and inner element grouping patterns.

Abstract: A rocket engine having a co-axial, bidirectional flow arrangement is described herein. The rocket engine receives fuel and an oxidizer into the rocket engine in a first direction, whereby a portion of the fuel is combusted in a pre-burner. The flow direction of the partially combusted fuel/oxidizer mixture is reversed, whereby the mixture is introduced into a combustion chamber. The fuel and oxidizer are combusted in the combustion chamber. The combustion products exit a throat and an expansion plenum in a direction similar to the first direction, whereby the combustion products exit a nozzle of the rocket engine, providing thrust.

Abstract: Solid propellant systems include a main propellant and a secondary propellant in contact with the first propellant that exhibits autoignition temperatures of at least about 100° F. lower than the autoignition temperature of the main propellant. The secondary propellant of the present invention is most advantageously employed with conventional AP-containing solid propellant formulations as the main propellant, especially formulations containing both AP, an energetic solid, and a binder. In especially preferred forms, the secondary propellant will include a nitramine which is at least one selected from nitroguanidine (NQ), cyclotrimethylene trinitramine (RDX) and cyclotetramethylenetetranitramine (HMX), and a binder which is at least one selected from HTPB, HTPE or glycidyl azide polymer (GAP). Most preferably, the secondary propellant will include a combination of nitramines which includes NQ and one of RDX or HMX.

Abstract: A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

Abstract: A nacelle inlet is provided for a turbo-fan engine for an aircraft. The nacelle inlet includes a lip skin having first and second pad-ups, an inner barrel including a forward flange and a bulkhead. The bulkhead, the lip skin and the inner barrel are attached together at mating surfaces. The bulkhead and the lip skin are attached together at mating surfaces. Each mating surface defines a surface geometry which approximates a developable surface having parallel ruling lines. The developable surfaces are formed by machining.

Abstract: A bowed rotor prevention system for a gas turbine engine includes a bowed rotor prevention motor operable to drive rotation of the gas turbine engine through an engine accessory gearbox. The bowed rotor prevention system also includes a controller operable to engage the bowed rotor prevention motor and drive rotation of the gas turbine engine below an engine starting speed until a bowed rotor prevention threshold condition is met.

Abstract: A turbojet is combined with a co-axially integrated rotary rocket to form a propulsion system called a Rotary Turbo Rocket that can function as a turbojet, as an afterburning turbojet, as an Air Turbo Rocket, or as a rotary rocket. The Rotary Turbo Rocket can operate in any of these propulsion modes singularly, or in any combination of these propulsion modes, and can transition continuously or abruptly between operating modes. The Rotary Turbo Rocket can span the zero to orbital flight velocity speed range and/or operate continuously as it transitions from atmospheric to space flight by transitioning between operating modes.

Abstract: A gas turbine engine has a fairing and an air intake that includes an air inlet embedded within the fairing for supplying free stream atmospheric air to a gas generator.

Abstract: A system includes a turbine combustor that includes a head end portion having a head end chamber, a combustion portion having a combustion chamber disposed downstream from the head end chamber, a cap disposed between the head end chamber and the combustion chamber, and a flow distributor configured to distribute at least one of an exhaust flow, an oxidant flow, an oxidant-exhaust mixture, or any combination thereof circumferentially around the head end chamber.

Abstract: The invention relates to a continuous detonation wave engine and aircraft provided with such an engine. The continuous detonation wave engine (1) operates with a detonation mixture of fuel and oxidant and includes, in particular, a detonation chamber (3) comprising an injection base (10), the length of which is defined along an open line (17), such as to form a detonation chamber (3) having an elongate form in a transverse plane, as well as an injection system (4) arranged such as to inject the fuel/oxidant mixture into the detonation chamber (3) at at least one segment of the injection base (10).

Abstract: A fuel nozzle for use in a turbine engine is provided. The fuel nozzle includes at least one premixer tube including a tube wall and a plurality of perforations defined therein and extending through the tube wall. The plurality of perforations are configured to channel a flow of air therethrough. The fuel nozzle also includes a liquid fuel plenum positioned upstream from the premixer tube, and at least one fuel injector coupled in flow communication with the liquid fuel plenum and the at least one premixer tube. The at least one fuel injector is configured to channel a flow of liquid fuel from the liquid fuel plenum into the premixer tube.

Abstract: A modular fuel nozzle tip for a gas turbine engine includes a body defining one or more fuel conveying passages extending therethrough, an annular cap having a radially inner shoulder surface interfacing with the peripheral end surface of the body to define a plurality of air channels extending through the head portion of the modular fuel nozzle tip. At least two fasteners fasten the annular cap to the body.

Abstract: An aircraft nacelle comprising a first duct secured to an air intake and a second duct secured to a powerplant, the first duct comprising an exterior surface and an end portion connected to the second duct. The nacelle includes a hollow section piece in contact with the exterior surface of the first duct which extends over at least part of the circumference of that end portion of the first duct that is connected to the second duct.

Abstract: A sandwich structure includes a first skin, a second skin, a first hinge member, and a second hinge member. The first hinge member may be movably coupled to the first skin and a first skin joint. The second hinge member may be movably coupled to the second skin and second skin joint. The first hinge member and the second member may be movably coupled to one another and a member joint located between the first and second skin joint.

Abstract: A method of inhibiting vanadic corrosion of a hot part of a gas turbine system is provided. The method includes introducing, in the combustor, a first oxide comprising magnesium oxide (MgO) and at least one second oxide from among Al2O3, Fe2O3, TiO2 and SiO2. A ratio “m” of a number of moles of MgO to a number of moles of V2O5 and a ratio “a” of a total number of moles of the at least one second oxide to the number of moles of V2O5 satisfy two conditions based on a firing temperature of the expansion turbine, an average density of one or more double oxides formed by a reaction between MgO and the at least one second oxide, and an average Knoop hardness of the one or more double oxides formed by the reaction between MgO and the at least one second oxide.

Abstract: An attachment structure for attaching a power plant including unducted propellers to a fuselage of an aircraft, the attachment structure being connected by first and second fasteners to respective front and rear spars penetrating into the fuselage and fastened to a carrier structure of the aircraft, and including a central longitudinal beam secured at its front end to a first frame, arranged in the plane of the front spar, that is secured at its rear end to a second rear frame, cantilevered out and arranged upstream from the unducted propellers, and that is secured to a third intermediate frame, arranged in the plane of the rear spar between the first and second frames. The first, second, and third frames are connected together by force-takeup beams, an assembly as constituted in this way forming a cradle for receiving the power plant via a standard attachment.

Abstract: One embodiment of the present disclosure is a gas turbine engine. Another embodiment is a unique combustion system. Another embodiment is a unique engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for employing continuous detonation combustion processes. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.