CN106885685A - A kind of dual airfoil test model for flow transition detection - Google Patents
A kind of dual airfoil test model for flow transition detection Download PDFInfo
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- CN106885685A CN106885685A CN201710220490.XA CN201710220490A CN106885685A CN 106885685 A CN106885685 A CN 106885685A CN 201710220490 A CN201710220490 A CN 201710220490A CN 106885685 A CN106885685 A CN 106885685A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
- G01M9/08—Aerodynamic models
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Abstract
The present invention relates to a kind of dual airfoil test model for flow transition detection, belong to wind tunnel test technical field, a containing groove is provided with the flow transition area of observation coverage of the upper surface of the test model, whole test model is only filled with heat insulation layer in the containing groove.The present invention to test model by way of using local insulation, only it is provided for accommodating the destruction of finely ground particles during the containing groove of the adiabatic paint to be sprayed avoids wind-tunnel air to adiabatic lacquer painting in flow transition observation area, in addition, painted because model integral spray is adiabatic, it is difficult to control to adiabatic paint thickness and the uniformity, and also it is difficult to ensure that the final type face precision after model spray painting, this model accommodates adiabatic paint by designing containing groove in observation area, and during the adiabatic lacquer painting of polishing, using the thickness of containing groove severity control thermal insulation enamelled coating, it is easy to the final lacquer painting for ensureing observation area to reach the type face requirement on machining accuracy of model.
Description
Technical field
The present invention relates to wind tunnel test technical field, more particularly to a kind of dual airfoil experiment for flow transition detection
Model.
Background technology
When seating plane cruising flight, air flows through passenger plane body surface, and the windage drag effect of generation is in winged
On machine.Air is divided into laminar flow and turbulent flow in the flow regime of body surface.Usually, near the leading edge of a wing very little region
Interior, air flow condition is laminar flow, then turns to twist and is changed into turbulent flow.Local windage resistance ratios turbulent flow produced by laminar flow
It is small.By aerodynamic design means, postpone air flow and turn to twist position, increase laminar flow zone area, it is possible to decrease Quan Jikong
Gas frictional resistance, and then reduce aircraft oil consumption.
In the prior art, laminar flow wing is the wing with large area laminar flow during a kind of aircraft cruising flight, is being tested
In the wind tunnel test of card laminar flow wing design or flight test, aerofoil surface air stream is detected usually using infrared imagery technique
It is dynamic to turn to twist position.Specifically, there is the principle of temperature difference by using adiabatic wall laminar flow zone and turbulent region, it is infrared into
As technology can wall flowing by laminar flow be changed into turbulent flow turn twist position and record.To meet the thermal insulation of test model wall
It is required that, i.e., prevent model from producing heat exchange with air as far as possible in process of the test, one layer must be sprayed on model metal surface
Thermal insulation paint.The spraying operation of thermal insulation paint is entirely to be completed by by hand.After spray painting, also need to beat lacquer painting repeatedly
Mill, to ensure the type face precision and surface roughness of the model after spray painting.Additionally, the lacquer painting of adiabatic paint is more crisp, by small particle
Small rut is also easy to produce after shock and cannot be repaired.In wind tunnel test, these small ruts often lure the layer on model wall into
Stream turns to twist in advance, disturbs the judgement to flow transition position.
Therefore, a kind of dual airfoil test model for flow transition detection how is provided, to solve in the prior art
Small rut is produced due to the lacquer painting of the adiabatic paint of small particle shock during wind tunnel test, causes the laminar flow on model wall to carry
Forward is twisted, technical problem of the interference to the judgement of flow transition position.
The content of the invention
It is an object of the invention to propose a kind of dual airfoil test model for flow transition detection, can solve the problem that existing
There is the above-mentioned technical problem in technology.
It is that, up to this purpose, the present invention uses following technical scheme：
A kind of dual airfoil test model for flow transition detection, in the flow transition of the upper surface of the test model
A containing groove is provided with the area of observation coverage, whole test model is only filled with heat insulation layer in the containing groove.
Further, the upper surface of test model, lower surface, leading edge and trailing edge are distributed with and connect with electric pressure scaner
The pressure tap for connecing, and be arranged at the chord length of 0.65~0.75 times of upper surface, the quantity of the pressure tap of leading edge and trailing edge accounts for pressure measurement
The 60%~70% of hole total quantity.
Further, the central axis of the pressure tap for being set on the upper surface, lower surface, leading edge and trailing edge is in the examination
Test on the same oblique section of model, the oblique section is 10 °~15 ° with the tangential angle of the test model.
Further, the pressure measurement hole number that the end of the trailing edge is set is one.
Further, little groove is provided near the pressure measurement section of the lower surface of test model, is installed in the little groove
There is incidence vane.
Further, the lower surface of the test model is provided with the big groove for accommodating pressure-measuring pipe, the pressure tap
Scanned with electron pressure by the pressure-measuring pipe and be connected；
The little groove is located in the big groove.
Further, the depth of the containing groove is 70mm~120mm.
Further, the two ends of the test model upper surface are equipped with a leveling groove, the leveling groove
Bottom surface as a reference plane, for the leveling of incidence vane.
Further, the lower surface two ends of the test model are equipped with interface, and the interface is articulated with vertical pole one
End；
The vertical pole other end is articulated with rotating mechanism.
Further, the test model is respectively arranged at two ends with left auricle and auris dextra piece, the left auricle and auris dextra piece
It is connected to the driven end and drive end of rotating mechanism.
Beneficial effect：
The present invention sets in flow transition observation area by way of using local insulation to test model, only and uses
The destruction of finely ground particles in the containing groove for accommodating the adiabatic paint to be sprayed avoids wind-tunnel air to adiabatic lacquer painting, this
Outward, painted because model integral spray is adiabatic, it is difficult to adiabatic paint thickness and the uniformity of control, and after being also difficult to ensure that model spray painting
Final type face precision, this model accommodates adiabatic paint by designing containing groove in observation area, and in the adiabatic lacquer painting process of polishing
In, using the thickness of containing groove severity control thermal insulation enamelled coating, it is easy to the final lacquer painting for ensureing observation area to reach model
Type face requirement on machining accuracy.
Brief description of the drawings
Fig. 1 is the structural representation of the dual airfoil test model provided by the present invention for flow transition detection；
Fig. 2 is the schematic diagram of the upper surface of the dual airfoil test model provided by the present invention for flow transition detection；
Fig. 3 is the schematic diagram of the lower surface of the dual airfoil test model provided by the present invention for flow transition detection；
Fig. 4 is the pressure tap provided by the present invention for the dual airfoil test model of flow transition detection in test model
On distribution schematic diagram.
In figure：
100th, test model；1st, upper surface；2nd, containing groove；3rd, pressure tap；4th, lower surface；5th, big groove；6th, little groove；
7th, leveling groove；8th, vertical pole；9th, left auricle；10th, auris dextra piece；11st, driven end；12nd, drive end.
Specific embodiment
In order that those skilled in the art more fully understand technical scheme, below in conjunction with the accompanying drawings and by specific
Implementation method further illustrates technical scheme.
A kind of dual airfoil test model for flow transition detection is present embodiments provided, such as Fig. 1-Fig. 4 is shown to be somebody's turn to do
Dual airfoil test model can be used to design the band aerofoil profile lifting surface model that all research Natural Laminar Flows turn twist detection test, like flying
The lifting surface of machine, propeller blade, helicopter blade, turbine impellers and wind power generator oar blade etc., as depicted in figs. 1 and 2, at this
A containing groove 2 is provided with the flow transition area of observation coverage of the upper surface 1 of test model 100, whole test model 100 is only described
Containing groove 2 is filled with heat insulation layer, specifically, the heat insulation layer used in the present embodiment is painted for thermal insulation.
As shown in Figures 2 and 3, X-direction is tangential for test model 100, Z-direction for test model 100 exhibition to.Due to
Flow transition area of observation coverage during experiment is located at the upper surface 1 of test model 100, and for natural laminar flow aerofoil profile, test model 100
It is turbulent flow by laminar flow transition at 30% to 60% chord length that the flowing of upper surface 1 is general, that is, turn to twist line positioned at 30% to 60% string
Strong point.As long as ensureing that the wall of test model 100 for turning to twist line region is adiabatic, infrared camera just can be by differentiating layer
Temperature difference handle turn is twisted line and is recorded between stream region and turbulent region.Thus, the present embodiment is used by test model 100
The mode of local insulation, i.e., be only provided for accommodating the containing groove of the adiabatic paint to be sprayed in flow transition observation area
2, the destruction this avoid the finely ground particles in wind-tunnel air to adiabatic lacquer painting, while using the thermal insulation of local insulation
Effect is also reached using overall adiabatic insulation effect.
Further, since the adiabatic paint of model integral spray, it is difficult to adiabatic paint thickness and the uniformity of control, and it is difficult to ensure that model
Final type face precision after spray painting.This model accommodates adiabatic paint by designing containing groove 2 in observation area, and adiabatic in polishing
During lacquer painting, using the thickness of the severity control of containing groove 2 thermal insulation enamelled coating, it is easy to the final lacquer painting for ensureing observation area to reach
To the type face requirement on machining accuracy of model；Meanwhile, it also avoid the metal of test model 100 in the coat of paint in process of the test
There is heat exchange with air in surface.Wherein, the depth of the containing groove 2 for being set in the present embodiment is 70~120mm, due to accommodating
The depth of groove 2 is less than normal to influence heat-insulating property and the depth of containing groove 2 is bigger than normal that process time can be caused to extend, processing cost
Can increase, so, the depth of containing groove 2 is preferably 100mm in the present embodiment.
With reference to Fig. 4, the pressure tap 3 is in the upper surface 1 of the experimental model, lower surface 4, leading edge and non-homogeneous point of trailing edge
Cloth；And be arranged at 0.65~0.75 times of chord length of upper surface 1, that the quantity of the pressure tap 3 of leading edge and trailing edge accounts for pressure tap 3 is total
The 60%~70% of quantity.；The quantity of the pressure tap 3 that the end of the trailing edge is set is one.By pressure tap 3 in the present embodiment
It is arranged near the shock-wave spot of the leading edge of test model 100, trailing edge and upper surface 1, pressure change is all relatively more violent, this survey
The pressure arrangement of hole 3 allows that the pressure distribution curve that experiment is obtained reduces real pressure point well in these positions
Cloth, improves the calculating accuracy of section lift.In particular, because the space of model trailing edges is very narrow and small, it is impossible to which arrangement is too many
The place of pressure tap 3, the present embodiment is also disposed with pressure tap 3 so that pressure distribution curve can be closed at trailing edge, it is to avoid
In the prior art due to the feelings of the error for introducing new can only be calculated to lift by interpolation to obtain the pressure value of trailing edge ends
Condition.
The pressure measurement that the pressure tap 3 set on the lower surface 4, leading edge and trailing edge of the test model 100 is set with upper surface 1
On the same oblique section of the test model 100, the oblique section angle tangential with experimental model is the central axis in hole 3
10 °~15 °, it is to avoid the pressure testing results of the caused disturbing influence downstream pressure tap 3 in flow field of upstream pressure tap 3.
Distribution of the pressure tap 3 on the upper surface 1 by containing groove 2, thus, is provided with portion in containing groove 2
Divide pressure tap 3.It should be noted that when adiabatic paint spray to containing groove 2, the pressure tap 3 on containing groove 2 is connected
Pressure-measuring pipe, pressure-measuring pipe is connected with outside air pump, and outside air pump output pressure is 0.11~0.15 MPa of air, now, air
Sprayed via pressure tap 3, it is to avoid during the adiabatic paint of spraying, the pressure tap 3 being arranged in containing groove 2 occurs stifled
Plug.
As shown in figure 3, the lower surface 4 of the test model 100 is provided with the big groove 5 for accommodating pressure-measuring pipe, pressure measurement
Hole 3 is connected by the pressure-measuring pipe with electric pressure scaner, additionally, the segmentation of structures line of most test model 100 also divides
It is distributed in the lower surface 4 of test model 100 so that the upper surface 1 of test model 100 is clean to do to observation area flowing reducing
Disturb.
Wherein, by big groove 5, the side pressure hole 3 being distributed on lower surface 4 is through excessive for distribution of the pressure tap 3 on lower surface 4
One is provided with groove 5, the big groove 5 is used to install the little groove 6 of incidence vane, and little groove 6 is arranged at test model 100
Lower surface 4 side pressure section near.The dual airfoil test model that this implementation is provided is a kind of conventional model in wind tunnel,
The characteristics of this class model, is chord length identical aerofoil profile for any cross section of model.Usually, the two ends of this class model respectively by
On the rotating mechanism of Rectangular Wind Tunnel test section side wall, it is easy to the angle of attack of adjustment model in experiment.The side of test model 100
Rotating mechanism motor is installed, be referred to as drive end 12；Opposite side is fitted without motor, is referred to as driven end
11.In wind tunnel test, dual airfoil model under lift and pitching moment collective effect and produce extend to torsional deflection,
The angle of attack of any cross section of test model 100 is caused to differ.And typically by calculating drive end 12 and the corner of driven end 11
Average value determine whole Model angle of attack, this just causes the angle of attack of whole model to have one with the local angle of attack of observation area
Determine deviation, and deviation changes with test load size.And this implementation installs an incidence vane by little groove 6, just
In the in test local angle of attack of measurement pressure measurement section in real time, the measuring accuracy to test model 100 is improve, it is to avoid model
The angle of attack measures introduced error indirectly.
The two ends of the upper surface 1 of the test model 100 are equipped with a leveling groove 7, the bottom surface conduct of the leveling groove 7
Reference planes, for the Levelling operation of incidence vane, the leveling groove 7 is symmetrical at the two ends of the upper surface 1 of test model 100
Set.Wherein, it is provided with lid on two leveling grooves 7.The bottom surface of leveling groove 7 is the machining benchmark of whole test model 100,
The plane is parallel with the string plane of test model 100 in theory, thus using the whole mould of bottom surface leveling of leveling groove 7
Type.Levelling operation is to find the position that Model angle of attack is zero degree.During leveling, an electrolevel is first placed on the leveling
On the bottom surface of groove 7, then according to the reverse angle of attack of Adjustment Tests model 100 of reading of display, until electrolevel reading is
Zero, i.e. Model angle of attack are zero, that is to say, that model is now in horizontality.
The two ends of the lower surface 4 of the test model 100 are equipped with interface, and the interface is articulated with the one end of vertical pole 8,
The other end of the vertical pole 8 is articulated with the rotating mechanism (not shown) of wind-tunnel bottom；Reference picture 1, the test model
100 are respectively arranged at two ends with left auricle 9 and auris dextra piece 10, and the left auricle 9 and auris dextra piece 10 be respectively fixedly connected with whirler
The driven end 11 and drive end 12 of structure.Wherein, driven end 11 and drive end 12 are circular and driven end 11 and drive end 12
Upper circumference is distributed with connecting hole, and driven end 11 and drive end 12 are connected by connecting hole thereon with rotating mechanism, and drive end
12 are connected with motor.
When the model provided in using the present embodiment is tested, by the left auricle 9 and auris dextra piece 10 of test model 100
It is connected on the driven end 11 and drive end 12 of rotating mechanism；The one end of vertical pole 8 is articulated with experimental model again
Interface on surface 4, the other end is articulated with the rotating mechanism of wind-tunnel bottom；Start the drive end 12 of rotating mechanism, drive end
12 are continuously rotated, and motoring ring test model 100 and driven end 11 are rotated together with, and change the angle of attack of model；At the same time, by peace
Infrared camera outside test chamber moves phenomenon to the transition flow of model surface and is observed, records.
Note, general principle of the invention and principal character and advantages of the present invention has been shown and described above.The industry
Technical staff it should be appreciated that the present invention is not limited to the embodiments described above, described in above-mentioned implementation method and specification
Merely illustrate the principles of the invention, without departing from the spirit and scope of the present invention, the present invention also have various change and
Improve, these changes and improvements all fall within the protetion scope of the claimed invention, and claimed scope of the invention is by appended
Claims and its equivalent thereof.
Claims (10)
1. it is a kind of for flow transition detection dual airfoil test model, it is characterised in that in the upper of the test model (100)
A containing groove (2) is provided with the flow transition area of observation coverage on surface (1), whole test model is only filled out in the containing groove (2)
Filled with heat insulation layer.
2. it is according to claim 1 for flow transition detection dual airfoil test model, it is characterised in that experiment mould
The upper surface (1) of type (100), lower surface (4), leading edge and trailing edge are distributed with the pressure tap being connected with electric pressure scaner
(3), and be arranged at 0.65~0.75 times of chord length of upper surface (1), the quantity of the pressure tap (3) of leading edge and trailing edge accounts for pressure measurement
The 60%~70% of hole (3) total quantity.
3. it is according to claim 2 for flow transition detection dual airfoil test model, it is characterised in that the survey
The central axis of hole (3) is pressed on the same oblique section of the test model (100), the oblique section and the test model
(100) tangential angle is 10 °~15 °.
4. it is according to claim 3 for flow transition detection dual airfoil test model, it is characterised in that after described
Pressure tap (3) quantity that the end of edge is set is one.
5. it is according to claim 1 for flow transition detection dual airfoil test model, it is characterised in that experiment mould
Little groove (6) is provided near the pressure measurement section of the lower surface (4) of type (100), angle of attack sensing is installed in the little groove (6)
Device.
6. it is according to claim 5 for flow transition detection dual airfoil test model, it is characterised in that the examination
The lower surface (4) for testing model (100) is provided with big groove (5) for accommodating pressure-measuring pipe, and the pressure tap (3) is by the survey
Pressure pipe is connected with electric pressure scaner；
The little groove (6) is in the big groove (5).
7. according to any described dual airfoil test models detected for flow transition of claim 1-6, it is characterised in that
The depth of the containing groove (2) is 70mm~120mm.
8. it is according to claim 5 for flow transition detection dual airfoil test model, it is characterised in that the examination
The two ends for testing model (100) upper surface (1) are equipped with a leveling groove (7), and the bottom surface of the leveling groove (7) is used as with reference to flat
Face, for the leveling of incidence vane.
9. it is according to claim 8 for flow transition detection dual airfoil test model, it is characterised in that the examination
Lower surface (4) two ends for testing model (100) are equipped with interface, and the interface is articulated with vertical pole (8) one end；
Vertical pole (8) other end is articulated with rotating mechanism.
10. it is according to claim 9 for flow transition detection dual airfoil test model, it is characterised in that it is described
Test model (100) is respectively arranged at two ends with left auricle (9) and auris dextra piece (10), the left auricle (9) and auris dextra piece (10) point
The driven end (11) and drive end (12) of rotating mechanism are not connected to.
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Cited By (9)
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| CN108303233A (en) * | 2017-08-08 | 2018-07-20 | 北京空天技术研究所 | A high-speed aircraft wind tunnel test device and method for device are twisted in a kind of band pressure turn |
| CN108303229A (en) * | 2017-08-08 | 2018-07-20 | 北京空天技术研究所 | A high-speed aircraft inlet characteristic apparatus for evaluating and method for device is twisted in a kind of band pressure turn |
| CN108304603A (en) * | 2017-08-16 | 2018-07-20 | 北京空天技术研究所 | A kind of high-speed aircraft is forced to turn to twist device verification method |
| CN110589020A (en) * | 2019-09-24 | 2019-12-20 | 中国航空工业集团公司沈阳飞机设计研究所 | Structure and method for determining influence of surface quality defects of airplane |
| CN110879128A (en) * | 2019-11-22 | 2020-03-13 | 北京空天技术研究所 | Test model and method for obtaining front edge heat flux density |
| CN111207903A (en) * | 2020-03-02 | 2020-05-29 | 北京空天技术研究所 | Transition measuring method suitable for sub-transonic wind tunnel |
| CN112304563A (en) * | 2020-10-30 | 2021-02-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Wind tunnel test method for researching influence of transition on aerodynamic characteristics of hypersonic aircraft |
| CN112964452A (en) * | 2021-02-08 | 2021-06-15 | 中国科学院力学研究所 | Variable-attack-angle hydrofoil experiment model for cutting free surface |
| CN114185265A (en) * | 2022-02-15 | 2022-03-15 | 中国空气动力研究与发展中心高速空气动力研究所 | Large-scale opening jet flow wind tunnel ultrasonic speed constant total pressure continuous variable Mach number control method |
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| CN108303233A (en) * | 2017-08-08 | 2018-07-20 | 北京空天技术研究所 | A high-speed aircraft wind tunnel test device and method for device are twisted in a kind of band pressure turn |
| CN108303229A (en) * | 2017-08-08 | 2018-07-20 | 北京空天技术研究所 | A high-speed aircraft inlet characteristic apparatus for evaluating and method for device is twisted in a kind of band pressure turn |
| CN108303229B (en) * | 2017-08-08 | 2019-06-07 | 北京空天技术研究所 | A high-speed aircraft inlet characteristic assessment device and method for device is twisted in a kind of band pressure turn |
| CN108303233B (en) * | 2017-08-08 | 2019-06-11 | 北京空天技术研究所 | A high-speed aircraft wind tunnel test device and method for device are twisted in a kind of band pressure turn |
| CN108304603A (en) * | 2017-08-16 | 2018-07-20 | 北京空天技术研究所 | A kind of high-speed aircraft is forced to turn to twist device verification method |
| CN110589020A (en) * | 2019-09-24 | 2019-12-20 | 中国航空工业集团公司沈阳飞机设计研究所 | Structure and method for determining influence of surface quality defects of airplane |
| CN110879128A (en) * | 2019-11-22 | 2020-03-13 | 北京空天技术研究所 | Test model and method for obtaining front edge heat flux density |
| CN111207903A (en) * | 2020-03-02 | 2020-05-29 | 北京空天技术研究所 | Transition measuring method suitable for sub-transonic wind tunnel |
| CN112304563A (en) * | 2020-10-30 | 2021-02-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Wind tunnel test method for researching influence of transition on aerodynamic characteristics of hypersonic aircraft |
| CN112964452A (en) * | 2021-02-08 | 2021-06-15 | 中国科学院力学研究所 | Variable-attack-angle hydrofoil experiment model for cutting free surface |
| CN112964452B (en) * | 2021-02-08 | 2022-10-21 | 中国科学院力学研究所 | Variable-attack-angle hydrofoil experiment model for cutting free surface |
| CN114185265A (en) * | 2022-02-15 | 2022-03-15 | 中国空气动力研究与发展中心高速空气动力研究所 | Large-scale opening jet flow wind tunnel ultrasonic speed constant total pressure continuous variable Mach number control method |
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