US3471107A - Stabilizing the vortices over a thin delta wing - Google Patents

Stabilizing the vortices over a thin delta wing Download PDF

Info

Publication number
US3471107A
US3471107A US615230A US3471107DA US3471107A US 3471107 A US3471107 A US 3471107A US 615230 A US615230 A US 615230A US 3471107D A US3471107D A US 3471107DA US 3471107 A US3471107 A US 3471107A
Authority
US
United States
Prior art keywords
wing
vortices
vortex
suction surface
attack
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US615230A
Other languages
English (en)
Inventor
Kjell Torsten Ornberg
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Saab AB
Original Assignee
Saab AB
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Saab AB filed Critical Saab AB
Application granted granted Critical
Publication of US3471107A publication Critical patent/US3471107A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
    • B64C23/06—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
    • B64C23/005—Influencing air flow over aircraft surfaces, not otherwise provided for by other means not covered by groups B64C23/02 - B64C23/08, e.g. by electric charges, magnetic panels, piezoelectric elements, static charges or ultrasounds
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C5/00—Stabilising surfaces
    • B64C5/10—Stabilising surfaces adjustable
    • B64C5/12—Stabilising surfaces adjustable for retraction against or within fuselage or nacelle
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/10—Drag reduction

Definitions

  • FIGURES 1, 5 and 11 of the accompanying drawings One expedient for accomplishing this is disclosed in Swedish Patent No. 160,134, and is illustrated in FIGURES 1, 5 and 11 of the accompanying drawings. It consists in providing one or more plates or fences 5 that project downwardly from the pressure surface of the wing 7 near each of its leading edges 8, such plates being located intermediate the wing root 9 and the wing tip 10 and edgewise aligned with the normal direction of flight. Each such plate produces a discontinuity in the normal outwardly directed flow at the pressure surface of the wing, and thereby generates an outboard leading edge vortex 11 which is additional to the inboard vortex 12 that originates near the apex 13 of the wing.
  • a plurality of vortices is also generated at each leading edge of a wing having a so called broken delta planform, wherein each leading edge has a forward portion which is disposed at one acute angle to the longitudinal centerline of the wing and a rearward portion which is at a diflerent acute angle to said centerline.
  • the break between the forward and rearward portions of each leading edge can define either an outward or an inward obtuse angle, the former being illustrated in FIGURE 2 and the latter being illustrated in FIGURE Patented Oct. 7, 1969 3.
  • the change of the sweep back angle of the wing at the break 14 and 14 produces a discontinuous change in the increase of the shearing strain of the vortex sheet separating from the leading edge, so that if the change in sweep-back angle at the break is sufficiently large, the vortex sheet can be rolled up into two separate leading edge vortices 11 and 12 at each side of the centerline, as illustrated in FIGURES 2 and 3, instead of only one vortex being formed at each side as in the case of a delta wing having straight leading edges.
  • a plurality of vortices at each side of the centerline is also produced over the main wing 107 of a canard aircraft, as illustrated in FIGURE 4, wherein a smaller secondary wing 207 is located in front of the main wing in accordance with the principles of US. Patent No, 3,188,- 022.
  • the two leading edge vortices 12' of the forward wing are bound to the two leading edge vortices 11' of the main wing as they pass over the latter.
  • the general object of the present invention is to provide a simple but effective means for preventing such interference between adjacent corotating leading edge vortices at each side of the centerline of a thin, sharply sweptback delta wing, to thereby prevent the above-mentioned undesirable loss of lift and disturbances to pitch and roll trim.
  • Another object of this invention is to provide a thin, sharply swept-back wing having smooth curves of lift coeflicient vs. angle of attack and moment coefficient vs. angle of attack through a substantially large range of angles of attack.
  • FIGURE 1 is a plan view of a thin delta wing, seen from above, having plates or fences as hereinabove described for generating a plurality of leading edge vortices, the vortices thus generated being shown diagrammatically;
  • FIGURE 2 is a view similar to FIGURE 1 but showing an outwardly broken delta wing and the vortices that obtain over its suction surface;
  • FIGURE 3 is a view similar to FIGURE 1 but illustrating an inwardly broken delta wing and the vortices that obtain over its suction surface;
  • FIGURE 4 is a top plan view of a delta wing canard aircraft embodying the principles of Patent No. 3,188,022, the vortices generated by the wing arrangement being illustrated diagrammatically;
  • FIGURE 5 is a diagrammatic sectional view through a thin delt wing, taken on a plane normal to the longitudinal centerline of the wing, illustrating a system of pairs of corotating vortices over its suction surface when the wing is at a low angle of attack and indicating the relative forces acting upon such vortices by reason of their mutual interference;
  • FIGURE 6 is a view similar to FIGURE 5 but illustrating a condition, such as occurs at high angles of attack, wherein mutual interference between adjacent corotating vortices has caused them to roll up on one another and to rotate about a common axis;
  • FIGURE 7 is a view corresponding to FIGURE 5 but showing the conditions that exist over a wing embodying the present invention at low angles of attack;
  • FIGURE 8 is a view corresponding to FIGURE 6 but showing the conditions that exist over a wing embodying the present invention at high angles of attack;
  • FIGURE 9 is a diagram comparing performance characteristics of a heretofore conventional inwardly broken delta wing with those of a similar wing embodying the principles of this invention.
  • FIGURE 10 is a diagram similar to that of FIGURE 9 but wherein rolling moment coeflicient is plotted against sideslip at subcritical and super-critical angles of attack;
  • FIGURE 11 is a side perspective view of an aircraft having an inwardly broken delta wing embodying the principles of this invention, the vortices over the suction surface of the wing being illustrated diagrammatically;
  • FIGURE 12 is a fragmentary vertical sectional view taken on a plane extending spanwise along the wing of the aircraft shown in FIGURE 11;
  • FIGURE 13 is a more or less diagrammatic perspective view of a delta wing canard aircraft embodying the principles of this invention.
  • FIGURE 14 is a sectional view taken on the plane of the line 14-14 in FIGURE 13;
  • FIGURE 15 is a view similar to FIGURE 12 but illustrating another modified embodiment of the invention.
  • FIGURE 16 is a sectional view taken on the plane of the line 1616 in FIGURE 15;
  • FIGURES 17-19 are views similar to FIGURE 16 but respectively illustrating other modified embodiments of the invention.
  • the numeral 7 designates generally a thin, sharply swept-back wing, so arranged, according to any of the above-described known expedients, that two or more laterally adjacent vortices 11 and 12 that rotate in the same direction are formed over the upper or suction surface 15 of the wing at each side of its longitudinal centerline.
  • means are provided on such a wing for introducing between a pair of laterally adjacent corotating vortices a vortex which rotates in the opposite direction from them and which thus cooperates with them in a manner somewhat analogous to the meshing of a train of gears.
  • the oppositely rotating vortex reinforces and stabilizes the corotating vortices at opposite sides of it, and, in a manner of speaking, holds them in check by preventing their deflection away from the suction surface of the wing.
  • FIGURES 7 and 8 illustrate the effect of introducing such an oppositely rotating vortex 18 between a pair of corotating vortices 11 and 12 at each side of the centerline of a delta wing.
  • FIGURE 7 illustrates the vortex system over the suction surface 15 of a wing embodying the present invention at angles of attack corresponding to and below the above-mentioned value a and is thus to be compared with FIGURE 5; while FIG- URE 8 illustrates the vortex system over the same wing at 0: and should therefore be compared with FIGURE 6.
  • FIGURE 11 illustrates an aircraft having a wing 7 of inwardly broken delta planform and which is provided with plates or fences 5 that project downwardly from its pressure surface in accordance with the teachings of Swedish Patent No. 160,134 to generate over the suction surface 15 of the wing, at each side of its longitudinal centerline, a plurality of outboard vortices 111, 211, 311 and 411 that all rotate in the same direction.
  • the wing also generates an inboard vortex 12 which rotates in the direction of the outboard vortices and which has its origin at the apex of the wing.
  • the outboard vortex 111 originates at the break 14' in the leading edge of the wing, while the other outboard vortices 211, 311 and 411 originate at the respective plates 5, which thus break up into four smaller corotating vortices what would otherwise be one large outboard vortex, and thereby improve the wing tip flow at transonic speeds.
  • the wing is provided, at each side of its centerline, with a vortex generating means comprising an upright triangular wall or plate 20 having its apex positioned inboard of the break in the leading edge of the wing and substantially opposite the same laterally, and extending rearwardly along the suction surface of the wing directly adjacent to the inboard vortex 12, in the outward flow which is induced by the latter at high angles of attack.
  • a vortex generating means comprising an upright triangular wall or plate 20 having its apex positioned inboard of the break in the leading edge of the wing and substantially opposite the same laterally, and extending rearwardly along the suction surface of the wing directly adjacent to the inboard vortex 12, in the outward flow which is induced by the latter at high angles of attack.
  • This plate produces a vortex 18 that lies outwardly adjacent to the inboard vortex 12 and the energy for production of the vortex 18 is derived from the inboard vortex 12 so that the inboard vortex 12 is somewhat weakened as it moves rearwardly and cannot exert a disturbing influence upon the outer small vortices 111, 211, 311 and 411, the positions and lift of which are decisive of the stability and control characteristics of the aircraft.
  • the plate or vortex generating means 20 is arranged to be retracted at high speeds so that the stability of the aircraft is not disturbed by its shock system, particularly at transonic speeds.
  • the plate can have a pivote dconnection 21 to the wing structure, with the pivot axis of the plate extending along its lower edge, at its junction with the suction surface 15 of the wing.
  • the plate is swingable flatwise between an operative position in which it projects above the suction surface of the wing and a retracted position in which its outboard surface lies flush with the suction surface of the wing.
  • the plate can be actuated to one or the other of these positions as by means of a link 23 rigidly connected to the plate and pivotally connected to generally conventional hydraulic mechanism 24 or the like.
  • FIGURE 13 illustrates a generally similar vortex generating arrangement embodied in a canard aircraft having a straight delta main wing 107 and auxiliary wing 207, and wherein the auxiliary wing generates an inboard vortex 12 over the main wing at each side of its longitudinal centerline.
  • a vortex generating plate 20' is located on the main wing at each side of its longitudinal centerline, each such plate being located about midway between the apex of the main wing and its trailing edge, and inboard from the leading edge a distance such as to be outwardly adjacent to the inboard vortex 12.
  • the plate 20 has a pivotal connection 21 with the wing structure at its apex or front angle, as illustrated in FIGURE 14, so as to be swingable edgewise between an extended operative position, projecting upwardly from the suction surface of the wing, and a retracted position in which its upper edge lies flush with the suction surface of the wing.
  • suitable hydraulic mechanism 24 provides for actuation of the plate between its extended and retracted positions.
  • FIGURE 15 illustrates another form of means for producing a vortex which is located between a pair of adjacent corotating vortices, and which rotates in the direction opposite to theirs and cooperates with them to bind the vortex system to the suction surface of the wing.
  • the Wing is provided with at least one slot 26 at each side of its longitudinal centerline, elongated in the direction parallel to said centerline.
  • Each slot opens through the wing from its pressure side to its suction side.
  • the generally upright wall surfaces 27 and 28 which define each slot converge upwardly so as to accelerate the air that flows upwardly through the slot in response to the pressure differential between the opposite surfaces of the wing.
  • the outboard one 28 of these two wall surfaces makes a substantially sharp angled junction with the suction surface 15 of the wing, as at 29, to provide a vortex generating separation of air flow.
  • the junction of the inboard slot defining wall 27 with the suction surface of the wing is smoothly rounded, as at 30, so as to substantially prevent the formation of a vortex at the inboard edge of the slot.
  • the spanwise location of the slots 26 and 26 is adjacent to the inboard vortex.
  • each of the slots 26 and 26' can be closed, if desired, by means of edgewise slidable cover plates 32 and 33 which are substantially flush, respectively, with the pressure and suction surfaces of the wing, and which can be operated by any suitable mechanism (not shown). Closure of these cover plates serves the same purpose as retraction of the upright triangular plates in the first described embodiments of the invention, that is, it renders the vortex generating means inoperative for purposes of high speed flight.
  • cover plates 32' and 33 can be arranged to swing fla-twise outwardly, from positions in which they close the slot 26 to operative positions in which they project outwardly from the suction and pressure surfaces, respectively, and are aligned with the direction of flight.
  • the lower plate 32 when extended, serves to increase the flow through the slot by deflecting upwardly the strong outward flow of air that exists under the pressure side of the wing, especially at high values of angle of attack a; While the upper plate serves as a vortex generating means operating similarly to the plate 20 in the first described embodiment of the invention.
  • air from a suitable pressurized source thereof within the aircraft can be directed through a duct 34 communicated with a slot 126 that opens to the suction surface of the wing, as illustrated in FIG- URE 19.
  • the strongly energized air which is emitted from the slot 126 produces and intensifies the velocity discontinuities in the flow over the suction surface of the wing and thereby produces and intensifies the counterrotating vortex 18.
  • FIGURE 9 illustrates the improvement in the curves of lift coeflicient C and pitching moment coeflicient C in a wing having inwardly broken leading edges, of the type shown in FIGURE 3 and embodying the present invention, as compared with the same characteristics of an identical prior wing not incorporating counterrotating vortex generating means.
  • the solid line 40 represents C plotted against a for the prior wing, and it will be observed that lift coefiicient increases substantially steadily with increasing angle of attack on until a critical angle of attack a is attained. With further increase in angle of attack beyond m the value of C remains constant, at best, until the wing reaches a supercritical angle of attack 01 after which C increases sharply with further increasing angle of attack.
  • the C curve 41 for the prior wing (solid line) has a sharp deflection, corresponding to a severe disturbance of the pitching moment, through the range of angles of attack between a and a
  • the curves of C and C will correspond to the broken lines 40' and 41, respectively, through the range of angles of attack from or; through and beyond a and will thus vary at a substantially steady rate through the entire useful range of angles of attack.
  • FIGURE 10 illustrates, on the basis of a comparison similar to that of FIGURE 9, the improvement in roll characteristics brought about by the present invention.
  • the solid lines 43 and 44 in FIGURE 10 represent rolling moment coeificient C as a function of the angle of side slip 13 for a prior inwardly broken delta wing, at a subcritical angle of attack a and a supercritical angle of attack a respectively.
  • the broken lines 43' and 44 represent the same respective characteristics of a similar wing but embodying the present invention.
  • this invention provides simple and effective means on a thin swept-back wing for preventing or minimizing deflection away from its suction surface of adjacent vortices that are located at the same side of the longitudinal centerline of the wing and have the same direction of rotation, to thereby bring about marked improvement of the performance and stability characteristics of the wing at high angles of attack.
  • said third vortex generating means being (a) located a substantial distance forwardly of the trailing edge of the wing and spanwise intermediate the vortices of said pair thereof, and
  • said surface defining means comprising a plate normally projecting upwardly from thesuction surface of the wing, said upper edge thereof being rearwardly and upwardly inclined.
  • the wing of claim 6 further characterized by:
  • said surface defining means comprising one wall of a slot in the wing which opens to its suction surface, said upper edge thereof being defined by a sharply angled junction of said wall with the suction surface;
  • (B) further characterized by means in the wing for expelling from said slot air at a pressure higher than that which obtains above its adjacent portion of the suction surface of the wing.
  • the wing of claim 8 further characterized by:
  • said means for expelling air from said slot comprising means communicating said slot with an opening in the pressure surface of the wing.
  • said surface defining means comprising (1) one wall of a slot in the wing which opens to its suction surface and (2) a plate having a surface normally coplanar with said slot wall and projecting above the suction surface, said upper edge being defined by the upper edge of said plate;
  • (B) further characterized by means in the wing for expelling from said slot air at a pressure higher than that which obtains above its adjacent portion of the suction surface of the wing;
  • (C) means mounting said plate for flatwise swinging motion to a position in which it substantially closes said slot and lies substantially flush with the suction surface of the wing.
  • a second plate mounted for flatwise swinging motion between an extended position in which said second plate is coplanar with said slot wall and projects downwardly from the pressure surface of the wing, and a retracted position wherein said second plate closes the opening of the slot to the pressure surface of the wing and lies substantially flush with said surface.
  • said third vortex generating means being (a) located a substantial distance forwardly of the trailing edge of the wing and spanwise intermediate the vortices of said pair thereof, and

Landscapes

  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Tires In General (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US615230A 1966-02-11 1967-02-10 Stabilizing the vortices over a thin delta wing Expired - Lifetime US3471107A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
SE1813/66A SE301089B (it) 1966-02-11 1966-02-11

Publications (1)

Publication Number Publication Date
US3471107A true US3471107A (en) 1969-10-07

Family

ID=20258881

Family Applications (1)

Application Number Title Priority Date Filing Date
US615230A Expired - Lifetime US3471107A (en) 1966-02-11 1967-02-10 Stabilizing the vortices over a thin delta wing

Country Status (6)

Country Link
US (1) US3471107A (it)
AT (1) AT280059B (it)
CH (1) CH470287A (it)
FR (1) FR1511212A (it)
GB (1) GB1179568A (it)
SE (1) SE301089B (it)

Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3744745A (en) * 1971-09-30 1973-07-10 Mc Donnell Douglas Corp Liftvanes
US3892075A (en) * 1973-10-29 1975-07-01 Michael Edward Tibbett Apparatus for vortex generation to precipitate suspended particles in fluid bodies
US4017041A (en) * 1976-01-12 1977-04-12 Nelson Wilbur C Airfoil tip vortex control
FR2405366A1 (fr) * 1977-10-05 1979-05-04 Rolls Royce Perfectionnements aux dispositifs de deviation d'ecoulement
US4293110A (en) * 1979-03-08 1981-10-06 The Boeing Company Leading edge vortex flap for wings
US4323209A (en) * 1977-07-18 1982-04-06 Thompson Roger A Counter-rotating vortices generator for an aircraft wing
WO1982004426A1 (en) * 1981-06-10 1982-12-23 Co Boeing Leading edge vortex flap for wings
US4429843A (en) 1978-11-13 1984-02-07 Thompson Roger A Counter-rotating vortices generator for an aircraft wing
US4485992A (en) * 1981-09-10 1984-12-04 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Leading edge flap system for aircraft control augmentation
US4569494A (en) * 1982-12-23 1986-02-11 The Boeing Company Pitch control of swept wing aircraft
US4643376A (en) * 1982-09-30 1987-02-17 The Boeing Company Shock inducing pod for causing flow separation
US4655419A (en) * 1984-12-31 1987-04-07 The Boeing Company Vortex generator
US4685643A (en) * 1983-08-04 1987-08-11 The Boeing Company Nacelle/wing assembly with vortex control device
US4739957A (en) * 1986-05-08 1988-04-26 Advanced Aerodynamic Concepts, Inc. Strake fence flap
US5062595A (en) * 1990-04-26 1991-11-05 University Of Southern California Delta wing with lift enhancing flap
US5094411A (en) * 1990-10-19 1992-03-10 Vigyan, Inc. Control configured vortex flaps
US5255881A (en) * 1992-03-25 1993-10-26 Vigyan, Inc. Lift augmentation for highly swept wing aircraft
WO1993022196A1 (en) * 1992-04-28 1993-11-11 British Technology Group Usa Inc. Lifting body with reduced-strength trailing vortices
US5901925A (en) * 1996-08-28 1999-05-11 Administrator, National Aeronautics And Space Administration Serrated-planform lifting-surfaces
US6095459A (en) * 1997-06-16 2000-08-01 Codina; George Method and apparatus for countering asymmetrical aerodynamic process subjected onto multi engine aircraft
US6138955A (en) * 1998-12-23 2000-10-31 Board Of Supervisors Of Louisiana State University And Agricultural And Mechanical College Vortical lift control over a highly swept wing
GB2428459A (en) * 2005-07-13 2007-01-31 Univ City An Element For Generating A Fluid Dynamic Force
US8292220B1 (en) * 2009-03-19 2012-10-23 Northrop Grumman Corporation Flying wing aircraft with modular missionized elements

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5230486A (en) * 1992-05-22 1993-07-27 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Underwing compression vortex attenuation device
GB9401691D0 (en) * 1994-01-28 1994-03-23 Hannay Ian Foils
US5806807A (en) * 1995-10-04 1998-09-15 Haney; William R. Airfoil vortex attenuation apparatus and method
GB2466478A (en) * 2008-12-02 2010-06-30 Aerovortex Mills Ltd Suction generation device

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2125738A (en) * 1936-12-17 1938-08-02 William K Rose Airfoil
US2163655A (en) * 1938-04-02 1939-06-27 Charles H Zimmerman Slotted airplane wing tip
US2885161A (en) * 1948-08-11 1959-05-05 Douglas Aircraft Co Inc Stability control device for aircraft
CA595877A (en) * 1960-04-12 G. Gould Donald Vortex generator
GB890418A (en) * 1958-11-18 1962-02-28 Secr Aviation Lateral control means for aeroplanes
US3188022A (en) * 1963-12-05 1965-06-08 Svenska Aeroplan Ab Delta wing canard aircraft
US3237892A (en) * 1962-08-29 1966-03-01 Power Jets Res & Dev Ltd Aircraft aerodynamic lifting member

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CA595877A (en) * 1960-04-12 G. Gould Donald Vortex generator
US2125738A (en) * 1936-12-17 1938-08-02 William K Rose Airfoil
US2163655A (en) * 1938-04-02 1939-06-27 Charles H Zimmerman Slotted airplane wing tip
US2885161A (en) * 1948-08-11 1959-05-05 Douglas Aircraft Co Inc Stability control device for aircraft
GB890418A (en) * 1958-11-18 1962-02-28 Secr Aviation Lateral control means for aeroplanes
US3237892A (en) * 1962-08-29 1966-03-01 Power Jets Res & Dev Ltd Aircraft aerodynamic lifting member
US3188022A (en) * 1963-12-05 1965-06-08 Svenska Aeroplan Ab Delta wing canard aircraft

Cited By (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3744745A (en) * 1971-09-30 1973-07-10 Mc Donnell Douglas Corp Liftvanes
US3892075A (en) * 1973-10-29 1975-07-01 Michael Edward Tibbett Apparatus for vortex generation to precipitate suspended particles in fluid bodies
US4017041A (en) * 1976-01-12 1977-04-12 Nelson Wilbur C Airfoil tip vortex control
US4323209A (en) * 1977-07-18 1982-04-06 Thompson Roger A Counter-rotating vortices generator for an aircraft wing
FR2405366A1 (fr) * 1977-10-05 1979-05-04 Rolls Royce Perfectionnements aux dispositifs de deviation d'ecoulement
US4232516A (en) * 1977-10-05 1980-11-11 Rolls-Royce Limited Flow deflecting devices
US4429843A (en) 1978-11-13 1984-02-07 Thompson Roger A Counter-rotating vortices generator for an aircraft wing
US4293110A (en) * 1979-03-08 1981-10-06 The Boeing Company Leading edge vortex flap for wings
WO1982004426A1 (en) * 1981-06-10 1982-12-23 Co Boeing Leading edge vortex flap for wings
US4485992A (en) * 1981-09-10 1984-12-04 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Leading edge flap system for aircraft control augmentation
US4643376A (en) * 1982-09-30 1987-02-17 The Boeing Company Shock inducing pod for causing flow separation
US4569494A (en) * 1982-12-23 1986-02-11 The Boeing Company Pitch control of swept wing aircraft
US4685643A (en) * 1983-08-04 1987-08-11 The Boeing Company Nacelle/wing assembly with vortex control device
US4655419A (en) * 1984-12-31 1987-04-07 The Boeing Company Vortex generator
US4739957A (en) * 1986-05-08 1988-04-26 Advanced Aerodynamic Concepts, Inc. Strake fence flap
US5062595A (en) * 1990-04-26 1991-11-05 University Of Southern California Delta wing with lift enhancing flap
US5094411A (en) * 1990-10-19 1992-03-10 Vigyan, Inc. Control configured vortex flaps
US5255881A (en) * 1992-03-25 1993-10-26 Vigyan, Inc. Lift augmentation for highly swept wing aircraft
WO1993022196A1 (en) * 1992-04-28 1993-11-11 British Technology Group Usa Inc. Lifting body with reduced-strength trailing vortices
US5492289A (en) * 1992-04-28 1996-02-20 British Technology Group Usa Inc. Lifting body with reduced-strength trailing vortices
US5901925A (en) * 1996-08-28 1999-05-11 Administrator, National Aeronautics And Space Administration Serrated-planform lifting-surfaces
US6095459A (en) * 1997-06-16 2000-08-01 Codina; George Method and apparatus for countering asymmetrical aerodynamic process subjected onto multi engine aircraft
US6138955A (en) * 1998-12-23 2000-10-31 Board Of Supervisors Of Louisiana State University And Agricultural And Mechanical College Vortical lift control over a highly swept wing
GB2428459A (en) * 2005-07-13 2007-01-31 Univ City An Element For Generating A Fluid Dynamic Force
GB2428459B (en) * 2005-07-13 2009-10-21 Univ City An element for generating a fluid dynamic force
US8292220B1 (en) * 2009-03-19 2012-10-23 Northrop Grumman Corporation Flying wing aircraft with modular missionized elements

Also Published As

Publication number Publication date
FR1511212A (fr) 1968-01-26
CH470287A (de) 1969-03-31
SE301089B (it) 1968-05-20
GB1179568A (en) 1970-01-28
AT280059B (de) 1970-03-25

Similar Documents

Publication Publication Date Title
US3471107A (en) Stabilizing the vortices over a thin delta wing
US5772155A (en) Aircraft wing flaps
US5156358A (en) Aircraft outboard control
US2846165A (en) Aircraft control system
US4479620A (en) Wing load alleviation system using tabbed allerons
US3188022A (en) Delta wing canard aircraft
US4108403A (en) Vortex reducing wing tip
US5102068A (en) Spiroid-tipped wing
US4739957A (en) Strake fence flap
US6138957A (en) Swept-back wings with airflow channeling
DE69326253T2 (de) Überschallflugzeug mit hochem wirkungsgrad
US4466586A (en) Directional control device for aircraft
US3478988A (en) Stol aircraft having by-pass turbojet engines
US4050397A (en) Foil fence for hydrofoil craft
US3126173A (en) Alvarez-calderdn
US3706430A (en) Airfoil for aircraft
US3738595A (en) Delta-wing aircraft
JPH0316319B2 (it)
US2885161A (en) Stability control device for aircraft
US3942746A (en) Aircraft having improved performance with beaver-tail afterbody configuration
US4132375A (en) Vortex-lift roll-control device
US5655737A (en) Split rudder control system aerodynamically configured to facilitate closure
US3371888A (en) Inverting flap system
CN110546067A (zh) 飞机的空气动力表面
US3848831A (en) Fuselage flaps for an aircraft