CN109436293A - A kind of shock wave control device - Google Patents
A kind of shock wave control device Download PDFInfo
- Publication number
- CN109436293A CN109436293A CN201811389511.1A CN201811389511A CN109436293A CN 109436293 A CN109436293 A CN 109436293A CN 201811389511 A CN201811389511 A CN 201811389511A CN 109436293 A CN109436293 A CN 109436293A
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- China
- Prior art keywords
- shock wave
- control device
- wave control
- convex slot
- static pressure
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Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
- B64C3/14—Aerofoil profile
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C21/00—Influencing air flow over aircraft surfaces by affecting 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
- B64C2003/148—Aerofoil profile comprising protuberances, e.g. for modifying boundary layer flow
Abstract
The invention discloses a kind of shock wave control devices.Described device is arranged on aircraft wing, and convex slot is opened up on aircraft wing；Porous plate is arranged in the opening of convex slot；Flexible covering is covered on porous plate；Flexible covering and the wing cover of convex slot surrounding are smoothly connected；Multiple static pressure holes are additionally provided on aircraft wing；Static pressure hole is connected to convex slot.Pressure change of the device by perception aerofoil surface after there is shock wave, passive type can be changed according to state of flight and change bulge shape automatically, the automatic control of drag due to shock wave when realizing aircraft flight, to effectively increase robustness of the shock wave control bulge under different flight state；And the characteristic for flow field being utilized itself reduces the manufacture and maintenance cost of wing and damping device without additional energy input.
Description
Technical field
The present invention relates to shock wave control technology fields, more particularly to a kind of shock wave control device.
Background technique
With the fast development of Civil Aviation Industry, the concept of green aviation increasingly by the concern of civil aviaton circle, how
Researching and developing more efficient, energy-saving and environment-friendly type becomes a popular subject under discussion of civil aircraft design.Studies have shown that certain in voyage
Under conditions of, increase by 1 resistance unit (Δ Cd=0.0001), being equivalent to will reduce by 8 passengers, therefore pneumatic drag reduction is civilian
One of vital task of airplane design directly influences flight and flight oil consumption, and then is related to the operation of aircraft
The market competitiveness of cost and the design type.
It is limited by drag divergence Mach number, the main cruising flight range of Mach numbers of modern large-sized civil passenger plane is
0.785 to 0.85.At this point, the upper surface of wing will appear normal shock wave due to there is local supersonic zone.The appearance of normal shock wave
It can cause quickling increase for aircraft drag, if the intensity of shock wave is excessive, effect be interfered with each other due to shock wave and boundary layer, very
The chattering phenomenon for seriously threatening aircraft safety can extremely be caused.Therefore, it can be said that the normal shock intensity of wave of upper surface of the airfoil is largely
On limit the flying quality of large-sized civil passenger plane.
In order to control shock wave, to weaken shock strength, a variety of shock wave control methods have been proposed at present, wrap
Include actively air blowing and air-breathing, porous cavity, contour bump etc..From the point of view of drag reduction, wherein contour bump technology is that most have
A kind of passive control methods of effect, basic scheme is: installing entity drum additional by occurring the region of normal shock wave in upper surface of the airfoil
Packet, changes the local form of wing, compresses air-flow in advance in bulge position, to reduce the intensity of shock wave, reduces
Drag due to shock wave and drag divergence Mach number is postponed.
Although the drag due to shock wave in flight under upper surface of the airfoil installation contour bump reduces desin speed, inclined
In the case where from desin speed, due to the variation of shock-wave spot and intensity, the increase of aircraft drag can be caused instead.This is because
Best bulge shape is sensitive to shock-wave spot and high strength, and contour bump lacks robustness for velocity variations.In order to change
Drawbacks described above, existing another kind thinking are actively to change the shape of bulge by increasing mechanical device, and then improve shock wave control
The robustness of bulge processed.But the addition of self-reacting device increases the weight and complexity of wing structure, manufacturing cost and dimension
Shield cost is also increase accordingly, and the energy outside the amount of imports is needed to maintain bulge shape, actively adaptive for total income
Shock wave control bulge is lost more than gain.
Summary of the invention
Based on this, it is necessary to a kind of adaptive shock wave control device of passive type is provided, to overcome existing shock wave control bulge
The relatively narrow limitation of drag reduction sphere of action is realizing the adaptive change of bulge shape, while improving the robustness of damping device,
Reduce existing active adaptive mechanism to the input of wing bring additional energy, weight and complexity, reduces wing and drag reduction dress
The manufacture and maintenance cost set.
To achieve the above object, the present invention provides following schemes:
A kind of shock wave control device, described device are arranged on aircraft wing, and described device includes: porous plate and flexible illiteracy
Skin；
Convex slot is opened up on the aircraft wing；The porous plate is arranged in the opening of the convex slot；The porous plate
On be covered with flexible covering；The flexible covering and the wing cover of the convex slot surrounding are smoothly connected；The aircraft wing
On be additionally provided with multiple static pressure holes；The static pressure hole is connected to the convex slot, for perceiving when the plane wing surfaces produce
The pressure at shock wave rear when raw shock wave, and the pressure perceived is transferred to the convex slot, so that the flexible covering produces
Raw deformation.
Optionally, the porous plate is arranged in close to the side of wing tail portion in the static pressure hole.
Optionally, the distance between static pressure hole described in every two is equal.
Optionally, the distance between static pressure hole described in every two is 2 δ, and δ indicates the thickness of local boundary layer.
Optionally, the shape of the static pressure hole is circle, the diameter value of the static pressure hole and the thickness value of local boundary layer
It is equal.
Optionally, the depth of the convex slot is that 5 δ -10 δ, δ indicate the thickness of local boundary layer.
Optionally, the convex slot is opened in the aircraft wing middle section along tangential region.
Optionally, the material of the porous plate is aviation titanium alloy material.
Optionally, the material of the flexible covering is silastic material or honeycomb composite material.
Compared with prior art, the beneficial effects of the present invention are:
The invention proposes a kind of shock wave control device, described device is arranged on aircraft wing, opens up on aircraft wing
Convex slot；Porous plate is arranged in the opening of convex slot；Flexible covering is covered on porous plate；Flexible covering and convex slot surrounding
Wing cover is smoothly connected；Multiple static pressure holes are additionally provided on aircraft wing；Static pressure hole is connected to convex slot.The device is in shock wave
The arranged beneath static pressure hole of control area perceives flow field pressure, and is transferred to upstream (flexible covering), basis by convex slot
The position of shock wave and Strength Changes realize the pressure difference of adaptive change, improve the robustness of damping device, and stream is utilized
The characteristic of field itself reduces the manufacture and maintenance cost of wing and damping device without additional energy input.
Detailed description of the invention
It in order to more clearly explain the embodiment of the invention or the technical proposal in the existing technology, below will be to institute in embodiment
Attached drawing to be used is needed to be briefly described, it should be apparent that, the accompanying drawings in the following description is only some implementations of the invention
Example, for those of ordinary skill in the art, without any creative labor, can also be according to these attached drawings
Obtain other attached drawings.
Fig. 1 is the schematic diagram that conventional entity bulge controls shock strength；
Fig. 2 is a kind of overall structure figure of shock wave control device of the embodiment of the present invention；
Fig. 3 is a kind of schematic diagram of internal structure of shock wave control device of the embodiment of the present invention；
Fig. 4 is the schematic diagram of shock wave control device in the case of having shock wave；
Fig. 5 for shock wave control device in the case of no shock wave schematic diagram；
Fig. 6 is schematic view of the mounting position of the shock wave control device in aircraft wing.
Specific embodiment
Following will be combined with the drawings in the embodiments of the present invention, and technical solution in the embodiment of the present invention carries out clear, complete
Site preparation description, it is clear that described embodiments are only a part of the embodiments of the present invention, instead of all the embodiments.It is based on
Embodiment in the present invention, it is obtained by those of ordinary skill in the art without making creative efforts every other
Embodiment shall fall within the protection scope of the present invention.
The object of the present invention is to provide one kind can be carried out automatically according to the shock-wave spot and intensity passive type of aerofoil surface
The shock wave control device of deformation is deteriorated rapidly, robust with solving existing shock wave control bulge performance near deviation design point
The insufficient problem of property.
In order to make the foregoing objectives, features and advantages of the present invention clearer and more comprehensible, with reference to the accompanying drawing and specific real
Applying mode, the present invention is described in further detail.
When aircraft is in transonic flight, upper surface of the airfoil will appear shock wave, has sudden change of pressure surface after shock wave front, thus generates
Biggish drag due to shock wave seriously affects the flight efficiency of aircraft.In order to reduce flight resistance, contour bump method be it is a kind of more
Mature shock wave control method, Fig. 1 is the schematic diagram that conventional entity bulge controls shock strength, referring to Fig. 1, in aircraft wing 1
The traditional contour bump 2 of upper setting, Ma indicate that Mach number, traditional contour bump 2 generate disturbance to 1 upper surface air-flow of wing
Geometry changes, the normal shock wave that 1 upper surface of (clean configuration) wing generates when acting on compared to no bulge, has under bulge effect, on
Surface normal shock wave can be changed into λ type shock wave 3, and λ type shock wave 3 is a series of oblique shock waves, and under identical pressure change, this is a series of
Pitot loss caused by oblique shock wave reduces, and wave resistance also accordingly reduces.Since traditional 2 position shape of contour bump immobilizes,
Although can reach drag-reduction effect under design condition, flight status changes, and the effect of contour bump 2, which can reduce, even loses
Go effect.
It is adaptive within the scope of certain flying speed in order to realize that bulge shape changes automatically according to different flying conditions
Shock strength is controlled, the invention proposes a kind of shock wave control devices.
Embodiment 1:
Fig. 2 is a kind of overall structure figure of shock wave control device of the embodiment of the present invention；Fig. 3 swashs for one kind of the embodiment of the present invention
The schematic diagram of internal structure of wave control device.
Referring to figs. 2 and 3, the shock wave control device of embodiment is arranged on aircraft wing 1, and described device includes: porous
Plate 4 and flexible covering 5；Convex slot 6 is opened up on the aircraft wing 1；The porous plate 4 is arranged in the opening of the convex slot 6；
Flexible covering 5 is covered on the porous plate 4；The flexible covering 5 and the wing cover 7 of 6 surrounding of convex slot smoothly connect
It connects；Multiple static pressure holes 8 are additionally provided on the aircraft wing 1；The static pressure hole 8 is connected to the convex slot 6；The flexible illiteracy
Skin 5, for realizing the auto Deformation of bulge；The static pressure hole 8 is for perceiving when 1 surface of aircraft wing generates shock wave (λ type
Shock wave 3) rear Shi Jibo pressure, and the pressure perceived is transferred to the convex slot 6, so that the flexible covering 5
Deformation is generated, achievees the purpose that self adaptive control shock strength；The porous plate 4, for guaranteeing that air-flow is unimpeded and can bear machine
Aerodynamic loading on the wing.
In the present embodiment, the porous plate 4 is arranged in close to the side of wing tail portion in the static pressure hole 8；Every two institute
It states the distance between static pressure hole 8 to be equal, specifically, the distance between static pressure hole 8 described in every two is 2 δ, rand is worked as in δ expression
The thickness of interlayer.As an alternative embodiment, the shape of the static pressure hole 8 can be circle, to reduce stress concentration,
The diameter value of the circular static pressure hole 8 is equal with the thickness value of local boundary layer.
In the present embodiment, the depth of the convex slot 6 is 5 δ -10 δ；The convex slot 6 is opened in the aircraft wing 1
Section is along tangential region；The material of the porous plate 4 is aviation titanium alloy material；The material of the flexible covering 5 is silicon rubber
Glue material or honeycomb composite material.
The specific working principle is as follows for the shock wave control device of embodiment:
Fig. 4 is the schematic diagram of shock wave control device in the case of having shock wave, and Fig. 5 is shock wave control device in the case of no shock wave
Schematic diagram, the right directional arrow in figure indicate direction of flow.When aircraft is in transonic flight, aircraft wing upper surface, which generates, swashs
Wave (λ type shock wave 3), referring to Fig. 3, static pressure hole 8 is by the pressure P at shock wave rear2It is transmitted into convex slot 6, due to pressure P in front of shock wave1
Than shock wave rear pressure P2It is small, the pressure difference Δ P to flexible covering 5 is just generated in convex slot 6, so that flexible covering 5 generates shape
Become, forms bulge, the generation of bulge influences shock wave form again, so that changing correspondingly in cabin to the pressure difference of flexible covering 5, so
Reciprocal variation, shape when finally bulge being made to reach equilibrium state, and then achieve the purpose that self adaptive control shock strength；When winged
In low-speed operations, upper surface of the airfoil generates machine without shock wave, referring to fig. 4, the interior pressure difference very little to flexible covering of convex slot 6,
Therefore it not will form bulge, the low-speed performance of aircraft will not be adversely affected at this time.The shock wave control device it is entire
Control process is not necessarily to the input of additional energy.
The shock wave control device of the present embodiment, has the advantage that
(1) adaptive bulge shape can be realized according to different flying conditions, improve bulge control shock wave damping device
Robustness, flight efficiency is high, and flight range is also wider.
(2) simple and reliable for structure, reaction speed is fast, not complicated mechanical device, without additional energy input, simultaneously
Structure light weight compact to design mitigates carrying wing burden, is conducive to aircraft loss of weight.
(3) smaller for space requirement, space utilization rate is high, this is particularly important for the trailing edge of narrow space, phase
Bulge concept adaptive for existing supercharging device, saves the weight and cost of supercharging equipment.
(4) flexible covering and the double design with porous plate convex slot realize the dual of auto Deformation and structural bearing
Target ensure that the continuous and derivable of aerofoil surface, will not cause additional viscous drag.
Embodiment 2:
In practical applications, the shock wave control device of embodiment can be applied to aircraft surfaces in the presence of the area compared with intense shock wave
Domain, predominantly airliner upper surface, as shown in Figure 6.First according to the type of aircraft, the main shock wave control zone of aircraft is determined
Domain, referring to Fig. 6, the place that the airliner in the present embodiment mainly generates shock wave is the aircraft connecting with airliner fuselage 9
1 middle section of wing is along the region of tangential 50%-80%, shock wave control area 10 as shown in FIG. 6；Then to shock wave control device
Specifically it is arranged, specifically, 1) static pressure hole 8 is arranged using discrete way, is placed in the most rear of shock wave control area 10, example
Such as, row's static pressure hole 8 can be designed in shock wave rear (downstream), wherein static pressure hole 8 uses round to reduce stress concentration, hole
Diameter is taken as local boundary layer thickness δ, and the distance between two holes are taken as 2 δ；2) bottom and the side use of convex slot 6 can meet
The opening setting of the metal material that strength and stiffness require, generally general aviation duralumin, hard alumin ium alloy material, convex slot 6 is porous
Plate 4 plays the role of carrying flexible covering 5 and is connected to convex slot 6, and convex slot 6 is closed using the very big aviation titanium of intensity and toughness
Gold, depth are taken as the δ of 5 δ~10；3) flexible covering 5 uses the silicon rubber for meeting deformation requirements and capable of carrying default normal load
Material or the honeycomb composite material that deformation design can be carried out.
The present embodiment perceives flow field pressure by designing row's static pressure hole 8 in shock wave downstream, and is transmitted by convex slot 6
To shock wave upstream, the exhibition for redesigning flexible covering 5 is drawn using the pressure difference in flow field itself in shock wave to rigidity and tangential rigidity
Flexible covering 5 realizes the adaptive deformation of bulge under the differential pressure action risen, is finally reached control shock strength and reduces shock wave resistance
The purpose of power.
Used herein a specific example illustrates the principle and implementation of the invention, and above embodiments are said
It is bright to be merely used to help understand method and its core concept of the invention；At the same time, for those skilled in the art, foundation
Thought of the invention, there will be changes in the specific implementation manner and application range.In conclusion the content of the present specification is not
It is interpreted as limitation of the present invention.
Claims (9)
1. a kind of shock wave control device, which is characterized in that described device is arranged on aircraft wing, and described device includes: porous
Plate and flexible covering；
Convex slot is opened up on the aircraft wing；The porous plate is arranged in the opening of the convex slot；It is covered on the porous plate
It is stamped flexible covering；The flexible covering and the wing cover of the convex slot surrounding are smoothly connected；On the aircraft wing also
It is provided with multiple static pressure holes；The static pressure hole is connected to the convex slot, is swashed for perceiving when the plane wing surfaces generate
The pressure at shock wave rear when wave, and the pressure perceived is transferred to the convex slot, so that the flexible covering generates shape
Become.
2. a kind of shock wave control device according to claim 1, which is characterized in that the static pressure hole is arranged described porous
Side of the plate close to wing tail portion.
3. a kind of shock wave control device according to claim 2, which is characterized in that between static pressure hole described in every two away from
From being equal.
4. a kind of shock wave control device according to claim 3, which is characterized in that between static pressure hole described in every two away from
The thickness of local boundary layer is indicated from for 2 δ, δ.
5. a kind of shock wave control device according to claim 1, which is characterized in that the shape of the static pressure hole is circle,
The diameter value of the static pressure hole and the thickness value of local boundary layer are equal.
6. a kind of shock wave control device according to claim 1, which is characterized in that the depth of the convex slot is 5 δ -10
The thickness of δ, δ expression local boundary layer.
7. a kind of shock wave control device according to claim 1, which is characterized in that the convex slot is opened in the aircraft
Wing middle section is along tangential region.
8. a kind of shock wave control device according to claim 1, which is characterized in that the material of the porous plate is aviation titanium
Alloy material.
9. a kind of shock wave control device according to claim 1, which is characterized in that the material of the flexible covering is silicon rubber
Glue material or honeycomb composite material.
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| Application Number | Priority Date | Filing Date | Title |
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| CN201811389511.1A CN109436293B (en) | 2018-11-21 | 2018-11-21 | Shock wave control device |
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| CN201811389511.1A CN109436293B (en) | 2018-11-21 | 2018-11-21 | Shock wave control device |
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| CN109436293A true CN109436293A (en) | 2019-03-08 |
| CN109436293B CN109436293B (en) | 2020-05-22 |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111255777A (en) * | 2020-02-14 | 2020-06-09 | 上海交通大学 | Method suitable for controlling dynamic skin flow of object |
| CN111619789A (en) * | 2020-05-08 | 2020-09-04 | 中国科学院空天信息创新研究院 | Blade upper surface airflow control device and method |
| CN113148110A (en) * | 2021-05-28 | 2021-07-23 | 西北工业大学 | Wing deformation device based on shock wave control bulge and wide-speed-range hypersonic aircraft |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| TR202021171A2 (en) * | 2020-12-22 | 2021-06-21 | Msg Teknoloji̇ Li̇mi̇ted Şi̇rketi̇ | Partially flexible wing profile made with silicone based flexible material |
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| CN111255777A (en) * | 2020-02-14 | 2020-06-09 | 上海交通大学 | Method suitable for controlling dynamic skin flow of object |
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| CN111619789A (en) * | 2020-05-08 | 2020-09-04 | 中国科学院空天信息创新研究院 | Blade upper surface airflow control device and method |
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| CN113148110A (en) * | 2021-05-28 | 2021-07-23 | 西北工业大学 | Wing deformation device based on shock wave control bulge and wide-speed-range hypersonic aircraft |
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|---|---|
| CN109436293B (en) | 2020-05-22 |
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