CN109649666A - Boundary layer guiding and suction combined hypersonic air inlet channel flow control method - Google Patents
Boundary layer guiding and suction combined hypersonic air inlet channel flow control method Download PDFInfo
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- CN109649666A CN109649666A CN201811603785.6A CN201811603785A CN109649666A CN 109649666 A CN109649666 A CN 109649666A CN 201811603785 A CN201811603785 A CN 201811603785A CN 109649666 A CN109649666 A CN 109649666A
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- 238000000034 method Methods 0.000 title claims abstract description 42
- 239000002243 precursor Substances 0.000 claims abstract description 15
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- 238000004088 simulation Methods 0.000 claims description 5
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- 230000006835 compression Effects 0.000 description 3
- 238000007906 compression Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 3
- 239000011229 interlayer Substances 0.000 description 3
- 238000007443 liposuction Methods 0.000 description 3
- 244000062793 Sorghum vulgare Species 0.000 description 2
- 238000009434 installation Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C30/00—Supersonic type aircraft
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C2230/00—Boundary layer controls
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Abstract
The invention discloses a hypersonic air inlet channel flow control method combining boundary layer guiding and suction, which comprises the following steps of: forming a boundary layer guide profile on the hypersonic air inlet channel precursor to enable the flowing direction of air flow flowing through the boundary layer guide profile to be directionally changed through a transverse pressure gradient generated on the boundary layer guide profile; controlling a transverse pressure gradient acting on the boundary layer guiding profile to guide a boundary layer to a local part of the boundary layer guiding profile; and a suction hole is arranged at a part of the boundary layer guide profile. The scheme solves the problem of low suction performance in the prior art, realizes targeted suction and improves suction performance.
Description
Technical field
The present invention relates to hypersonic inlet technical field, especially a kind of boundary layer guides and aspirates the superb of combination
Velocity of sound inlet duct flow flowing control method.
Background technique
Boundary layer suction is considered as a kind of very effective inlet boundary layer flow control method.Boundary layer suction is
Suction hole is arranged in finger on wall surface, and the pressure difference at through hole both ends extracts boundary layer airflow out, to keep boundary layer thinning.Boundary
Layer suction is made of suction hole, suction chamber and bleed-off system etc., and the course of work enters suction chamber through suction hole for boundary layer, then
It is discharged by bleed-off system.Suction substantially reduces boundary layer thickness, and air intake duct capture flow quality is improved；Another party
Face, since boundary layer is thinning, the interference of air intake duct Shock/Boundary-Layer weakens, and intake duct starting performance also greatly improves.Currently, side
Interlayer liposuction technique has been widely used in surpassing/hypersonic inlet.
The boundary layer suction of the prior art includes suction hole, suction chamber, bleed-off system etc., and structure is relative complex, is increased simultaneously
Aircraft weight is added, suction chamber and bleed-off system also occupy the space of Vehicle nose.On the other hand, aircraft fuselage one
As use composite material, large area punching can destroy the structural strength of material on wall surface.
For hypersonic inlet application No. is: 201710784957.3 Chinese patent literature proposes one kind
Precursor and Bump have been carried out integrated design by Bump/ precursor integrated design method, this method, are solved Bump substantially and are existed
The excessive problem of flow losses in hypersonic air-flow.Bump/ precursor integration air intake duct passes through the cross generated on Bump type face
It is realized to barometric gradient and the row in boundary layer is moved, achieve preferable effect.But due to hypersonic flowing excessive velocities, laterally
Row moves apart from too long, and designs obtained Bump limited height, therefore this method moves the row of hypersonic boundary layer airflow
Ability is weaker.
Pneumatic research and experiment, 2010,28 (4): 1-6 is " based on numerical simulation air intake duct suction flowing control optimization design
[J] " Bump air intake duct is combined with boundary layer suction technology, the different installation site of suction socket is had studied to air intake duct
The influence of performance.Also there is the scheme that suction hole is arranged in the bottom side Bump in actual engineering design.These schemes are all by Bump
Combine with liposuction technique, to further increase inlet flow capture quality.The method that Bump and air intake duct suction combine
Ability is moved to the row in boundary layer to be obviously improved, but the mode of the two combination is also rested on mechanically combine at present
Stage.The installation site for studying different suction holes in Bump air intake duct for the first selects one in multi-scheme of only comforming
A preferably scheme, does not account for the flow feature of Bump air intake duct, has blindness.Suction hole is arranged in the bottom side Bump
Scheme is taken away boundary layer by suction hole in view of being accumulated in the bottom side Bump in boundary layer, and this scheme is applicable in
In moving more powerful supersonic speed Bump air intake duct to boundary layer row, and it is hypersonic under the conditions of Bump type face tend not to big portion
Point boundary layer row moves on to bottom side, therefore suction hole cannot play swabbing action well.
Summary of the invention
The present invention provides the hypersonic inlet flow control method that a kind of guidance of boundary layer and suction combine, for gram
Boundary layer airflow row moves the defects of performance is weaker to clothes in the prior art, realizes the effect for effectively playing aspirating hole, improves boundary layer
Air-flow row moves the purpose of performance.
To achieve the above object, the present invention proposes a kind of hypersonic inlet boundary layer flow flowing control method, including with
Lower step:
Step 1, boundary layer guide profile is formed on hypersonic inlet precursor, by the guide profile of boundary layer
Directed change occurs for the transverse-pressure gradient of generation, the flow direction for flowing through the air-flow of boundary layer guide profile,；
Step 2, boundary layer is directed to by the transverse-pressure gradient that acts in the boundary layer guide profile described
The part of boundary layer guide profile；
Step 3, in the local setting aspirating hole of the boundary layer guide profile.
Hypersonic inlet flow control method provided by the invention passes through what is be distributed to boundary layer guide profile pressure
It is well-designed, so that boundary layer guide profile is generated transverse-pressure gradient, to carry out active guidance to boundary layer airflow, makes boundary
Layer be directed separately under the action of transverse-pressure gradient boundary layer guide profile part (such as easily operated middle part and
Bottom), the position due to concentrating on boundary layer guide profile can form thicker boundary layer, in the thicker place arrangement in boundary layer
Suction hole, realization targetedly effectively aspirate boundary layer.
Detailed description of the invention
In order to more clearly explain the embodiment of the invention or the technical proposal in the existing technology, to embodiment or will show below
There is attached drawing needed in technical description to be briefly described, it should be apparent that, the accompanying drawings in the following description is only this
Some embodiments of invention for those of ordinary skill in the art without creative efforts, can be with
The structure shown according to these attached drawings obtains other attached drawings.
Fig. 1 is the hypersonic inlet flowing control that the boundary layer guidance that the embodiment of the present invention one provides and suction combine
Lower jaw formula axialsymmetrical inlet schematic diagram in method；
Fig. 2 is Fig. 1 left view；
Fig. 3 is the top view of Fig. 1；
Fig. 4 is the hypersonic inlet flowing control that the boundary layer guidance that the embodiment of the present invention one provides and suction combine
K in methodθWith theta function relation curve schematic diagram；
Fig. 5 is the hypersonic inlet flowing control that the boundary layer guidance that the embodiment of the present invention one provides and suction combine
X=x in method0Section inner boundary layer extends to distribution and puff profile band and determines schematic diagram.
The embodiments will be further described with reference to the accompanying drawings for the realization, the function and the advantages of the object of the present invention.
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 embodiment is only a part of the embodiments of the present invention, instead of all the embodiments.Base
Embodiment in the present invention, it is obtained by those of ordinary skill in the art without making creative efforts it is all its
His embodiment, shall fall within the protection scope of the present invention.
It is to be appreciated that the directional instruction (such as up, down, left, right, before and after ...) of institute is only used in the embodiment of the present invention
In explaining in relative positional relationship, the motion conditions etc. under a certain particular pose (as shown in the picture) between each component, if should
When particular pose changes, then directionality instruction also correspondingly changes correspondingly.
In addition, the description for being such as related to " first ", " second " in the present invention is used for description purposes only, and should not be understood as
Its relative importance of indication or suggestion or the quantity for implicitly indicating indicated technical characteristic.Define as a result, " first ",
The feature of " second " can explicitly or implicitly include at least one of the features.In the description of the present invention, " multiple " contain
Justice is at least two, such as two, three etc., unless otherwise specifically defined.
In the present invention unless specifically defined or limited otherwise, term " connection ", " fixation " etc. shall be understood in a broad sense,
For example, " fixation " may be a fixed connection, it may be a detachable connection, or integral；It can be mechanical connection, be also possible to
Electrical connection can also be physical connection or wireless communication connection；It can be directly connected, the indirect phase of intermediary can also be passed through
Even, the connection inside two elements or the interaction relationship of two elements be can be, unless otherwise restricted clearly.For this
For the those of ordinary skill in field, the specific meanings of the above terms in the present invention can be understood according to specific conditions.
It in addition, the technical solution between each embodiment of the present invention can be combined with each other, but must be general with this field
Based on logical technical staff can be realized, it will be understood that when the combination of technical solution appearance is conflicting or cannot achieve this
The combination of technical solution is not present, also not the present invention claims protection scope within.
The present invention proposes the hypersonic inlet flow control method of a kind of boundary layer guidance and suction combination.
Embodiment one
Referring to Fig. 1-5, the embodiment of the present invention provides the hypersonic inlet stream of a kind of boundary layer guidance and suction combination
Flowing control method, comprising the following steps:
Step 1, boundary layer guide profile is formed on hypersonic inlet precursor, thus in the guide profile of boundary layer
Transverse-pressure gradient is generated, is oriented by the flow direction that transverse-pressure gradient flows through the air-flow of boundary layer guide profile
Change；
By the working principle of supersonic speed Bump, the flow direction for flowing through the air-flow of boundary layer guide profile is oriented
Change, precursor type face generates transverse-pressure gradient, to carry out active guidance to boundary layer airflow, the direction of guidance can be pre-
Fixed, theoretically any direction is ok, and is also possible to being set according to actual design condition or that analysis obtains.Draw in boundary layer
Conductivity type face is a part of hypersonic aircraft precursor, and its role is to flow through the flowing of boundary layer guide profile overdraught
Directed change occurs for direction, and directed change here refers to flowing towards scheduled direction；And then boundary layer is influenced on boundary
The distribution of thickness, is adjusted boundary layer thickness under the action of guide profile in boundary layer in layer guide profile；
Step 2, boundary layer is directed to by the transverse-pressure gradient that acts in the boundary layer guide profile described
The part of boundary layer guide profile；
It can be any part in this local theory, be by being distributed to boundary layer guide profile pressure in the present embodiment
It is well-designed, so that boundary layer is directed separately to the middle part and bottom of boundary layer guide profile under the action of transverse-pressure gradient
Then portion is respectively arranged suction hole in the two regions, realizes the suction to boundary layer.It is optimal for being arranged in the two positions
Because being typically all only to be arranged in bottom, but hypersonic condition boundary layer is difficult bottom, place at middle part and
Bottom has carried out row respectively and has moved.This is also invention distinguishes in other importances.
In boundary layer, the distribution of boundary layer thickness can change under the action of guide profile.According to boundary layer in boundary layer
Characteristic distributions in guide profile determine puff profile band by given threshold, then on each puff profile band respectively
Suction hole is set.
Step 3, in the local setting aspirating hole of the boundary layer guide profile.In boundary layer, the part of guide profile can
Thicker boundary layer is formed, in the thicker place setting aspirating hole in boundary layer, the quantity of aspirating hole can be reduced, while improving pumping
Absorption energy.
Preferably, include: the step of formation boundary layer guide profile in the step 1
Step 11, the pressure on each osculating plane is defined by functional relation using variable θ and x as independent variable；θ be exhibition to
Variable, x are to flow to variable, the i.e. axial deflection of air intake duct.The ginseng on a two-dimensional surface is assured that by the two variables
Number, flow field symmetrical for outer shaft, the osculating plane, which refers to, flows to section with the plane of symmetry angle o degree；Referring to Fig. 1, region 1 is side
Interlayer guide profile；Boundary layer guide profile is a part of hypersonic aircraft precursor, and its role is to flow through boundary
Directed change occurs for the flow direction of layer guide profile overdraught, and the plane where face 2 is the plane of symmetry, and the plane where face 3 is
It is in the osculating plane of θ with the plane of symmetry, point 4 is coordinate origin, while being also boundary layer guide profile starting point, and point 5 is at x=L
Point, while being also boundary layer guide profile terminating point, each osculating plane is defined with variable θ, for Tidal Flow, closely
Face refers to parallel with the plane of symmetry, and distance is that θ flows to section；Axisymmetric flow field is turned for inside/outside, osculating plane refers to is in the plane of symmetry
θ angle flows to section.
Step 12, the functional relation of the pressure on each osculating plane and independent variable obtains each and opens up to scale factor
Pressure distribution curve on osculating plane；
Step 13, the corresponding flow field of pressure distribution curve is obtained based on the method for characteristic curves in each osculating plane；
Step 14, the flow field wall surface line in all osculating planes is subjected to curved surface setting-out, the type face of acquisition is drawn as boundary layer
Conductivity type face.
Preferably, the step 11 includes:
The intersection point 4 that coordinate origin is boundary layer guide profile start line and the plane of symmetry is defined, following functional relation is passed through
Define the pressure on each osculating plane:
P (θ, x)=kθy(x)+P0 (1)
Wherein, P0It is given value for the pressure value of coordinate origin；Independent variable x is abscissa, and value range is 0 < x < L,
Middle L is boundary layer guide profile terminated line abscissa, that is, puts 5 abscissa；Y (x) is the function of x, meets the following conditions: with x
It is increased monotonically；It is 0 in coordinate origin functional value, i.e. y (0)=0；The symmetrical flow field of the outer shaft of θ indicates osculating plane locating for the flow field
Angular range with the plane of symmetry is-θc<θ<θc, wherein θcFor given value；kθIt is the function of θ to open up to scale factor；Due to height
Supersonic speed precursor needs to undertake a part of compression duty, therefore boundary layer guide profile flow field edge flows to pressure and constantly increases
(boundary layer guide profile plays the role of compression to air-flow, so wall surface constantly increases along stroke pressure).
The step 12 includes:
By changing kθ, controlling each and opening up to scale factor is kθOsculating plane in pressure distribution curve, realize to pressure
Power be distributed in exhibition to control, in boundary layer, guide profile forms transverse-pressure gradient, realizes guidance to boundary layer.
Preferably, the step 12 includes:
kθAs θ increases monotone decreasing, by second dervative by kθCurve is divided into following three sections:
θ is located at [0, θ1] section when, kθ" (θ) > 0, and kθ(0)=1；
θ is located at [θ1,θ2] section when, kθ"(θ)<0；
θ is located at [θ2,θc] section when, kθ" (θ) > 0, and kθ(θc)=kmin, wherein kminFor specified value, range is 0 < kmin
<1；
kθCurve first derivative at θ=0 is zero, θ=θ1With θ=θ2It is continuous to locate first derivative.
Preferably, k in each sectionθSpecific functional relation includes the letters such as given polynomial type, trigonometric function type, exponential type
The specific analytic expression of at least one of number relationship.
As a preferred embodiment of the present invention: for example given y (x) is the functional relation of parabolic, a in formula
The compression of air-flow can be required to give according to practical precursor, value range is generally [0.01,1].
P (θ, x)=kθ(ax2)+P0 (2)
kθIt is the function of θ to open up to scale factor.The present invention is by changing kθ, realize to pressure be distributed in exhibition to control
System to form transverse-pressure gradient in boundary layer guide profile, and then realizes the guidance to boundary layer.Below to kθWith θ's
Functional relation is described.kθAlong the direction θ monotone decreasing, curve can be divided into three sections by second dervative is positive and negative.First segment is
[0,θ1] section, there is k in this sectionθ" (θ) > 0, meets kθ(0)=1.Second segment is [θ1,θ2] section, k in this sectionθ"(θ)<
0.Third section is [θ2,θc] section, there is k in this sectionθ" (θ) > 0, meets kθ(θc)=kmin, wherein kminFor specified value, range
For 0 < kmin<1.In addition, kθCurve first derivative at θ=0 is zero, θ=θ1With θ=θ2It is continuous to locate first derivative, to guarantee
Line smoothing.Concrete functional form can give polynomial type, trigonometric function type, exponential type etc. and specifically parse in each section
Formula.
One specific implementation case: the given quadratic function distribution form in three sections takes θ1=10, θ2=25, θc
=60；θ1Corresponding kθTake 0.8, kmin=0.6.Simultaneously curve meet it is above-mentioned required, equation number etc. under these conditions
In variable number, the quadratic function expression formula in three sections can be thus solved.(directly solve equation or be converted into line
Property algebraic solution can obtain totally 9 variables and 9 equations)
By k in each sectionθValue is discrete along the direction θ, and discrete accuracy rating is [1 °, 10 °].It is obtained according to the solution of formula 1
It is k that each, which is opened up to scale factor,θOsculating plane in pressure distribution curve.
To osculating plane pressure controllable flow field calculation, given pressure distribution is solved based on the method for characteristic curves in each osculating plane
The corresponding flow field of curve.This method is this field publicly-owned technology, solution procedure can refer to application No. is: 201710784957.3
Chinese patent literature and Beijing Higher Education Publishing House -2012 " aerodynamics [M] ".
It is that known two phases are solved using the method estimated-corrected based on there is the rotation method of characteristic curves in specific solution procedure
The downstream wall millet cake these two types unit process that the downstream inner of adjacent internal point selects, known upstream wall millet cake and adjacent inner are selected.So
The corresponding entire flow field of pressure distribution curve and wall surface molded line are solved according to space step-by-step system afterwards.
The corresponding wall surface of pressure distribution curve in osculating plane is thus obtained.Then it solves according to the method all close
The corresponding wall surface of pressure distribution curve in section.
The flow field wall surface line in all osculating planes is finally subjected to curved surface setting-out, obtained type face is boundary layer leading type
Face.
Preferably, the step 2 includes:
Step 21, numerical simulation is carried out to boundary layer guide profile, obtains boundary layer guide profile top thickness of boundary layer point
Cloth.In k of the present inventionθSetting boundary layer close to θ1<θ<θ2And θ=θcRegion, boundary layer thickness extend to formed pair
Answer the first protrusion and the second protrusion.
The step 3 includes:
Step 31, the characteristic distributions according to boundary layer on boundary layer type face determine puff profile band by threshold value；
Step 32, suction hole is respectively set on each puff profile band.
Preferably, the step 31 includes:
Step 311, in the case that boundary layer leading type is not added in boundary layer, x=x0The boundary layer thickness conduct in section
T0；
Step 312, for the first elevated regions of boundary layer, give proportional factor r 1, when boundary layer thickness meet T > r1 ×
T0When, assert the position on the first puff profile band；
For the second elevated regions of boundary layer, proportional factor r 2 is given, the boundary layer thickness at the position meets T > r2
×T0When, assert the position on the second puff profile band；
R1 and r2 is given value, and r1, r2 are all larger than zero, and r1 < r2；
Step 313, puff profile band position is carried out really perpendicular to the section in the direction x to each according to step 312
Recognize, obtains the first puff profile band and the second puff profile band in entire boundary layer guide profile.
As a specific embodiment of the invention, (axial direction of air intake duct) arranges one in the guide profile of boundary layer in the x-direction
Serial equidistantly to open up to section, section spacing value range is [10,100].
This, which sentences x=x, is confirmed to puff profile band position into section in each open up0It is illustrated for section.
With no addition boundary layer guide profile, identical x=x in the case of the Free Development of boundary layer0The boundary layer thickness in section is as base
Quasi- thickness, is denoted as To。
For the first elevated regions of boundary layer, proportional factor r is given1, when boundary layer thickness meets T > r1·To, it is assumed that
The position is on the first puff profile band.For the second elevated regions of boundary layer, proportional factor r is given2, side at the position
Thickness of boundary layer meets T > r2·To, it is assumed that the position is on the second puff profile band.r1And r2For given value, passed through according to design
It tests, the two value range takes [0.5,2] proper, in general meets r1<r2。
Each is opened up to section according to the method and carries out confirmation of the puff profile with position, available entire side
Puff profile band in interlayer guide profile.
The first puff profile band and the second puff profile band are obtained in the guide profile of boundary layer as a result,.
One specific implementation case: x=x is provided0Section inner boundary layer extend to distribution illustrate puff profile
The confirmation process of band.Boundary layer thickness extends apparent raised to forming two in figure.Numerical simulation obtains T in the sectiono=
0.58.Take r1=0.9, take r2=1.04, the puff profile band range for finally obtaining the first elevated regions is [10.2,28.8], is obtained
Puff profile band range to the second elevated regions is [48.0,58.6].
Preferably, the step 32 specifically includes:
Step 321, according to the area of given shape, suction area ratio and single suction hole, the sum of aspirating hole is obtained
Amount；Suction area ratio refers to that the area of the total suction hole percentage with area that accounts for puff profile, total area that aspirates are taken out divided by single
Total hole count can be obtained in sucker area.
Step 322, the distribution form in hole can be by the way of array equidistant on puff profile band, it is of course possible to root
It is arranged according to actual demand.Suction hole is laid respectively on each puff profile band according to the distribution form in hole.
The arrangement of suction hole is the publicly-owned technology in this field, and suction hole arrangement can be divided into two: determining suction hole shape,
Quantity；The distribution form of given suction hole.
The shape of given bore first, suction hole can be circle, triangle, rectangle etc..Then give suction area ratio and
The area of single suction hole, suction area ratio refers to the area of the total suction hole percentage with area that accounts for puff profile, total to aspirate
Total hole count can be obtained divided by single suction hole area in area.
The distribution form in hole can be by the way of array equidistant on puff profile band, it is of course possible to according to practical need
It asks and is arranged.
The working principle of supersonic speed Bump is introduced hypersonic precursor design by the present invention, and precursor type face is made to generate laterally pressure
Force gradient is distributed in boundary layer thickness in the guide profile of boundary layer and forms two to carry out active guidance to boundary layer airflow
A protrusion, then in the place arrangement suction hole of protrusion, to realize that a kind of boundary layer suitable for hypersonic inlet is drawn
Lead the boundary layer flow flowing control method of cooperation suction.
Compared with traditional liposuction technique, under identical suction effect, suction hole of the present invention is less, suction chamber
Also smaller, be conducive to the damaged condition for reducing fuselage material structural strength, while also saving the space of aircraft interior；With
Bump air intake duct is compared, and the present invention proposes boundary layer aiming at the problem that boundary layer row moves relative difficulty in hypersonic flowing
The boundary layer controllable way that guidance and suction match, solves the problems, such as this substantially；Cooperate with existing Bump and suction
Control mode is compared, and the present invention guides boundary layer, then targetedly taken out by precursor type face special designing
It inhales, the effect of suction hole can be played to a greater extent.By numerical simulation, it was demonstrated that the program is feasible.
The above description is only a preferred embodiment of the present invention, is not intended to limit the scope of the invention, all at this
Under the inventive concept of invention, using equivalent structure transformation made by description of the invention and accompanying drawing content, or directly/use indirectly
It is included in other related technical areas in scope of patent protection of the invention.
Claims (9)
1. the hypersonic inlet flow control method that a kind of guidance of boundary layer and suction combine, which is characterized in that including with
Lower step:
Step 1, boundary layer guide profile is formed on hypersonic inlet precursor, by generating in the guide profile of boundary layer
Transverse-pressure gradient, flow through the air-flow of boundary layer guide profile flow direction occur directed change；
Step 2, boundary layer is directed to by the boundary by the transverse-pressure gradient acted in the boundary layer guide profile
The part of layer guide profile；
Step 3, in the local setting aspirating hole of the boundary layer guide profile.
2. the hypersonic inlet flow control method that boundary layer guidance as described in claim 1 and suction combine, special
The step of sign is, boundary layer guide profile is formed in the step 1 include:
It step 11, is that independent variable defines the pressure on each osculating plane by functional relation with variable x, θ；X is air intake duct axis
To variable, flow field symmetrical for outer shaft, the osculating plane, which refers to, flows to section with the plane of symmetry angle o degree；
Step 12, it according to the functional relation of pressure and independent variable on each osculating plane, obtains each and opens up to scale factor
Pressure distribution curve on osculating plane；
Step 13, the corresponding flow field of pressure distribution curve is obtained based on the method for characteristic curves in each osculating plane；
Step 14, the flow field wall surface line in all osculating planes is subjected to curved surface setting-out, the type face of acquisition is as boundary layer leading type
Face.
3. the hypersonic inlet flow control method that boundary layer guidance as claimed in claim 2 and suction combine, special
Sign is that the step 11 includes:
The intersection point that coordinate origin is boundary layer guide profile start line and the plane of symmetry is defined, is defined by following functional relation every
Pressure on a osculating plane:
P (θ, x)=kθy(x)+P0 (1)
Wherein, P0It is given value for the pressure value of coordinate origin；Independent variable x is abscissa, and value range is 0 < x < L, and wherein L is
Boundary layer guide profile terminated line abscissa；Y (x) is the function of x, meets the following conditions: being increased monotonically with x；In coordinate origin
Functional value is 0, i.e. y (0)=0；θ indicates the angle of osculating plane locating for the flow field and the plane of symmetry, model for the symmetrical flow field of outer shaft
It encloses for-θc<θ<θc, wherein θcFor given value；kθIt is the function of θ to open up to scale factor；
The step 12 includes:
By changing kθ, it is k that each, which is opened up to scale factor,θOsculating plane in pressure distribution curve, realization pressure is distributed in
Open up to control, in boundary layer, guide profile forms transverse-pressure gradient, realizes guidance to boundary layer.
4. the hypersonic inlet flow control method that boundary layer guidance as claimed in claim 3 and suction combine, special
Sign is that the step 12 includes:
kθAs θ increases monotone decreasing, by second dervative by kθCurve is divided into following three sections:
θ is located at [0, θ1] section when, kθ" (θ) > 0, and kθ(0)=1；
θ is located at [θ1,θ2] section when, kθ"(θ)<0；
θ is located at [θ2,θc] section when, kθ" (θ) > 0, and kθ(θc)=kmin, wherein kminFor specified value, range is 0 < kmin<1；
kθCurve first derivative at θ=0 is zero, θ=θ1With θ=θ2It is continuous to locate first derivative.
5. the hypersonic inlet flow control method that boundary layer guidance as claimed in claim 4 and suction combine, special
Sign is, k in each sectionθSpecific functional relation includes at least one of given polynomial type, trigonometric function type, exponential type tool
Body analytic expression.
6. the hypersonic inlet flow control method that boundary layer guidance as claimed in claim 5 and suction combine, special
Sign is, boundary layer is directed to the part of the boundary layer guide profile and includes: in the step 2
The variation in the boundary layer thickness direction concentrate on the boundary layer guide profile extend to middle part and bottom；
The step 3 includes:
Step 31, the characteristic distributions according to boundary layer on boundary layer type face determine puff profile band by threshold value；
Step 32, suction hole is respectively set on each puff profile band.
7. the hypersonic inlet flow control method that boundary layer guidance as claimed in claim 6 and suction combine, special
Sign is that the step 2 includes:
Step 21, numerical simulation is carried out to boundary layer guide profile, obtained close to θ1<θ<θ2And θ=θcRegion, boundary
Layer thickness profile extend to be correspondingly formed first protrusion and second protrusion.
8. the hypersonic inlet flow control method that boundary layer guidance as claimed in claim 7 and suction combine, special
Sign is that the step 31 includes:
Step 311, in the case that boundary layer leading type is not added in boundary layer, x=x0When the section perpendicular to the direction x side
Thickness of boundary layer is as T0；
Step 312, for the first elevated regions of boundary layer, proportional factor r 1 is given, when boundary layer thickness meets T > r1 × T0When,
Assert the position on the first puff profile band；
For the second elevated regions of boundary layer, proportional factor r 2 is given, the boundary layer thickness at the position meets T > r2 × T0
When, assert the position on the second puff profile band；
R1 and r2 is given value, and r1, r2 are all larger than zero, and r1 < r2；
Step 313, confirmation of the puff profile with position is carried out to each section perpendicular to the direction x according to step 312, obtained
Obtain the first puff profile band and the second puff profile band in entire boundary layer guide profile.
9. being guided according to the described in any item boundary layers of claim 6~8 and the hypersonic inlet of suction combination flowing control
Method processed, which is characterized in that the step 32 includes:
Step 321, according to the area of given shape, suction area ratio and single suction hole, the total quantity of aspirating hole is obtained；
Step 322, suction hole is laid on each puff profile band according to the distribution form in hole respectively.
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110990955A (en) * | 2019-12-12 | 2020-04-10 | 中国人民解放军国防科技大学 | Hypersonic speed Bump air inlet channel design method and system |
Citations (16)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4815279A (en) * | 1985-09-27 | 1989-03-28 | The United States Of America As Represented By The National Aeronautics And Space Administration | Hybrid plume plasma rocket |
| US6216982B1 (en) * | 1998-05-06 | 2001-04-17 | Daimlerchrysler Aerospace Airbus Gmbh | Suction device for boundary layer control in an aircraft |
| RU2385820C1 (en) * | 2009-05-04 | 2010-04-10 | Государственное образовательное учреждение высшего профессионального образования Томский государственный университет (ТГУ) | Cooling method of taper head end of flying machine |
| CN102562461A (en) * | 2010-12-21 | 2012-07-11 | 通用电气公司 | System and method of operating an active flow control system for manipulating a boundary layer across a rotor blade of a wind turbine |
| CN104890887A (en) * | 2015-04-20 | 2015-09-09 | 南京航空航天大学 | Supersonic-velocity hypersonic-velocity gas inlet duct adopting pneumatic unstart control method |
| CN104975950A (en) * | 2015-06-16 | 2015-10-14 | 南京航空航天大学 | Method for determining binary hypersonic inlet passage based on appointed wall pressure distribution |
| CN105956286A (en) * | 2016-05-06 | 2016-09-21 | 北京航空航天大学 | Prediction method of forecabin thermal protection system whole trajectory temperature boundary of hypersonic velocity aircraft |
| US20170240271A1 (en) * | 2015-11-11 | 2017-08-24 | The Arizona Board Of Regents On Behalf Of The University Of Arizona | Control of hypersonic boundary layer transition |
| US20170313413A1 (en) * | 2000-05-31 | 2017-11-02 | Kevin Kremeyer | Shock Wave Modification Method and System |
| CN107590330A (en) * | 2017-09-04 | 2018-01-16 | 中国人民解放军国防科技大学 | Design method of two-dimensional pre-compressed precursor with boundary layer displacement |
| CN107628266A (en) * | 2017-09-04 | 2018-01-26 | 中国人民解放军国防科技大学 | Design method of axisymmetric pre-compression precursor with boundary layer displacement |
| US20180043996A1 (en) * | 2016-08-11 | 2018-02-15 | General Electric Company | Inlet assembly for an aircraft aft fan |
| US10071798B2 (en) * | 2012-11-19 | 2018-09-11 | The Regents Of The University Of California | Hypersonic laminar flow control |
| US10118696B1 (en) * | 2016-03-31 | 2018-11-06 | Steven M. Hoffberg | Steerable rotating projectile |
| CN108999704A (en) * | 2018-08-17 | 2018-12-14 | 中国人民解放军国防科技大学 | Hypersonic air inlet starting method and starting device |
| CN109033525A (en) * | 2018-06-27 | 2018-12-18 | 浙江大学 | A kind of hypersonic transition prediction method based on simplified three equation transition models |
-
2018
- 2018-12-26 CN CN201811603785.6A patent/CN109649666B/en active Active
Patent Citations (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4815279A (en) * | 1985-09-27 | 1989-03-28 | The United States Of America As Represented By The National Aeronautics And Space Administration | Hybrid plume plasma rocket |
| US6216982B1 (en) * | 1998-05-06 | 2001-04-17 | Daimlerchrysler Aerospace Airbus Gmbh | Suction device for boundary layer control in an aircraft |
| US20180170525A1 (en) * | 2000-05-31 | 2018-06-21 | Kevin Kremeyer | Shock Wave Modification Method and System |
| US20170313413A1 (en) * | 2000-05-31 | 2017-11-02 | Kevin Kremeyer | Shock Wave Modification Method and System |
| RU2385820C1 (en) * | 2009-05-04 | 2010-04-10 | Государственное образовательное учреждение высшего профессионального образования Томский государственный университет (ТГУ) | Cooling method of taper head end of flying machine |
| CN102562461A (en) * | 2010-12-21 | 2012-07-11 | 通用电气公司 | System and method of operating an active flow control system for manipulating a boundary layer across a rotor blade of a wind turbine |
| US10071798B2 (en) * | 2012-11-19 | 2018-09-11 | The Regents Of The University Of California | Hypersonic laminar flow control |
| CN104890887A (en) * | 2015-04-20 | 2015-09-09 | 南京航空航天大学 | Supersonic-velocity hypersonic-velocity gas inlet duct adopting pneumatic unstart control method |
| CN104975950A (en) * | 2015-06-16 | 2015-10-14 | 南京航空航天大学 | Method for determining binary hypersonic inlet passage based on appointed wall pressure distribution |
| US20170240271A1 (en) * | 2015-11-11 | 2017-08-24 | The Arizona Board Of Regents On Behalf Of The University Of Arizona | Control of hypersonic boundary layer transition |
| US10118696B1 (en) * | 2016-03-31 | 2018-11-06 | Steven M. Hoffberg | Steerable rotating projectile |
| CN105956286A (en) * | 2016-05-06 | 2016-09-21 | 北京航空航天大学 | Prediction method of forecabin thermal protection system whole trajectory temperature boundary of hypersonic velocity aircraft |
| US20180043996A1 (en) * | 2016-08-11 | 2018-02-15 | General Electric Company | Inlet assembly for an aircraft aft fan |
| CN107628266A (en) * | 2017-09-04 | 2018-01-26 | 中国人民解放军国防科技大学 | Design method of axisymmetric pre-compression precursor with boundary layer displacement |
| CN107590330A (en) * | 2017-09-04 | 2018-01-16 | 中国人民解放军国防科技大学 | Design method of two-dimensional pre-compressed precursor with boundary layer displacement |
| CN109033525A (en) * | 2018-06-27 | 2018-12-18 | 浙江大学 | A kind of hypersonic transition prediction method based on simplified three equation transition models |
| CN108999704A (en) * | 2018-08-17 | 2018-12-14 | 中国人民解放军国防科技大学 | Hypersonic air inlet starting method and starting device |
Non-Patent Citations (1)
| Title |
|---|
| 陈坚强: "高超声速边界层转捩研究现状与发展趋势", 《空气动力学学报》 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110990955A (en) * | 2019-12-12 | 2020-04-10 | 中国人民解放军国防科技大学 | Hypersonic speed Bump air inlet channel design method and system |
| CN110990955B (en) * | 2019-12-12 | 2024-05-28 | 中国人民解放军国防科技大学 | Hypersonic speed Bump air inlet channel design method and hypersonic speed Bump air inlet channel design system |
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