CN102323961A - Asymmetric supersonic velocity spray pipe and design method thereof - Google Patents

Abstract

The invention provides an asymmetric supersonic velocity spray pipe and a design method thereof. The design method of the supersonic velocity spray pipe comprises the following steps: according to an inlet coordinate and a throat area, determining a subsonic velocity segment curve; according to the subsonic velocity segment curve and a throat radius of curvature, determining a characteristic line in a transonic velocity area; according to one end point of the characteristic line and a corresponding supersonic velocity area outlet end point, determining a first wall surface curve; and according to an initial characteristic line and the Mach number distribution of the first wall surface curve, determining a second wall surface curve by mass conservation with a characteristic line method. With the design method of the asymmetric supersonic velocity spray pipe, concentrated compression wave in the flow field of the asymmetric supersonic velocity spray pipe can be reduced or eliminated, the quality of the flow field of the spray pipe is greatly improved, equipment performance is improved, and research funds can be saved. According to the asymmetric supersonic velocity spray pipe disclosed by the invention, the design method of the asymmetric supersonic velocity spray pipe is adopted.

Description

Asymmetric supersonic nozzle and method for designing thereof

Technical field

The present invention relates to field of fluid power, in particular to a kind of asymmetric supersonic nozzle and method for designing thereof.

Background technology

Supersonic nozzle is widely used in the equipment such as high-speed aircraft, rocket, supersonic wind tunnel, high-energy laser, injection vacuum pump, and the performance of nozzle flow field article confrontation equipment has significant effects.Obtain suitable jet pipe wall surface curve through the certain designed technology, can improve the nozzle flow field quality greatly, improve equipment performance, save reasearch funds.Supersonic nozzle generally is made up of contraction section and expansion segment; Under certain pressure drove, gas quickened at contraction section gradually, and near throat, reaches the velocity of sound; Continue to quicken at expansion segment then, until going out the supersonic flow that needed Mach number of interruption-forming and flow direction angle distribute.

Offset nozzle is a kind of special supersonic speed/hypersonic nozzle, and two walls that are characterized in shrinkage expansion are asymmetrical, the air-flow acceleration of in the pipeline of bending, expanding, and the entrance and exit direction has certain included angle.This jet pipe has good geometric compliance, can be widely used in the generation and the control in supersonic speed in the restricted clearance/hypersonic flow field.

Though existing multiple at present comparatively ripe two dimension or axisymmetric nozzle method for designing can be obtained higher exit flow field quality, because the singularity of offset nozzle, existing method can not directly expand in the design of asymmetric surface curve.Its subject matter that faces comprises the following aspects: the one, and current of spring district hypothesis is no longer set up; The 2nd, lack ripe asymmetric wave absorption technology, can't guarantee flow field quality; Three to be based on the linear boundary layer modification method of revising no longer suitable.

Monograph " supersonic speed free-vortex aerodynamic window's and optical quality thereof " (Yi Shi and etc.; Publishing house of the National University of Defense technology; 2005) the free-vortex jet pipe method for designing (a kind of typical offset nozzle) found the solution based on the subregion of shortest length jet pipe (MLN) has been proposed, this method design process is following:

1. adopt the MLN method for designing to obtain a symmetrical jet pipe profile curve.

2. extract the flow parameter on MLN nozzle exit border, as asymmetric section inlet boundary condition.

3. be divided into several zones such as equal uniform flow district, simple wave district and non-simple wave district with asymmetric section.

4. utilize free-vortex relational expression, Prandtl-mayer's relation and mass conservation relation to confirm the wall curve in asymmetric district.

Adopt this method can access needed free-vortex jet pipe profile curve, numerical value checking result shows that the jet pipe profile that is designed can generate needed free-vortex flow field basically.

In some equipment less demanding to work efficiency, offset nozzle can design by the use experience curve.Some punching engine jet pipe also belongs to offset nozzle, and its target is to realize maximum thrust, and is indifferent to the details of flow field structure.Along with development of modern computer technology, also can be applied in the offset nozzle design based on the aerodynamic configuration designing technique of CFD optimized Algorithm.

The method of finding the solution based on the MLN subregion lacks adaptability, can only be used to design the free-vortex jet pipe at present.This type jet pipe will comprise the rhomboid of MLN, and the jet pipe length of being designed is long.Asymmetric wall is not done the viscosity correction, be difficult to accurately assess the flow field structure variation that the adsorption plane effect layer brings.MLN expands too fast, and precision own is not high, is not suitable for designing high Mach number jet pipe.

Based on the offset nozzle method for designing of empirical curve, parameters such as nozzle exit Mach number and flow direction angle can't design as requested.

Based on the offset nozzle designing technique of CFD optimized Algorithm, calculated amount is very big, and design efficiency is lower, and precision does not obtain too big improvement.

Summary of the invention

The present invention aims to provide a kind of asymmetric supersonic nozzle and method for designing thereof, can reduce or eliminate the concentrated wave of compression in the flow field of asymmetric supersonic nozzle, improves the nozzle flow field quality greatly, improves equipment performance, saves reasearch funds.

To achieve these goals, according to an aspect of the present invention, a kind of asymmetric supersonic nozzle method for designing is provided, has it is characterized in that, having comprised: confirmed subsonic speed section curve according to inlet coordinate and throat opening area; Confirm the transonic speed interior initial characteristics line in district according to subsonic speed section curve and throat's radius-of-curvature; An end points according to the initial characteristics line is confirmed the first side wall surface curve with corresponding supersonic region outlet end points; Mach Number Distribution according to initial characteristics line and the first side wall surface curve is confirmed the second side wall surface curve through the mass conservation and method of characteristic.

Further, the method for characteristic iterative formula is:

Wherein, x is a horizontal ordinate, and r is an ordinate, and θ is local flow direction angle, and M is local Mach number and M＞1, and δ is the pattern of flow parameter, for two dimension δ=0 of flowing, and rotational symmetry δ=1 of flowing, r ≠ 0, γ is the specific heat at constant pressure of gas and the specific heat ratio of specific heat at constant volume.

Further, after confirming the second side wall surface curve, also comprise: adopt the reference temperature solution of momentum integral relational expression to find the solution boundary layer displacement thickness, carry out the boundary layer and revise, obtain the wall curve of actual use, the momentum integral relational expression is:

$\frac{\mathrm{d\&tau;}}{\mathrm{dx}}+\mathrm{\&tau;}[\frac{2-{\mathrm{<figure-callout\; id="2"\; label="M"\; filenames="HDA0000062089930000041.png"\; state="\{\{state\}\}">M</figure-callout>}}^2+H}{M(1+\frac{\mathrm{\&gamma;}-1}{2}{M}^2)}\frac{\mathrm{dM}}{\mathrm{dx}}+\frac{1}{y}\frac{\mathrm{dy}}{\mathrm{dx}}]=\frac{{\mathrm{<figure-callout\; id="2"\; label="C\; f"\; filenames="HDA0000062089930000041.png"\; state="\{\{state\}\}">C</figure-callout>}}_f}{2}\sec \mathrm{\&phi;}$

$\mathrm{\&phi;}={\tan }^{-1}\left(\frac{\mathrm{dy}}{\mathrm{dx}}\right)$

$H=\frac{{\mathrm{\&delta;}}^{*}}{\mathrm{\&tau;}}$

Wherein, τ is a momentum loss thickness, δ ^*Be boundary layer displacement thickness, φ is the flow direction angle, and H is the boundary layer form factor, C _fFor pressing friction factor.

Further, confirming also to comprise before the subsonic speed section curve based on inlet coordinate and throat opening area: confirm said throat opening area based on the isentropic relation formula, the isentropic relation formula is:

$\frac{A}{A^{*}}=\frac{1}{M_t}{\left[\left(\frac{2}{\mathrm{\&gamma;}+1}\right)(1+\frac{\mathrm{\&gamma;}-1}{2}{M}_t^2)\right]}^{(\mathrm{\&gamma;}+1)/2(\mathrm{\&gamma;}-1)}$

Wherein, A is the nozzle exit area, A ^*Be nozzle throat area, M _tBe the nozzle throat Mach number, γ is the specific heat ratio of gas.

Further, confirming that according to the isentropic relation formula said throat opening area also comprises before: the boundary curve that requires to confirm the entrance and exit position according to the jet pipe geometry.

Further, the two circular fitting methods of subsonic speed section curve negotiating are confirmed.

Further, characteristic curve confirms that according to the Sauer method the corresponding transonic speed initial value line equation of Sauer method is provided by following formula:

$x=-\frac{(\mathrm{\&gamma;}+1)\mathrm{\&alpha;}}{8\mathrm{<figure-callout\; id="2"\; label="L\; y"\; filenames="HDA0000062089930000041.png"\; state="\{\{state\}\}">L</figure-callout>}}{y}^{}{}^2$

Wherein, $\mathrm{\&alpha;}={\left[\frac{2}{(\mathrm{\&gamma;}+1)r_t{\mathrm{\&rho;}}_t}\right]}^{1/2}L$

L＝ρ _t+r _t

ρ _tAnd r _tBe respectively the radius-of-curvature and half height of nozzle throat.

Further, the first side wall surface curve confirms with corresponding supersonic region outlet end points through an end points that utilizes the quadratic spline curve to connect the initial characteristics line, and through parameter adjustment the first side wall surface curve to the shape that satisfies Structural Design Requirement.

Further; An end points according to the initial characteristics line is confirmed also to comprise after the first side wall surface curve with corresponding supersonic region outlet end points: according to the nozzle structure designing requirement; Confirm the Mach Number Distribution of the first side wall surface curve through cubic spline curve, continuous with the second derivative that guarantees Mach number.

According to a further aspect in the invention; A kind of asymmetric supersonic nozzle is provided; Comprise the asymmetric supersonic nozzle wall that the first side wall surface curve and the second side wall surface curve form; An end points of the initial characteristics line in the transonic speed district that the first side wall surface curve is confirmed according to subsonic speed section curve and throat's radius-of-curvature confirms that with corresponding supersonic region outlet end points the second side wall surface curve is definite through the mass conservation and method of characteristic according to the Mach Number Distribution of said initial characteristics line and the first side wall surface curve.

Further, asymmetric supersonic nozzle wall has and adopts the reference temperature solution of momentum integral relational expression to find the solution the determined layer that revises the boundary of boundary layer displacement thickness.

Further, asymmetric supersonic nozzle wall comprises according to inlet coordinate and the definite formed subsonic speed section of the subsonic speed section curve wall of throat opening area.

According to technical scheme of the present invention; The supersonic nozzle wall of asymmetric supersonic nozzle is confirmed according to the mass conservation and method of characteristic; Can pass through the method for characteristic wave absorption, making does not have the wave of compression of concentrating in the flow field, and flow field parameter distributes and can be optimized according to practical application; Can effectively guarantee the homogeneity at flow field Mach number and flow direction angle, improve flow field quality greatly.Adopt the momentum integral relational expression to calculate the boundary layer displacement thickness of supersonic nozzle; Confirm the boundary layer curve; Can there be the boundary layer near having overcome the wall curve that causes owing to gas viscosity; Thereby influence the problem of flow field quality, further improved the precision and the quality of asymmetric supersonic nozzle.Through quadratic spline curve adjustment the first side wall surface curve, make the first side wall surface curve to adjust according to the Mach Number Distribution of design, can control the Mach Number Distribution of the supersonic region of jet pipe more accurately, make it satisfy design demand.

Description of drawings

The accompanying drawing that constitutes a part of the present invention is used to provide further understanding of the present invention, and illustrative examples of the present invention and explanation thereof are used to explain the present invention, does not constitute improper qualification of the present invention.In the accompanying drawings:

Fig. 1 shows first design procedure according to asymmetric supersonic nozzle of the present invention, confirms the configuration line segment synoptic diagram of entrance and exit position;

Fig. 2 shows second design procedure according to asymmetric supersonic nozzle of the present invention, confirms subsonic speed section curve configuration synoptic diagram;

Fig. 3 shows the 3rd design procedure according to asymmetric supersonic nozzle of the present invention, confirms initial characteristics line configuration synoptic diagram;

Fig. 4 shows the 4th design procedure according to asymmetric supersonic nozzle of the present invention, confirms the first side wall surface curve configuration synoptic diagram;

Fig. 5 shows the 5th design procedure according to asymmetric supersonic nozzle of the present invention, confirms the second side wall surface curve configuration synoptic diagram;

Fig. 6 shows the 6th design procedure according to asymmetric supersonic nozzle of the present invention, confirms revised actual wall surface curve configuration synoptic diagram; And

Fig. 7 shows the structural representation according to asymmetric supersonic nozzle of the present invention.

Embodiment

Hereinafter will and combine embodiment to specify the present invention with reference to accompanying drawing.Need to prove that under the situation of not conflicting, embodiment and the characteristic among the embodiment among the application can make up each other.

In the present invention; The subsonic speed section is meant that air-flow gets into after the jet pipe contraction section its flow velocity less than the part of the velocity of sound, and the Asia is also i.e. section transonic speed of section transonic speed, is meant that air-flow gets into the boundary part between the supersonic speed state from the subsonic speed state; Usually; The Mach number of subsonic speed section is less than 0.8, and the Mach number of inferior transonic speed section is between 0.8 to 1.2, and the Mach number of supersonic speed section is greater than 1.2.

As shown in Figure 1, according to asymmetric supersonic nozzle method for designing of the present invention, at first need confirm inlet (A according to the geometry designing requirement of jet pipe ₁, A ₂) the configuration line segment A of position ₁A ₂And outlet (A ₅, A ₆) the configuration line segment A of position ₅A ₆, confirm the jet pipe flow according to outlet density, speed and area then; Nozzle flowmeter is calculated formula:

$\stackrel{\&CenterDot;}{m}=\mathrm{\&rho;VA}$

Wherein: is the jet pipe flow; ρ is a nozzle exit density; V is a nozzle velocity, and A is the nozzle exit area.

Confirm the area of throat's anchor ring afterwards according to the isentropic relation formula; The isentropic relation formula is:

$\frac{A}{A^{*}}=\frac{1}{M_t}{\left[\left(\frac{2}{\mathrm{\&gamma;}+1}\right)(1+\frac{\mathrm{\&gamma;}-1}{2}{M}_t^2)\right]}^{(\mathrm{\&gamma;}+1)/2(\mathrm{\&gamma;}-1)}$

Wherein, A ^*Be nozzle throat area, M _tBe the nozzle throat Mach number, γ is the specific heat at constant pressure of gas and the specific heat ratio of specific heat at constant volume.

As shown in Figure 2, after confirming inlet geometric parameter and nozzle throat area, confirm subsonic speed section curve according to inlet geometric parameter and nozzle throat area, subsonic speed section curve has two sections, is distributed in the configuration line segment A of entry position respectively ₁A ₂Two ends.Adopt two arc methods to obtain subsonic speed section curve A in the present embodiment ₁A ₃And A ₂A ₄, two circular arcs can connect with straight-line segment, and wherein the center of circle and radius are confirmed according to Structural Design Requirement, the starting point and the slope thereof of two-end-point corresponding jet pipe wall entrance of difference and slope thereof, the actual wall of jet pipe supersonic speed section.This structure can guarantee that the subsonic speed section curve of designing can have good transition and natural being connected with supersonic speed section curved portion; Make asymmetric supersonic nozzle have the pneumatic profile of jet pipe of continuous curvature, thereby make the air-flow that gets into the supersonic speed section obtain better fluidised form.

As shown in Figure 3, after definite subsonic speed section curve,, confirm that by subsonic speed section curve and throat's radius-of-curvature transonic speed the district is interior through A according to the Sauer method ₃The characteristic curve A of point ₃A ₄

The corresponding transonic speed initial value line equation of Sauer method is provided by following formula:

$x=-\frac{(\mathrm{\&gamma;}+1)\mathrm{\&alpha;}}{8\mathrm{<figure-callout\; id="2"\; label="L\; y"\; filenames="HDA0000062089930000041.png"\; state="\{\{state\}\}">L</figure-callout>}}{y}^{}{}^2$

Wherein:

$\mathrm{\&alpha;}={\left[\frac{2}{(\mathrm{\&gamma;}+1)r_t{\mathrm{\&rho;}}_t}\right]}^{1/2}L$

L＝ρ _t+r _t

ρ _tAnd r _tBe respectively the radius-of-curvature and half height of nozzle throat.

According to transonic speed initial value line and symmetric condition, obtain the initial characteristics line then through the method for characteristic iteration.

The method of characteristic iterative formula is:

Wherein, x is a horizontal ordinate, and Δ x is the variation numerical value on horizontal ordinate, and r is an ordinate; Δ r is the variation numerical value on ordinate, and θ is local flow direction angle, and Δ θ is that local flow direction angle changes numerical value, and M is local Mach number and M＞1; Δ M is that local Mach number changes numerical value, and δ is the pattern of flow parameter, for two dimension δ=0 of flowing, and rotational symmetry δ=1 of flowing; R ≠ 0, in the present embodiment, owing to be asymmetric supersonic nozzle, when it is two-dimentional flowing; Select δ=0, when it be three-dimensional asymmetric when mobile, δ and r are definite according to actual conditions.

The process of finding the solution a characteristic curve unit is shown in the following figure:

Known (x ₁, r ₁, M ₁, θ ₁), (x ₂, r ₂, M ₂, θ ₂), find the solution (x ₃, r ₃, M ₃, θ ₃)

Preferably, it is symmetrical jet pipe that jet pipe is transonic speed distinguished in the Asia, is convenient to design so more, can effectively improve design efficiency, and obtains satisfied project organization.

As shown in Figure 4, after confirming the characteristic curve of transonic speed distinguishing, an end points that utilizes the quadratic spline curve to connect the initial characteristics line is confirmed the first side wall surface curve A with corresponding supersonic region outlet end points ₃A ₆, and, make the first side wall surface curve A according to jet pipe design parameter adjustment quadratic spline curve ₃A ₆To the shape that satisfies Structural Design Requirement.According to designing requirement, confirm the first side wall surface curve A then ₃A ₆Mach Number Distribution, and guarantee that through cubic spline curve the second derivative of Mach number is continuous.Referring to Fig. 5, segment of curve GH wherein, EF, DA ₅Be initial characteristics line A ₃A ₄Etc. the Mach Number Distribution line.

As shown in Figure 5, after confirming that transonic speed the mach line such as grade of section distributes, confirm the second side wall surface curve A through the mass conservation and method of characteristic according to the Mach Number Distribution of initial characteristics line and the first side wall surface curve ₄A ₅, because the initial characteristics line is confirmed the some A at initial characteristics line two ends ₃And A ₄Confirmed that also through utilizing the initial characteristics line and waiting mach line, according to the Mach Number Distribution of the first side wall surface curve, and the mass distribution on the initial characteristics line calculates, and confirms at A according to the mass conservation ₄A ₅Between the position distribution of each unique point finally confirm the second side wall surface curve A ₄A ₅Curve, then according to the asymmetric supersonic nozzle of above-mentioned resulting each parameter and curve design processing needs.Method of characteristic can be referring to the characteristic curve iterative equation in the initial characteristics line solution procedure.

As shown in Figure 6, because can there be the boundary layer in the existence of gas viscosity near the desirable wall curve of jet pipe, thereby influences the nozzle flow field quality, therefore need the desirable wall curve of jet pipe supersonic speed section is carried out the viscosity correction, obtain jet pipe supersonic speed section actual wall surface curve A ₄A ₇And A ₃A ₈

The step of viscosity correction:

A. find the solution coefficient of viscosity:

$\frac{\mathrm{\&mu;}}{{\mathrm{\&mu;}}_0}={\left(\frac{T}{T_0}\right)}^{1.5}\left(\frac{T_0+T_s}{T+T_s}\right)$

T wherein ₀=273.16K, μ ₀Be T under the atmospheric pressure ₀The dynamics coefficient of viscosity of gas during=273.16K, T _sBe the Sutherland constant, relevant with the character of gas, for air, μ ₀=1.7161 * 10 ^-5, T _s=124K, T represent local observed temperature.

B. find the solution static temperature:

$T_e=T_0(1+\frac{\mathrm{\&gamma;}-1}{2}{M}^2)$

C. find the solution static pressure:

$p_e=p_0{(1+\frac{\mathrm{\&gamma;}-1}{2}{M}^2)}^{\frac{\mathrm{\&gamma;}}{1-\mathrm{\&gamma;}}}$

D. find the solution density:

${\mathrm{\&rho;}}_e=\frac{p_e}{R{T}_e}$

For air:

R＝287J/(kg·mol)

E. find the solution the velocity of sound:

$a_e=\sqrt{\mathrm{\&gamma;}{\mathrm{RT}}_e}$

F. find the solution speed:

u _e＝M _e*a _e

G. find the solution adiabatic wall temperature:

$T_{\mathrm{aw}}\&ap;T_e(1+\frac{\mathrm{\&gamma;}-1}{2}{\mathrm{Pr}}^{1/3}{M}_e^2)$

H. find the solution the reference length of Re number:

$x=\sqrt{\frac{\mathrm{\&gamma;}+1}{2}{r}^{*}{R}^{*}}$

R wherein ^*Be throat's half height, R ^*Be the nozzle throat radius-of-curvature.

I. find the solution the Re number:

${\mathrm{Re}}_x=\frac{{\mathrm{\&rho;}}_e{u}_{e}x}{{\mathrm{\&mu;}}_e}$

J. find the solution reference temperature:

T′＝0.5(T _w+T _e)+0.22(T _aw-T _e)

T wherein _wRepresent local actual measurement ST.

That k. finds the solution correspondence can not press friction factor:

L. can not press the relation of form factor and friction factor to be:

$H_i=\frac{1}{1-7\sqrt{C_{\mathrm{fi}}/2}}$

M. can press form factor and can not press the relation of form factor to be:

$H=\frac{T_w}{T_e}{H}_i+\frac{T_{\mathrm{aw}}}{T_e}-1$

N. can press friction factor and can not press the relation of friction factor to be:

With the C that tries to achieve _FiBe updated to the momentum integral relational expression with H:

$\frac{\mathrm{d\&tau;}}{\mathrm{dx}}+\mathrm{\&tau;}[\frac{2-{\mathrm{<figure-callout\; id="2"\; label="M"\; filenames="HDA0000062089930000041.png"\; state="\{\{state\}\}">M</figure-callout>}}^2+H}{M(1+\frac{\mathrm{\&gamma;}-1}{2}{M}^2)}\frac{\mathrm{dM}}{\mathrm{dx}}+\frac{1}{y}\frac{\mathrm{dy}}{\mathrm{dx}}]=\frac{{\mathrm{<figure-callout\; id="2"\; label="C\; f"\; filenames="HDA0000062089930000041.png"\; state="\{\{state\}\}">C</figure-callout>}}_f}{2}\sec \mathrm{\&phi;}$

$\mathrm{\&phi;}={\tan }^{-1}\left(\frac{\mathrm{dy}}{\mathrm{dx}}\right)$

$H=\frac{{\mathrm{\&delta;}}^{*}}{\mathrm{\&tau;}}$

Wherein, τ is a momentum loss thickness, δ ^*Be boundary layer displacement thickness, φ is the flow direction angle, and H is the boundary layer form factor.This is an ordinary differential equation group, adopts four step Runge-Kutta methods to find the solution, and obtains boundary layer displacement thickness, displacement thickness is appended to desirable wall curve obtain the actual wall of jet pipe.Obtain the revised supersonic nozzle curved wall of viscosity profile, can have the boundary layer near having overcome the wall curve that causes owing to gas viscosity, thereby influence the problem of flow field quality, further improved the precision and the quality of asymmetric supersonic nozzle.

Fig. 7 is according to asymmetric supersonic nozzle of the present invention; The asymmetric supersonic nozzle wall 2 of the first side wall surface curve that end points that this supersonic nozzle comprises transonic speed initial characteristics line in the district and corresponding supersonic region outlet end points are confirmed and the second side wall surface curve formation confirmed through the mass conservation and method of characteristic according to the Mach Number Distribution of initial characteristics line and the first side wall surface curve, this initial characteristics line is definite according to subsonic speed section curve and throat's radius-of-curvature.The asymmetric supersonic nozzle wall of supersonic nozzle also has the reference temperature solution of employing momentum integral relational expression and finds the solution the determined layer that revises the boundary of boundary layer displacement thickness, and according to inlet coordinate and the definite formed subsonic speed section of the subsonic speed section curve wall 1 of throat opening area.

From above description; Can find out that the above embodiments of the present invention have realized following technique effect: the supersonic nozzle wall of asymmetric supersonic nozzle is confirmed according to the mass conservation and method of characteristic, can be passed through the method for characteristic wave absorption; Make and do not have the wave of compression of concentrating in the flow field; Flow field parameter distributes and can be optimized according to practical application, can effectively guarantee the homogeneity at flow field Mach number and flow direction angle, improves flow field quality greatly.Adopt the momentum integral relational expression to calculate the boundary layer displacement thickness of supersonic nozzle; Confirm the boundary layer curve; Can there be the boundary layer near having overcome the wall curve that causes owing to gas viscosity; Thereby influence the problem of flow field quality, further improved the precision and the quality of asymmetric supersonic nozzle.Through quadratic spline curve adjustment the first side wall surface curve, make the first side wall surface curve to adjust according to the Mach Number Distribution of design, can control the Mach Number Distribution of the supersonic region of jet pipe more accurately, make it satisfy design demand.

The above is merely the preferred embodiments of the present invention, is not limited to the present invention, and for a person skilled in the art, the present invention can have various changes and variation.All within spirit of the present invention and principle, any modification of being done, be equal to replacement, improvement etc., all should be included within protection scope of the present invention.

Claims (12)

1. an asymmetric supersonic nozzle method for designing is characterized in that, comprising:

Confirm subsonic speed section curve according to inlet coordinate and throat opening area;

Confirm the transonic speed interior initial characteristics line in district according to said subsonic speed section curve and throat's radius-of-curvature;

An end points according to said initial characteristics line is confirmed the first side wall surface curve with corresponding supersonic region outlet end points;

Mach Number Distribution according to said initial characteristics line and said the first side wall surface curve is confirmed the second side wall surface curve through the mass conservation and method of characteristic.

2. asymmetric supersonic nozzle method for designing according to claim 1 is characterized in that the iterative formula of said method of characteristic is:

Wherein, x is a horizontal ordinate, and r is an ordinate, and θ is local flow direction angle, and M is local Mach number and M＞1, and δ is the pattern of flow parameter, for the two dimension δ=O that flows, and rotational symmetry δ=1 of flowing, r ≠ 0, γ is the specific heat at constant pressure of gas and the specific heat ratio of specific heat at constant volume.

3. asymmetric supersonic nozzle method for designing according to claim 1; It is characterized in that; After confirming the second side wall surface curve, also comprise: adopt the reference temperature solution of momentum integral relational expression to find the solution boundary layer displacement thickness; Carry out the boundary layer and revise, obtain the wall curve of actual use, said momentum integral relational expression is:

$\frac{\mathrm{d\&tau;}}{\mathrm{dx}}+\mathrm{\&tau;}[\frac{2-M^2+H}{M(1+\frac{\mathrm{\&gamma;}-1}{2}{M}^2)}\frac{\mathrm{dM}}{\mathrm{dx}}+\frac{1}{y}\frac{\mathrm{dy}}{\mathrm{dx}}]=\frac{C_f}{2}\sec \mathrm{\&phi;}$

$\mathrm{\&phi;}={\tan }^{-1}\left(\frac{\mathrm{dy}}{\mathrm{dx}}\right)$

$H=\frac{{\mathrm{\&delta;}}^{*}}{\mathrm{\&tau;}}$

Wherein, τ is a momentum loss thickness, δ ^*Be boundary layer displacement thickness, φ is the flow direction angle, and H is the boundary layer form factor, C _fFor pressing friction factor, M is local Mach number and M＞1, and γ is the specific heat at constant pressure of gas and the specific heat ratio of specific heat at constant volume.

4. asymmetric supersonic nozzle method for designing according to claim 1 is characterized in that, confirming also to comprise before the subsonic speed section curve according to inlet coordinate and throat opening area: confirm said throat opening area according to the isentropic relation formula, said isentropic relation formula is:

$\frac{A}{A^{*}}=\frac{1}{M_t}{\left[\left(\frac{2}{\mathrm{\&gamma;}+1}\right)(1+\frac{\mathrm{\&gamma;}-1}{2}{M}_t^2)\right]}^{(\mathrm{\&gamma;}+1)/2(\mathrm{\&gamma;}-1)}$

Wherein, A is the nozzle exit area, A ^*Be nozzle throat area, M _tBe the nozzle throat Mach number, γ is the specific heat at constant pressure of gas and the specific heat ratio of specific heat at constant volume.

5. asymmetric supersonic nozzle method for designing according to claim 4 is characterized in that, is confirming that according to the isentropic relation formula said throat opening area also comprises before: the boundary curve that requires to confirm entrance and exit according to the jet pipe geometry.

6. asymmetric supersonic nozzle method for designing according to claim 1 is characterized in that, the two arc methods of said subsonic speed section curve negotiating are confirmed.

7. asymmetric supersonic nozzle method for designing according to claim 6 is characterized in that, said initial characteristics line is confirmed according to following transonic speed initial value line equation:

$x=-\frac{(\mathrm{\&gamma;}+1)\mathrm{\&alpha;}}{8L}{y}^2$

Wherein, $\mathrm{\&alpha;}={\left[\frac{2}{(\mathrm{\&gamma;}+1)r_t{\mathrm{\&rho;}}_t}\right]}^{1/2}L$

L＝ρ _t+r _t

ρ _tAnd r _tBe respectively the radius-of-curvature and half height of nozzle throat, γ is the specific heat at constant pressure of gas and the specific heat ratio of specific heat at constant volume.

8. asymmetric supersonic nozzle method for designing according to claim 1; It is characterized in that; Said the first side wall surface curve confirms with corresponding supersonic region outlet end points through the end points that utilizes the quadratic spline curve to connect said initial characteristics line, and through the said the first side wall surface curve of parameter adjustment to the shape that satisfies Structural Design Requirement.

9. asymmetric supersonic nozzle method for designing according to claim 8; It is characterized in that; An end points according to the initial characteristics line is confirmed also to comprise after the first side wall surface curve with corresponding supersonic region outlet end points: according to said nozzle structure designing requirement; Confirm the Mach Number Distribution of said the first side wall surface curve through cubic spline curve, continuous with the second derivative that guarantees said Mach number.

10. asymmetric supersonic nozzle; It is characterized in that; Comprise the asymmetric supersonic nozzle wall that the first side wall surface curve and the second side wall surface curve form; An end points of the initial characteristics line in the transonic speed district that said the first side wall surface curve is confirmed according to subsonic speed section curve and throat's radius-of-curvature confirms that with corresponding supersonic region outlet end points the said second side wall surface curve is definite through the mass conservation and method of characteristic according to the Mach Number Distribution of said initial characteristics line and said the first side wall surface curve.

11. asymmetric supersonic nozzle according to claim 10 is characterized in that, said asymmetric supersonic nozzle wall has the reference temperature solution of employing momentum integral relational expression and finds the solution the determined layer that revises the boundary of boundary layer displacement thickness.

12. asymmetric supersonic nozzle according to claim 10 is characterized in that, said asymmetric supersonic nozzle wall comprises according to inlet coordinate and the definite formed subsonic speed section of the subsonic speed section curve wall of throat opening area.

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