CN102607799A - Device for changing Mach number in supersonic velocity wind tunnel model experiment and working method - Google Patents
Device for changing Mach number in supersonic velocity wind tunnel model experiment and working method Download PDFInfo
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Abstract
The invention discloses a device for changing Mach number in a supersonic velocity wind tunnel model experiment and a working method, which relate to the technical field of supersonic velocity wind tunnel experiments. The invention comprises an experiment model and a wind tunnel support element. The wind tunnel support element comprises a support base and a wind tunnel support. The experiment model is fixed on the wind tunnel support through the support base. The wind tunnel support element further comprises a shock wave plate fixed on the wind tunnel support element. The device for changing the Mach number in the supersonic velocity wind tunnel model experiment solves the problem that the model experiment at certain Mach numbers cannot be performed by the supersonic velocity wind tunnel and avoids the problem that parameter interpolation brings errors under two experimental states where the Mach numbers are more than or less than the set Mach number, so that the actual aerodynamic performance of the experiment model at a special Mach number can be obtained and half of the cost for the wind tunnel experiment is reduced.
Description
Technical field
What the present invention relates to is a kind of experimental technique that changes supersonic wind tunnel model experiment Mach number, belongs to supersonic speed air inlet road technical field, aerodynamics experimental technique field.
Background technology
Wind-tunnel is the major equipment that carries out the aerodynamics experiment.So-called wind-tunnel is exactly by the custom-designed a kind of pipeline of requirement necessarily; In this special pipeline; Adopt suitable propulsion system in pipeline, to produce the air-flow that to regulate; Can simulate in the experimental section or the situation of object simulating in atmospheric flow field basically, be used for carrying out various types of aerodynamics experiments.Be fixed on model in the experimental section with pole or support during experiment, when air-flow was blown over model, the pressure parameter on the model outputed on the surveying instrument through sensor or conduit, thereby obtained the experimental data on the model.
Press the size of gas velocity in the wind tunnel experiment section or Mach number, wind-tunnel can be divided into following several types: several types of low-speed wind tunnel (
), subsonic wind tunnel (
), transonic wind tunnel (
), supersonic wind tunnel (
), hypersonic wind tunnel (
), hypervelocity wind tunnels etc.
For supersonic wind tunnel; The gasflow mach number of its experimental section can not be regulated, and changes the gasflow mach number of experimental section if desired, then must change corresponding jet pipe; Each jet pipe is corresponding to a fixing gasflow mach number, like Ma1.5, Ma1.8, Ma2.0, Ma2.5 etc.Because jet pipe profile limited amount; Supersonic wind tunnel can only carry out the experiment of limited several gasflow mach numbers; For the situation in the wind tunnel experiment range of Mach numbers not,, generally obtain through the parameter interpolation under Ma1.5 and two kinds of experimental states of Ma1.8 like Model Design Mach 2 ship Ma1.6.
Therefore; For designing the not situation in the wind tunnel experiment range of Mach numbers of Mach number; Can not obtain the performance parameter of model through wind tunnel experiment, can't check design and numerical simulation results, and can't verify its accuracy through the performance parameter that linear interpolation obtains in design point.
Summary of the invention
The object of the invention provides a kind of experimental technique that changes supersonic wind tunnel model experiment Mach number, and the shock wave that the shock wave plate through a wedge shape produces changes the gasflow mach number before the empirical model.This method can be applicable to the supersonic speed air inlet road, and empirical model can be full machine model, also can be only to comprise forebody/air intake duct at half interior machine model.The present invention can solve the problem that present supersonic wind tunnel can't be known the performance parameter of empirical model under some specific Mach number, can be used for the aerodynamic characteristic research of aircraft or air intake duct scheme, thereby obtains the actual performance under specific Mach number.
The present invention adopts following technical scheme for realizing above-mentioned purpose:
A kind of device that changes supersonic wind tunnel model experiment Mach number; Comprise empirical model and wind-tunnel supporting member; Described wind-tunnel supporting member comprises that base for supporting and wind-tunnel support; Described empirical model is fixed on described wind-tunnel through described base for supporting and supports, and it is characterized in that: also comprise a shock wave plate, this shock wave plate is fixed on the described wind-tunnel supporting member
Empirical model axis of the present invention is parallel with shock wave plate upper surface, and the shock wave plate comes flow path direction to become an angle of attack δ with wind-tunnel, and this angle of attack δ satisfies:
A kind of method of work that changes the device of supersonic wind tunnel model experiment Mach number comprises the steps:
The first step: according to the Mach number of wind tunnel experiment section gasflow mach number and model needs, calculate the angle of shock wave plate and air-flow, make that the gasflow mach number behind the shock wave plate equals the Mach number that model needs;
Second step: through wind-tunnel angle of attack governor motion adjustment shock wave plate angle is the calculated value of the first step;
The 3rd step: carry out wind tunnel experiment by conventional wind tunnel methods, obtain the performance parameter of empirical model.
The present invention adopts technique scheme, compared with prior art has following advantage:
1) utilize the present invention can solve the problem that existing supersonic wind tunnel can not carry out the model experiment under some Mach number.
2) utilize the present invention to experimentize to have avoided through greater than with the error problem that brings less than the following two kinds of experimental states of design Mach number parameter interpolation down, and reduced half wind tunnel experiment expense.
3) shock wave that produces of the shock wave plate of the present invention through a wedge shape changes the Mach number before the empirical model, and is simple in structure, easy to process and cost is lower.
Description of drawings
Fig. 1 adopts shock wave plate of the present invention and the scheme of installation of empirical model in supersonic wind tunnel.
Fig. 2 is the shock wave plate produces oblique shock wave in supersonic flow a schematic three dimensional views.
Fig. 3 is the shock wave plate produces oblique shock wave in supersonic flow a two-dimentional wave system synoptic diagram.
Fig. 4 is the vertical view of shock wave plate.
Among the figure: 1, supersonic wind tunnel incoming flow, 2, the shock wave plate, 3, the oblique shock wave that produces of shock wave plate, 4, empirical model, 5, the oblique shock wave that produces of empirical model head, 6, base for supporting, 7, wind-tunnel supports, 8, the model axis direction.
Embodiment
The present invention will contrast accompanying drawing below and give more fully to explain, given among each figure is an application example of the present invention, only is not confined to said application example and should not be construed to the present invention.Institute is the airplane intake model, does not comprise aircraft rear body and wing to empirical model among the figure, and is suitable too for other supersonic aircraft models such as full machine model, independent air intake duct models.Should use instance to test as airplane intake, for aircraft outflow experiment, the present invention also can implement.
A kind of device that changes supersonic wind tunnel model experiment Mach number is characterized in that comprising shock wave plate 2, empirical model 4, supports the 7 wind-tunnel supporting members that constitute by base for supporting 6 and wind-tunnel; Empirical model 4 is installed on the base for supporting 6, and is rigidly connected through base for supporting 6 and shock wave plate 2, is fixed on wind-tunnel again and supports on 7.
The axis of empirical model 4 is parallel with the upper surface of shock wave plate 2, and shock wave plate 2 becomes certain angle of attack with the direction of supersonic wind tunnel incoming flow 1.
Based on a kind of device that changes supersonic wind tunnel model experiment Mach number of the present invention, comprise the steps:
The first step: according to wind tunnel experiment section gasflow mach number Ma
1Mach number Ma with the model needs
D, calculate the angle δ of shock wave plate and air-flow according to the oblique shock wave relational expression, make gasflow mach number Ma behind the shock wave plate
2Equal the Mach number Ma that model needs
D
Wherein, Mach number Ma behind the oblique shock wave
2Computing formula following:
In the formula, Ma
1Be wind tunnel experiment section gasflow mach number, Ma
2Be the gasflow mach number behind the shock wave plate, β is the shock wave angle, and k is a specific heats of gases ratios, for air k=1.4.
Oblique shock wave wave angle β, flow-deviation angle δ and incoming flow Mach number Ma
1Calculation relational expression do
This formula is the implicit function about shock wave angle β, can look into oblique shock wave table or oblique shock wave figure line, or adopts numerical method directly to ask approximate solution from this formula.
Second step: the angle of attack size through wind-tunnel angle of attack governor motion adjustment shock wave plate and air-flow is the calculated value δ of the first step;
The 3rd step: carry out wind tunnel experiment by conventional wind tunnel methods, obtain the performance parameter of empirical model.
Fig. 1 illustrates the scheme of installation of supersonic speed air inlet road empirical model in wind-tunnel of an employing both sides of the present invention air inlet.Supersonic wind tunnel incoming flow 1 produces oblique shock wave 3 through shock wave plate 2 one at shock wave plate 2 heads, and the supersonic flow behind the oblique shock wave 3 produces forward shock 5 at empirical model 4 heads.Empirical model 4 is installed on the base for supporting 6, and is rigidly connected through base for supporting 6 and shock wave plate 2, is fixed on wind-tunnel again and supports on 7.
Fig. 2 illustrates one and adopts the three-dimensional flow synoptic diagram of shock wave plate of the present invention in supersonic flow.Supersonic wind tunnel incoming flow 1 produces plane oblique shock wave 3 through shock wave plate 2 one at shock wave plate 2 heads.
Fig. 3 illustrates one and adopts the two-dimentional flow schematic diagram of shock wave plate of the present invention in supersonic flow.Supersonic speed incoming flow 1 produces plane oblique shock wave 3 through shock wave plate 2 one at shock wave plate 2 heads, and the angle of shock wave plate 2 upper surfaces and supersonic wind tunnel incoming flow 1 is δ, and the shock wave angle of oblique shock wave 3 is β.
Fig. 4 illustrates one and adopts the vertical view synoptic diagram of shock wave plate of the present invention in supersonic flow.Shock wave plate 2 anterior width are greater than the rear portion width, and side is Mach angle μ with coming the angle of flow path direction, and the computing formula of μ does
The present invention can change the experiment Mach number of model in existing supersonic wind tunnel.Can be used for the experiment of supersonic speed air inlet road, also can be used for the aircraft experiment.Stop up than satisfy wind tunnel experiment requires except that the area that need to guarantee whole experiment device during experiment; Be also noted that the oblique shock wave 3 that shock wave plate 2 heads produce can not interfere with empirical model 4, and the reflection wave of empirical model 4 forward shocks does not influence experiment measuring.
The foregoing description just is used for explanation of the present invention, and can not be as limitation of the present invention.Therefore the embodiment that mentality of designing every and of the present invention is identical is all in protection scope of the present invention.
Claims (5)
1. device that changes supersonic wind tunnel model experiment Mach number; Comprise empirical model (4) and wind-tunnel supporting member; Described wind-tunnel supporting member comprises that base for supporting (6) and wind-tunnel support (7); Described empirical model (4) is fixed on described wind-tunnel through described base for supporting (6) and supports on (7), and it is characterized in that: also comprise a shock wave plate (2), this shock wave plate (2) is fixed on the described wind-tunnel supporting member.
2. a kind of device that changes supersonic wind tunnel model experiment Mach number according to claim 1 is characterized in that: following relational expression is satisfied at the shock wave angle of said shock wave plate:
In the formula, Ma
1Be wind tunnel experiment section gasflow mach number, Ma
2Be the gasflow mach number behind the shock wave plate, β is the shock wave angle, and k is a specific heats of gases ratios.
3. a kind of device that changes supersonic wind tunnel model experiment Mach number according to claim 1 and 2 is characterized in that: the axis of empirical model (4) is parallel with the upper surface of shock wave plate (2).
4. a kind of device that changes supersonic wind tunnel model experiment Mach number according to claim 3 is characterized in that shock wave plate (2) and wind-tunnel come flow path direction to become an angle of attack δ, and this angle of attack δ satisfies:
5. based on the described a kind of method of work that changes the device of supersonic wind tunnel model experiment Mach number of claim 1, it is characterized in that comprising the steps:
The first step:,, make that the gasflow mach number behind the shock wave plate (2) equals the Mach number that model needs according to the angle δ of oblique shock wave relational expression calculating shock wave plate (2) and air-flow according to the Mach number of wind tunnel experiment section gasflow mach number (1) and model needs;
Second step: the angle of attack size through wind-tunnel angle of attack governor motion adjustment shock wave plate (2) and air-flow is the calculated value δ of the first step;
The 3rd step: carry out wind tunnel experiment by conventional wind tunnel methods, obtain the performance parameter of model.
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| CN102788677A (en) * | 2012-09-03 | 2012-11-21 | 中国科学院力学研究所 | Hypersonic mach-number-variable wind tunnel nozzle |
| CN103134657A (en) * | 2012-12-19 | 2013-06-05 | 中国空气动力研究与发展中心高速空气动力研究所 | Rear space flow field optimizing method for flexible wall spray pipe supersonic velocity first diamond area |
| CN104132811A (en) * | 2014-05-04 | 2014-11-05 | 中国航天空气动力技术研究院 | Ramjet air inlet starting hysteresis characteristics test device |
| CN104483093A (en) * | 2014-12-15 | 2015-04-01 | 中国燃气涡轮研究院 | Variable mach number transonic rigid free jet nozzle |
| CN105628086A (en) * | 2014-10-29 | 2016-06-01 | 北京临近空间飞行器系统工程研究所 | Supersonic speed flight inflow parameter solving method based on conical surface pressure distribution |
| CN106679932A (en) * | 2017-01-23 | 2017-05-17 | 厦门大学 | Attack angle measurement method based on micro thermal film sensor array |
| CN107806977A (en) * | 2017-11-29 | 2018-03-16 | 中国航空工业集团公司沈阳空气动力研究所 | A kind of high enthalpy impulse wind tunnel pipe structure of the wide Mach number of combined type |
| CN108007667A (en) * | 2017-11-20 | 2018-05-08 | 北京航天长征飞行器研究所 | A kind of high-temperature fuel gas wind-tunnel Mach number measuring device and method |
| CN108168831A (en) * | 2017-12-15 | 2018-06-15 | 中国航空工业集团公司沈阳空气动力研究所 | A kind of continuous change Mach number experiment supersonic wind tunnel |
| CN108195553A (en) * | 2016-12-08 | 2018-06-22 | 中国航空工业集团公司沈阳空气动力研究所 | A kind of supersonic aircraft sonic boom token test measuring device |
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| CN109632867A (en) * | 2018-12-28 | 2019-04-16 | 中国航天空气动力技术研究院 | It is a kind of for examine the hypersonic Burning corrosion resistance of material can pilot system and method |
| CN110361156A (en) * | 2019-08-22 | 2019-10-22 | 湖北三江航天红阳机电有限公司 | A kind of test chamber inner core that Mach number is continuously adjustable |
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