WO2009106870A2 - Shock bump - Google Patents
Shock bump Download PDFInfo
- Publication number
- WO2009106870A2 WO2009106870A2 PCT/GB2009/050151 GB2009050151W WO2009106870A2 WO 2009106870 A2 WO2009106870 A2 WO 2009106870A2 GB 2009050151 W GB2009050151 W GB 2009050151W WO 2009106870 A2 WO2009106870 A2 WO 2009106870A2
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- shock
- bump
- bumps
- edge
- aerofoil
- Prior art date
Links
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
- B64C23/04—Influencing air flow over aircraft surfaces, not otherwise provided for by generating shock waves
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
- B64C3/14—Aerofoil profile
- B64C2003/148—Aerofoil profile comprising protuberances, e.g. for modifying boundary layer flow
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
- B64C3/14—Aerofoil profile
- B64C2003/149—Aerofoil profile for supercritical or transonic flow
Definitions
- the present invention relates to a shock bump; and a method of operating an aerodynamic structure comprising a shock bump extending from its surface.
- VVGs vane vortex generators
- US 2006/0060720 uses a shock control protrusion to generate a shock extending away from the lower surface of a wing.
- a first aspect of the invention provides a shock bump comprising a diverging nose and a converging tail, wherein the tail has at least one plan- form contour line with a pair of concave opposite sides.
- the shock bump of the first aspect of the invention provides an improved shape with relatively low drag. Furthermore, the concave shape of the tail tends to promote the development of longitudinal vortices which can reduce shock induced buffet in certain operating conditions.
- plan- form contour line may become convex and meet each other head-on at the trailing edge of the shock bump, or may meet at a cusp-like point.
- the shock bump has a leading edge, a trailing edge, an inboard edge and an outboard edge.
- the bump may merge gradually into the surface at its edges or there may be an abrupt concave discontinuity at one or more of its edges.
- shock bump has substantially no sharp convex edges or points.
- a second aspect of the invention provides an aerodynamic structure comprising one or more shock bumps of the type described above extending from its surface.
- each shock bump is shaped and positioned so as to modify the structure of a shock which would form adjacent to the surface of the structure in the absence of the shock bump(s) when the structure is moved at transonic speeds.
- US 2006/0060720 which uses a shock control protrusion to generate a shock which would not otherwise exist in the absence of the shock control protrusion.
- a third aspect of the invention provides a method of operating an aerodynamic structure, the structure comprising a shock bump extending from its surface, the method comprising:
- the second condition is one involving a higher flow speed and/or a higher lift coefficient than the first condition.
- the structure may comprise an aerofoil such as an aircraft wing, horizontal tail plane or control surface; an aircraft structure such as a nacelle, pylon or fin; or any other kind of aerodynamic structure such as a turbine blade.
- an aerofoil such as an aircraft wing, horizontal tail plane or control surface
- an aircraft structure such as a nacelle, pylon or fin
- any other kind of aerodynamic structure such as a turbine blade.
- the shock bump may be located on a high pressure surface of the aerofoil (that is, the lower surface in the case of an aircraft wing) but more preferably the surface is a low pressure surface of the aerofoil (that is, the upper surface in the case of an aircraft wing).
- the shock bump typically has an apex which is positioned towards the trailing edge of the aerofoil, in other words it is positioned aft of 50% chord.
- the apex of the bump may be a single point, or a plateau. In the case of a plateau then the leading edge of the plateau is positioned towards the trailing edge of the aerofoil.
- Figure 1 is a plan view of the top of an aircraft wing carrying an array of shock bumps according to a first embodiment of the invention, operating at its "design" operating condition;
- Figures 2 is a longitudinal cross-sectional view through the centre of one of the bumps taken along a line A-A, with the wing in its "design" operating condition;
- Figure 3 is a plan view of the top of the aircraft wing of Figure 1, with the wing in an "off-design" operating condition;
- Figure 4 is a longitudinal cross-sectional view through the centre of one of the bumps taken along a line B-B, with the wing in an "off-design" operating condition;
- Figure 5 is a transverse cross-sectional view through the centre of one of the bumps taken along a line C-C;
- Figure 6 is a plan view of one of the bumps showing a series of contour lines.
- Figure 7 is a plan view of the top of an aircraft wing carrying an array of shock bumps according to a second embodiment of the invention.
- Figure 1 is a plan view of the upper surface of an aircraft wing.
- the wing has a leading edge 1 and a trailing edge 2, each swept to the rear relative to the free stream direction.
- the upper surface of the wing carries an array of shock bumps extending from its surface.
- the array comprises a first series of shock bumps 3; and a second series of shock bumps 10 positioned aft of the first series.
- Each bump 3, 10 protrudes from a nominal surface 8 of the wing, and meets the nominal surface 8 at a leading edge 3a, 10a; a trailing edge 3b, 10b; an inboard edge 3c, 10c; and an outboard edge 3d, 1Od.
- the lower portions of the sides of bump are concave and merge gradually into the nominal surface 8.
- the lower portion 9 of the front side of the bump merges gradually into the nominal surface 8 at leading edge 3 a.
- the lower portion of the front side of the bump may be planar as illustrated by dashed line 9a. In this case the front side 9a of the shock bump meets the nominal surface 8 with an abrupt discontinuity at the leading edge 3 a.
- Figure 2 is a cross-sectional view through the centre of one of the bumps 3 taken along a line A-A parallel with the free stream direction.
- the apex point 7 of the fore/aft cross-section A-A is offset aft of the centre 6 of the bump.
- the apex 7 of each bump 3 is positioned aft of 50% chord, typically between 60% and 65% chord.
- FIGS 1 and 2 show the position 4 of the shock when the aircraft is operated with a Mach number and lift coefficient which together define a "design" operating condition (generally associated with the cruise phase of a flight envelope).
- a "design" operating condition generally associated with the cruise phase of a flight envelope.
- the shock bumps 3 are positioned so as to induce a smeared foot 5 in the shock 4 with a lambda like wave pattern as shown in Figure 2, and the flow over the second series of shock bumps 10 is fully attached.
- the smeared foot 5 When the shock bumps 3 are operated at their optimum with the shock 4 just ahead of the apex 7 of the bump as shown in Figure 2, the smeared foot 5 has a lambda-like wave pattern with a single forward shock 5 a towards the leading edge of the bump and a single rear shock 5b positioned slightly forward of the apex 7.
- the smeared foot may have a lambda-like wave pattern with a fan- like series of forward shocks.
- the second series of shock bumps 10 is positioned to modify the structure of a shock 11 which forms adjacent to the surface of the wing when the aerofoil is operated at a higher Mach number or lift coefficient associated with an "off-design" operating condition as shown in Figures 3 and 4.
- the shock moves aft to a position 11 shown in Figure 3, and the shock bumps 10 are positioned so as to induce a smeared shock foot 15 with a lambda like wave pattern as shown in Figure 4.
- the bumps have no sharp convex edges or points so the flow remains attached over the bumps when they are operated at their optimum (i.e. when the shock is positioned on the bump just ahead of its apex).
- a characteristic of three-dimensional shock bumps is that when operated away from their optimum i.e. when the shock is positioned on the bump but not just ahead of the apex of the bump, the flow at the rear of the bump tends to detach.
- This rear bump separation is exploited to form a pair of counter rotating longitudinal vortices 12,13 aligned with the flow direction that will have a similar positive impact on high speed buffet as VVGs.
- These vortices are embedded in or just above the boundary layer.
- the shock bumps 10 provide an improved flight envelope and speed range or reduced loads at high speed.
- the second series of shock bumps is offset slightly relative to the first series, so that none of the shock bumps 10 in the second series are positioned directly aft of any of the shock bumps 3 in the first series.
- Figure 5 is a lateral cross-section through the centre of one of the bumps 10, and Figure 6 shows a series of plan- form contour lines (equivalent to contour lines in a map) including a footprint contour line in solid line where the shock bump merges into the upper surface of the wing; an intermediate contour line 25; and an upper contour line 24.
- the footprint contour line comprises a diverging nose 20 and a converging tail with concave opposite sides 22,23 which meet at a cusp-like point 21 at the trailing edge of the bump.
- the tail of the intermediate contour line 25 has a pair of concave sides which become convex and meet head-on at the trailing edge of the contour line 25.
- the shock bump 10 is laterally symmetric about its fore-and-aft centre line 26.
- each individual shock bump 10 can be adjusted from the shape illustrated such that at the "design" operating condition the flow over the bump is fully attached as shown in Figure 1.
- some beneficial modification of the shock foot will take place in addition to the formation of a pair of longitudinal vortices.
- the upper surface of the wing carries an array of shock bumps comprising a first series of shock bumps 3 with an elliptical footprint, and a second series of cusp-shaped shock bumps 10 positioned aft of the first series.
- shock bumps comprising a first series of shock bumps 3 with an elliptical footprint, and a second series of cusp-shaped shock bumps 10 positioned aft of the first series.
- various other embodiments fall within the scope of the invention, including: • a single cusp-shaped shock bump
- Figure 7 is a plan view of the upper surface of an aircraft wing according to a second embodiment of the present invention.
- the embodiment of Figure 7 is identical to the embodiment of Figure 1, except in this case the forward series has ten shock bumps 3, whereas there is only a single rear shock bump 10.
- Figure 10 shows the span- wise extent of the shocks 4, 11. It can be seen that the shock 4 extends over a significant span- wise portion of the wing, whereas the shock 11 is relatively short so only a small number of rear shock bumps 10 (in this case only one) is needed.
Abstract
Description
Claims
Priority Applications (7)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP2010548186A JP5481394B2 (en) | 2008-02-29 | 2009-02-17 | Shock bump |
RU2010139001/11A RU2503587C2 (en) | 2008-02-29 | 2009-02-17 | Ledge to vary shock wave structure |
BRPI0908335-9A BRPI0908335A2 (en) | 2008-02-29 | 2009-02-17 | Shock control bulge |
US12/735,540 US8302912B2 (en) | 2008-02-29 | 2009-02-17 | Shock bump |
CN200980106626.8A CN101959756B (en) | 2008-02-29 | 2009-02-17 | Shock bump |
EP09714018.0A EP2250086B1 (en) | 2008-02-29 | 2009-02-17 | Shock bump |
CA2713362A CA2713362C (en) | 2008-02-29 | 2009-02-17 | Shock bump |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GBGB0803722.8A GB0803722D0 (en) | 2008-02-29 | 2008-02-29 | Shock bump |
GB0803722.8 | 2008-02-29 |
Publications (2)
Publication Number | Publication Date |
---|---|
WO2009106870A2 true WO2009106870A2 (en) | 2009-09-03 |
WO2009106870A3 WO2009106870A3 (en) | 2009-10-22 |
Family
ID=39315674
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/GB2009/050151 WO2009106870A2 (en) | 2008-02-29 | 2009-02-17 | Shock bump |
Country Status (9)
Country | Link |
---|---|
US (1) | US8302912B2 (en) |
EP (1) | EP2250086B1 (en) |
JP (1) | JP5481394B2 (en) |
CN (1) | CN101959756B (en) |
BR (1) | BRPI0908335A2 (en) |
CA (1) | CA2713362C (en) |
GB (1) | GB0803722D0 (en) |
RU (1) | RU2503587C2 (en) |
WO (1) | WO2009106870A2 (en) |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP2011513115A (en) * | 2008-02-29 | 2011-04-28 | エアバス・ユ―ケ―・リミテッド | Shock bump |
US20210197952A1 (en) * | 2019-12-27 | 2021-07-01 | Bombardier Inc. | Variable wing leading edge camber |
Families Citing this family (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB0803730D0 (en) * | 2008-02-29 | 2008-04-09 | Airbus Uk Ltd | Shock bump array |
US20110006165A1 (en) * | 2009-07-10 | 2011-01-13 | Peter Ireland | Application of conformal sub boundary layer vortex generators to a foil or aero/ hydrodynamic surface |
US8939410B2 (en) * | 2013-02-06 | 2015-01-27 | Reginald J Exton | Boundary layer flow disruptors for delaying transition to turbulent flow |
US10093411B1 (en) * | 2014-03-18 | 2018-10-09 | The Boeing Company | Apparatus and method for reducing a sonic boom signature |
US9810099B2 (en) | 2015-06-29 | 2017-11-07 | Siemens Energy, Inc. | Turbine exhaust cylinder strut strip for shock induced oscillation control |
CN112124561B (en) * | 2020-09-27 | 2022-02-25 | 中国商用飞机有限责任公司 | Aerodynamic drag reduction structure for wingtip winglet of aircraft and aircraft |
GB2612766A (en) * | 2021-10-27 | 2023-05-17 | Airbus Operations Ltd | Improved lifting surface |
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US2532753A (en) * | 1947-07-05 | 1950-12-05 | Lockheed Aircraft Corp | Transonic airfoil design |
US2800291A (en) * | 1950-10-24 | 1957-07-23 | Stephens Arthur Veryan | Solid boundary surface for contact with a relatively moving fluid medium |
US2898059A (en) * | 1957-09-11 | 1959-08-04 | Richard T Whitcomb | Fuselage shaping to reduce the strength of the initial shock wave on lifting airplane wings |
US3741285A (en) * | 1968-07-09 | 1973-06-26 | A Kuethe | Boundary layer control of flow separation and heat exchange |
US3578264A (en) | 1968-07-09 | 1971-05-11 | Battelle Development Corp | Boundary layer control of flow separation and heat exchange |
US4067518A (en) * | 1976-05-20 | 1978-01-10 | Lockheed Corporation | Drag reducer for lift surface of aircraft |
US4354648A (en) * | 1980-02-06 | 1982-10-19 | Gates Learjet Corporation | Airstream modification device for airfoils |
US4643376A (en) * | 1982-09-30 | 1987-02-17 | The Boeing Company | Shock inducing pod for causing flow separation |
JPS63207796A (en) * | 1987-02-25 | 1988-08-29 | 三菱重工業株式会社 | Missile with auxiliary booster |
US5058837A (en) * | 1989-04-07 | 1991-10-22 | Wheeler Gary O | Low drag vortex generators |
GB9116787D0 (en) * | 1991-08-01 | 1991-09-18 | Secr Defence | Article having an aerofoil section with a distensible expansion surface |
JPH07149299A (en) * | 1993-11-29 | 1995-06-13 | Mitsubishi Electric Corp | Atmosphere reentering aircraft |
US5692709A (en) * | 1994-11-01 | 1997-12-02 | Condor Systems, Inc. | Shock wave stabilization apparatus and method |
DE4446031C2 (en) | 1994-12-23 | 1998-11-26 | Deutsch Zentr Luft & Raumfahrt | Wing with means for changing the profile |
GB9814122D0 (en) | 1998-07-01 | 1998-08-26 | Secr Defence | Aerofoil having improved buffet performance |
RU2150403C1 (en) * | 1999-10-28 | 2000-06-10 | Таланов Борис Петрович | Method of enhancing hydro- and aerodynamic properties of object moving in amorphous medium |
US6929214B2 (en) * | 2003-07-22 | 2005-08-16 | Northrop Grumman Corporation | Conformal airliner defense (CAD) system |
US7118071B2 (en) * | 2004-03-31 | 2006-10-10 | The Boeing Company | Methods and systems for controlling lower surface shocks |
US20060021560A1 (en) * | 2004-05-02 | 2006-02-02 | Mcmillan David W | Tail fairing designed with features for fast installation and/or for suppression of vortices addition between fairings, apparatus incorporating such fairings, methods of making and using such fairings and apparatus, and methods of installing such fairings |
US20090294596A1 (en) * | 2005-03-29 | 2009-12-03 | Sinha Sumon K | Method of Reducing Drag and Increasing Lift Due to Flow of a Fluid Over Solid Objects |
US7300021B2 (en) * | 2005-05-20 | 2007-11-27 | The Boeing Company | Aerospace vehicle fairing systems and associated methods |
US20070018055A1 (en) * | 2005-07-11 | 2007-01-25 | Schmidt Eric T | Aerodynamically efficient surface |
US7784737B2 (en) * | 2005-09-19 | 2010-08-31 | The Boeing Company | Drag reduction fairing |
US8016245B2 (en) * | 2006-10-18 | 2011-09-13 | The Boeing Company | Dynamic bumps for drag reduction at transonic-supersonic speeds |
GB0803722D0 (en) * | 2008-02-29 | 2008-04-09 | Airbus Uk Ltd | Shock bump |
GB201002281D0 (en) * | 2010-02-11 | 2010-03-31 | Univ Sheffield | Apparatus and method for aerodynamic drag reduction |
-
2008
- 2008-02-29 GB GBGB0803722.8A patent/GB0803722D0/en not_active Ceased
-
2009
- 2009-02-17 CN CN200980106626.8A patent/CN101959756B/en not_active Expired - Fee Related
- 2009-02-17 US US12/735,540 patent/US8302912B2/en not_active Expired - Fee Related
- 2009-02-17 CA CA2713362A patent/CA2713362C/en not_active Expired - Fee Related
- 2009-02-17 BR BRPI0908335-9A patent/BRPI0908335A2/en not_active IP Right Cessation
- 2009-02-17 RU RU2010139001/11A patent/RU2503587C2/en not_active IP Right Cessation
- 2009-02-17 JP JP2010548186A patent/JP5481394B2/en not_active Expired - Fee Related
- 2009-02-17 WO PCT/GB2009/050151 patent/WO2009106870A2/en active Application Filing
- 2009-02-17 EP EP09714018.0A patent/EP2250086B1/en not_active Not-in-force
Non-Patent Citations (1)
Title |
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None |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP2011513115A (en) * | 2008-02-29 | 2011-04-28 | エアバス・ユ―ケ―・リミテッド | Shock bump |
US20210197952A1 (en) * | 2019-12-27 | 2021-07-01 | Bombardier Inc. | Variable wing leading edge camber |
Also Published As
Publication number | Publication date |
---|---|
GB0803722D0 (en) | 2008-04-09 |
EP2250086B1 (en) | 2013-07-17 |
CN101959756B (en) | 2013-08-07 |
BRPI0908335A2 (en) | 2015-07-28 |
WO2009106870A3 (en) | 2009-10-22 |
CA2713362A1 (en) | 2009-09-03 |
CA2713362C (en) | 2016-04-05 |
CN101959756A (en) | 2011-01-26 |
US8302912B2 (en) | 2012-11-06 |
JP5481394B2 (en) | 2014-04-23 |
RU2010139001A (en) | 2012-04-10 |
JP2011513115A (en) | 2011-04-28 |
US20100314500A1 (en) | 2010-12-16 |
EP2250086A2 (en) | 2010-11-17 |
RU2503587C2 (en) | 2014-01-10 |
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