CN110990955B - Hypersonic speed Bump air inlet channel design method and hypersonic speed Bump air inlet channel design system - Google Patents
Hypersonic speed Bump air inlet channel design method and hypersonic speed Bump air inlet channel design system Download PDFInfo
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Abstract
The invention discloses a hypersonic speed Bump air inlet channel design method and a hypersonic speed Bump air inlet channel design system, wherein the hypersonic speed Bump air inlet channel design method comprises the following steps: optimizing the profile expansion parametrization of the hypersonic axisymmetric air inlet, taking the flow coefficient range of the air inlet as constraint and taking the maximum total pressure recovery coefficient and the minimum internal contraction ratio of the throat of the air inlet as optimization targets, and optimizing to obtain a series of wall surface lines of the two-dimensional axisymmetric air inlet with different internal contraction ratios in the flow direction tangential plane; the two-dimensional axisymmetric air inlet channel is driven by a polynomial function, and the wall surface lines of the obtained two-dimensional axisymmetric air inlet channel in the flow direction tangential plane are circumferentially arranged to form a Bump profile; the polynomial function takes the center angle of the air inlet channel as an independent variable and the internal shrinkage ratio as a function value. The problems of low total pressure recovery performance, poor starting capability and the like of an air inlet channel in the prior art are solved.
Description
Technical Field
The invention relates to the technical field of aircrafts, in particular to a hypersonic speed Bump air inlet channel design method and system.
Background
The last nineties of the united states rocheid martin corporation proposed the concept of a dump inlet that enabled the lateral displacement of boundary layer air flow by adding a three-dimensional convex hull-like device at the inlet of the inlet, thereby allowing high energy air flow into the inlet. The Bump air inlet is always an important research direction of researchers at home and abroad. The current research on the Bump air inlet channel mainly focuses on the displacement performance of the Bump on a boundary layer under the supersonic incoming flow condition, and the Bump air inlet channel is applied to various fighter planes including F-35 and J-20. On the other hand, the Bump profile design is always the focus of the study of the Bump air inlet, and it is now common practice to use the wave-taking principle, i.e. to trace a wave-taking profile on a streamline in a reference flow field, as a Bump configuration, and then attach or directly install on the air inlet.
However, the design method of the Bump air inlet under hypersonic incoming flow condition is less researched at present. However, the direct application of the supersonic Bump design method to the hypersonic inlet design causes serious flow loss, resulting in low quality of the captured airflow. Therefore, the design method of the Bump air inlet under the hypersonic speed condition has an important effect on widening the application range of the Bump and improving the performance of the hypersonic speed air inlet, and is an important research field to be developed urgently at present. Patent "a design method of axisymmetric pre-compression precursor with boundary layer displacement", patent No.: ZL 201710784957.3' provides a precompressed precursor design method with boundary layer displacement suitable for high ultrasonic speed conditions, the method realizes the integrated design of a Bump and an air inlet channel, effectively reduces flow loss, and simultaneously the Bump molded surface obtained by design produces the displacement effect on the boundary layer. However, the method only provides a Bump/precursor integrated design thought, the molded surface of the Bump air inlet is not carefully designed, and the total pressure recovery performance of the air inlet is not high.
Disclosure of Invention
The invention provides a hypersonic speed Bump air inlet channel design method and system, which are used for overcoming the defects of low total pressure recovery performance, poor starting performance and the like of an air inlet channel in the prior art, realizing the integrated design of a Bump molded surface and the air inlet channel under hypersonic speed incoming flow conditions, and improving the total pressure recovery performance and the starting performance of the air inlet channel.
In order to achieve the above purpose, the present invention provides a hypersonic speed Bump air inlet design method, including:
optimizing the profile expansion parametrization of the hypersonic axisymmetric air inlet, taking the flow coefficient range of the air inlet as constraint and taking the maximum total pressure recovery coefficient and the minimum internal contraction ratio of the throat of the air inlet as optimization targets, and optimizing to obtain a series of wall surface lines of the two-dimensional axisymmetric air inlet with different internal contraction ratios in the flow direction tangential plane;
The two-dimensional axisymmetric air inlet channel is driven by a polynomial function, and the wall surface lines of the obtained two-dimensional axisymmetric air inlet channel in the flow direction tangential plane are circumferentially arranged to form a Bump profile; the polynomial function takes the center angle of the air inlet channel as an independent variable and the internal shrinkage ratio as a function value.
In order to achieve the above object, the present invention further provides a hypersonic speed Bump air inlet design system, which includes a memory and a processor, wherein the memory stores a hypersonic speed Bump air inlet design program, and the processor executes the steps of the method when running the hypersonic speed Bump air inlet design program.
According to the hypersonic speed Bump air inlet design method and system, parameterization optimization is carried out on the air inlet molded surface determined by hypersonic speed incoming flow conditions and design conditions, the air inlet wall surface line is optimally designed by taking the range of the flow coefficient of the air inlet as a constraint and taking the maximum total pressure recovery coefficient and the minimum internal contraction ratio of the throat of the air inlet as optimization targets, a series of wall surface lines of two-dimensional axisymmetric air inlets with different internal contraction ratios in the flow direction tangential plane are obtained, then the two-dimensional axisymmetric air inlets are driven according to the functional relation of the central angle of the air inlet and the internal contraction ratio, the obtained wall surface lines of the two-dimensional axisymmetric air inlets in the flow direction tangential plane are unfolded and arranged in the circumferential direction, and finally the Bump molded surface is obtained. The wall surface lines of the air inlet on the profile flow direction tangential plane are all obtained by taking the range of the flow coefficient of the air inlet as constraint and taking the maximum total pressure recovery coefficient and the minimum internal contraction ratio of the throat of the air inlet as optimization targets under hypersonic speed conditions, so that the Bump air inlet design with high total pressure recovery capability and high starting performance can be realized.
Drawings
In order to more clearly illustrate the embodiments of the invention or the technical solutions in the prior art, the drawings that are required in the embodiments or the description of the prior art will be briefly described, it being obvious that the drawings in the following description are only some embodiments of the invention, and that other drawings may be obtained from the structures shown in these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a schematic view of wall line parameterization optimization of a two-dimensional axis symmetric air inlet in a flow direction section in a hypersonic speed Bump air inlet design method according to an embodiment of the invention;
FIG. 2 is a graph showing the distribution of available sample points in a target space optimized according to the first embodiment;
Fig. 3 is a schematic front view of a hypersonic speed Bump air intake duct obtained by circumferential arrangement in embodiment one.
The achievement of the objects, functional features and advantages of the present invention will be further described with reference to the accompanying drawings, in conjunction with the embodiments.
Detailed Description
The following description of the embodiments of the present invention will be made clearly and fully with reference to the accompanying drawings, in which it is evident that the embodiments described are only some, but not all embodiments of the invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
It should be noted that all directional indicators (such as up, down, left, right, front, and rear … …) in the embodiments of the present invention are merely used to explain the relative positional relationship, movement, etc. between the components in a particular gesture (as shown in the drawings), and if the particular gesture changes, the directional indicator changes accordingly.
Furthermore, descriptions such as those referred to as "first," "second," and the like, are provided for descriptive purposes only and are not to be construed as indicating or implying a relative importance or number of features indicated. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include at least one such feature. In the description of the present invention, the meaning of "plurality" means at least two, for example, two, three, etc., unless specifically defined otherwise.
In the present invention, unless specifically stated and limited otherwise, the terms "connected," "affixed," and the like are to be construed broadly, and for example, "affixed" may be a fixed connection, a removable connection, or an integral body; the device can be mechanically connected, electrically connected, physically connected or wirelessly connected; either directly or indirectly, through intermediaries, or both, may be in communication with each other within the two elements or in interaction with each other, unless expressly defined otherwise. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
In addition, the technical solutions of the embodiments of the present invention may be combined with each other, but it is necessary to be based on the fact that those skilled in the art can implement the present invention, and when the technical solutions are contradictory or cannot be implemented, the combination of the technical solutions should be considered as not existing, and not falling within the scope of protection claimed by the present invention.
Example 1
1-3, A hypersonic speed Bump air inlet design method comprises the following steps:
Step S1, performing parameterized optimization on the profile expansion of a hypersonic axisymmetric air inlet channel, taking the flow coefficient range of the air inlet channel as constraint and taking the maximum total pressure recovery coefficient and the minimum internal shrinkage ratio of the throat of the air inlet channel as optimization targets, and obtaining a series of wall surface lines of the two-dimensional axisymmetric air inlet channel with different internal shrinkage ratios in the flow direction section through optimization; the wall surface line of the two-dimensional axisymmetric air inlet in the flow direction section refers to a two-dimensional molded surface of the air inlet, and the integral molded surface of the air inlet is formed by a plurality of two-dimensional molded surfaces;
And (3) carrying out expansion optimization calculation on the profile of the air inlet channel based on an optimization algorithm to obtain the high-performance air inlet channel with different internal contraction ratios. The optimization of the air inlet profile comprises an air inlet profile parameterization method, optimization target and optimization constraint determination, and an air inlet profile optimization process based on an optimization platform.
And (5) expanding and parameterizing and optimizing the axisymmetric air inlet channel. Because the air inlet configuration has axisymmetry, only the two-dimensional molded surface of the air inlet in the flow direction tangential plane (namely the wall surface line of the two-dimensional axisymmetric air inlet in the flow direction tangential plane) can be solved, and the complete axisymmetric molded surface of the air inlet can be obtained by carrying out rotary transformation on the obtained two-dimensional molded surface.
The step S1 comprises the following steps:
step S11, modeling the wall surface in the flow direction section of the air inlet channel under a two-dimensional coordinate system based on design conditions, and constructing the wall surface line of the air inlet channel by adopting a third-order rational B-spline curve to obtain the geometric configuration of the wall surface line of the air inlet channel;
Referring to fig. 1, the wall line of the inlet in the flow direction section is known to include the relative coordinates of any two points of the inlet wall line AC and the lip shroud line BD, A, B, C, D; a is the initial point of the air inlet channel wall surface line, B is the lip point, C is the lower throat point of the air inlet channel, and D is the upper throat point of the air inlet channel; the above points are given points, and the lip mask line BD is given.
Taking the point A as a coordinate origin, taking the horizontal direction as an x axis, taking the vertical direction as an r axis, establishing a two-dimensional coordinate system, and placing the point B, C, D and the lip cover line BD in the two-dimensional coordinate system;
By varying the position of point P 1、P2 in the middle of A, C, an inlet wall line AC geometry is constructed in which point P 1 is near point A, point P 2 is near point C, the AP 1 angle, and the CP 2 angle are determined.
As an embodiment of the present invention:
taking A as an origin of a coordinate system, taking the horizontal direction as an x axis and taking the vertical direction as an r axis. Coordinates of C (1200,162), coordinates of B point (750,197), and coordinates of D point (1200,212). The points B and D of the lip cover are given points, the BD section adopts an arc-shaped line, the curve is alpha=4 degrees with the horizontal direction at the point B, and the curve is tangent with the horizontal direction at the point D, so that the molded line of the lip cover BD can be determined.
The problem to be solved in the step S1 of the invention is the optimal design of the inlet channel wall line AC.
The AC curves were constructed using 3-order rational B-splines. Rational B-splines are a common technique, and a 3-order rational B-spline can form a smooth curve by giving four driving points. In AC curve parameterization, points a and C are fixed points, and different curves can be achieved by changing the positions of intermediate points P 1 and P 2. The moving direction of P 1 is delta from the horizontal direction. Delta is given as the tangential direction of the precursor at point a, thereby ensuring a smooth transition between the aircraft precursor and the inlet profile. The length of AP 1 is defined as L1.P 2 is equal in height to point C and moves in a horizontal direction, thereby ensuring that the inlet wall line is horizontal at the outlet, defining the length of P 2 C as L2. In the parameterization of the AC curve, L1 and L2 are variables. Given the range of variation of L1 and L2, the AC curve can be flexibly adjusted within a certain range. One example is: δ=6.5 °, L1 ranges from [100,450], L2 ranges from [560,770].
S12, obtaining an internal contraction ratio of the air inlet channel according to the geometrical configuration of the air inlet channel wall surface line, and obtaining a total pressure recovery coefficient of the air inlet channel throat and a flow coefficient of the air inlet channel by carrying out flow field numerical simulation calculation on the geometrical configuration of the air inlet channel wall surface line;
The intake passage internal contraction ratio CR in can be obtained by the following formula:
CRin＝rB 2-rE 2/cosα(rD 2-rC 2) (1)
wherein BE is a straight line passing through point B and perpendicular to the inlet wall line AC, r is the vertical coordinate of the lower punctuation, α is the angle between BE and the r-axis direction, and the meaning of the angle between the upper curve BD and the horizontal direction at point B is the same, where α=4°.
And S13, performing multi-objective global optimization by taking a given flow coefficient range of the air inlet channel as a constraint condition and taking the maximum total pressure recovery coefficient of the throat part of the air inlet channel and the minimum internal contraction ratio of the air inlet channel as optimization targets to obtain wall surface lines of a plurality of two-dimensional axisymmetric air inlet channels in the flow direction tangential plane.
Determining an optimization target and an optimization constraint:
In axisymmetric inlet configuration optimization, the invention comprehensively considers the airflow flow, quality and inlet starting characteristics of the inlet. And (3) calculating the internal contraction ratio of the air inlet channel in the geometric configuration obtained by parameterization, and respectively reading the total pressure recovery coefficient of the throat of the air inlet channel and the flow coefficient of the air inlet channel in the flow field obtained by numerical simulation calculation. Taking the total pressure recovery coefficient sigma max of the throat of the air inlet and the internal contraction ratio CR in minimum of the air inlet as optimization targets, and simultaneously giving the flow coefficient range of the air inlet as constraint. The total pressure recovery coefficient of the air inlet and the flow coefficient of the air inlet can BE directly read out in flow field calculation software, CR in can BE obtained through calculation according to a formula (1), E is a point on a curve AC, BE is a straight line passing through the point B and perpendicular to the wall line AC of the air inlet, and the length of the line CD is the height of the throat of the air inlet.
As an embodiment of this step:
The optimization problem of the present invention can thus be described as:
optimization target: the total pressure recovery coefficient is maximum, and the internal shrinkage ratio is minimum;
Optimization constraint 1: l1[100,450], L2[560,770];
optimization constraint 2: the flow coefficient is not less than 0.95;
wherein the optimization constraint 1 is a constraint on an optimization variable, the optimization constraint 2 is a constraint on an optimization target, and the sample points conforming to the optimization constraint 2 are called available samples.
Preferably, the multi-target global optimization is carried out through a genetic algorithm, and a Pareto front edge is formed in a two-dimensional coordinate system which takes the total pressure recovery coefficient of the throat of the air inlet channel as a horizontal coordinate and the internal contraction ratio of the air inlet channel as an ordinate;
Fitting a trend line by using the Pareto front edge points to form a Pareto front edge line;
And finding out a corresponding wall line AC of the air inlet according to the total pressure recovery coefficient of the air inlet throat part on the Pareto front edge line and the internal contraction ratio distribution of the air inlet, and obtaining the wall line of the two-dimensional axisymmetric air inlet in the flow direction tangential plane.
Optimizing a wall line (two-dimensional profile) of the two-dimensional axisymmetric air inlet channel in a flow direction tangential plane based on an optimization platform:
The optimization platform is built in commercial software ISIGHT, ISIGHT is mature software in the optimization field, various optimization algorithms are provided in the software, and other software can be called in the software through script files. In the invention, the optimization platform comprises the following modules: the system comprises a geometric configuration generating module, a grid dividing and calculating simulation module and a data extracting module. The geometric configuration generating module is used for generating geometric configuration of the air inlet channel through the parameterization method; the grid division and calculation simulation module is used for carrying out grid division on the generated geometric configuration and expanding numerical simulation calculation to obtain the flow field of the air inlet channel. The grid division and numerical calculation method is a mature known technology in the field, and the invention adopts ICEM under open commercial software ANSYS company to carry out grid division and FLUENT to carry out numerical calculation of a flow field. The numerical result of the flow field of the air inlet channel can be obtained through the software; the data extraction module is used for reading the flow coefficient and the total pressure recovery coefficient of the throat CD of the air inlet in the obtained air inlet flow field, obtaining the internal contraction ratio CR in of the air inlet through a formula (1) in the generated geometric configuration, and counting the flow of the throat of the air inlet to obtain the flow coefficient of the air inlet.
A multi-objective global optimization algorithm (such as a genetic algorithm) is adopted to develop multi-objective optimization for the optimization problem.
As an example of this step: the NSGA-II algorithm in the genetic algorithm is adopted, the population is 20, the algebra is 30, and the crossing rate is 0.9. Through optimization calculation, 600 sample points are obtained.
For the multi-objective optimization problem, a Pareto front is formed in the optimization, the points on the front have a dominant effect on other points, and the points on the front are in a non-dominant solution relation. That is, the sample point on the Pareto leading edge can be considered to be the profile corresponding to the maximum total pressure recovery performance at different levels of the internal shrinkage ratio.
And circumferentially arranging a series of optimized two-dimensional axisymmetric air inlets on a wall line (two-dimensional molded surface) of a flow direction tangential plane to form a Bump molded surface.
In the optimization problem of the present invention, the points on the Pareto leading edge whose parameterized geometric variables L1 and L2 are also continuously variable, meaning that the sample configuration on the Pareto leading edge is continuously transitional.
The central angle theta of the air inlet is used as an independent variable, the internal shrinkage ratio is used as a function, and the distribution of the high-performance air inlet configurations with different internal shrinkage ratios in the circumferential direction is driven through a polynomial function, so that the distribution of different air inlet flow direction configurations in the circumferential direction is realized. Through polynomial function, the two-dimensional axisymmetric air inlet with small internal shrinkage ratio is arranged near the symmetrical plane at the wall line (two-dimensional profile) of the flow direction tangential plane, and the two-dimensional axisymmetric air inlet with large internal shrinkage ratio is arranged at two sides at the wall line (two-dimensional profile) of the flow direction tangential plane, so that a Bump configuration with high middle and low two sides is formed at the air inlet. Since the sample point argument distribution on the Pareto leading edge is smooth and continuous, the resulting Bump inlet channel is also smooth and continuous in the circumferential direction.
As an example of this step: the configuration of the air inlet channel with different internal contraction ratios is driven to be distributed in the circumferential direction through a quadratic function. Where θ 1 is a given value representing the distribution range of the lamp in the circumferential direction, which is given as 50 ° in the example. CR in,max and CR in,min are ranges of air inlets selected in Pareto leading edge, and CR in,max＝1.85,CRin,min =1.65 is selected in the example.
Fig. 2 is a graph of the distribution of available sample points in the target space, wherein triangle points are common sample points, circle points represent pareto front edge points, and black solid lines are pareto front edge lines obtained by performing polynomial fitting on the pareto front edge points four times.
S2, driving by a polynomial function, arranging the obtained two-dimensional axisymmetric air inlet channel in the circumferential direction of a wall line of a flow direction tangential plane to form a Bump profile; the polynomial function takes the center angle of the air inlet channel as an independent variable and the internal shrinkage ratio as a function value.
In step S2, the polynomial function using the inlet channel center angle θ as an argument and the internal contraction ratio as a function value is:
Wherein θ is the center angle of the air inlet, and represents the angle between the tangential plane and the symmetrical plane of the flow direction; θ 1 is a given value representing the distribution range of the ramp in the circumferential direction, and CR in,max and CR in,min are ranges of intake port inside contraction ratios selected from the optimized intake port wall lines.
Arranging the wall lines of the air inlet channel in the direction from the approaching symmetrical surface to the separating symmetrical surface on the flow direction tangential surface according to the order of the internal contraction ratio from small to large; and (3) carrying out curved surface lofting on the wall surface lines of the air inlet channel in all the flow direction tangential planes which are circumferentially arranged, and finally obtaining the three-dimensional Bump profile with high middle and low two sides.
Fig. 3 is a front view of a dump air intake duct, in which the vertical plane in which S0 is located is a symmetrical plane, the planes in which S1 to Sn are located are a tangential plane, and θ is the angle from the symmetrical plane to a different tangential plane, which is also referred to as the air intake duct center angle. The shape of the wall surface in the flow direction tangential plane of S0 to Sn is the wall surface line (two-dimensional surface) of the high-performance two-dimensional axisymmetric air inlet channel with different internal shrinkage ratios in the flow direction tangential plane, which is driven according to the formula (2); fig. 3 is a schematic front view of a hypersonic speed Bump air inlet obtained through circumferential arrangement, and curves S0, S1, S2, … … to Sn (curves of arc edge points in the flow direction tangential plane in fig. 3) in fig. 3 are air inlet wall surface lines of two-dimensional axisymmetric air inlets with different high performance internal shrinkage ratios in the flow direction tangential plane. θ is the angle between the different tangential flow planes and the vertical plane of symmetry S0. The plane where S0 to Sn are located is the flow direction tangential plane. CL is obtained by rotating the point C by theta 1 according to the symmetry plane S0 in FIG. 1, and DL is obtained by rotating the point D in FIG. 1 by the center angle of the air inlet channel according to the symmetry plane S0. And (3) lofting the S0-Sn curves to obtain the complete hypersonic speed Bump profile.
The invention provides an optimization design method of a Bump air inlet channel. Firstly, a series of high-performance two-dimensional axisymmetric air inlets with different internal shrinkage ratios are obtained through optimal design, and then the two-dimensional axisymmetric air inlets with different internal shrinkage ratios are arranged along the circumferential direction of the wall surface line (two-dimensional surface) of the flow direction tangential plane through function driving, so that the air inlets form a buffer surface with high middle and low two sides. The Bump configuration and the precursor profile are in smooth transition, namely the Bump air inlet channel and the precursor are completely integrated, so that unnecessary wave trains and resistance in a flow field are reduced, and the aerodynamic performance of the aircraft is improved. The two-dimensional profile of the air inlet channel in each flow direction tangential plane has good total pressure recovery performance, so that the obtained Bump air inlet channel has high total pressure recovery performance.
Example two
Based on the first embodiment, the present invention provides a hypersonic speed Bump air inlet design system, which includes a memory and a processor, wherein the memory stores a hypersonic speed Bump air inlet design program, and the processor executes steps of a method in any embodiment when the hypersonic speed Bump air inlet design program is executed.
The foregoing description is only of the preferred embodiments of the present invention and is not intended to limit the scope of the invention, and all equivalent structural changes made by the present description and drawings or direct/indirect application in other related technical fields are included in the scope of the present invention under the inventive concept of the present invention.
Claims (6)
1. The hypersonic speed Bump air inlet channel design method is characterized by comprising the following steps of:
Optimizing the profile expansion parametrization of the hypersonic axisymmetric air inlet, taking the flow coefficient range of the air inlet as constraint and taking the maximum total pressure recovery coefficient and the minimum internal contraction ratio of the throat of the air inlet as optimization targets, and optimizing to obtain a series of wall surface lines of the two-dimensional axisymmetric air inlet with different internal contraction ratios in the flow direction tangential plane; modeling the wall surface in the flow direction section of the air inlet under a two-dimensional coordinate system based on design conditions, and constructing the air inlet wall surface by adopting a third-order rational B-spline curve to obtain the geometrical configuration of the air inlet wall surface;
Obtaining an internal contraction ratio of the air inlet according to the geometrical configuration of the air inlet wall line, and obtaining a total pressure recovery coefficient of the throat of the air inlet and a flow coefficient of the air inlet by carrying out flow field numerical simulation calculation on the geometrical configuration of the air inlet wall line;
taking a given flow coefficient range of the air inlet as a constraint condition, and taking the maximum total pressure recovery coefficient of the throat part of the air inlet and the minimum internal contraction ratio of the air inlet as optimization targets to carry out multi-target global optimization to obtain wall lines of a plurality of two-dimensional axisymmetric air inlets in a flow direction tangential plane;
The wall line of the two-dimensional axisymmetric air inlet in the flow direction section comprises the relative coordinates of any two points of an air inlet wall line AC and a lip cover line BD, A, B, C, D, and the line type of the lip cover line BD is known; a is the initial point of the air inlet channel wall surface line, B is the lip point, C is the lower throat point of the air inlet channel, and D is the upper throat point of the air inlet channel;
Establishing a two-dimensional coordinate system by taking the point A as a coordinate origin, taking the horizontal direction as an x axis and taking the vertical direction as an r axis, and placing the point B, C, D and the lip cover line BD in the two-dimensional coordinate system;
By changing the position of point P 1、P2 in the middle of A, C, an inlet wall line AC geometry is constructed in which point P 1 is near point a, point P 2 is near point C, the AP 1 angle and the CP 2 angle are determined;
The two-dimensional axisymmetric air inlet channel is driven by a polynomial function, and the wall surface lines of the obtained two-dimensional axisymmetric air inlet channel in the flow direction tangential plane are circumferentially arranged to form a Bump profile; the polynomial function takes the center angle of the air inlet channel as an independent variable and the internal shrinkage ratio as a function value.
2. The hypersonic speed Bump inlet design method as set forth in claim 1, wherein the inlet internal contraction ratio CR in is obtained by the following formula:
CRin＝rB 2-rE 2/cosα(rD 2-rC 2)
Where r represents the vertical coordinate of the lower punctuation, E is the point on the curve AC, BE is the straight line passing through the point B and perpendicular to the inlet wall line AC, and α is the angle between BE and the r axis.
3. The hypersonic speed Bump air inlet design method of claim 2, wherein a genetic algorithm is used for carrying out multi-objective global optimization, and a Pareto front edge is formed in a two-dimensional coordinate system taking an air inlet throat total pressure recovery coefficient as a horizontal coordinate and an air inlet internal contraction ratio as a vertical coordinate;
Fitting a trend line by using the Pareto front edge points to form a Pareto front edge line;
And finding out a corresponding wall line AC of the air inlet according to the total pressure recovery coefficient of the air inlet throat part on the Pareto front edge line and the internal contraction ratio distribution of the air inlet, and obtaining the wall line of the two-dimensional axisymmetric air inlet in the flow direction tangential plane.
4. A hypersonic speed Bump air intake duct design method as set forth in claim 3, wherein a polynomial function using an air intake duct center angle θ as an independent variable and an internal contraction ratio CR in as a function value is:
wherein θ is the center angle of the air inlet, and represents the angle between the tangential plane and the symmetrical plane of the flow direction; θ 1 is a given value, which represents the distribution range of the Bump in the circumferential direction, and CR in,max and CR in,min are the ranges of the intake port internal contraction ratios selected from the optimized intake port wall lines.
5. The hypersonic speed Bump air inlet design method as set forth in claim 4, wherein the air inlet wall lines are arranged in the direction from the approaching symmetry plane to the separating symmetry plane on the flow direction tangential plane in the order of the decreasing internal contraction ratio;
And (3) carrying out curved surface lofting on the wall surfaces of the air inlet channels in all the flow direction tangential planes which are circumferentially arranged, and finally obtaining the three-dimensional Bump profile with high middle and low two sides.
6. A hypersonic speed Bump air inlet design system, comprising a memory and a processor, wherein the memory stores a hypersonic speed Bump air inlet design program, and the processor performs the steps of the method according to any one of claims 1 to 5 when running the hypersonic speed Bump air inlet design program.
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| CN113221483B (en) * | 2021-05-28 | 2023-04-18 | 西北工业大学 | Binary free-form air inlet channel optimization design method and system based on class shape function |
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Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6793175B1 (en) * | 1999-08-25 | 2004-09-21 | The Boeing Company | Supersonic external-compression diffuser and method for designing same |
| CN101813027A (en) * | 2010-03-29 | 2010-08-25 | 南京航空航天大学 | Bump air inlet method for realizing integration of unequal-strength wave system with forebody |
| CN107191272A (en) * | 2017-06-08 | 2017-09-22 | 南京航空航天大学 | A kind of internal channel method for designing profile of rectangle hypersonic inlet |
| CN109649666A (en) * | 2018-12-26 | 2019-04-19 | 中国人民解放军国防科技大学 | Boundary layer guiding and suction combined hypersonic air inlet channel flow control method |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8041545B2 (en) * | 2005-04-28 | 2011-10-18 | Vladimir Sevastyanov | Gradient based methods for multi-objective optimization |
-
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Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6793175B1 (en) * | 1999-08-25 | 2004-09-21 | The Boeing Company | Supersonic external-compression diffuser and method for designing same |
| CN101813027A (en) * | 2010-03-29 | 2010-08-25 | 南京航空航天大学 | Bump air inlet method for realizing integration of unequal-strength wave system with forebody |
| CN107191272A (en) * | 2017-06-08 | 2017-09-22 | 南京航空航天大学 | A kind of internal channel method for designing profile of rectangle hypersonic inlet |
| CN109649666A (en) * | 2018-12-26 | 2019-04-19 | 中国人民解放军国防科技大学 | Boundary layer guiding and suction combined hypersonic air inlet channel flow control method |
Non-Patent Citations (1)
| Title |
|---|
| Bump/前体一体化对高速进气道起动特性的影响研究;徐尚成;推进技术;20181208;全文 * |
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