12results about "Influencers by shock waves" patented technology

A wing structure for a supersonic aircraft including a pair of supersonic wave-tripping channels formed on each of two laterally extending wings of the supersonic aircraft, wherein each of the pair of supersonic wave-tripping channels extend in a span-wise direction of the wings respectively, wherein an upper supersonic wave-tripping channel of the pair of supersonic wave-tripping channels is disposed on an upper surface of each of the wings and a lower supersonic wave-tripping channel of the pair of supersonic wave-tripping channels is disposed on a lower surface of each of the wings, wherein the upper supersonic wave-tripping channel and the lower supersonic wave-tripping channel are set back from a leading edge of the wings respectively.

A supersonic flow system in which a supersonic flow is encountered includes a surface configured to be exposed to a supersonic fluid flow, wherein the surface includes an upstream end and a downstream end opposite the upstream end, a wave generator coupled to the surface and including one or more actuators configured to selectably induce strain in the surface, and a wave controller in signal communication with the wave generator and configured to activate the one or more actuators to induce one or more travelling waves configured to travel along the surface between the upstream end and the downstream end of the surface.

A plasma-combined trailing edge suction flow control device for rotor blades is provided, comprising a plurality of blades (1), a nanosecond pulse dielectric barrier discharge plasma actuator (2) mounted at the leading edge of the blades and covering the entire blade extension; an air suction device (3) is mounted at the tail end of the blades, and the center of the air suction port is located at a position A%c from the leading edge. The tail end suction device (3) comprises a tail end suction port (3-1), a low-pressure air chamber (3-2) and a connecting pipe (3-3). When the dynamic stall of the rotor airfoil is controlled, a high-frequency nanosecond pulse power supply (4) is used to power the nanosecond pulse dielectric barrier discharge plasma actuator (2). A plasma-combined trailing edge suction flow control method for rotor blades is also provided. The present invention can effectively interfere with the flow field structure on the upper surface of the airfoil by combining plasma excitation with trailing edge suction, reduce the degree of flow separation, and thereby increase the average lift of the rotor blades, thereby improving the performance of the rotor blades.

[Object] To realize an improvement in design accuracy and a reduction in design time in a process of matching an equivalent cross-sectional area of a design shape of a supersonic aircraft to a target equivalent cross-sectional area in a sonic boom reduction method based on an equivalent cross-sectional area. [Solving Means] The technique includes: setting an initial shape of the airframe and a target equivalent cross-sectional area of the airframe; estimating a near field pressure waveform for the initial shape of the airframe assuming that the supersonic aircraft flies at a cruising speed; evaluating an equivalent cross-sectional area from the estimated near field pressure waveform for the initial shape of the airframe; and setting a Mach plane corresponding to the cruising speed, and setting a design curve on the Mach plane, the design curve corresponding to an initial curve at which the initial shape of the airframe and the Mach plane intersect so that the equivalent cross-sectional area approaches the target equivalent cross-sectional area. Then, the shape of the airframe is designed based on the design curve.

A cavity system for use on a vehicle is provided. The cavity system comprises a cavity having at least one of a front edge and a rear edge arranged orthogonally to the actual or intended direction of travel of the vehicle, and longitudinal edges. The front and / or rear and longitudinal edges define an opening. The cavity system also comprises at least one shock stabiliser protruding from a surface adjacent each longitudinal edge of the cavity into the ambient flow outside of the cavity. An aircraft comprising the cavity system is also provided.

The invention discloses a reconfigurable hypersonic aircraft head and a deformation control strategy thereof. Relates to the technical field of resistance and heat reduction of aircrafts, a supporting frame is arranged in the end of the aircraft, a hollow rod is arranged in the supporting frame, and a shock wave needle is slidably connected into the hollow rod; a pneumatic assembly used for driving the shock wave needle to slide is further arranged in the aircraft end. A through hole is formed in the shock wave needle and communicates with a reverse jet flow assembly. The end of the shock wave needle is provided with a boss and an aircraft end concave cavity, and the aircraft end is further provided with an ejection assembly. Through the arrangement of the reconfigurable head, the aircraft can be matched with the flight incoming flow condition and change the shape of the head, so that the requirement of the real-time working condition is met, the end heat flow or the aerodynamic resistance of the aircraft is reduced, and the optimal flight effect is achieved; meanwhile, the invention further provides a control strategy of the structure. And carrying out collaborative optimization on the continuously changing head configuration and the incoming flow parameters updated in real time to ensure that an optimal head configuration switching strategy is obtained.

An aircraft wing (104) is disclosed herein having a fixed-location trip device (108) placed along the span of the wing to transition airflow from laminar flow (302) to turbulent flow (306) 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.

The application discloses a reconfigurable hypersonic vehicle head and a deformation control strategy thereof, relates to the technical field of drag reduction and heat reduction of a vehicle, and is characterized in that a support frame is arranged in the inside of the vehicle end head, a hollow rod is arranged in the support frame, a shock wave needle is slidably connected in the inside of the hollow rod, a pneumatic assembly for driving the shock wave needle to slide is further arranged in the inside of the vehicle end head, a through hole is arranged in the inside of the shock wave needle, a reverse jet flow assembly is communicated with the through hole, a boss is arranged at the end of the shock wave needle, a concave cavity is arranged in the vehicle end head, and an ejection assembly is further arranged on the vehicle end head. The reconfigurable head is arranged, the vehicle can match the flight flow conditions, the head shape can be changed, the requirements of real-time working conditions can be met, the end head heat flow or the aerodynamic drag of the vehicle can be reduced, the best flight effect can be realized, and the control strategy of the structure is further provided. The continuously changed head configuration and the real-time updated flow parameters are cooperatively optimized, and the optimal head configuration switching strategy is ensured.

The application provides a tailless wide-speed-range aircraft based on vortex wave effect and a control method thereof. The aircraft comprises a body, the body comprises a fuselage and a wing, the wing and the fuselage form a blended wing body layout, the wing comprises an inner wing and an outer wing, the inner wing is provided with an inboard elevator and an outboard elevator, the outer wing is provided with an aileron, a wing trailing edge resistance rudder and a wing embedded resistance rudder, and the tail of the fuselage is provided with a fuselage lift resistance rudder. The application optimizes the characteristics of the wing windward shock wave and the leeward vortex, realizes wide-speed-range lift generation and wave resistance reduction. The tailless wing layout reduces the radar reflection cross section, and the smooth blended wing body design improves the stealth performance. The multi-surface coordinated control strategy solves the lateral static instability problem of the tailless aircraft, and meets the wide-speed-range control demand. The layout flexibility is improved, the functions of the inner and outer wings and the surfaces are divided, and the aerodynamic control and structural performance of the aircraft are optimized.