54 results about "Leading-edge extension" patented technology

An aerodynamic system providing an airfoil having a chord length, a leading edge, and a trailing edge. The airfoil further includes a first airfoil surface extending from the leading edge to the trailing edge, a second airfoil surface opposite the first airfoil surface, extending from the leading edge to the trailing edge, an injection opening in the first airfoil surface, and a recovery opening in the first airfoil surface located between the injection opening and the trailing edge. A pressurized fluid source is in fluid communication with the injection opening and a vacuum source is in fluid communication with the recovery opening. An exemplary use of the aerodynamic system of the present invention provides the ejection of a mass of fluid out of the injection opening along a surface of the airfoil and drawing a mass of fluid into the recovery opening.

A method of reducing crack propagation in an airfoil includes: providing an airfoil having a root spaced apart from a tip, spaced-apart leading and trailing edges, a suction side extending from the leading edge to the trailing edge, and an opposed pressure side extending from the leading edge and the trailing edge, supporting the airfoil against bending loads; and burnishing the airfoil using a burnishing element, so as to create at least one burnished section of residual compressive stress, the at least one burnished section being located adjacent the leading edge and spaced from the leading edge by an offset distance selected so as to avoid deformation of the leading edge.

A method of reducing crack propagation in an airfoil includes: providing an airfoil having a root spaced apart from a tip, spaced-apart leading and trailing edges, a suction side extending from the leading edge to the trailing edge, and an opposed pressure side extending from the leading edge and the trailing edge, supporting the airfoil against bending loads; and burnishing the airfoil using a burnishing element, so as to create at least one burnished section of residual compressive stress, the at least one burnished section being located adjacent the leading edge and spaced from the leading edge by an offset distance selected so as to avoid deformation of the leading edge.

The invention relates to multimode aircraft which can be operated at supersonic and subsonic flying speeds in a wide range of flying altitudes. The primary field of use of the invention is multimode super-manoeuvrable aircraft capable of cruising at supersonic speed and having low radar visibility. The invention is intended to achieve the technical result of producing an aircraft which has low radar visibility, is highly manoeuvrable at large angles of attack, has a high aerodynamic quality at supersonic speeds while retaining a high aerodynamic quality in subsonic modes, and is capable of accommodating a large load in its internal holds. The aircraft with an integral aerodynamic configuration comprises a fuselage (1) with a leading-edge extension (2), a wing, the main planes (3) of which are continuously adjacent to the fuselage (1), a fully rotatable horizontal tail unit (FRHTU) (4), and a fully rotatable vertical tail unit (FRVTU) (5). The mid-fuselage is flattened and formed longitudinally by a set of aerodynamic profiles. The engines are arranged in engine nacelles (6) spaced apart from one another horizontally, and the axes of the engines are oriented at an acute angle to the plane of symmetry of the aircraft in the flying direction. The leading-edge extension (2) comprises controllable rotatable parts (8).

A vertical take-off and landing (VTOL) aircraft comprises (1) a fuselage having a front end, a rear end and two lateral sides, the fuselage defining a substantially horizontal central longitudinal axis of the aircraft; (2) an aircraft tail arranged at the rear end of the fuselage and including a rudder and an elevator on each side of the fuselage with movable surfaces for controlling the aircraft; and (3) a wing on each side of the fuselage having a front edge, a trailing edge and an upper surface extending from the front edge to the trailing edge. According to the invention, means are provided for increasing the speed of an airstream flowing over the upper surface of each wing, and an air deflector is disposed on each side of the fuselage between the trailing edge of each wing and the aircraft tail. These deflectors each have a deflection surface to deflect downward the airstream flowing over the upper surface of the wing to act as a “brake” to the forward motion of the aircraft and to provide additional lift so that the aircraft may hover.

An aircraft includes an elongated wing having a leading edge extending to a trailing edge, an upper surface extending from the leading edge to the trailing edge, and a lower surface extending from the leading edge to the trailing edge; an air passage forming an entrance opening at the leading edge and configured to extend to an exit opening positioned at the upper surface; and an airflow stop configured to control airflow movement through the air passage.

An aeroplane turbojet low-pressure compressor vane includes a leading edge, a trailing edge, a surface, and an extrados surface which extend from the leading edge to the trailing edge. To combat the presence and the formation of ice, the vane is provided with an electric de-icing device with a thermistor. The thermistor forms a heating electrical track suitable for de-icing the vane. The present application also proposes a method for producing a turbine engine vane.

A vertical take-off and landing (VTOL) aircraft comprises (1) a fuselage having a front end, a rear end and two lateral sides, the fuselage defining a substantially horizontal central longitudinal axis of the aircraft; (2) an aircraft tail arranged at the rear end of the fuselage and including a rudder and an elevator on each side of the fuselage with movable surfaces for controlling the aircraft; and (3) a wing on each side of the fuselage having a front edge, a trailing edge and an upper surface extending from the front edge to the trailing edge. According to the invention, means are provided for increasing the speed of an airstream flowing over the upper surface of each wing, and an air deflector is disposed on each side of the fuselage between the trailing edge of each wing and the aircraft tail. These deflectors each have a deflection surface to deflect downward the airstream flowing over the upper surface of the wing to act as a “brake” to the forward motion of the aircraft and to provide additional lift so that the aircraft may hover.

A fairing for streamlining the junction between the horizontal stabilizer and the fuselage or the fin of an aircraft, comprising an upper fairing shell, a lower fairing shell, a leading edge extension, and several fittings to join the upper and lower fairing shells to the horizontal stabilizer. The upper and lower fairing shells are solid laminates of composite material that are designed with a shape, composition, and thickness such that they can be elastically deformed when they are fitted to the horizontal stabilizer, providing a contact force permitting them to lay permanently in contact with the fuselage or fin. A process for manufacturing the fairing shells.

A boundary layer control apparatus including a duct having a streamlined shape which is mounted in spaced apart relation upwind from the wing and extending generally parallel to the leading edge of the wing. The duct has a thickness substantially less than the thickness of the wing to which it is attached. The duct of the present invention includes an opening on the downwind side of the duct for injecting compressed air toward a stagnation line associated with the aircraft wing. The duct according to the present invention can be pivoted during flight so that the air injected from the duct is always directed at the stagnation line, which moves relative to the aircraft wing as aircraft speed increases.

An aerofoil and an aerofoil assembly, in particular an aerofoil with improved stagnation zone cooling and an aerofoil assembly comprising such an aerofoil. The aerofoil is an aerofoil for a gas turbine engine comprising a pressure surface, a suction surface, a leading edge, a trailing edge, a stagnation zone located in the region of the leading edge, and an elongate channel running along the leading edge at the stagnation zone.

In a featured embodiment, a gas turbine engine component comprises an airfoil having a leading edge, a trailing edge, and pressure and suction side walls extending from the leading edge to the trailing edge. The airfoil extends from a base to a tip. A shelf is formed in the tip, and extends from the pressure side wall, around the leading edge, to the suction side wall.

A system for detecting misalignment of an aero surface relative to other aero surfaces during simultaneous deployment of said aero surfaces is disclosed. It comprises a generator configured to shine a laser through aligned apertures in a series of aero surfaces towards a reflector during deployment of said aero surfaces, a receptor to detect reflection of said laser from said reflector back through said apertures to the generator and, a controller operable to terminate further deployment of said aero surfaces in the event that no reflection is detected by said receptor, or if an actual time taken for the reflected beam of light to be detected by the receptor differs from a predetermined time. In a preferred embodiment the aero sufaces are slats extending from the leading edge of an aircraft wing.

A fan section for a gas turbine engine is provided. The fan section having a fan hub with a slot and a fan blade with an airfoil extending from a root to a tip, the airfoil having a leading edge and the root is received in the slot, wherein a first platform is secured to the fan hub and arranged between adjacent fan blades of the fan section, the first platform having a first platform seal including a platform seal leading edge and a base is secured to a side of the first platform and a first winglet extends from the platform seal leading edge and contacts the airfoil leading edge.

An aircraft comprises a fuselage, a wing (2) attached to the fuselage, a wing tip device (6) attached to a wing end (4) of the wing (2), at least one additional wing-like element (14) having a wing root (22), a wing leading edge (18) and a wing trailing edge (20), and a torque control device having a rotatable interface means. The torque control device is adapted for rotatably supporting the wingroot of the wing-like element on the interface means under creation of a rotational axis (16) extending from the interface means into the wing-like element (14), about which rotational the wing-like element (14) is rotatable. The wing-like element (14) is adapted to induce a rotation around the rotational axis (16) in an air flow. The torque control device is adapted to limit the degree of rotation depending on a torque introduced into the interface means by the wing-like element (14) and the wing root (22) of the at least one wing-like element (14) is coupled with at least one of the wing tipdevice (6), the wing (2) and the fuselage through the torque control device such that the leading edge (18) extends into an airflow surrounding the aircraft.

In a first aspect, there is a method for improving a lift to drag ratio of a rotor blade, including providing a blade member having a leading edge and a trailing edge; providing a leading edge extension member; and coupling the leading edge extension member to a portion of the leading edge of the blade member to form the rotor blade. In a second aspect, there is a rotor blade including a blade member having a leading edge, and a trailing edge; and a leading edge extension member disposed on the leading edge of the blade member, wherein the leading edge extension member is configured to extend the chord length of at least a portion of the rotor blade. In a third aspect, there is a leading edge extension member for a rotor blade including a convex exterior surface configured to extend at least a portion of the chord length of the rotor blade.

A puncturing feature is provided on the deployable wings of a precision guidance kit of a rocket to facilitate breaking through the wing slot seal that protecting the wings prior to deployment. The puncturing feature may be a sharp region on the leading edge near the wing tip that first contacts the wing slot seal. The puncturing feature may be a sharp edge extended along the leading edge of the wing to provide a cutting action as the wing passes through the wing slot seal.

An aircraft includes a fuselage, a wing attached thereto, a wing tip device attached to a wing end of the wing (2), a wing-like element having a wing root, a wing leading edge and a wing trailing edge, and a torque control device having a rotatable interface means. The torque control device is adapted for rotatably supporting the wing root of the wing-like element on the interface means about a rotational axis extending from the interface means into the wing-like element and to limit the degree of rotation depending on a torque introduced into the interface means by the wing-like element. The wing-like element is adapted to induce a rotation around the rotational axis in an air flow. The wing root is coupled with the wing tip device, the wing or the fuselage through the torque control device such that the leading edge extends into an airflow surrounding the aircraft.

The invention belongs to the field of wind power generation, and relates to a wind turbine blade with a leading edge extension power increasing device, which comprises an original blade airfoil, the extension power increasing device is additionally arranged at the leading edge of the original blade airfoil, and the shape of the contact surface of the extension power increasing device and the leading edge of the original blade airfoil is the same as that of the leading edge of the original airfoil; the extension power increasing device is in a downward turning posture; the extension power increasing device and the original blade airfoil form a new airfoil profile, the lift coefficient of the new airfoil profile is larger than that of the original blade airfoil, and the resistance coefficient of the new airfoil profile is smaller than that of the original blade airfoil. The relative thickness of the airfoil is reduced, so that the aerodynamic performance of the airfoil is improved, the lift coefficient of the airfoil is increased, and the resistance coefficient is reduced; and meanwhile, the chord length of the airfoil is increased, the stress area is increased, and the blade output is improved.

The present invention involves a method for manufacturing a blade (1) for a propeller, which blade (1) has a leading edge (2) and a trailing edge, the method comprising the steps of: forming a conduit in the blade (1), making a plurality of holes (7) through which the conduit (6) communicates with the exterior of the blade (1), and providing a blade blank having an edge part receiving surface (4) extending along at least a major part of the leading edge (2) of the blade (1) to be manufactured, wherein forming a conduit (6) comprises building up an edge part (3) onto the edge part receiving surface (4) by a wire-based additive manufacturing process, wherein the additive manufacturing process is adapted to form the conduit (6) at least partly delimited by the edge part (3) and extending along the leading edge (2) of the blade (1) to be manufactured.

A tip structure for a turbine blade comprises a tip surface of an airfoil that extends from a leading edge to a trailing edge and from a pressure side to a suction side of the airfoil. A suction side rail protrudes from the tip surface and comprises: a trailing portion extending flush with the suction side of the airfoil from an origination location at or near the trailing edge to an intermediate location on the suction side; and a leading portion extending from the intermediate location across the tip surface so as to create a suction side shelf region between the leading portion of the suction side rail and the suction side of the airfoil. A pressure side rail protrudes from the tip surface and comprises: a trailing segment extending flush with the pressure side of the airfoil from an origination location at or near the trailing edge to a middle location on the pressure side; and a leading segment extending from the middle location across a camber line on the tip surface so as to create a pressure side shelf region between the leading segment of the pressure side rail and the pressure side of the airfoil.

A fairing for streamlining the junction between the horizontal stabilizer and the fuselage or the fin of an aircraft, comprising an upper fairing shell, a lower fairing shell, a leading edge extension, and several fittings to join the upper and lower fairing shells to the horizontal stabilizer. The upper and lower fairing shells are solid laminates of composite material that are designed with a shape, composition, and thickness such that they can be elastically deformed when they are fitted to the horizontal stabilizer, providing a contact force permitting them to lay permanently in contact with the fuselage or fin. A process for manufacturing the fairing shells.

The invention provides a design method of a composite airfoil fan blade. The method comprises the steps that a standard aerial aerodynamic force airfoil is adopted as an airfoil blade main body of a composite airfoil; the front edge of the airfoil blade main body extends forward for airfoil superposition; or the back edge of the airfoil blade main body extends backward for the airfoil superposition; or the front edge of the airfoil blade main body extends forward and the back edge of the airfoil blade main body extends backward for the airfoil superposition simultaneously; the airfoil superposition is conducted on the suction surface of the airfoil blade main body; and the novel airfoil is composited. The composite airfoil can meet aerodynamic force requirements of different impellers by cutting and trimming a front edge extension airfoil and a back edge extension airfoil, changing a blade angle, changing a relative torsion angle, adjusting blade numbers, changing chord length according to a used calculation flow pattern and a use occasion. The method realizes mass production of blades, is high in generality, is particularly applied to the various fan and minitype wind power industries, and has a very good popularization value.

The invention discloses a blended wing body aircraft including a center body having a lower side and an upper side opposed to the lower side. The center body has a central chord extending from a leading edge to a trailing edge of the center body. The lower side has a lowest point located between a first location forward of a pivot point about which the aircraft rotates during take-off and a secondlocation aft of the pivot point. The first location is at a first distance corresponding to about 10% of a length of the central chord forward of the pivot point.

The invention provides a design method of a composite airfoil fan blade. The method comprises the steps that a standard aerial aerodynamic force airfoil is adopted as an airfoil blade main body of a composite airfoil; the front edge of the airfoil blade main body extends forward for airfoil superposition; or the back edge of the airfoil blade main body extends backward for the airfoil superposition; or the front edge of the airfoil blade main body extends forward and the back edge of the airfoil blade main body extends backward for the airfoil superposition simultaneously; the airfoil superposition is conducted on the suction surface of the airfoil blade main body; and the novel airfoil is composited. The composite airfoil can meet aerodynamic force requirements of different impellers by cutting and trimming a front edge extension airfoil and a back edge extension airfoil, changing a blade angle, changing a relative torsion angle, adjusting blade numbers, changing chord length according to a used calculation flow pattern and a use occasion. The method realizes mass production of blades, is high in generality, is particularly applied to the various fan and minitype wind power industries, and has a very good popularization value.

The invention relates to multimode aircraft operated at supersonic and subsonic flight speeds in a wide range of flight altitudes. The invention is particularly applicable in multimode super-maneuverable aircraft capable of cruising at supersonic speed and having low radar signature. The technical effect to be achieved by the invention is to provide an aircraft which has low radar signature, is super-maneuverable at high angles of attack, has high aerodynamic efficiency at supersonic speeds while keeping high aerodynamic efficiency at subsonic modes, and can accommodate bulky cargo in its internal compartments. The integral aerodynamic configuration aircraft comprises fuselage (1) with strake (2), wing with panels (3) smoothly blended into fuselage (1), all-moving horizontal tail (AMHT) (4), and all-moving vertical tail (AMVT) (5). Mid-fuselage is flattened and formed longitudinally by a set of aerodynamic profiles. Engines are arranged in engine nacelles (6) spaced apart from each other horizontally, and centerlines of engines are oriented at an acute angle to the plane of symmetry of the aircraft in the direction of flight. Strake (2) comprises controlled pivoting portions (8).