A method of reducing low-energy flow in a flight vehicle engine includes an isolator of the engine having a swept-back wedge to improve flow mixing. The wedge includes forward shock-anchoring locations, such as edges or rapidly-curved portions, that anchor oblique shocks in situations where the isolator has sufficient back pressure. The swept-back wedge may also create swept oblique shocks along its length. Boundary layer flow streamlines are diverted running parallel to or parallel but moving outward conically to the swept-wedge leading edge moving outboard and upward. The non-viscous flow outside the boundary layer is processed through the swept-back ramp shock and diverted outboard and upward as well. The outboard aft portion of the wedge at the sidewall intersection may also induce shocks and divert flow near the walls closer toward the walls and upward, and/or improve flow mixing.

Provided herein is a fuel injector capable of providing fuel into a jet engine operating at hypersonic speeds. Embodiments may include a system for fuel injection for an engine traveling at supersonic speeds. The system may include a fuel injection strut extending between opposing walls of an inlet to the engine, and a porous surface extending across at least a portion of the fuel injection strut. The fuel may be introduced into the inlet of the engine through the porous surface of the fuel injection strut. The porous surface of the fuel injection strut may extend along a fuel injecting portion of the fuel injection strut spaced a predefined distance from the opposing walls of the inlet. The porous portion of the fuel injection strut may include a porosity of about 100 pores per square inch or lower porosities as dictated by the specific design considerations.

A rotary disk, useful for transferring kinetic energy or generating torque or electricity, is disclosed. The rotary disk includes a housing enclosing a manifold, an axle or shaft to which the housing is joined or fixed, conduits or passages within and/or extending from the housing, combustion chambers at distal ends of the conduits or passages and external to the housing, nozzles at distal ends of the combustion chambers, and a compressor in or upstream from the manifold. The conduits or passages are joined or fixed to the housing, and carry an oxygen-containing gas from the manifold to the nozzles. The nozzles direct the heated oxygen-containing gas and combustion gases in a predetermined direction. The compressor includes fins or blades joined or fixed to the axle or shaft, or to a different, colinear axle or shaft. The compressor is configured to increase a pressure of the oxygen-containing gas at entrances of the conduits or passages.

A rotary disk, useful for transferring kinetic energy or generating torque or electricity, is disclosed. The rotary disk includes a housing enclosing a manifold, an axle or shaft to which the housing is joined or fixed, conduits or passages within and/or extending from the housing, combustion chambers at distal ends of the conduits or passages and external to the housing, nozzles at distal ends of the combustion chambers, and a compressor in or upstream from the manifold. The conduits or passages are joined or fixed to the housing, and carry an oxygen-containing gas from the manifold to the nozzles. The nozzles direct the heated oxygen-containing gas and combustion gases in a predetermined direction. The compressor includes fins or blades joined or fixed to the axle or shaft, or to a different, colinear axle or shaft. The compressor is configured to increase a pressure of the oxygen-containing gas at entrances of the conduits or passages.

A flight vehicle engine includes an isolator with a swept-back wedge to improve flow mixing. The wedge includes forward shock-anchoring locations, such as edges or rapidly-curved portions, that anchor oblique shocks in situations where the isolator has sufficient back pressure. The swept-back wedge may also create swept oblique shocks along its length. Boundary layer flow streamlines are diverted running parallel to or parallel but moving outward conically to the swept-wedge leading edge moving outboard and upward. The non-viscous flow outside the boundary layer is processed through the swept-back ramp shock and diverted outboard and upward as well. The outboard aft portion of the wedge at the sidewall intersection may also induce shocks and divert flow near the walls closer toward the walls and upward, and/or improve flow mixing.

A flight vehicle engine includes an isolator with a swept-back wedge to improve flow mixing. The wedge includes forward shock-anchoring locations, such as edges or rapidly-curved portions, that anchor oblique shocks in situations where the isolator has sufficient back pressure. The swept-back wedge may also create swept oblique shocks along its length. Boundary layer flow streamlines are diverted running parallel to or parallel but moving outward conically to the swept-wedge leading edge moving outboard and upward. The non-viscous flow outside the boundary layer is processed through the swept-back ramp shock and diverted outboard and upward as well. The outboard aft portion of the wedge at the sidewall intersection may also induce shocks and divert flow near the walls closer toward the walls and upward, and/or improve flow mixing.

Described is a propulsion system (1) for hypersonic aircraft, having an air inlet (10) of a fluid (110), a containment duct (20) and an exhaust nozzle (30). The propulsion system (1) comprises a bypass duct (40) for a flow (100) of fluid (110), an air-breathing engine (22) and a rocket (23) configured for processing respective flows (22a, 23a) of fluid (110). The bypass duct (40), the air-breathing engine (22) and the rocket (23) are operatively associated with each other in such a way as to generate a thermodynamic-fluid interaction in a same portion of space (33) between the respective flows (40a, 22a, 23a) processed in an operating configuration of the propulsion system (1) and wherein the portion of space (33) is inside the containment duct (20).

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 so mixture into the detonation chamber (3) at at least one segment of the injection base (10).

A scramjet includes a converging inlet, a combustor configured to introduce a fuel stream into an air stream in a combustion chamber and to combust the fuel air mixture stream to create an exhaust stream, and a diverging exit nozzle configured to accelerate the exhaust stream. The combustor includes a fuel injection system including at least one arcjet. A method of creating thrust for an aircraft includes compressing a supersonic incoming air stream in a converging inlet, injecting a fuel stream into the air stream in a combustion chamber to create a fuel air mixture stream, igniting the fuel air mixture stream to create an exhaust stream, and exhausting the exhaust stream from a diverging exit nozzle. The injecting the fuel stream into the air stream includes injecting the fuel stream at a fuel speed sufficient to create shear between the fuel stream and the air stream.

A combustor of a jet engine includes a fuel injector, an igniter for igniting a gas mixture of air and fuel, and a flame holder. The igniter is disposed in the flame holder. After an activation of the igniter, the igniter disappears, and a space after the disappearance functions as a flame-holding space.