CN107352042A - A kind of drag reduction method in supersonic turbulent boundary layer - Google Patents
A kind of drag reduction method in supersonic turbulent boundary layer Download PDFInfo
- Publication number
- CN107352042A CN107352042A CN201710556494.5A CN201710556494A CN107352042A CN 107352042 A CN107352042 A CN 107352042A CN 201710556494 A CN201710556494 A CN 201710556494A CN 107352042 A CN107352042 A CN 107352042A
- Authority
- CN
- China
- Prior art keywords
- mrow
- msub
- surface temperature
- wall surface
- temperature
- 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.)
- Granted
Links
Classifications
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Manufacturing & Machinery (AREA)
- Transportation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Management, Administration, Business Operations System, And Electronic Commerce (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention discloses a kind of drag reduction method in supersonic turbulent boundary layer, is the temperature fluctuation striped of 100 sticky yardsticks by the width being superimposed on the turbulent region wall surface temperature of aircraft surface on a streamwise, obtains wall surface temperature distribution；And by printing the mesh resistive silk of minute yardstick on the turbulent area of aircraft surface, and thermocouple element is used for at-once monitor wall surface temperature in mesh node printing, the wall surface temperature for controlling the heating power wall surface temperature of each branch resistance silk on wall to meet to determine by computer is distributed.Compared with prior art, the positive effect of the present invention is：The present invention passes through the resistance wire circuit in object plane using minute yardstick in circuit printing techniques printing, and adopt the resistance wire heating power for each branch road that computerizeds control, so as to realize the accurate control of surface temperature distribution, pass through influencing each other for temperature fluctuation field and velocity fluctuation field, limit the development of speed band, turbulence pulsation can be suppressed, reduce resistance.
Description
Technical field
The present invention relates to field of fluid mechanics, and in particular to supersonic turbulent boundary layer flow, by introducing specific wall
Face Temperature Distribution realizes the control to turbulent flow in boundary layer, and realizes the reduction of frictional resistance.
Background technology
Speed can be improved by reducing the skin resistance of large transport airplane and long-range bomber, saved fuel, increased voyage, carry
Elevator power.In their drag overall, frictional resistance occupies very big proportion, such as the skin friction resistance of Typical Aircraft accounts for
The 50% of drag overall, body surface frictional resistance can be divided into two kinds of laminar flow frictional resistance and turbulent flow frictional resistance according to fluidised form, for most of
For aircraft, 80% surface is at turbulence state, its turbulent flow frictional resistance than about 1 order of magnitude under laminar flow conditions,
Therefore it is that one of most important work, its method is exactly that directly surface is fully sent out in the design of aircraft drag reduction to reduce turbulent flow frictional resistance
The turbulent flow of exhibition is controlled to reach the purpose that suppression turbulent flow reduces resistance.Research is found, not only can by the control to turbulent flow
Corresponding noise radiation can also largely be reduced by enough realizing the reduction of frictional resistance, can be improved the hidden ability of sound and be taken and relax
Adaptive.It is becoming tight in energy day, today of environmental pollution getting worse, turbulent flow drag reduction control technology causes extensively in the world
Attention, and one of aviation key technology of 21 century is classified as by US Airways and space travel office.
It is many in order to reduce the corresponding drag reduction technology of the resistance of body surface, it is to be directed to potential flows mostly, such as
Bubble is added in wall flowing, high molecular polymer is added, uses flexible wall, groove wall etc..This kind of technology is corresponding
Flowing velocity than relatively low, corresponding flowing control and drag reduction are easier to realize, but it is imitated for supersonic flows
Fruit is just had a greatly reduced quality.
The content of the invention
In order to overcome the disadvantages mentioned above of prior art, the present invention proposes a kind of drag reduction side in supersonic turbulent boundary layer
Method.
The technical solution adopted for the present invention to solve the technical problems is：A kind of drag reduction side in supersonic turbulent boundary layer
Method, comprise the following steps：
Step 1: the flow direction of the turbulent region of determination aircraft surface, thereon any p points(due to wall speed
Degree is zero, and flow direction here refers to that the limit flows to direction in fact) and wall surface temperature Tw(p)；
Step 2: in wall surface temperature Tw(p) temperature of the width on a flow direction for 100 sticky yardstick η is superimposed on
Degree fluctuation striped T ' (p), obtain wall surface temperature distribution T (p)=Tw(p)+T′(p)；
Step 3: the realization of wall surface temperature control：
The mesh resistive silk of minute yardstick, and the thermoelectricity in mesh node printing are printed on the turbulent area of aircraft surface
Even element is used for at-once monitor wall surface temperature, and the heating power wall of each branch resistance silk on wall is controlled by computer
Temperature meets the wall surface temperature distribution that step 2 determines.
Compared with prior art, the positive effect of the present invention is：The present invention in object plane using circuit printing techniques by being printed
The resistance wire circuit of minute yardstick is painted, and adopts the resistance wire heating power for each branch road that computerizeds control, so as to realize surface temperature
The control of distribution is spent, the speed of the temperature fringe distribution of flow direction, the width of fringe and turbulent boundary layer is finally realized on wall
Strip width yardstick is suitable, about 100 sticky yardsticks；Temperature fluctuation field and velocity fluctuation field are realized according to Reynolds analogy theory
Influence each other, limit the development of speed band, further according to turbulent flow from remaining theoretical, open up to the limited speed band of development
Turbulence pulsation can be suppressed, reduce resistance.
Brief description of the drawings
Examples of the present invention will be described by way of reference to the accompanying drawings, wherein：
Fig. 1 is the flow schematic diagram of aerofoil surface, wherein：1 far field flow, 2 wings, the speed point in 3 wall boundary layers
Cloth, 4 wall velocity attitudes, 5 turbulent areas, 6 laminar regions；
Fig. 2 is wall surface temperature distributed stripes schematic diagram；
Fig. 3 is wall surface temperature control schematic diagram, wherein：11. thermocouple, 12. micro resistance silks, 13. computers, 14. controls
Circuit, 15. temperature stripeds, 16. velocities；
Fig. 4 is the drag-reduction effect comparison diagram (direct Numerical result) of this method.
Embodiment
Aircraft its surface fluidised form in high-speed flight has two kinds of laminar flow and turbulent flow, and most regions are all in turbulent flow
(fuselage, the wing of such as aircraft), and the frictional resistance under turbulent flow fluidised form is than about 1 order of magnitude under laminar flow conditions,
Therefore the frictional resistance for reducing turbulent area is most important work in the design of aircraft drag reduction.The present invention intends real by following steps
Now reduce the purpose of turbulent area frictional resistance.
First, the determination of aircraft surface flowing.
The frictional resistance overwhelming majority of aircraft derives from the turbulent areas on surface, wants to reduce the resistance in the region first
It needs to be determined which is laminar region to flying surface, which is turbulent area, and further determines that surface arbitrfary point p flow directionWith wall surface temperature Tw(p), as shown in Figure 1.The research of this respect is always the focus and difficult point of hydrodynamics research, generally
There are three class methods to determine：One is exactly engineering experience method, such as it has been generally acknowledged that 5% region of leading edge is on conventional airfoil
Laminar region, remaining region are all turbulent flows, and fuselage then thinks entirely turbulent area, and flow direction is roughly the same with direction of flow,
Surface temperature is all incoming stagnation temperature.Second is exactly the method for wind tunnel test, can use naphthalene sublimation method, infrared measurement of temperature the methods of it is true
Determine the turbulent region and Temperature Distribution of aircraft surface, the flow direction on surface is determined by surface oil flow method.3rd class is just
It is method for numerical simulation, is exactly that Fluid Control Equation is solved by computer numerical, and by various turbulent flow forecast models, so that
The turbulent region of aircraft surface, flow direction and surface temperature distribution at judgement.Either which kind of method, its core are sought to
Determine turbulent region, the flow direction of surface p point of aircraft surfaceAnd wall surface temperature distribution Tw(p), to carry out
Corresponding turbulence control.
2nd, the design of wall surface temperature distribution.
When aircraft is with supersonic flight, energy equation is coupled with the equation of momentum and dominates flowing, at this moment conduct
The wall surface temperature distribution of the boundary condition of energy equation just embodies to the influence of wall flowing.According to the phase of turbulent flow
Theory is closed, certain negative correlativing relation is presented in temperature fluctuation and the velocity fluctuation of near wall region, and velocity fluctuation distribution determines
Speed band, therefore the purpose of turbulence control and drag reduction can be reached by the control to speed band.The present invention is reasonable in design
The distribution of wall surface temperature cause wall temperature to be pulsed, matched with the speed band distribution of near wall region, so as to influenceing speed band knot
The evolution of structure, to reach the purpose of drag reduction.
1st, wall surface temperature striped
It has been found that the Temperature Distribution T of wall in step 1w(p), a temperature fluctuation item is then being superimposed thereon
T ' (p), the fluctuation are can be represented along the striated structure flowed on direction, the striped by periodic function f (z)：
T ' (p)=f (zp) (1)
The cycle of the periodic function is s, that is, meets f (z)=f (z+s), and the average value in the cycle is zero
So a variety of periodic function forms can be selected, such as sin cos functionses, square wave, triangular wave function, as shown in Figure 2.Here z is
Curve in surface perpendicular to flowing is opened up to coordinate, and the curve of any p points is opened up on surface may be defined as to coordinate：
In formulaFor normal vector outside the unit of wall.Mathematically it is not difficult to verify, the temperature fluctuation striped normal direction and wall speed
It is vertical to spend direction, that is, has
, it is necessary to further determine that the width of striped after wall surface temperature stripe direction is determined, such as principle in front is chatted
To state, the width of temperature striped needs to be consistent with speed band, and the width of speed striped is usually 100 sticky yardstick η,
Take the cycle s=100 η in periodic function.
2nd, the estimation of sticky yardstick
Sticky yardstick η can be estimated according to equation below：
Wherein：
ρ in formula∞,u∞,μ∞,T∞The respectively density of incoming, speed, viscosity and temperature, M∞For flight Mach number,
For wall mean temperature,For theoretical constants.
So, wall surface temperature is set as
T (p)=Tw(p)+T′(p) (4)
Pass through above-mentioned steps, it is known that the Temperature Distribution that formula (4) provides there are following features：
1) mean temperature in wall is substantially free from change, the caused temperature fluctuations of only T ', therefore does not interfere near
The large scale fluidal texture in wall area；
2) wall surface temperature fluctuates to form striated structure, and the striped is parallel to wall flow direction；
3) width of fringe is suitable with the speed strip width of near wall region, about 100 sticky yardsticks.
3rd, the realization of wall surface temperature control.
The design method of wall surface temperature distribution is described in step 1 and step 2, and gives corresponding wall surface temperature
Distribution function.Sticky yardstick very little, it is micron dimension because the supersonic turbulent border of aircraft surface is than relatively thin, temperature strip
Line width is also submillimeter magnitude, and common means are difficult to realize that corresponding temperature is accurately controlled in the turbulent area of aircraft surface
System.It is horizontal according to current integrated circuit technique, it is possible to achieve the integrated circuit of micron dimension, therefore the present invention is proposed with integrated electricity
Road mode of printing realizes the control of wall surface temperature.
The mesh resistive silk of minute yardstick is printed on the turbulent area of aircraft surface using integrated circuit technique first, and
The upper thermocouple element of mesh node printing is used for at-once monitor wall surface temperature, as shown in Figure 3.The wall that will be provided in step 2
Temperature Distribution inputs computer, and the heating power of each branch resistance silk on wall is controlled by control circuit；By on wall
Thermocouple monitors the Temperature Distribution on wall in real time, and computer is fed back to from control circuit, a successive step branch resistance silk of going forward side by side
Heating power cause the temperature of wall to meet given distribution.
4th, temperature controlled drag principle
1st, temperature fluctuation field is with flowing to influencing each other for velocity fluctuation field
In supersonic speed compressible flows, because energy equation and momentum are coupled, pass through temperature field and velocity field
It must influence each other.Temperature striped in wall can produce the temperature fluctuation field of striated near wall region, according to Reynolds analogy
Theoretical Ru′T′≈ -1, for the temperature fluctuation T ' of near wall region with flowing to the negatively correlated relations of fluctuation velocity u ', i.e. temperature fluctuation T ' is small
It is big that place flows to fluctuation velocity u '.So near wall region form with temperature fluctuation field it is positive and negative it is opposite flow to velocity fluctuation field,
The speed strip width for flowing to velocity fluctuation place forming is also about 100 sticky yardsticks.So, speed band and wall are flowed to
Face temperature fringe distribution is closely related.
2nd, influence of the speed band to turbulent resistance
Because the speed band of near wall region is influenceed by wall surface temperature striped, its open up to development also by temperature fluctuation field
Limit, postponed the Instability of speed band, enhanced its stability, according to abundant development wall turbulent flow from maintaining to manage
By stable speed band can slow down Turbulence-bursting process, realize the suppression of turbulent flow, it is defeated in normal direction to reduce eddy stress
Fortune process, finally realize the reduction of turbulent skn friction resistance.
5th, direct Numerical is verified
In order to verify the correctness of this method and validity, comparative simulation has been carried out using the method for numerical simulation here,
Two kinds of supersonic turbulent boundary layers of uniform isothermal wall and temperature striped wall, corresponding temperature boundary condition are simulated respectively
Such as formula (5)：
Uniform isothermal wall：T=TW(
Temperature striped wall：T=TW+TW·0.3·cos(2π·z/s+) 5)
Fig. 4 gives the friction velocity u in the flow direction under two kinds of wall boundary conditionsτ(friction that wall is subject to should for contrast
Power isSo uτThe frictional resistance on the wall of boundary layer is directly reacted).It can be seen that after using this patent method,
The friction velocity of wall is reduced to 0.048 from original 0.051, reduces 6%, so this method is from feasible effective.
The present invention operation principle be：Minute yardstick resistance wire circuit is printed in body surface using circuit printing techniques, and
Realize that computer is controlled the resistance wire heating power of each subcircuits；By the control of each resistance wire heating power, in object
The temperature fringe distribution along flow direction is realized on surface, and temperature striped is about 100 sticky yardsticks；When supersonic flow flows through the table
Face, corresponding Temperature Distribution can boundary layer near wall region produce distributional class as temperature fluctuation field；According to the thunder of compressible turbulent flow
Promise analogy theory, what the temperature fluctuation field can influence boundary layer near wall region flows to velocity fluctuation field so that itself and temperature fluctuation field
It is distributed similar, obtain striated flows to velocity fluctuation field, i.e., the speed ribbon structure of turbulent flow is by the obvious of wall surface temperature striped
Control；The temperature striped obvious postpone of wall speed band Instability, so as to enhance the stability of speed band, according to
Fully development wall turbulent flow is theoretical from maintenance, and stable speed band can slow down Turbulence-bursting process, realizes and suppresses turbulent flow, reduction
The purpose of turbulent flow frictional resistance.
Claims (4)
- A kind of 1. drag reduction method in supersonic turbulent boundary layer, it is characterised in that：Comprise the following steps：Step 1: determine the turbulent region of aircraft surface, the flow direction of surface p pointAnd wall surface temperature Tw(p)；Step 2: in wall surface temperature Tw(p) temperature fluctuation of the width on a flow direction for 100 sticky yardstick η is superimposed on Striped T ' (p), obtain wall surface temperature distribution T (p)=Tw(p)+T′(p)；Step 3: the realization of wall surface temperature control：The mesh resistive silk of minute yardstick, and the thermocouple member in mesh node printing are printed on the turbulent area of aircraft surface Part is used for at-once monitor wall surface temperature, and the heating power wall surface temperature of each branch resistance silk on wall is controlled by computer Meet the wall surface temperature distribution that step 2 determines.
- A kind of 2. drag reduction method in supersonic turbulent boundary layer according to claim 1, it is characterised in that：The sticky chi Degree η is estimated according to equation below：<mrow> <mi>&eta;</mi> <mo>=</mo> <mfrac> <msub> <mi>&mu;</mi> <mi>&infin;</mi> </msub> <mrow> <msub> <mi>&rho;</mi> <mi>&infin;</mi> </msub> <msub> <mi>u</mi> <mi>&infin;</mi> </msub> </mrow> </mfrac> <mo>&CenterDot;</mo> <msubsup> <mover> <mi>T</mi> <mo>&OverBar;</mo> </mover> <mi>w</mi> <mrow> <mi>&omega;</mi> <mo>+</mo> <mn>1</mn> </mrow> </msubsup> <mo>&CenterDot;</mo> <msubsup> <mi>U</mi> <mrow> <mi>v</mi> <mi>d</mi> <mo>,</mo> <mi>w</mi> <mi>a</mi> <mi>k</mi> <mi>e</mi> </mrow> <mo>+</mo> </msubsup> <mo>&CenterDot;</mo> <mfrac> <mi>A</mi> <mrow> <mi>arcsin</mi> <mrow> <mo>(</mo> <mi>B</mi> <mo>+</mo> <mi>C</mi> <mo>)</mo> </mrow> <mo>-</mo> <mi>a</mi> <mi>r</mi> <mi>c</mi> <mi>s</mi> <mi>i</mi> <mi>n</mi> <mrow> <mo>(</mo> <mi>C</mi> <mo>)</mo> </mrow> </mrow> </mfrac> </mrow>Wherein：<mrow> <msub> <mover> <mi>T</mi> <mo>&OverBar;</mo> </mover> <mi>r</mi> </msub> <mo>=</mo> <mn>1</mn> <mo>+</mo> <mi>r</mi> <mfrac> <mrow> <mi>&gamma;</mi> <mo>-</mo> <mn>1</mn> </mrow> <mn>2</mn> </mfrac> <msubsup> <mi>M</mi> <mi>&infin;</mi> <mn>2</mn> </msubsup> <mo>,</mo> <msub> <mover> <mi>T</mi> <mo>&OverBar;</mo> </mover> <mi>w</mi> </msub> <mo>=</mo> <mfrac> <msub> <mi>T</mi> <mi>w</mi> </msub> <msub> <mi>T</mi> <mi>&infin;</mi> </msub> </mfrac> </mrow>ρ in formula∞,u∞,μ∞,T∞The respectively density of incoming, speed, viscosity and temperature, M∞For flight Mach number,For wall Face mean temperature, ω=0.7, r=0.89,
- A kind of 3. drag reduction method in supersonic turbulent boundary layer according to claim 1, it is characterised in that：The temperature wave Dynamic striped is represented using periodic function.
- A kind of 4. drag reduction method in supersonic turbulent boundary layer according to claim 3, it is characterised in that：The cycle letter Number includes：Sin cos functionses, square wave, triangular wave function.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710556494.5A CN107352042B (en) | 2017-07-10 | 2017-07-10 | Drag reduction method for supersonic turbulent boundary layer |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710556494.5A CN107352042B (en) | 2017-07-10 | 2017-07-10 | Drag reduction method for supersonic turbulent boundary layer |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN107352042A true CN107352042A (en) | 2017-11-17 |
| CN107352042B CN107352042B (en) | 2020-03-10 |
Family
ID=60292932
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201710556494.5A Active CN107352042B (en) | 2017-07-10 | 2017-07-10 | Drag reduction method for supersonic turbulent boundary layer |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN107352042B (en) |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108446419A (en) * | 2018-01-17 | 2018-08-24 | 天津大学 | A kind of supersonic boundary layer characteristic thickness evaluation method |
| CN109002572A (en) * | 2018-05-29 | 2018-12-14 | 南京航空航天大学 | A kind of array flows to the control of slit pressure-vaccum and reduces turbulent flow frictional resistance method |
| CN109505829A (en) * | 2018-11-28 | 2019-03-22 | 中国核电工程有限公司 | A kind of passive modular fluid resistance element |
| CN109682549A (en) * | 2019-03-01 | 2019-04-26 | 冀凯河北机电科技有限公司 | A kind of novel air pipe can be reduced pressure drop |
| CN110489823A (en) * | 2019-07-29 | 2019-11-22 | 天津大学 | The size design and arrangement of the miniature square column vortex generator of adaptive deformation |
| CN110531619A (en) * | 2019-08-28 | 2019-12-03 | 天津大学 | A kind of implementation method of flowing control |
| CN112395694A (en) * | 2020-12-03 | 2021-02-23 | 中国人民解放军国防科技大学 | Drag reduction control method for ultrahigh-speed turbulent boundary layer |
| CN113899657A (en) * | 2021-12-06 | 2022-01-07 | 中国空气动力研究与发展中心高速空气动力研究所 | Compressible Reynolds stress measurement method based on composite schlieren technology |
| CN116923689A (en) * | 2023-09-18 | 2023-10-24 | 中国空气动力研究与发展中心高速空气动力研究所 | Active flow control structure based on rapid height Wen Wenchang and preparation method |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6027078A (en) * | 1998-02-27 | 2000-02-22 | The Boeing Company | Method and apparatus using localized heating for laminar flow |
| CN101916314A (en) * | 2010-08-16 | 2010-12-15 | 北京理工大学 | High-speed aircraft lifting surface aerodynamic heating structure multidisciplinary optimization design platform |
| CN102872686A (en) * | 2011-07-14 | 2013-01-16 | 温岭安能节能科技有限公司 | Boundary layer control and mainstream perturbation coordinated integrated enhanced heat transfer method and system |
| CN106184743A (en) * | 2016-09-23 | 2016-12-07 | 中国人民解放军国防科学技术大学 | A kind of hypersonic aircraft fall by the use of thermal means controlled based on shock wave |
| US20170057644A1 (en) * | 2015-08-25 | 2017-03-02 | The Boeing Company | Synergetic noise absorption and anti-icing for aircrafts |
| CN106828887A (en) * | 2016-11-14 | 2017-06-13 | 中国航天空气动力技术研究院 | A kind of Vehicle nose's self adaptation thermal protection method |
-
2017
- 2017-07-10 CN CN201710556494.5A patent/CN107352042B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6027078A (en) * | 1998-02-27 | 2000-02-22 | The Boeing Company | Method and apparatus using localized heating for laminar flow |
| CN101916314A (en) * | 2010-08-16 | 2010-12-15 | 北京理工大学 | High-speed aircraft lifting surface aerodynamic heating structure multidisciplinary optimization design platform |
| CN102872686A (en) * | 2011-07-14 | 2013-01-16 | 温岭安能节能科技有限公司 | Boundary layer control and mainstream perturbation coordinated integrated enhanced heat transfer method and system |
| US20170057644A1 (en) * | 2015-08-25 | 2017-03-02 | The Boeing Company | Synergetic noise absorption and anti-icing for aircrafts |
| CN106184743A (en) * | 2016-09-23 | 2016-12-07 | 中国人民解放军国防科学技术大学 | A kind of hypersonic aircraft fall by the use of thermal means controlled based on shock wave |
| CN106828887A (en) * | 2016-11-14 | 2017-06-13 | 中国航天空气动力技术研究院 | A kind of Vehicle nose's self adaptation thermal protection method |
Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108446419A (en) * | 2018-01-17 | 2018-08-24 | 天津大学 | A kind of supersonic boundary layer characteristic thickness evaluation method |
| CN109002572A (en) * | 2018-05-29 | 2018-12-14 | 南京航空航天大学 | A kind of array flows to the control of slit pressure-vaccum and reduces turbulent flow frictional resistance method |
| CN109505829A (en) * | 2018-11-28 | 2019-03-22 | 中国核电工程有限公司 | A kind of passive modular fluid resistance element |
| CN109505829B (en) * | 2018-11-28 | 2021-12-03 | 中国核电工程有限公司 | Passive modularized fluid resistance element |
| CN109682549A (en) * | 2019-03-01 | 2019-04-26 | 冀凯河北机电科技有限公司 | A kind of novel air pipe can be reduced pressure drop |
| CN109682549B (en) * | 2019-03-01 | 2023-09-08 | 冀凯河北机电科技有限公司 | Novel trachea capable of reducing pressure drop |
| CN110489823A (en) * | 2019-07-29 | 2019-11-22 | 天津大学 | The size design and arrangement of the miniature square column vortex generator of adaptive deformation |
| CN110489823B (en) * | 2019-07-29 | 2023-06-09 | 天津大学 | Size design and arrangement mode of self-adaptive deformation miniature square column vortex generator |
| CN110531619B (en) * | 2019-08-28 | 2022-04-08 | 天津大学 | Method for realizing flow control |
| CN110531619A (en) * | 2019-08-28 | 2019-12-03 | 天津大学 | A kind of implementation method of flowing control |
| CN112395694A (en) * | 2020-12-03 | 2021-02-23 | 中国人民解放军国防科技大学 | Drag reduction control method for ultrahigh-speed turbulent boundary layer |
| CN113899657A (en) * | 2021-12-06 | 2022-01-07 | 中国空气动力研究与发展中心高速空气动力研究所 | Compressible Reynolds stress measurement method based on composite schlieren technology |
| CN116923689A (en) * | 2023-09-18 | 2023-10-24 | 中国空气动力研究与发展中心高速空气动力研究所 | Active flow control structure based on rapid height Wen Wenchang and preparation method |
Also Published As
| Publication number | Publication date |
|---|---|
| CN107352042B (en) | 2020-03-10 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN107352042A (en) | A kind of drag reduction method in supersonic turbulent boundary layer | |
| Yokokawa et al. | Experiment and CFD of a high-lift configuration civil transport aircraft model | |
| Lei et al. | Numerical study of separation on the trailing edge of a symmetrical airfoil at a low Reynolds number | |
| CN108009383A (en) | A kind of definite method and system of Natural Laminar Flow nacelle shape | |
| Dell’Orso et al. | Measurement of three-dimensional stall cells on a two-dimensional NACA0015 airfoil | |
| Hristov et al. | Poststall hysteresis and flowfield unsteadiness on a NACA 0012 airfoil | |
| Hahn et al. | Implicit large-eddy simulation of swept-wing flow using high-resolution methods | |
| Perez et al. | Instabilities on a hypersonic yawed straight cone | |
| Subramanian et al. | Investigation on the aerodynamics characteristics of dimple patterns on the aircraft wing | |
| Alaminos-Quesada et al. | Effect of the angle of attack on the transient lift during the interaction of a vortex with a flat plate. Potential theory and experimental results | |
| Goc et al. | Wind tunnel and grid resolution effects in large-eddy simulations of the high-lift common research model | |
| Sterrett | Transition fixing for hypersonic flow | |
| Murayama et al. | CFD Validation Study for a High-Lift Configuration of a Civil Aircraft Model | |
| Komuro et al. | Flow-control characteristics with nanosecond-pulse plasma actuator for different airfoil shapes | |
| Tasci et al. | The impact of the ground on flow structure and aerodynamic characteristics of a double delta wing | |
| Hariyadi | An analysis on Aerodynamics Performance Simulation of NACA 23018 Airfoil Wings on Cant Angles | |
| Aminaei et al. | Experimental and numerical study of wing boundary layer behavior in propeller flowfield | |
| Hinson et al. | Parametric exploration of wing–body junction flow using computational fluid dynamics | |
| Li et al. | Large-Eddy Simulation of Low-Reynolds-Number Flow Around Partially Porous Airfoils | |
| Yichen et al. | Near-wall topological patterns and flow structures over a simplified Danaus plexippus model | |
| Mi et al. | Numerical simulation of static and dynamic aerodynamics for formation flight with UCAVs. | |
| Ben-Gida et al. | Lift enhancement by a stationary leading-edge vortex over a high aspect ratio wing | |
| Samal et al. | Reduction of wingtip vortex from suction at wingtip | |
| Sogukpinar et al. | Numerical Investigation and Water Tunnel Experiment for F35 Fighter Jet | |
| Yi et al. | Design Optimization of Vortex Generators for a Junction Vortex of Wing-Body Configuration by Discrete Adjoint Approach |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |