90results about "Influencers by shock waves" patented technology

Porous surfaces on an aerodynamic structure driven with positive and negative pressures are used in an active control system for attenuating shock waves responsible for high-speed impulsive (HSI) noise. The control system includes an array of apertures in the outer skin of the structure providing fluid communication between the exterior flow stream and an interior volume of the structure. A movable diaphragm within the structure pushes air out of and pulls air in through the apertures under the action of a drive mechanism within the structure, thus creating oscillating air jets. The drive mechanism may be actuated by a controller based on information supplied by a sensor in the leading edge of the aerodynamic structure. The array of apertures may be spaced apart along the outer skin of the aerodynamic structure so as to span a distance of about 15% of the chord length. The oscillating airjets may be provided on multiple surfaces of the aerodynamic structure, including the upper and lower surfaces.

A porous surface on an aircraft structure driven with oscillating positive and negative pressures is used as an active control device for attenuating negative aerodynamic interactions. The porous surfaces can be driven with positive and negative pressures either continuously or when predetermined flight conditions are present. The porous surfaces can be used on rotor blades to reduce BVI noise in descent flight conditions. The porous surfaces can be configured on rotor blades for affecting blade variable twist in accordance with various flight conditions, and can further be incorporated for reducing rotor hub vibrations as well. Porous surfaces placed on aerodynamic surfaces below the rotor blades of a tiltrotor aircraft can attenuate or eliminate download and fountain flow conditions. When placed on the trailing edges of a tip jet-exhaust driven rotor blade, the porous surfaces of the present invention can supplement the tip jet momentum of the exhaust to thereby reduce an amount of exhaust needed to drive the rotor blade.

A method for inhibiting dynamic stall of an airfoil by causing a fluid to flow out of at least one location on the airfoil. This location may be anywhere on the airfoil; but if the location is within one-quarter of the airfoil chord from the leading edge and the fluid flow has non-zero net mass flux, then the fluid flow is modulated at a frequency described by a Strouhal ratio greater than one.

An aircraft thickness / camber control device mounts to the lower surface of a airfoil configuration, for example on a fuselage, and extends along a longitudinal axis. The device, when deployed, generates expansions ahead of compressions generated by off-design conditions, inlet spillage for example, and enables maintenance of a low boom signature. The device, when positioned at appropriate locations, may also be used as a drag reduction device. The thickness / camber control device comprises a structural member capable of coupling to the airfoil at a position forward of the concentrated source of added compression and a control element. The control element is coupled to the structural member and controls the structural member to adjust thickness / camber of the configuration to cancel the far-field effect of the extra compression or concentrated pressure source.

The present invention relates to apparatus for influencing fluid flow over a surface, and more particularly, but not exclusively, to turbulent boundary layer flow drag reduction for an aircraft. The present invention provides such apparatus including a plasma generator comprising first and second spaced-apart independently controllable electrodes operable to cause a change in direction of the flow of the fluid over the surface.

A shock wave in a gas is modified by emitting energy to form an extended path in the gas; heating gas along the path to form a volume of heated gas expanding outwardly from the path; and directing a path. The volume of heated gas passes through the shock wave and modifies the shock wave. This eliminates or reduces a pressure difference between gas on opposite sides of the shock wave. Electromagnetic, microwaves and / or electric discharge can be used to heat the gas along the path. This application has uses in reducing the drag on a body passing through the gas, noise reduction, controlling amount of gas into a propulsion system, and steering a body through the gas. An apparatus is also disclosed.

A system for modifying a shock wave formed in a fluid by a body to modify effects of the shock wave on an object released from the body including heating fluid along a path to form a volume of heated fluid expanding outwardly from the path, the path extending from the body and through the shock wave, releasing the object from the body, and timing a release of the object relative to the heating of the fluid along the path to modify certain effects of the shock wave on the object.

A method of controlling movement of a vehicle including generating a fluid jet adjacent an aerodynamic surface of the vehicle, and changing at least one of a strength and a position of a shock wave traveling over the aerodynamic surface of the vehicle by directing the jet into a boundary layer flow of air attached to the aerodynamic surface.

A supersonic aircraft comprises a fuselage extending forward and aft along a longitudinal axis, the fuselage having a lower surface and an upper surface, a highly swept low aspect ratio wing coupled to the fuselage and having a forward leading edge and an aft trailing edge, an effector flap coupled to the wing trailing edge, and a tail empennage. The tail empennage is coupled to the fuselage aft of the wing on the fuselage upper surface in a position high relative to the wing. The tail empennage forms a channel region subject to complex shock patterns at transonic conditions. The aircraft further comprises an effector coupled to the tail empennage and a controller coupled to the effector flaps and the effectors. The controller further comprises a control process that reduces drag through channel relief by deflecting both the effector flap down and the effector up.

A system for modifying a shock wave formed in a fluid by a body to modify effects of the shock wave on information transferred to or from the body. The system includes an element for heating fluid along a path to form a volume of heated fluid expanding outwardly from the path, the path extending from the body and through the shock wave; mechanism for transferring the information to or from the body; and device for timing the transferring of the information relative to the heating of the fluid along the path to modify certain effects of the shock wave on the information.

A supersonic aircraft comprises a fuselage extending forward and aft along a longitudinal axis, the fuselage having a lower surface and an upper surface, a highly swept low aspect ratio wing coupled to the fuselage and having a forward leading edge and an aft trailing edge, an effector flap coupled to the wing trailing edge, and a tail empennage. The tail empennage is coupled to the fuselage aft of the wing on the fuselage upper surface in a position high relative to the wing. The tail empennage forms a channel region subject to complex shock patterns at transonic conditions. The aircraft further comprises an effector coupled to the tail empennage and a controller coupled to the effector flaps and the effectors. The controller further comprises a control process that reduces drag through channel relief by deflecting both the effector flap down and the effector up.

A wing for use on a supersonic aircraft that includes an inboard section a central section of the wing outboard of the inboard portion, and an outboard section. The outboard section can be a winglet oriented anhedrally relative to a lateral axis of the supersonic aircraft. Leading edge segments on the inboard section, central section and outboard winglet may have mounted thereon leading-edge flaps. These flaps are adjusted by a control system operable to reposition the leading-edge flaps in order to improve the aerodynamic performance of the supersonic aircraft. This winglet promotes sonic boom minimization. Further, the wing tip anhedral allows greater inboard dihedral. This effectively pushes lift aft for sonic boom and control purposes while minimizing the movement of control surfaces.

A shock wave in a gas is modified by emitting energy to form an extended path in the gas; heating gas along the path to form a volume of heated gas expanding outwardly from the path; and directing a path. The volume of heated gas passes through the shock wave and modifies the shock wave. This eliminates or reduces a pressure difference between gas on opposite sides of the shock wave. Electromagnetic, microwaves and / or electric discharge can be used to heat the gas along the path. This application has uses in reducing the drag on a body passing through the gas, noise reduction, controlling amount of gas into a propulsion system, and steering a body through the gas. An apparatus is also disclosed.

The invention discloses a method for reducing heat of a hypersonic velocity aircraft based on shock wave control, and relates to the fields of flowing control of hydromechanics and plasma physical application. The method comprises the following steps of using a high-pressure gas source gas spraying device or a plasma synthesizing flow jetting device to control shock waves in a hypersonic velocity aircraft flow field, lifting the leading head shock wave of the hypersonic velocity aircraft, and eliminating the leading shock of side wings, or producing reciprocating swinging, so as to eliminate or control the shock wave disturbance area, reduce the heat flow in key periods and key areas, and realize the heat protection of the aircraft. The method has the advantages that by adopting the outer flow field control type, the root produced by high heat flow is prevented or controlled, and the heat reducing and protection can be realized by reducing the production of heat flow in source; compared with the traditional method of performing passive heat protection on the cause of high heat flow, the material revolution is realized, and a new active heat protection control technique is realized.

A trailing edge wedge for an aircraft wing is attached to a bottom downwardly facing wing surface along the trailing edge of the wing so that an inner edge wedge surface is fully covered by the bottom wing surface. When the edge wedge is secured to a bottom surface of a wing, the wedge tip faces in the flight direction. The inner edge wedge surface and an opposite outer edge wedge surface enclose a wedge angle within the range of about 5° to about 20°. The edge wedge compensates for asymmetric aircraft characteristics due to unavoidable manufacturing tolerances.

A system for reducing overall drag of a mobile platform includes a surface on which an airflow forms a boundary layer and a generally normal shockwave. The airflow is at a first velocity that is one of transonic and supersonic. An oscillating jet injects and extracts a jet flow through the surface. The jet flow is at a second velocity that is substantially less than the first velocity. A recirculation region is upstream of the normal shockwave and is disposed at least partially in the boundary layer. The recirculation region is established at least by the oscillating jet. A generally oblique wave is established by the recirculation region and weakens the normal shockwave to reduce the overall drag experienced by the surface.

An aerodynamic structure comprising an array shock bumps (3, 10) extending from its surface, the array comprising: a first series of shock bumps; and one or more shock bumps positioned aft of the first series. Preferably at least one of the one or more shock bumps positioned aft of the first series is offset so that it is not positioned directly aft of any of the shock bumps in the first series. By providing an array of shock bumps instead of a single line, the first series of shock bumps and the one or more shock bumps positioned aft of the first series can be positioned to modify the structure of a shock which forms under a different respective condition.

The present invention provides flow field control techniques that adapt the aft body region flow field to eliminate or mitigate the development of massive separated flow field zones and associated unsteady vortical flow field structures. Embodiments of the present invention use one or more distributed arrays of flow control devices (submerged in the boundary layer) to create disturbances in the flow field that inhibit the growth of larger vortical structures and / or to energize the aft body shear layer to keep the shear layer attached the aft body surface. These undesirable aerodynamic phenomena produce increased vehicle drag which harms vehicle range, persistence, and loiter capabilities. Additionally, the unsteady nature of the turbulent vortical structures shed in the aft body wake region may produce increased dynamic buffeting and aft body heating by entraining nozzle jet exhaust (a.k.a. jet wash)—requiring additional structural support, shielding, and vehicle weight.

A supersonic aircraft comprises the fuselage extending forward and aft along a longitudinal axis, a wing coupled to the fuselage, and a canard. The canard is coupled onto the fuselage forward of the wing at an elevated position that enables stretching of the aircraft lifting length and forms an effective area distribution to attain a shaped sonic boom signature.

The invention discloses a laminated plate aircraft skin with lift augmentation, resistance reduction, flow control and deicing prevention functions. The laminated plate aircraft skin comprises a glass fiber layer (2) and an aluminum alloy (1), wherein the glass fiber layer (2) and the aluminum alloy (1) are alternative; the surface of the skin is provided with an insulating material layer (4) and an electrode layer; the electrode layer is divided into an upper layer and a lower layer; the upper layer and the lower layer are respectively positioned on two sides of the insulating material layer (4) and are asymmetrically distributed; the upper electrode layer (3) is exposed in air and is connected with an alternating current power supply; and the lower electrode layer (5) is positioned below the insulating material layer (4) and is grounded. The laminated plate aircraft skin can be directly applied to aircraft manufacturing engineering, cannot be damaged by high-speed airflow and cannot bring parasite drag due to the damage of the original pneumatic shape. The deicing prevention function is realized by utilizing plasma heating on air and the pulse expansion wave function, the flow separation can be effectively controlled, the stall is delayed, the highest lift coefficient is increased, and the pressure difference resistance is reduced.

An airfoil system includes an airfoil body and at least one flexible strip. The airfoil body has a top surface and a bottom surface, a chord length, a span, and a maximum thickness. Each flexible strip is attached along at least one edge thereof to either the top or bottom surface of the airfoil body. The flexible strip has a spanwise length that is a function of the airfoil body's span, a chordwise width that is a function of the airfoil body's chord length, and a thickness that is a function of the airfoil body's maximum thickness.

An aircraft capable of supersonic flight comprises a body portion including a fuselage, a wing, and an engine nacelle mounted below the wing. The aircraft may also include a high-mounted aft, tail. The area / lift distribution of the body portion is tailored to reduce sonic boom disturbance. The body portion further includes a blunt nose and a gull dihedral wing configuration that further reduces sonic boom disturbance and eases constraints on area / lift distribution tailoring. The gull dihedral wing or tail is configured to carry lifting force to its trailing edge to create an expansion at the aft end of the aircraft that reduces aft sonic boom ground shock strength. The volume of the mid-portion of the fuselage can be reduced above the wing to create a sloped surface that generates an expansion fan over the wings. The expansion fan lowers the pressure above the wing in the area covered by the expansion to reduce the pressure required on the upper and lower surfaces of the wing to generate the same lifting force.

The present invention provides flow field control techniques that adapt the aft body region flow field to eliminate or mitigate the development of massive separated flow field zones and associated unsteady vortical flow field structures. Embodiments of the present invention use one or more distributed arrays of flow control devices (submerged in the boundary layer) to create disturbances in the flow field that inhibit the growth of larger vortical structures and / or to energize the aft body shear layer to keep the shear layer attached the aft body surface. These undesirable aerodynamic phenomena produce increased vehicle drag which harms vehicle range, persistence, and loiter capabilities. Additionally, the unsteady nature of the turbulent vortical structures shed in the aft body wake region may produce increased dynamic buffeting and aft body heating by entraining nozzle jet exhaust (a.k.a. jet wash) —requiring additional structural support, shielding, and vehicle weight.

An aircraft thickness / camber control device mounts to the lower surface of a airfoil configuration, for example on a fuselage, and extends along a longitudinal axis. The device, when deployed, generates expansions ahead of compressions generated by off-design conditions, inlet spillage for example, and enables maintenance of a low boom signature. The device, when positioned at appropriate locations, may also be used as a drag reduction device. The thickness / camber control device comprises a structural member capable of coupling to the airfoil at a position forward of the concentrated source of added compression and a control element. The control element is coupled to the structural member and controls the structural member to adjust thickness / camber of the configuration to cancel the far-field effect of the extra compression or concentrated pressure source.

Method and arrangement for reducing the effects of a sonic boom created by an aerospace vehicle when said vehicle is flown at supersonic speed. The method includes providing the aerospace vehicle with a first spike extending from the nose thereof substantially in the direction of normal flight of the aerospace vehicle, the first spike having a second section aft of a first section that is aft of a leading end portion, the first and second sections having a second transition region therebetween and each of the sections having different cross-sectional areas, the leading end portion of the first spike tapering toward a predetermined cross-section with a first transition region between the predetermined cross-section and the first section. The first transition region is configured so as to reduce the coalescence of shock waves produced by the first spike during normal supersonic flight of the aerospace vehicle. A spike may also be included that extends from the tail of the aerospace vehicle to reduce the coalescence of shock waves produced by the spike during normal supersonic flight of the aerospace vehicle.