CN109815549A - A kind of design method of single pair hypersonic flow to vortex generating device - Google Patents
A kind of design method of single pair hypersonic flow to vortex generating device Download PDFInfo
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Abstract
The invention discloses single pair hypersonic flows to the design method of vortex generating device.Wherein specify upper wall surface arc radius, vortex segment occurred is divided into the infinitesimal section of the deflection angles such as multiple, by solving supersonic speed circular arc Flow Field outside, since changeover portion lower wall surface terminal, downstream gradually solve streamline equation, initial lower wall surface molded line is obtained, local boundary layer displacement thickness is then superimposed to initial lower wall surface molded line, obtains final lower wall surface type face；Later, the equivalent angle of flare is determined, lower wall surface molded line on design experiment window.The design method utilizes supersonic speed circular arc Flow Field outside linear flow structure vortex segment occurred, avoids generating unnecessary background wave system in vortex segment occurred, it is only necessary to which changing vortex segment occurred radius of curvature can be realized to hypersonic flow to vortex intensity, the control of scale.A kind of practicable experimental rig design method is provided to the characteristic of vortex and its with the interference mechanism experimental study of downstream shock train for hypersonic flow in development scramjet engine.
Description
Technical field
The present invention relates to scramjet engine aerodynamic experiment field, especially a kind of design side for generating supersonic speed vortex
Method.
Background technique
Hypersonic flow is a kind of universal flow phenomenon to vortex, and especially in High Mach number air intake duct, flowing to whirlpool is
Its internal main flow characteristics.The separation stream that generates main and Shock/Boundary-Layer relevant induced synthesis of the hypersonic flow to whirlpool
Dynamic correlation, due to the needs of actual engineering design, Shock/Boundary-Layer interference is universal relatively strong in air intake duct, therefore this supersonic speed
Whirlpool is flowed to as accessory to be difficult to avoid.The presence for flowing to whirlpool, like a double-edged sword can be by stream for burning
Reinforce the blending of air-flow and fuel oil to whirlpool, improves efficiency of combustion；However, the presence for flowing to whirlpool may be more for air intake duct
More is " more harm than good ", and it is all low stagnation pressure air that this, which is primarily due to flow to assemble in whirlpool, this partial air is to inverse pressure ladder
It spends most sensitive, is the short slab of entire air intake duct cross section anti-reflective pressure energy power, once the high back pressure that fired downstream is formed is applied to
After these vortexs, it is likely that promote shock train to overflow in air intake duct pipeline, induce air intake duct Unstart phenomena, harm flight peace
Entirely.Therefore, the hypersonic flow studied in pipeline is most important to whirlpool.
For necessity of research, how to simulate in test hypersonic flow to whirlpool be the basic item for carrying out correlative study
Part.However, existing theoretical, test method is all based on vortex generator, this method can efficiently generate really
Whirlpool is flowed to, but introduces the oblique shock wave of vortex generator slope generation and the expansion of tail portion again while generating and flowing to whirlpool
Fan, the presence of these oblique shock waves and expansion fan can form reflection in the duct, when it is incident on hypersonic flow to whirlpool, convection current
The characteristics such as development, evolution, rupture to whirlpool itself cause irreversible influence, so if needing ultrasound in simple research pipeline
The coupled interference problem that the flow behavior and hypersonic flow in flow speed and direction whirlpool are flowed to whirlpool and downstream shock train, is occurred using vortex
Device scheme is inappropriate.
For this purpose, how by Pneumatic design method to design the hypersonic flow in pipeline to whirlpool, and avoid other oblique shock waves,
The interference of the backgrounds wave system such as dilatational wave is current critical issue urgently to be solved.
Summary of the invention
To solve the above problems, the present invention provides a kind of single pair hypersonic flows of no background wave system to vortex generating device
Design method, the fully developed single pair hypersonic flow of no background wave system can be generated to vortex, have good versatility.
In order to achieve the above object, The technical solution adopted by the invention is as follows:
A kind of single pair hypersonic flow is to the design method of vortex generating device, and the single pair hypersonic flow is to vortex generating device
Including sequentially connected Laval nozzle, changeover portion, vortex segment occurred, test window；Wherein, the design of vortex segment occurred includes
Following steps:
(1) passes through Prandtl-Mei Ye relationship according to Laval nozzle exit Mach number and test window Mach number demand
The total deflection angle theta of vortex segment occurred is calculated in formula；
(2) total deflection angle theta of vortex segment occurred is equally divided into n parts points of deflection angle Δ θ by, and every part is divided deflection angle Δ θ corresponding
One circular arc infinitesimal section；
(3) sets vortex segment occurred upper wall curvature radius as R；
(4) is with first of right lateral dilatational wave L in vortex segment occurred1Starting point A1Rectangular coordinate system is established for origin, the right angle
Two reference axis of coordinate system are x-axis and y-axis, using vortex segment occurred exit flow direction as x-axis forward direction；
(5) first of right lateral dilatational wave L is calculated according to Laval nozzle exit Mach number in1Linear equation, it is right
Row dilatational wave L1Starting point B is bent with changeover portion lower wall surface intersection point, that is, vortex segment occurred lower wall surface1；
(6) it is swollen according to Laval nozzle exit Mach number, total static temperature relational expression and energy equation to solve first of right lateral by
Swollen wave L1Flow field parameter afterwards；
(7) solves second right lateral dilatational wave L according to arc radius R and the geometrical relationship for dividing deflection angle Δ θ2
Starting point A2Coordinate；And the n-th right lateral dilatational wave L is solved according to the relationship2Starting point AnCoordinate；
(8) L that has been acquired according to (6) step1Flow field parameter afterwards solves second right lateral dilatational wave L2Linear equation；
(9) L that has been acquired according to (6) step1Flow field parameter afterwards solves streamline m2Linear equation；
(10) simultaneous L2、m2Linear equation obtains second right lateral dilatational wave L2With changeover portion lower wall surface intersection points B2；
(11) repeats (6)-(10) step, solves B respectively3, B4…BnCoordinate；Wherein BnIt is respectively the n-th right lateral
The intersection point of dilatational wave and changeover portion lower wall surface；
(12) flow field parameter of the based on solution calculates local boundary layer displacement thickness, amplifies A1~AnAnd B1~BnInstitute's shape
At the area of runner.
Single pair hypersonic flow provided by the invention is to the design method key of vortex generating device: being sprayed according to Lavalle
The demand of pipe exit Mach number and test window design Mach number, specifies upper wall surface arc radius, vortex segment occurred is divided into
The infinitesimal section of the deflection angles such as multiple, by solve supersonic speed circular arc Flow Field outside, since changeover portion lower wall surface terminal, downstream by
Step solves streamline equation, obtains lower wall surface molded line discrete point, constitutes lower wall surface by these discrete points；Later, it is based on vortex
Section outlet border layer displacement thickness, calculates the equivalent angle of flare, lower wall surface molded line on design experiment window.
Compared with the existing technology, the present invention can have the advantages that
The design method dexterously utilizes supersonic speed circular arc Flow Field outside linear flow structure vortex segment occurred, can effectively avoid vortex
Unnecessary background wave system is generated in segment occurred, and design method is easy, is easy to program realization, it is only necessary to change vortex segment occurred
Radius of curvature can be realized to hypersonic flow to vortex intensity, the control of scale.To carry out shock train in scramjet engine
Interference mechanism experimental study with from hypersonic flow to vortex provides a kind of practicable experimental rig design method.
Detailed description of the invention
Fig. 1 is the single pair hypersonic flow of no background wave system to vortex generating device flow field parameter and specification schematic diagram.
Fig. 2 is vortex intensity and vortex segment occurred upper wall curvature radius mapping relations schematic diagram, and wherein Q is that vortex is strong
Degree.
Fig. 3 is type face modification method schematic diagram, wherein CiPlace curve is to test the revised molded line of window.
Fig. 4 is the vortex segment occurred and test window molded line using the method for the present invention design.
Fig. 5 is the single pair hypersonic flow of no background wave system to vortex generating device vortex segment occurred design flow diagram.
Specific embodiment
The present invention discloses a kind of single pair hypersonic flow of no background wave system to the design method of vortex generating device.It please refers to
Shown in Fig. 1, the detailed implementation steps for designing the embodiment to the method for the present invention below are described.
(1) is based on the existing vacuum source in laboratory, volume 400m3, 3 vacuum pumps.According to the suction of vacuum pump
Flow is based on flow equation, considers test period, determines Laval nozzle exit Mach number M0It is 2, throat area is
23703mm2.According to Laval nozzle venturi and rate of discharge relation of equality, it may be determined that Laval nozzle exports high 200mm, wide
200mm；
(2) molded line of the from Laval nozzle throatpiston to outlet is designed using the method for characteristic curves, and sign collimation method design can
With reference to known existing technical literature Maurice J.Zucrow, Joe D.Hoffman, Gas dynamic, Volume II, pp
83；
(3) for lower wall surface there are micro- expansion, the angle of flare is 0.5 ° on changeover portion；
(4) the design detailed process of vortex segment occurred is as follows:
1) setting test window Mach number Mn=3, pass through Prandtl-Mei Ye relational expression:
θ=ν (M0)-ν(Mn)
Obtain the total deflection angle θ of vortex segment occurred.In above formula, γ is specific heat ratio, takes 1.4, Ma for expansion air γ
Mach number after wave wave；
2) the total deflection angle θ of vortex segment occurred is equally divided into 23 parts by, and every part of deflection angle Δ θ is approximately equal to 1 °；
3) it is R=1m that, which gives vortex segment occurred upper wall curvature radius,；
4) is with first of right lateral dilatational wave L in vortex segment occurred1Starting point A1Rectangular coordinate system is established for origin, is sent out with vortex
Raw 3 exit flow direction 7 of section is that x-axis is positive；
5) first of right lateral dilatational wave L is calculated according to Laval nozzle exit Mach number in1Linear equation L1: y
=-tan (α1+ θ) × x, wherein α1Calculation formula for the Mach angle of first of right lateral dilatational wave, Mach angle refers to following formula.L1With
6 intersection points of vortex segment occurred lower wall surface molded line, that is, vortex segment occurred lower wall surface is bent starting point B1, in point B1Before, under vortex segment occurred
Wall surface molded line 6 is consistent with 2 wall surface molded line of changeover portion:
In above formula, αi、MiMach angle and Mach number after respectively the i-th right lateral dilatational wave wave expand first of right lateral
Wave i=1；
6) gives total temperature T*=300K calculates the i-th right lateral dilatational wave LiFlow field parameter afterwards successively solves T, V, most
After solve right lateral dilatational wave LiVelocity component V afterwardsxi、Vyi, calculating process is as follows:
Vxi=Vcos Δ θ；
Vyi=Vsin Δ θ；
Wherein, T*, T be respectively total temperature, static temperature after the i-th right lateral dilatational wave wave, Vxi、VyiRespectively the i-th right lateral is swollen
Velocity component of the velocity vector V in the direction x and y in flow field after swollen wave wave；RgFor constant, it is taken as 287；
7) sets upper wall millet cake AiCoordinate is (xi, yi), lower wall millet cake BiCoordinate is (ki, ji), center of circle O coordinate is (xo=x1+
Rsinθ,yo=y1+ Rcos θ), i+1 road right lateral dilatational wave is solved according to the geometrical relationship of arc radius R and deflection angle Δ θ and is risen
Point Ai+1(xi+1,yi+1) coordinate；
xi+1=xO-Rsin(θ-iΔθ)；
yi+1=yO-Rcos(θ-iΔθ)；
8) is according to the velocity component V in the direction x, y after the i-th right lateral dilatational wave acquiredxi、VyiSolve straight line mi+1
Slope gi；
According to right lateral dilatational wave Mach angle αiStraight line L is solved with the relationship of deflection angle Δ θi+1Slope hi, it may be assumed that
hi=-tan (αi+1+θ-iΔθ)；
9) is according to right lateral dilatational wave terminal Bi(ki,ji) coordinate and straight line mi+1Slope giProvide straight line mi+1Equation,
According to right lateral dilatational wave starting point Ai+1(xi+1,yi+1) coordinate and straight line Li+1Slope hiProvide straight line Li+1Equation:
10) simultaneous straight line Li+1, straight line mi+1Equation seek its intersection points Bi+1Coordinate；
11) repeats above procedure, solves B respectively2, B3, B4…BnCoordinate；
12) is by required point BiIt is sequentially connected as vortex segment occurred lower wall surface molded line；
13) refers to Fig. 3, and the flow field parameter based on solution can obtain Bi-1And BiPoint coordinate crosses the vertical line that two o'clock makees its tangent line,
Local boundary layer displacement thickness δ is calculated, in Bi-1And BiWith B on the vertical line of point tangent linei-1、BiThe point for taking length to be equal to δ for starting point
Ci-1、Ci, should be noted at this time in Bi-1、BiThere is two o'clock respectively up and down, the point for expanding circulation passage should be taken to amplify A1-AnAnd B1-
BnThe area of formed runner, specific implementation method are to seek straight line Ci-1CiIntercept, take intercept lesser straight in channel wall face
Line.Wall surface takes the biggish straight line of intercept on a passage.Local boundary layer displacement thickness δ calculation is as follows:
Wherein, RexFor local Reynolds number；
(5) is referring to Fig. 4, devise vortex segment occurred and test Window-type face using the above method, in the amendment of type face
When, lower wall surface, which is tested on window, compared to vortex segment occurred exports upper 0.5 ° of the outside deviation of lower wall surface.Above-mentioned vortex segment occurred is set
The flow chart of meter method is as shown in Figure 5.
In addition, there are many concrete methods of realizing and approach of the invention, the above is only a preferred embodiment of the present invention.
It should be pointed out that for those skilled in the art, without departing from the principle of the present invention, can also do
Several improvements and modifications out, these modifications and embellishments should also be considered as the scope of protection of the present invention.
Claims (6)
1. a kind of single pair hypersonic flow is to the design method of vortex generating device, it is characterised in that: the single pair hypersonic flow is to rotation
Whirlpool generating device includes sequentially connected Laval nozzle (1), changeover portion (2), vortex segment occurred (3), test window；Wherein,
The design of vortex segment occurred (3) the following steps are included:
(1) passes through Prandtl-Mei Ye relational expression meter according to Laval nozzle exit Mach number and test window Mach number demand
Calculation obtains vortex segment occurred (3) total deflection angle theta；
(2) total deflection angle theta of vortex segment occurred (3) is equally divided into n parts points of deflection angle Δ θ by, and every part is divided deflection angle Δ θ corresponding
One circular arc infinitesimal section；
(3) sets vortex segment occurred upper wall surface (5) radius of curvature as R；
(4) is with first of right lateral dilatational wave L in vortex segment occurred1Starting point A1Rectangular coordinate system is established for origin, the rectangular co-ordinate
Two reference axis of system are x-axis and y-axis, are that x-axis is positive with vortex segment occurred (3) exit flow direction (7)；
(5) first of right lateral dilatational wave L is calculated according to Laval nozzle exit Mach number in1Linear equation, right lateral expansion
Wave L1Starting point B is bent with changeover portion (2) lower wall surface intersection point, that is, vortex segment occurred lower wall surface1；
(6) solves first of right lateral dilatational wave according to Laval nozzle exit Mach number, total static temperature relational expression and energy equation
L1Flow field parameter afterwards；
(7) solves second right lateral dilatational wave L according to arc radius R and the geometrical relationship for dividing deflection angle Δ θ2Starting point A2Coordinate；
And the n-th right lateral dilatational wave L is solved according to the relationship2Starting point AnCoordinate；
(8) L that has been acquired according to (6) step1Flow field parameter afterwards solves second right lateral dilatational wave L2Linear equation；
(9) L that has been acquired according to (6) step1Flow field parameter afterwards solves streamline m2Linear equation；
(10) simultaneous L2、m2Linear equation obtains second right lateral dilatational wave L2With changeover portion (2) lower wall surface intersection points B2；
(11) repeats (6)-(10) step, solves B respectively3, B4…BnCoordinate；Wherein BnIt is respectively the n-th right lateral dilatational wave
With the intersection point of changeover portion (2) lower wall surface；
(12) flow field parameter of the based on solution calculates local boundary layer displacement thickness, amplifies A1~AnAnd B1~BnFormed runner
Area.
2. design method according to claim 1, it is characterised in that: the geometric configuration of Laval nozzle is two-dimensional nozzle,
The throat area that Laval nozzle is determined according to laboratory gas source ability, according to test it needs to be determined that design point exports Mach
Number.
3. design method according to claim 1, it is characterised in that: Laval nozzle is from throatpiston to outlet
Molded line is designed using the method for characteristic curves.
4. design method according to claim 1, it is characterised in that: the width and Laval nozzle of changeover portion keep one
It causes, for lower wall surface there are micro- expansion, the angle of flare is 0~1 ° on changeover portion.
5. design method according to claim 1, it is characterised in that: lower wall surface is sent out compared to vortex in test window (4)
Raw section exports the upper outside deviation of lower wall surface, forms micro- expansion pipe, and unilateral angle of flare Δ β is 0.5 °~1.0 °.
6. design method according to claim 1, it is characterised in that: set up examination by emulation mode or test method
The mapping relations for testing window vortex intensity and vortex segment occurred upper wall surface (5) radius of curvature R, it is suitable to be matched according to mapping relations
Upper wall surface (5) radius of curvature R occurs for vortex.
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| CN112035952A (en) * | 2020-08-21 | 2020-12-04 | 南京航空航天大学 | Design method of ejector nozzle experimental device for simulating outflow of aircraft |
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| CN111159814A (en) * | 2019-12-19 | 2020-05-15 | 中国航天空气动力技术研究院 | Design method and configuration of rectangular supersonic velocity spray pipe with turning inlet and high slenderness ratio |
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