A method of supersonic thrust generation includes generating a thrust supersonic exhaust plume having a first average velocity from an engine, and expelling a bypass exhaust plume having a second average velocity from the engine, the first average velocity greater than the second average velocity, so that the bypass exhaust plume inhibits coalescence of an engine exhaust plume compression shockwave.

A nacelle defining an intake is provided for channelling and conditioning freestream airflow to a fan of a ducted fan gas turbine engine. The nacelle intake comprises in flow series an intake lip, and a diffuser. The intake lip has at its forwardmost end a highlight which is a closed loop defining a boundary between inner and outer surfaces of the nacelle, and the diffuser terminates at its rearwardmost end at a front face of the fan. The intake has a circumferentially extending shock control bump or a series of circumferentially spaced shock control bumps formed thereon. On longitudinal cross sections through the intake containing the bump or series of bumps: the bump or each bump of the series of bumps has a profile which forms in flow series an up-ramp, a maximum and a down-ramp, at least one of the up-ramp and the down-ramp forming an inflection point or curvature discontinuity point on the profile, and the inflection point or curvature discontinuity point being axially located between the highlight and a position which is axially rearward therefrom by a distance of no more than 0.6 L, where L is the axial distance between the highlight and the front fan face.

A nacelle defining an intake is provided for channelling and conditioning freestream airflow to a fan of a ducted fan gas turbine engine. The nacelle intake comprises in flow series an intake lip, and a diffuser. The intake lip has at its forwardmost end a highlight which is a closed loop defining a boundary between inner and outer surfaces of the nacelle, and the diffuser terminates at its rearwardmost end at a front face of the fan. The intake has a circumferentially extending shock control bump or a series of circumferentially spaced shock control bumps formed thereon. On longitudinal cross sections through the intake containing the bump or series of bumps: the bump or each bump of the series of bumps has a profile which forms in flow series an up-ramp, a maximum and a down-ramp, at least one of the up-ramp and the down-ramp forming an inflection point or curvature discontinuity point on the profile, and the inflection point or curvature discontinuity point being axially located between the highlight and a position which is axially rearward therefrom by a distance of no more than 0.6 L, where L is the axial distance between the highlight and the front fan face.

A flight vehicle has a propulsion system that includes an air inlet, an isolator (or diffuser) downstream of the air inlet, and a combustor downstream of the isolator. The isolator includes an obstruction that protrudes inwardly from an inner wall of the isolator, into the flow channel in which air flows through the isolator. The obstruction diverts the flow to either side of it. Downstream of the obstruction the flow on either side of the obstruction comes together again, leading to mixing of the flow, for example including mixing of low energy and boundary layer flow with high energy flow. This mixing of flow may make for a more uniform flow at the exit of the isolator. In addition the obstruction may help fix the location of shocks within the isolator, providing longer flow mixing length in the isolator.

A flight vehicle has a propulsion system that includes an air inlet, an isolator (or diffuser) downstream of the air inlet, and a combustor downstream of the isolator. The isolator includes an obstruction that protrudes inwardly from an inner wall of the isolator, into the flow channel in which air flows through the isolator. The obstruction diverts the flow to either side of it. Downstream of the obstruction the flow on either side of the obstruction comes together again, leading to mixing of the flow, for example including mixing of low energy and boundary layer flow with high energy flow. This mixing of flow may make for a more uniform flow at the exit of the isolator. In addition the obstruction may help fix the location of shocks within the isolator, providing longer flow mixing length in the isolator.

A flow control actuator includes at least one side wall, an upstream wall coupled to an upstream end of the side wall, a downstream cap coupled to a downstream end of the side wall, the downstream cap comprising at least one orifice disposed therein, at least one fuel injector disposed in at least one of the upstream wall, and the sidewall, the fuel injector dispersing fuel into the interior of the flow control actuator, and at least one oxidizer inlet disposed in at least one of the upstream wall and the sidewall, the at least one oxidizer inlet introducing an oxidizer into the interior of the flow control actuator. The flow control actuator includes at least one external fuel injector disposed adjacent to the side wall. The fuel from the fuel injector and oxidizer from the oxidizer inlet ignite in the interior of the flow control actuator.

By manipulating the fluid flow in the proximity of an object such as a fuselage, a wing, or the hull of ship, the wave drag associated with this object can be substantially reduced. This can be accomplished by locally changing both the fluid flow velocity and the pressure of the fluid flow.

An aircraft wing is disclosed herein having a fixed-location trip device placed along the span of the wing to transition airflow from laminar flow to turbulent flow so that potential load increases are limited and flight performance uncertainties associated with laminar flow wings are reduced. Wings designed for extended laminar flow offer the potential to significantly reduce airplane drag and fuel consumption. A collateral impact of a laminar flow wing is the generation of elevated wing loads at critical load conditions. This impact is mitigated by controlling the downstream limit of transition at these critical load conditions.

An aircraft wing is disclosed herein having a fixed-location trip device placed along the span of the wing to transition airflow from laminar flow to turbulent flow so that potential load increases are limited and flight performance uncertainties associated with laminar flow wings are reduced. Wings designed for extended laminar flow offer the potential to significantly reduce airplane drag and fuel consumption. A collateral impact of a laminar flow wing is the generation of elevated wing loads at critical load conditions. This impact is mitigated by controlling the downstream limit of transition at these critical load conditions.

A propulsion system includes at least one rotating detonation actuator comprising: a flow path extending from an inlet end to an outlet end; an inner wall defining a radially inner boundary of the flow path; an outer wall defining a radially outer boundary of the flow path; and at least one aircraft wing. The rotating detonation actuator is disposed in the aircraft wing. At least one rotating detonation wave travels through the flow path from the inlet end to the outlet end.