CN115828418A - Strong interference area profile design method based on two-dimensional bending characteristic line theory - Google Patents
Strong interference area profile design method based on two-dimensional bending characteristic line theory Download PDFInfo
- Publication number
- CN115828418A CN115828418A CN202211448970.9A CN202211448970A CN115828418A CN 115828418 A CN115828418 A CN 115828418A CN 202211448970 A CN202211448970 A CN 202211448970A CN 115828418 A CN115828418 A CN 115828418A
- Authority
- CN
- China
- Prior art keywords
- shock wave
- flow field
- shape
- wall
- flow
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Classifications
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention develops a strong interference area profile design method based on a two-dimensional bending characteristic line theory, which comprises the following steps: step (1), acquiring a preset shock wave shape; step (2), acquiring flow field data i; step (3), determining the shape i of the wall surface according to the preset shock wave shape and the flow field data i; step (4), performing flow field simulation according to the shape i of the wall surface to obtain flow field data i +1; step (5), judging whether the similarity of the simulated shock wave shape i and the preset shock wave shape meets the requirement or not according to the flow field parameters corresponding to the flow field data i +1 and the wall surface shape i; and (6) outputting the wall shape i if the requirement is met, and if the requirement is not met, i = i +1 and re-executing the steps (2) to (6). Therefore, the technical blank in the field of reverse design of the appearance of the aircraft in the shock wave strong interference area in China is filled, and support is provided for aerodynamic thermal protection of the high-speed aircraft and analysis of the aerodynamic thermal environment of strong shock wave interference in the future.
Description
Technical Field
The application relates to the technical field of combined power aircrafts, in particular to a strong interference area profile design method based on a two-dimensional bending characteristic line theory.
Background
The lip front edge of the air inlet of the hypersonic internal and external flow integrated aircraft has a quasi-two-dimensional characteristic in a symmetrical plane, the solution of a flow field in the symmetrical plane is generally regarded as a two-dimensional problem in the early stage of engineering application, and the influence brought by the three-dimensional effect of the flow field is only considered in a boundary condition. In order to ensure that the air intake system of the aircraft normally works under the preset flight condition, the profile of the strong interference area of the lip mouth needs to be designed, and the shock wave interference characteristic is weakened (no shock wave interference under the ideal condition) as far as possible on the premise of meeting the working condition, so that the effect of reducing the wall surface heat flow is achieved.
At present, the method for designing the profile of the shock wave interference at home and abroad is mainly based on an Oswatitsch wave matching theory and an isentropic compression theory, and is used for solving a post-shock wave flow field through the alternate propulsion of oblique shock wave characteristic lines and finally designing the profile of the wall surface center line through repeated iteration. The method is commonly used for the design of the traditional binary air inlet, but a new design method of the profile of the interference area needs to be researched in consideration of the fact that most of the appearance of the air inlet of the hypersonic aircraft is a curved profile and the shock wave form is a curved compression shock wave. Meanwhile, the Oswatitsch wave matching theory has a limited application range, and the original formula is not applicable under the condition of hypersonic velocity.
Disclosure of Invention
The application provides a strong interference area profile design method based on a two-dimensional bending characteristic line theory, and aims to solve the problem of two-dimensional design of a strong interference area profile by combining bending shock wave compression and isentropic compression based on the characteristic line theory. Therefore, the technical blank in the field of reverse design of the appearance of the aircraft in the shock wave strong interference area in China is filled, and support is provided for aerodynamic thermal protection of the high-speed aircraft and analysis of the strong shock wave interference aerodynamic thermal environment in the future.
In a first aspect, a wall design method is provided, including:
step (1), acquiring a preset shock wave shape;
step (2), acquiring flow field data i;
step (3), determining a wall surface shape i according to the preset shock wave shape and the flow field data i, wherein the wall surface shape if is W,i () Satisfies the following conditions:
y W,i ＝f W,i (x W,i ),0≤x W,i ≤L，
x W,i is the x coordinate, y, of the wall surface W,i Is the wall y coordinate, y st0,i Is the streamline origin y coordinate, M x,i Mach number of cross section x, q ∞ For standard heat flow, q x,i Is the heat flow of section x, mu is the Mach angle of section x, theta x,i The flow angle of the gas stream being the cross section x;
step (4), flow field simulation is executed according to the wall surface shape i, and flow field data i +1 are obtained;
step (5), judging whether the similarity between the simulated shock wave shape i and the preset shock wave shape meets the requirement or not according to the flow field parameters corresponding to the flow field data i +1 and the wall surface shape i;
and (6) outputting the wall shape i if the requirement is met, and if the requirement is not met, i = i +1 and re-executing the steps (2) to (6).
Compared with the prior art, the scheme provided by the application at least comprises the following beneficial technical effects:
(1) By enhancing the shock wave capturing stability and the high-resolution characteristic of the boundary layer, the aerodynamic heat accuracy of the complex flow area predicted by the conventional numerical method is effectively improved.
(2) Basic characteristics to be satisfied based on a hypersonic aerodynamic heat prediction flux function: shock wave stability and boundary layer solving capability, a novel flux function is developed and constructed, and a high-precision heat flow prediction method under micro-scale, high-energy area and strong shock wave interference is obtained.
(3) And providing a rough zone flow regulation and control heat reduction scheme and a flow field regulation and control heat reduction mechanism. By changing the size of the separation bubble, the flow is reattached at a proper angle under the condition of corresponding Mach number incoming flow, and the jet flows along the outer edge of the separation bubble in a shear layer mode and is reattached at a small angle finally.
With reference to the first aspect, in certain implementations of the first aspect, the preset shock wave shape is used to indicate at least one of: pneumatic parameter distribution of a compression surface, outlet parameter distribution and bending shock wave form.
The preset shock wave shape is set through a plurality of angles, so that the wall surface shape obtained through inversion can meet application requirements.
With reference to the first aspect, in certain implementations of the first aspect, the flow field data i includes at least one of: parameters before and after shock waves, wall parameters, outlet parameters and streamline data in a flow field.
The flow data volume is relatively large, and the wall surface shape can be deduced more accurately.
With reference to the first aspect, in certain implementations of the first aspect, for a case where the same-side oblique shock wave 1 intersects with the oblique shock wave 2, the streamline data i satisfies:
discrete expansion wave turning angle
Wherein M represents Mach number, p represents pressure, k is gas specific heat ratio, and subscript BC represents incoming flow passing through turning angle delta BC ＝δ B +δ C The region B is a region between the oblique shock wave 1 and the oblique shock wave 2, the region B is located at the upstream of the oblique shock wave 2, and the region C is a downstream region of the oblique shock wave 2.
Before the wall shape is deduced, streamline data is determined according to the intersection condition of the oblique shock wave 1 and the oblique shock wave 2 on the same side, so that the input data of wall inversion is more close to the actual condition, and the deduction process is more accurate.
With reference to the first aspect, in certain implementations of the first aspect, the wall shape i is also f W,i () Satisfies the following conditions:
The wall shape i also needs to satisfy some of the flow field constraints to make the deduction process more accurate.
With reference to the first aspect, in certain implementations of the first aspect, the wall shape i is a wall shape of a supersonic curved leading edge section and a subsonic root region, the supersonic curved leading edge section and the subsonic root region being determined by a curved shock wave shape and a wall heat flow.
The method and the device can be designed aiming at the local wall surface, are favorable for reducing the data volume introduced during deduction of the shape of the wall surface, and improve the data processing efficiency.
With reference to the first aspect, in certain implementations of the first aspect, the wall shape i further satisfies:
θ st ＝θ W
δ st ＝(θ W -2δ exp )-(θ W,st0 -2δ exp,st0 )
where θ is the airflow angle, δ represents the turning angle, subscript st represents a point on the streamline, subscript w represents a point on the wall, subscript w, st0 represents the start of the streamline on the wall, subscript exp represents the expansion wave, subscript exp, st0 represents the start of the streamline in the expansion wave zone.
Therefore, the influence of expansion wave reflected on the shock wave and the re-reflection of the expansion wave on the wall surface can be considered, and the deduction process is more accurate.
With reference to the first aspect, in certain implementations of the first aspect, the flow field parameter includes a mach number Ma st And pressure p st ：
M st ＝f P-M,M (M st0 ,δ st )
p st ＝f P-M,p (Ma st0 ,p st0 ,δ st )
M represents Mach number, p represents pressure, delta represents turning angle, subscript st represents point on streamline, subscript st0 represents streamline starting point, f P-M,p () Representing the Prandtl-Meyer flow function.
And the Mach number and the pressure are adopted to verify whether the deduction of the wall shape is reasonable or not, so that the design verification process can be improved more reasonably.
With reference to the first aspect, in certain implementations of the first aspect, the wall shape i satisfies isentropic fan separation, cross-sectional averaging, and two-dimensional isentropic steady flow assumptions.
And the calculation model is reasonably simplified by assumption, so that the data processing efficiency is improved.
With reference to the first aspect, in certain implementations of the first aspect, the determining the wall shape i includes:
and calculating shock waves, expansion waves and post-wave airflow parameters in different flow field domains by using a chord section method until all internal units and intersection points of incident shock waves and slip lines are calculated, and tracking a streamline emitted from a shock wave starting point to obtain the wall surface shape i.
With reference to the first aspect, in certain implementations of the first aspect, the performing flow field simulation includes:
generating a CFD flow field mesh topology based on the wall surface obtained by inverse design;
introducing the mesh topology into a flow field solver, establishing a CFD simulation model, and selecting a shock wave stabilization method, a time propulsion method, a turbulence model and a control equation;
setting initial conditions, and performing steady simulation on the CFD simulation model to obtain the flow field data i +1.
In a second aspect, an electronic device is provided, which is configured to perform the method as described in any one of the implementations of the first aspect.
Drawings
Fig. 1 is a flow chart of a two-dimensional profile design method.
Fig. 2 is a schematic diagram of a two-dimensional profile design near-wall flow field.
FIG. 3 is a schematic diagram of a oblique shock wave on a wedge.
Fig. 4 is a schematic diagram of the intersection of oblique shock waves and ipsilateral oblique shock waves.
Fig. 5 is a schematic illustration of the reflection of an expansion wave.
FIG. 6 is a schematic representation of a curved leading edge flow field partitioning inversion.
Fig. 7 is a schematic diagram of a preset shock wave, an inversion wall and a numerical verification wall.
FIG. 8 is a schematic diagram comparing near-wall numerical simulation Mach number cloud charts before and after inversion design.
Detailed Description
The present application is described in further detail below with reference to the accompanying drawings and specific embodiments.
The invention discloses a strong interference area profile design method based on a two-dimensional bending characteristic line theory. Fig. 1 shows a schematic flow diagram of the method. Fig. 2 is a schematic diagram of a two-dimensional profile design near-wall flow field. The method specifically comprises the following steps.
S1, based on the characteristics of two-dimensional shock wave interference, establishing the following simplifying assumptions of equal-entropy wave fan separation, section averaging and two-dimensional equal-entropy steady flow.
(1) Wave fan separation
When the turning angle caused by the reflected wave is calculated, the scattered wave fan expansion wave is simplified into a discrete expansion wave which is concentrated by the compression wave fan and is intersected with the shock wave to be reflected, and the calculation is approximate to a discrete expansion wave.
(2) Cross section averaging
The outlet cross section is defined as a left-line characteristic line emitted from the tail end of the wall surface in the approximate calculation. Mach number, pressure, total pressure are taken as the average of the parameters of the wall, after the shock wave and at the tail end of the streamline:
in the formula, n is the number of streamlines, and i is the serial number. M denotes mach number, P denotes pressure, P x denotes total pressure, subscript x denotes cross section x, subscript W denotes wall surface, subscript st, i denotes ith streamline, and subscript s denotes shock wave. Therefore, the approximate values of the performance parameters such as the Mach number of the outlet section, the pressure ratio, the total pressure recovery coefficient and the like can be calculated.
(3) Two-dimensional isentropic steady flow
In order to solve the flow under the interference of the two-dimensional shock wave, a characteristic line equation set (comprising a characteristic line equation, a streamline equation and a corresponding compatible equation) is derived from an Euler equation, and then an ordinary differential equation set is solved by using a finite difference method. Therefore, the design method is suitable for two-dimensional steady isentropic flow, and if a flow field with stronger shock waves (the pressure ratio exceeds 2.0) appears inside, the calculation is still continued according to the isentropic condition.
And S2, establishing an approximate method for calculating parameters before and after shock waves, wall parameters, outlet parameters and streamlines in a flow field according to the pneumatic parameter distribution, the outlet parameter distribution and the bending shock wave form of the compression surface required by the specification, wherein the calculated relational expressions are all given in the form of explicit analytical solutions.
(1) Oblique shock wave
As shown in fig. 3, the shock angle β can be determined using the oblique shock front-to-back relationship:
in the formula, M represents mach number, p represents pressure, δ represents a turning angle (representing a turning change in the direction of the gas flow), β represents a shock angle (an angle between a shock wave and a wall surface), k represents a gas specific heat ratio, and subscripts a and B represent regions shown in fig. 3. The region a may be a region upstream of the shock wave. The B area is the downstream area of the A area and is positioned at the downstream of the shock wave.
After some algebraic transformations, an explicit expression of the exact solution to the shock tangent can be derived:
wherein
All are variables determined by incoming flow parameters.
(2) The oblique shock waves on the same side are intersected with each other
As shown in FIGS. 2 and 4, oblique shock wave 1 (airflow turning angle and shock wave angle are delta, respectively) B 、β B ) And oblique shock wave 2 (airflow turning angle and shock wave angle are delta respectively) C 、β C ) A new oblique shock wave 3 (the shock angle is beta) is formed after the intersection F ) And produces reflected wave trains and slip stream discontinuities. At this time, the bending angle of the discrete extensional wave simplified from the fan extensional wave is as follows:
wherein M represents Mach number, p represents pressure, k is gas specific heat ratio, subscript C represents C region in FIG. 4, D represents expansion wave D region, and subscript BC represents incoming flow passing through turning angle delta BC ＝δ B +δ C The area of (a). The region B is a region between oblique shock wave 1 and oblique shock wave 2. The region B is located upstream of the oblique shockwave 2 and is an upstream region of the region C. The region C is a downstream region of the oblique shock wave 2. Oblique shock wave 1 and oblique shock wave 2 can form oblique shock wave 3 and slip line after intersecting, and the slip line is the slip of oblique shock wave 1The line shift, region F, is located between oblique shock wave 3 and slip line | (shown as SL in fig. 2). Oblique shock wave 1 and oblique shock wave 2 will also form an expansion wave after intersecting, and the expansion sector is indicated by the D area. The region located between the D region and the F region is an E region. A jet can be formed between zone E and zone F.
(3) Expansion wave reflection at wall surface
As shown in fig. 5, the discrete expansion wave with the expansion angle δ is reflected on the wall surface parallel to the incoming flow, and is easy to analyze, and the gas flow passes through the region B from the region a to the region C, and the expansion angle is twice of the original expansion angle. The region a is an upstream region of the incident expansion wave, the region B is a region between the incident expansion wave and the reflected expansion wave, and the region C is a downstream region of the reflected expansion wave.
And S3, according to the bending shock wave form and the wall surface heat flow, giving a bending shock wave, taking a streamline emitted from the starting point of the bending shock wave as a boundary, and decomposing the bending front edge section flow field into a bending shock wave section flow field (the regional streamlines pass through the bending shock wave) and a straight front edge section flow field (the regional streamlines pass through the straight front edge shock wave), as shown in figure 6.
(1) And dividing the flow field into an ultrasonic bending front edge section and a subsonic root area, and then performing partition inversion on the wall surface. And at the flow field boundary point, the shock wave at the upstream of the subsonic region is intersected with the shock wave at the supersonic bending front edge section, and the shock wave angles of the shock wave and the shock wave are equal.
(2) Due to the fact that a theoretical solving method is lacked in the subsonic velocity region, a flow field with a proper root shock wave form is selected from an existing sRR configuration, and the subsonic velocity flow field which is not affected by shock wave interference is intercepted to serve as the subsonic velocity root region. The Mach number, pressure and flow direction angle of incoming flow are respectively M ∞ 、p ∞ 、θ ∞ The curved compression profile (i.e., the wall at the supersonic curved leading edge section and subsonic root region) is determined by the following equation:
y W ＝f W (x W ),0≤x W ≤L
where x and y represent coordinates, the f-function represents the profile function of the curved wall, the subscript W represents the wall, and L is the wall length. The coordinate system is selected so that the abscissa direction is the same as the incoming flow direction, and the origin is the wall surface starting point. The distribution of wall surface angles can be obtained:
wherein delta W δ denotes the turning angle at the wall, x denotes the coordinate, f denotes the profile function of the curved wall, and the subscript w denotes the wall.
(3) According to the matching condition of the existing subsonic velocity section, by referring to a supersonic air inlet channel/special-shaped supersonic inner flow channel reverse design method, a bending shock wave section flow field and a straight front edge section flow field section shock wave with reasonable forms are respectively constructed. Dispersing the bending shock waves, and calculating shock waves, expansion waves and post-wave airflow parameters in different flow field areas until all internal units and intersection points of incident shock waves and slip lines are calculated.
And S4, calculating shock waves, expansion waves and post-wave airflow parameters in different flow field domains by using a chord section method until all internal units and intersection points of incident shock waves and slip lines are calculated, and tracking streamlines emitted from shock wave starting points, namely the wall surface required by inverse design. The wall inversion design method is summarized as follows.
And calculating coordinates of the shock waves at each interference point according to the parameters of each area, reversely deducing an intersection point of the compression waves and the wall surface from the shock wave interference point, and tracking a streamline sent out from the shock wave initial point, namely the target wall surface form obtained by inversion. From the fact that the flow in the cross section (the cross section is perpendicular to the X direction) in the X direction (i.e. the X abscissa direction in S3) is equal to the inflow capture cross section flow, the X-ray flow can be obtained w Streamline equation for independent variable:
where X and y represent coordinates, μ is the Mach angle, θ is the airflow flow angle, M is the Mach number, q is the heat flow, subscript st represents a point on the streamline, subscript st0 represents the streamline start, subscript w represents a point on the wall, subscript X represents the cross-section X, subscript X may be a discrete value in the X direction, q is the discrete value in the X direction, and ∞ is standard heat flow, is used forThe standard value of the incoming flow is defined and can be artificially determined according to requirements, for example, the peak heat flow of the sharp front edge is used as a reference value.
Referring to FIG. 2, the flow direction θ of a point on the streamline is considered in consideration of the influence of the expansion wave reflected by the shock wave and the re-reflection of the expansion wave on the wall surface st And a turning angle delta compressed from the starting point of the streamline st Are respectively approximated as:
θ st ＝θ W
δ st ＝(θ W -2δ exp )-(θ W,st0 -2δ exp,st0 )
where θ is the airflow angle, δ represents the turning angle, subscript st represents a point on the streamline, subscript w represents a point on the wall, subscript w, st0 represents the start of the streamline on the wall, subscript exp represents the expansion wave, subscript exp, st0 represents the start of the streamline in the expansion wave zone.
Solving corresponding Mach number M from streamline starting point parameter (i.e. shock wave back parameter) st And pressure p st :
M st ＝f P-M,M (M st0 ,δ st )
p st ＝f P-M,p (M st0 ,p st0 ,δ st )
Where M represents Mach number, p represents pressure, delta represents turning angle, subscript st represents a point on the streamline, subscript st0 represents the start of the streamline, f P-M Representing the Prandtl-Meyer flow function. Finally, the product can be obtained by x w And as an expression of the independent variable, solving the coordinate on the streamline, and simultaneously obtaining the Mach number, the pressure and the flow direction of the corresponding position. That is, even, the flow near the wall surface satisfies Prandtl-Meyer flow, namely the air flow presents two-dimensional constant isentropic flow accelerated by supersonic expansion around the external obtuse angle.
And S5, generating a CFD flow field mesh topology based on the wall surface obtained by inverse design. The method comprises the following steps: selecting the characteristic size of the strong shock wave interference area of the aircraft as a length unit, and carrying out mesh topology division on each geometric surface without gaps and overlapping areas.
And S6, introducing the mesh topology into a flow field solver, and establishing a CFD simulation model. The constructed high-resolution flux function is in a RoeMAS format; the shock wave stabilization method adopts a MUSCL format and a high-precision WENO format to be mixed; the unsteady coupling heat transfer simulation method is a radial basis function interpolation method; the unsteady time advancing method comprises an explicit Runge-Kutta format and an implicit post-difference time format; the turbulence model adopts a two-equation SST k-turbulence model; the control equation adopts a three-dimensional compressible Reynolds average N-S equation, which is concretely as follows:
wherein x, y, z represent three-dimensional coordinates,
is a conservation variable;
no sagittal flux for three directions;
is the flux of the viscosity vector in three directions.
S7, setting initial conditions, performing steady simulation on the CFD simulation model, and verifying the shock wave structure of the reverse design wall surface through numerical simulation results such as flow field pneumatic parameter distribution, flow field outlet parameter distribution and shock wave shape.
The simulation results may include the simulated mach number and pressure. Comparing Mach number obtained by simulation with Mach number M in the preceding st Comparing the simulated pressure with the pressure p in the preceding text st And judging whether the similarity of the simulated shock wave shape and the preset shock wave shape meets the requirement or not. And if the requirements are met, determining the shape of the wall surface obtained by inversion as the design shape of the final wall surface. And if the simulation result does not meet the requirement, re-executing the S1 to S7 according to the flow field data obtained by simulation and the preset shock wave shape until the similarity between the simulation shock wave shape and the preset shock wave shape meets the requirement.
Fig. 7 is a schematic diagram showing that the similarity of the simulated shock wave shape and the preset shock wave shape meets the requirement. The simulated shock shape, the preset shock shape, and the wall surface in fig. 7 may all refer to a central simulated shock shape, a central preset shock shape, and a central wall surface.
Fig. 8 shows a verification calculation example of a two-dimensional bending shock flow field, and the form of a two-dimensional bending shock is constructed by an elliptic parameter equation (incoming flow Ma =6, shock expression: x = cos (θ), y =2sin (θ), θ = -60 ° -15 °). In fig. 8, the predetermined gray scale of the shock wave is the same as the gray scale of the wall surface simulation result. As can be seen from the figure, the shock wave form obtained by numerical simulation of the inversion wall surface is in good accordance with the preset shock wave form, and the method is shown to have better performance in solving the two-dimensional problem.
According to the method, a two-dimensional bow shock wave and oblique shock wave interference theoretical model is used for decomposing a curved leading edge section flow field into a curved shock wave section flow field and a straight leading edge section flow field, and solving is carried out respectively, so that two-dimensional curved profile flow field parameters of a shock wave interference area are obtained. The application provides a flow field iteration method of two-dimensional bow shock wave and bending shock wave interference under the shock wave same-side interference, and the flow field parameter calculation of the bending shock wave and bow shock wave interference is realized. According to the method, a flow field is recurred according to the preset shape of a bending shock wave and the three-dimensional effect boundary condition of the flow field, the aerodynamic parameters and the coordinate conditions of the flow field of a bending shock wave section and a straight front edge section are obtained through solving, the profile flow line is tracked according to the characteristic line theory, and the high-interference-area two-dimensional profile after heat reduction design is formed. The application relates to a strong interference area profile design method based on a two-dimensional bending characteristic line theory, a lip mouth reverse design method is developed through the bending shock wave characteristic line theory, the method can adapt to the wall surface boundary design of two groups of bow shock wave interference, and can be widely applied to the fields of high-speed aircraft aerodynamic thermal environment assessment, aerodynamic thermal protection design and the like.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications without departing from the spirit and scope of the present invention.
Claims (12)
1. A wall surface designing method, comprising:
step (1), acquiring a preset shock wave shape;
step (2), acquiring flow field data i;
step (3), determining a wall surface shape i according to the preset shock wave shape and the flow field data i, wherein the wall surface shape is if W,i () Satisfies the following conditions:
y W,i ＝f W,i (x W,i ),0≤x W,i ≤L，
x W,i is the x coordinate, y, of the wall surface W,i Is the wall y coordinate, y st0,i Is the streamline origin y coordinate, M x,i Mach number, q, of cross section x ∞ For standard heat flow, q x,i Is the heat flow of section x, mu is the Mach angle of section x, theta x,i The flow angle of the gas stream being the cross section x;
step (4), flow field simulation is executed according to the wall surface shape i, and flow field data i +1 are obtained;
step (5), judging whether the similarity between the simulated shock wave shape i and the preset shock wave shape meets the requirement or not according to the flow field data i +1 and the flow field parameters corresponding to the wall surface shape i;
and (6) outputting the wall shape i if the requirement is met, and if the requirement is not met, i = i +1 and re-executing the steps (2) to (6).
2. The method of claim 1, wherein the preset shock shape is indicative of at least one of: pneumatic parameter distribution of a compression surface, outlet parameter distribution and bending shock wave form.
3. The method of claim 1, wherein the flow field data i comprises at least one of: parameters before and after the shock wave, wall parameters, outlet parameters and streamline data in the flow field.
4. The method according to claim 1, wherein the streamline data i satisfies, for the case where the same-side oblique shock wave 1 intersects with the oblique shock wave 2:
discrete expansion wave turning angle
Wherein M represents Mach number, p represents pressure, k is gas specific heat ratio, subscript BC represents incoming flow passing through turning angle delta BC ＝δ B +δ C The region B is a region between the oblique shock wave 1 and the oblique shock wave 2, the region B is located at the upstream of the oblique shock wave 2, and the region C is a downstream region of the oblique shock wave 2.
5. Method according to claim 1, characterized in that the wall shape i is also f W,i () Satisfies the following conditions:
6. The method of claim 1, wherein the wall shape i is a wall shape of a supersonic curved leading edge section and a subsonic root region, the supersonic curved leading edge section and the subsonic root region being determined by a bending shock wave shape and a wall heat flow.
7. The method of claim 1, wherein the wall shape i further satisfies:
θ st ＝θ W
δ st ＝(θ W -2δ exp )-(θ W,st0 -2δ exp,st0 )
where θ is the airflow angle, δ represents the turning angle, subscript st represents a point on the streamline, subscript w represents a point on the wall, subscript w, st0 represents the streamline start on the wall, subscript exp represents the expansion wave, subscript exp, st0 represents the streamline start of the expansion wave zone.
8. The method of claim 1, wherein the flow field parameter comprises a mach number Ma st And pressure p st ：
M st ＝f P-M,M (M st0 ,δ st )
p st ＝f P-M,p (Ma st0 ,p st0 ,δ st )
M represents Mach number, p represents pressure, delta represents turning angle, subscript st represents point on streamline, subscript st0 represents streamline starting point, f P-M,p () Representing the Prandtl-Meyer flow function.
9. The method of claim 1, wherein the wall shape i satisfies isentropic fan separation, cross-sectional averaging, and two-dimensional isentropic steady flow assumptions.
10. The method of claim 1, wherein the determining the wall shape i comprises:
and calculating shock waves, expansion waves and post-wave airflow parameters in different flow field domains by using a chord section method until all internal units and intersection points of incident shock waves and slip lines are calculated, and tracking a streamline emitted from a shock wave starting point to obtain the wall surface shape i.
11. The method of claim 1, wherein the performing a flow field simulation comprises:
generating a CFD flow field mesh topology based on the wall surface obtained by inverse design;
introducing the mesh topology into a flow field solver, establishing a CFD simulation model, and selecting a shock wave stabilization method, a time propulsion method, a turbulence model and a control equation;
setting initial conditions, and performing steady simulation on the CFD simulation model to obtain the flow field data i +1.
12. An electronic device, characterized in that the electronic device is configured to perform the method of any of claims 1 to 11.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211448970.9A CN115828418A (en) | 2022-11-18 | 2022-11-18 | Strong interference area profile design method based on two-dimensional bending characteristic line theory |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211448970.9A CN115828418A (en) | 2022-11-18 | 2022-11-18 | Strong interference area profile design method based on two-dimensional bending characteristic line theory |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN115828418A true CN115828418A (en) | 2023-03-21 |
Family
ID=85529290
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211448970.9A Pending CN115828418A (en) | 2022-11-18 | 2022-11-18 | Strong interference area profile design method based on two-dimensional bending characteristic line theory |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN115828418A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116384291A (en) * | 2023-06-06 | 2023-07-04 | 中国航天空气动力技术研究院 | Method for improving applicability of inverse characteristic line method by using expansion flow |
| CN117034470A (en) * | 2023-09-08 | 2023-11-10 | 北京流体动力科学研究中心 | Aircraft appearance rapid reverse design method based on high-performance numerical calculation |
-
2022
- 2022-11-18 CN CN202211448970.9A patent/CN115828418A/en active Pending
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116384291A (en) * | 2023-06-06 | 2023-07-04 | 中国航天空气动力技术研究院 | Method for improving applicability of inverse characteristic line method by using expansion flow |
| CN116384291B (en) * | 2023-06-06 | 2023-08-29 | 中国航天空气动力技术研究院 | Method for improving applicability of inverse characteristic line method by using expansion flow |
| CN117034470A (en) * | 2023-09-08 | 2023-11-10 | 北京流体动力科学研究中心 | Aircraft appearance rapid reverse design method based on high-performance numerical calculation |
| CN117034470B (en) * | 2023-09-08 | 2024-03-29 | 北京流体动力科学研究中心 | Aircraft appearance rapid reverse design method based on high-performance numerical calculation |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN115828418A (en) | Strong interference area profile design method based on two-dimensional bending characteristic line theory | |
| CN101477710B (en) | Body propelling integrated outer appearance modeling process for supersonic aircraft | |
| Huang et al. | Design method of internal waverider inlet under non-uniform upstream for inlet/forebody integration | |
| Rodi | The osculating flowfield method of waverider geometry generation | |
| CN105069221B (en) | For the critical Calculation Methods for Performance of Supersonic Inlet optimal design | |
| KR102616901B1 (en) | Aerodynamic layout design method and system for wide-speed-range hypersonic aircraft | |
| Boiko et al. | Determination of the laminar-turbulent transition location in numerical simulations of subsonic and transonic flows past a flat plate | |
| Slater | SUPIN: A computational tool for supersonic inlet design | |
| CN116070538A (en) | Interference area wall inversion method based on three-dimensional bending shock wave interference theory | |
| CN111780949B (en) | CFD analysis-based total pressure correction method for high-speed air inlet channel precursor wind tunnel experiment | |
| CN105138787A (en) | Supersonic flow field design method based on characteristic line tracing | |
| Boiko et al. | Numerical prediction of laminar-turbulent transition on an airfoil | |
| CN105574294A (en) | Method for rapidly determining curved shock wave compression flow field air flow parameters | |
| Saha et al. | Computational analysis on flow through transition S-diffusers: Effect of inlet shape | |
| Boiko et al. | Laminar-turbulent transition prediction module for LOGOS package | |
| CN109815549B (en) | Design method of single-pair supersonic flow direction vortex generating device | |
| Gounko et al. | Numerical modeling of the conditions for realization of flow regimes in supersonic axisymmetric conical inlets of internal compression | |
| CN115358101B (en) | Jet pipe design method based on sound velocity solution and characteristic line reverse thrust | |
| CN116976241A (en) | Wide-speed-range mixed-pressure type air inlet channel simplified modeling method | |
| CN115879216A (en) | Flow field reconstruction design method under internal flow channel strong wave system interference control | |
| Kuzmin | Shock wave instability in a bent channel with subsonic/supersonic exit | |
| Povinelli et al. | Viscous analyses for flow through subsonic and supersonic intakes | |
| Ni | A rational analysis of periodic flow perturbation in supersonic two-dimensional cascade | |
| Yuriy et al. | Numerical study of flow in the stage of an axial compressor with different topology of computational grid | |
| Borisov et al. | Numerical simulation of two streamwise supersonic vortices interaction |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |