| 序号 | 专利名 | 申请号 | 申请日 | 公开（公告）号 | 公开（公告）日 | 发明人 |
|---|---|---|---|---|---|---|
| 121 | Aircraft structure to reduce sonic boom intensity | US3794274D | 1969-12-19 | US3794274A | 1974-02-26 | EKNES O |
| The structure of a fuselage of an aircraft having a plurality of apertures therein which permit high pressure air to be received within the confines of the exterior portion of said fuselage, and means within said fuselage adapted to direct and control the flow of said air until the energy thereof has been substantially reduced or same is emitted through carefully situated ports. The effective cross-sectional area of the nose portion of said fuselage is thereby substantially reduced so as to cause a similar reduction of the drag upon said aircraft.
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| 122 | Sonic boom eliminator | US3776489D | 1972-01-07 | US3776489A | 1973-12-04 | WEN L; BEARD M |
| A structure for eliminating sonic booms developed by the major sub-structures of a supersonic aircraft in accordance with the selective positioning of the exhaust gases of the aircraft, air inlets and outlets and conduits between the inlets and outlets. A supersonic aircraft using the structure is disclosed wherein the major sub-structures of the aircraft incorporate exhaust gas positioning, inlets, outlets and conduits between the inlets and outlets.
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| 123 | Device for reducing the supersonic boom caused by aircraft | US3655147D | 1968-04-25 | US3655147A | 1972-04-11 | PREUSS HEINZ |
| The invention describes means for reducing the supersonic boom caused by aircraft. Underneath the aircraft reflecting surfaces are provided in the area of the maximum pressure difference of the Mach cone created by the aircraft.
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| 124 | Air deflector for supersonic aircraft | US3425650D | 1967-10-02 | US3425650A | 1969-02-04 | SILVA JOSEPH |
| 125 | SYSTEM AND METHOD FOR CONTROLLING A PRESSURE FIELD AROUND AN AIRCRAFT IN FLIGHT | EP15766673.6 | 2015-02-26 | EP3180244A1 | 2017-06-21 | CONNERS, Timothy; KNIGHT, Michael; COWART, Robert |
| A system for controlling a pressure field around an aircraft in flight is disclosed herein. In a non-limiting embodiment, the system includes, but is not limited to, a plurality of pressure sensors that are arranged on the aircraft to measure the pressure field. The system further includes, but is not limited to, a controller that is communicatively coupled with the plurality of pressure sensors. The controller is configured to receive information that is indicative of the pressure field from the plurality of pressure sensors. The controller is also configured to determine when the pressure field deviates from a desired pressure field based on the information. The controller is also configured to transmit an instruction to a movable component onboard the aircraft that will cause the movable component to move in a manner that reduces the deviation. | ||||||
| 126 | SWEPT GRADIENT BOUNDARY LAYER DIVERTER | EP16001034.4 | 2016-05-06 | EP3109153A1 | 2016-12-28 | Troia, Trajaen J.; Mangus, John F. |
A swept gradient air boundary layer diverter for an aircraft. The aircraft includes a fuselage and an air inlet for an engine of the aircraft, where the air inlet includes a cowl at a leading edge of the inlet. The diverter includes a V-shaped ramp portion formed in the fuselage in an area proximate to and in front of the cowl where the ramp portion extends downward away from an outer surface of the fuselage towards an inside of the aircraft. The diverter also includes a V-shaped trough portion formed into the fuselage and being positioned adjacent to and integral with the ramp portion between the ramp portion and the air inlet. Air flowing over the fuselage towards the cowl is expanded and compressed by the ramp portion and the trough portion so as to create pressure gradients that generate vortices to redirect boundary layer airflow around the air inlet. |
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| 127 | Synthetic jet muffler | EP14183188.3 | 2014-09-02 | EP2873609A1 | 2015-05-20 | Griffin, Steven F. |
A synthetic jet muffler (200) includes an exit end (210), a propagation path (215) for conducting a first sound wave (220) emitted by a synthetic jet generator (225) to the exit end (210), and a shroud (240) for conducting a second sound wave (245) emitted from the synthetic jet generator (225) in a direction opposite to the first sound wave (220) to the exit end (210), wherein the shroud (240) is disposed so that the first and second sound waves (220, 245) travel different distances to effect noise cancellation at the exit end (210).
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| 128 | System and method for minimizing wave drag through bilaterally asymmetric design | EP13191646.2 | 2013-11-05 | EP2738092A2 | 2014-06-04 | Pflug, William; Tillotson, Brian J. |
An air vehicle (100, 102) having a bilaterally asymmetric configuration may include a body (104) having a longitudinal axis (106). The air vehicle (100, 102) may further include longitudinally offset engine nacelles (234, 236), asymmetrically lengthened engine nacelles (232), and/or longitudinally offset protruding aerodynamic surfaces (190). ( |
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| 129 | Airfraft with a flow control system around an engine exhaust and Method | EP06425228.1 | 2006-04-03 | EP1710156B1 | 2014-03-05 | Miller, Daniel N.; Young, David D. |
| 130 | METHOD AND APPARATUS FOR CONTROLLING FLOW ABOUT A TURRET | EP11727877.0 | 2011-06-13 | EP2595880A1 | 2013-05-29 | ULLMAN, Alan, Z. |
| Methods and apparatus are provided to control flow separation of an ambient flow along a surface and about a turret (12), such as by reducing flow separation aft of the turret. By reducing flow separation, the resulting turbulence may be similarly reduced such that the performance of a system, such as a laser system, housed by the turret may be improved. To reduce flow separation, a motive flow may be provided by ejector nozzles (32) that open through the surface and are positioned proximate to and aft of the turret relative to the ambient flow. The motive flow has a greater velocity than the ambient flow to thereby create a region aft of the turret of reduced pressure relative to an ambient pressure. Within this region of reduced pressure aft of the turret, a portion of the ambient flow mixes with the motive flow, thereby reducing or eliminating flow separation. | ||||||
| 131 | PROPULSION DEVICE USING FLUID FLOW | EP10815596.1 | 2010-09-07 | EP2479100A2 | 2012-07-25 | Kim, Nak Hwe |
The present invention relates to a propulsion device using fluid flow, which quickly discharges the vortex flow generated on an upper surface of the propulsion device to the outside to improve the propulsion of the product provided with the propulsion device. For this purpose, the propulsion device of the present invention comprises: a fluid storage unit in which a downwardly curved fluid storage surface is formed between a first inlet line and a first outlet line such that a fluid storage space is formed on the fluid storage surface; and a fluid flow unit in which a downwardly curved fluid flow surface is formed between a second inlet line and a second outlet line which are outwardly and backwardly inclined such that a fluid flow space is formed on the fluid flow surface, wherein the fluid flow surface adjacent to the second outlet line becomes gradually flattened as it extends outwardly. The above-described configuration of the propulsion device of the present invention is advantageous in that the fluid introduced into the fluid storage space and the fluid flow space flows in a vortex to increase pressure, and the fluid flow space gradually narrows as it extends toward an end of the fluid flow surface so as to quickly discharge the fluid to the end of the fluid flow surface, thus increasing fluid flow velocity and improving the propulsion and thrust of transportation means provided with the propulsion device. |
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| 132 | AERODYNAMIC STRUCTURE WITH NON-UNIFORMLY SPACED SHOCK BUMPS | EP09715682.2 | 2009-02-17 | EP2250088A2 | 2010-11-17 | WOOD, Norman |
| An aerodynamic structure comprising a series of shock bumps (3) extending from its surface. The shock bumps are distributed across the structure with a non-uniform spacing (dl, d2) between the centres and/or edges of adjacent bumps. The non-uniform spacing between the bumps can be arranged to give maximum wave drag alleviation for the minimum number of bumps as a function of the shock strength across the span, leading to minimum wing weight penalties for a given amount of wave drag alleviation. | ||||||
| 133 | AERODYNAMIC STRUCTURE WITH SERIES OF SHOCK BUMPS | EP09715558.4 | 2009-02-17 | EP2250087A2 | 2010-11-17 | WOOD, Norman |
| An aerodynamic structure (1) comprising a series of shock bumps (3a, 3b, 3c) extending from its surface. The shock bumps are distributed along a line (7) with a smaller mean angle of sweep than an unperturbed shock (4) which would form adjacent to the surface during transonic movement of the structure in the absence of the shock bumps. Instead of being distributed along the line of the unperturbed shock, the shock bumps are distributed along a line which is less swept than the mean angle of sweep of the unperturbed shock. When the structure is moved at a transonic speed; a shock forms adjacent to its surface and the shock bumps perturb the shock (9) so as to reduce its angle of sweep. | ||||||
| 134 | FUSELAGE SHAPING AND INCLUSION OF SPIKE ON A SUPERSONIC AIRCRAFT FOR CONTROLLING AND REDUCING SONIC BOOM | EP03707587.6 | 2003-01-30 | EP1478569B1 | 2009-09-23 | HENNE, Preston; HOWE, Donald; WOLZ, Robert; HANCOCK, Jimmy, Jr. |
| 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. | ||||||
| 135 | WING GULL INTEGRATION FOR NACELLE CLEARANCE, COMPACT LANDING GEAR STOWAGE, AND SONIC BOOM REDUCTION | EP04788816 | 2004-09-17 | EP1692038A4 | 2009-08-05 | QUAYLE BRIAN; MORGENSTERN JOHN M; ARSLAN ALAN E |
| A supersonic cruise configuration aircraft (100) comprises a fuselage (142) extending on a longitudinal axis from a forward nose (110) to an aft tail (114), and a wing (104) coupled at an inboard section to the fuselage (142) and extending to an outboard tip, and having a leading edge and a trailing edge. The aircraft (100) further comprises a landing gear (146) that is coupled to the wing (104) and capable of stowing into the wing (104) and fuselage (142) on retraction. The landing gear (146) has a landing gear strut. The wing (104) is gulled with a dihedral (152) at an angle that is increased inboard and aligns with the retracted landing gear (146). The wing (104) has a minimum thickness sufficient to enclose the landing gear (146). | ||||||
| 136 | Conformal aero-adaptive nozzle / aftbody | EP06425228.1 | 2006-04-03 | EP1710156A2 | 2006-10-11 | Miller, Daniel N.; Young, David D. |
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.
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| 137 | AIRCRAFT WING AND FUSELAGE CONTOURS | EP00919440.8 | 2000-03-16 | EP1169224B1 | 2006-01-11 | Tracy, Richard R. |
| A wing in combination with a fuselage having a body which is elongated in the direction of flight, the wing having physical parameters [comprising a wing having a relatively unswept and sharp leading edge, smooth convex chordwise contour over a majority of its surface from the leading edge, and a thickness to chord ratio of about 2% or less as a spanwise average, beyond a spanwise distance from the fuselage centerline of not more than about C/2 beta on each side of the body, where beta = 2ROOT +E,rad M2-1+EE , M=cruise Mach number, C=wing chord at centerline, and wherein the thickness to chord ratio over the above referenced spanwise distance is increased substantially over 2%,] to benefit strength, stiffness, weight, interior volume, and laminar boundary layer stability, and limited only by the extent to which the increase in volume drag which would otherwise occur is substantially eliminated by the body having indentation proximate the wing, the wing thickening and body indentation being characterized by one of the following: a) selection of a wing planform airfoil and thickness distribution, the fuselage then indented to minimize or reduce the combined volume wave drag, or optimize a design figure of merit, b) selection of a fuselage longitudinal distribution of cross section areas, the wing thickness then distributed spanwise so as to reduce the combined volume wave drag, or optimize a design figure of merit, c) reduction of drag, or optimization of other figure of merit such as weight or cost, in accordance with variation in both the fuselage longitudinal distribution of cross section area and wing planform airfoil and spanwise thickness distribution. | ||||||
| 138 | AIRCRAFT WING AND FUSELAGE CONTOURS | EP00919440 | 2000-03-16 | EP1169224A4 | 2004-11-10 | TRACY RICHARD RIPLEY |
| A wing in combination with a fuselage having a body which is elongated in the direction of flight, the wing having physical parameters [comprising a wing having a relatively unswept and sharp leading edge, smooth convex chordwise contour over a majority of its surface from the leading edge, and a thickness to chord ratio of about 2% or less as a spanwise average, beyond a spanwise distance from the fuselage centerline of not more than about C/2 beta on each side of the body, where beta = 2ROOT +E,rad M2-1+EE , M=cruise Mach number, C=wing chord at centerline, and wherein the thickness to chord ratio over the above referenced spanwise distance is increased substantially over 2%,] to benefit strength, stiffness, weight, interior volume, and laminar boundary layer stability, and limited only by the extent to which the increase in volume drag which would otherwise occur is substantially eliminated by the body having indentation proximate the wing, the wing thickening and body indentation being characterized by one of the following: a) selection of a wing planform airfoil and thickness distribution, the fuselage then indented to minimize or reduce the combined volume wave drag, or optimize a design figure of merit, b) selection of a fuselage longitudinal distribution of cross section areas, the wing thickness then distributed spanwise so as to reduce the combined volume wave drag, or optimize a design figure of merit, c) reduction of drag, or optimization of other figure of merit such as weight or cost, in accordance with variation in both the fuselage longitudinal distribution of cross section area and wing planform airfoil and spanwise thickness distribution. | ||||||
| 139 | AIRCRAFT WING AND FUSELAGE CONTOURS | EP00919440.8 | 2000-03-16 | EP1169224A2 | 2002-01-09 | TRACY, Richard Ripley |
| A wing in combination with a fuselage having a body which is elongated in the direction of flight, the wing having physical parameters [comprising a wing having a relatively unswept and sharp leading edge, smooth convex chordwise contour over a majority of its surface from the leading edge, and a thickness to chord ratio of about 2% or less as a spanwise average, beyond a spanwise distance from the fuselage centerline of not more than about C/2 beta on each side of the body, where beta = 2ROOT +E,rad M2-1+EE , M=cruise Mach number, C=wing chord at centerline, and wherein the thickness to chord ratio over the above referenced spanwise distance is increased substantially over 2%,] to benefit strength, stiffness, weight, interior volume, and laminar boundary layer stability, and limited only by the extent to which the increase in volume drag which would otherwise occur is substantially eliminated by the body having indentation proximate the wing, the wing thickening and body indentation being characterized by one of the following: a) selection of a wing planform airfoil and thickness distribution, the fuselage then indented to minimize or reduce the combined volume wave drag, or optimize a design figure of merit, b) selection of a fuselage longitudinal distribution of cross section areas, the wing thickness then distributed spanwise so as to reduce the combined volume wave drag, or optimize a design figure of merit, c) reduction of drag, or optimization of other figure of merit such as weight or cost, in accordance with variation in both the fuselage longitudinal distribution of cross section area and wing planform airfoil and spanwise thickness distribution. | ||||||
| 140 | VIBRATION-DRIVEN ACOUSTIC JET CONTROLLING BOUNDARY LAYER SEPARATION | EP00915876.7 | 2000-02-25 | EP1159534A1 | 2001-12-05 | Miller, Robin Mihekun; Tunkel, Roman N. |
| An acoustic jet disposed within an aerodynamic surface (9), such as a wing or a blade, has a resilient wall (16) supporting a mass (17). Vibrations of the blade cause oscillatory pressure waves within the acoustic jet, the nozzle (20) of which directs fluid particles having high momentum flux essentially tangentially into the boundary layer of the suction surface (22) of the blade, the resonant cavity of the synthetic jet being replenished with particles having low momentum flux drawn from the flow in a direction normal to the surface, thereby to provide a net time averaged flow of fluid particles of increasing momentum flux into the boundary layer to defer or prevent the onset of boundary layer separation. Single and double chambers drive nozzles separated streamwise or spanwise on airfoils (blades, wings) and fuselages. Applications include helicopters, airplanes, air moving machines and wind energy electric power generators. | ||||||
