CN109649666B - A Hypersonic Inlet Flow Control Method Combining Boundary Layer Guidance and Suction - Google Patents

A Hypersonic Inlet Flow Control Method Combining Boundary Layer Guidance and Suction Download PDF

Info

Publication number
CN109649666B
CN109649666B CN201811603785.6A CN201811603785A CN109649666B CN 109649666 B CN109649666 B CN 109649666B CN 201811603785 A CN201811603785 A CN 201811603785A CN 109649666 B CN109649666 B CN 109649666B
Authority
CN
China
Prior art keywords
boundary layer
suction
profile
hypersonic
distribution
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.)
Active
Application number
CN201811603785.6A
Other languages
Chinese (zh)
Other versions
CN109649666A (en
Inventor
王翼
徐尚成
王振国
范晓樯
苏丹
赵星宇
闫郭伟
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
National University of Defense Technology
Original Assignee
National University of Defense Technology
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by National University of Defense Technology filed Critical National University of Defense Technology
Priority to CN201811603785.6A priority Critical patent/CN109649666B/en
Publication of CN109649666A publication Critical patent/CN109649666A/en
Application granted granted Critical
Publication of CN109649666B publication Critical patent/CN109649666B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/02—Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C30/00—Supersonic type aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
    • B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C2230/00—Boundary layer controls

Landscapes

  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Manufacturing & Machinery (AREA)
  • Transportation (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)

Abstract

The invention discloses a hypersonic air inlet channel flow control method combining boundary layer guiding and suction, which comprises the following steps of: forming a boundary layer guide profile on the hypersonic air inlet channel precursor to enable the flowing direction of air flow flowing through the boundary layer guide profile to be directionally changed through a transverse pressure gradient generated on the boundary layer guide profile; controlling a transverse pressure gradient acting on the boundary layer guiding profile to guide a boundary layer to a local part of the boundary layer guiding profile; and a suction hole is arranged at a part of the boundary layer guide profile. The scheme solves the problem of low suction performance in the prior art, realizes targeted suction and improves suction performance.

Description

Boundary layer guiding and suction combined hypersonic air inlet channel flow control method
Technical Field
The invention relates to the technical field of hypersonic air inlets, in particular to a flow control method of a hypersonic air inlet with boundary layer guiding and sucking combined.
Background
Boundary layer suction is considered a very effective method of controlling the flow of the boundary layer of the air intake. Boundary layer suction is to provide suction holes in a wall surface, and to draw out a boundary layer air flow by a pressure difference between both ends of the holes, thereby thinning the boundary layer. The boundary layer suction comprises a suction hole, a suction cavity, a discharge system and the like, and the boundary layer enters the suction cavity through the suction hole in the working process and is discharged through the discharge system. The thickness of a boundary layer is greatly reduced by suction, and the quality of air flow captured by an air inlet channel is improved; on the other hand, because the boundary layer becomes thinner, the shock wave/boundary layer interference of the air inlet channel is weakened, and the starting performance of the air inlet channel is greatly improved. Boundary layer suction technology is widely applied to ultra/hypersonic inlet channels at present.
The boundary layer suction of the prior art comprises suction holes, suction chambers, bleed systems and the like, the structure is relatively complex, the weight of the aircraft is increased, and the suction chambers and the bleed systems also occupy the space of the aircraft head. On the other hand, aircraft fuselages are generally made of composite materials, and the structural strength of the materials is damaged by perforating the wall surface in a large area.
The application number for the hypersonic inlet channel is as follows: 201710784957.3, the method integrates the precursor and the Bump, and basically solves the problem of excessive flow loss of the Bump in the hypersonic airflow. The Bump/precursor integrated air inlet realizes the displacement of the boundary layer through the transverse pressure gradient generated on the Bump profile, and achieves better effect. However, the hypersonic velocity flow speed is too high, the transverse displacement distance is too long, and the designed Bump height is limited, so that the displacement capacity of the method on hypersonic velocity boundary layer airflow is weak.
Pneumatic research and experiments, 2010,28(4):1-6 optimization design for controlling suction flow based on numerical simulation air inlet duct suction [ J ], combined Bump air inlet duct and boundary layer suction technology, and studied the influence of different installation positions of suction grooves on the performance of the air inlet duct. In practical engineering designs, there are also solutions in which the suction holes are arranged on the underside of the Bump. Both of these schemes combine Bump with suction to further improve intake air flow capture quality. The displacement capacity of the boundary layer is obviously improved by the combined method of Bump and air inlet suction, but the current combined mode of the Bump and the air inlet suction also stays at the stage of mechanically combining the Bump and the air inlet suction. For the first method, the installation positions of different suction holes are researched in the Bump air inlet, only a better scheme is selected from a plurality of schemes, the flow characteristics of the Bump air inlet are not considered, and the method is blind. The scheme of arranging the suction holes on the bottom side of the Bump takes the boundary layer accumulated on the bottom side of the Bump into consideration, so that the boundary layer is sucked away through the suction holes, the scheme is suitable for a supersonic Bump air inlet channel which is more powerful in displacement of the boundary layer, and the Bump profile cannot always displace most of the boundary layer to the bottom side under the high supersonic speed condition, so that the suction holes cannot well exert the suction effect.
Disclosure of Invention
The invention provides a hypersonic air inlet flow control method combining boundary layer guiding and suction, which is used for overcoming the defects of weak boundary layer air flow displacement performance and the like in the prior art, and achieving the purposes of effectively playing the role of an air suction hole and improving the boundary layer air flow displacement performance.
In order to achieve the above object, the present invention provides a hypersonic air inlet boundary layer flow control method, which comprises the following steps:
step 1, forming a boundary layer guide profile on a precursor of a hypersonic air inlet channel, and directionally changing the flow direction of air flow flowing through the boundary layer guide profile through a transverse pressure gradient generated on the boundary layer guide profile;
step 2, guiding the boundary layer to the local part of the boundary layer guiding profile through a transverse pressure gradient acting on the boundary layer guiding profile;
and 3, arranging an air exhaust hole at a local part of the boundary layer guide molded surface.
The hypersonic air inlet channel flow control method provided by the invention has the advantages that the boundary layer guide profile generates a transverse pressure gradient through the careful design of the pressure distribution of the boundary layer guide profile, so that the boundary layer air flow is actively guided, the boundary layer is respectively guided to the local parts (such as the middle part and the bottom part which are easy to operate) of the boundary layer guide profile under the action of the transverse pressure gradient, a thicker boundary layer can be formed due to the concentration of the boundary layer guide profile, and the suction holes are arranged at the thicker part of the boundary layer, so that the targeted effective suction on the boundary layer is realized.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the structures shown in the drawings without creative efforts.
FIG. 1 is a schematic view of a mandibular axisymmetric inlet in a boundary layer guiding and suction combined hypersonic inlet flow control method according to an embodiment of the present invention;
FIG. 2 is a left side view of FIG. 1;
FIG. 3 is a top view of FIG. 1;
FIG. 4 shows a boundary layer guiding and suction combined hypersonic inlet flow control method according to an embodiment of the present inventionθThe curve is shown in the relation with theta;
fig. 5 shows a hypersonic inlet channel flow control method combining boundary layer guiding and suction according to an embodiment of the present invention, where x is equal to x0The boundary layer in the cross section is distributed along the spanwise direction and the suction distribution zone determines a schematic diagram.
The implementation, functional features and advantages of the objects of the present invention will be further explained with reference to the accompanying drawings.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that all the directional indicators (such as up, down, left, right, front, and rear … …) in the embodiment of the present invention are only used to explain the relative position relationship between the components, the movement situation, etc. in a specific posture (as shown in the drawing), and if the specific posture is changed, the directional indicator is changed accordingly.
In addition, the descriptions related to "first", "second", etc. in the present invention are only for descriptive purposes and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless specifically limited otherwise.
In the present invention, unless otherwise expressly stated or limited, the terms "connected," "secured," and the like are to be construed broadly, and for example, "secured" may be a fixed connection, a removable connection, or an integral part; the connection can be mechanical connection, electrical connection, physical connection or wireless communication connection; they may be directly connected or indirectly connected through intervening media, or they may be connected internally or in any other suitable relationship, unless expressly stated otherwise. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of those skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination of technical solutions should not be considered to exist, and is not within the protection scope of the present invention.
The invention provides a hypersonic air inlet channel flow control method combining boundary layer guiding and suction.
Example one
Referring to fig. 1 to 5, an embodiment of the present invention provides a hypersonic inlet flow control method combining boundary layer guiding and suction, including the following steps:
step 1, forming a boundary layer guide profile on a hypersonic air inlet channel precursor, so as to generate a transverse pressure gradient on the boundary layer guide profile, and directionally changing the flow direction of air flow flowing through the boundary layer guide profile through the transverse pressure gradient;
by means of the working principle of supersonic Bump, the flow direction of the airflow flowing through the boundary layer guiding profile is directionally changed, the front body profile generates a transverse pressure gradient, and accordingly active guiding is conducted on the boundary layer airflow, the guiding direction can be preset, can be any direction theoretically, and can be set or obtained through analysis according to actual design conditions. The boundary layer guide profile is a part of a hypersonic aircraft precursor and is used for directionally changing the flow direction of the airflow flowing on the boundary layer guide profile, wherein the directional change refers to the airflow facing to a preset direction; further influencing the thickness distribution of the boundary layer on the boundary layer guide molded surface, and adjusting the thickness of the boundary layer under the action of the boundary layer guide molded surface;
step 2, guiding the boundary layer to the local part of the boundary layer guiding profile through a transverse pressure gradient acting on the boundary layer guiding profile;
the local area can be any part theoretically, in the embodiment, the boundary layer is guided to the middle part and the bottom part of the boundary layer guiding profile under the action of the transverse pressure gradient through the careful design of the pressure distribution of the boundary layer guiding profile, and then the suction holes are respectively arranged in the two areas to realize the suction of the boundary layer. The arrangement is optimal in these two positions because it is generally arranged only at the bottom, but in the hypersonic condition it is difficult for the boundary layer to reach the bottom, where it is displaced in the middle and at the bottom, respectively. This is also an important aspect of the patent that distinguishes it from others.
The distribution of boundary layer thickness will vary under the influence of the boundary layer guide profile. According to the distribution characteristics of the boundary layer on the boundary layer guiding profile, the suction distribution zones are determined by setting a threshold value, and then suction holes are respectively arranged on each suction distribution zone.
And 3, arranging an air exhaust hole at a local part of the boundary layer guide molded surface. The thick boundary layer can be formed at the local of boundary layer guide profile, sets up the aspirating hole in the thick place of boundary layer, can reduce the quantity of aspirating hole, improves suction performance simultaneously.
Preferably, the step of forming a boundary layer guide profile in step 1 comprises:
step 11, defining the pressure on each section of the dense section by taking the variables theta and x as independent variables through a functional relation; theta is a spanwise variable, and x is a flow direction variable, i.e., an axial variable of the inlet duct. Parameters on a two-dimensional plane can be determined according to the two variables, and for an external rotation axis symmetric flow field, the tangent plane refers to a flow direction tangent plane forming an angle theta with the symmetric plane; referring to fig. 1, region 1 is the boundary layer guide profile; the boundary layer guide profile is a part of a hypersonic aircraft precursor and has the function of enabling the flow direction of airflow flowing through the boundary layer guide profile to be directionally changed, the plane where the plane 2 is located is a symmetrical plane, the plane where the plane 3 is located is a tangential plane which is theta with the symmetrical plane, the point 4 is a coordinate origin and is also a boundary layer guide profile starting point, the point 5 is a point where x is equal to L and is also a boundary layer guide profile terminating point, each tangential plane is defined by a variable theta, and for a planar two-dimensional flow field, the tangential plane is a flow direction tangential plane which is parallel to the symmetrical plane and has a distance of theta; for an internal/external axisymmetric flow field, the tangential section refers to a flow direction section forming an angle theta with the symmetric plane.
Step 12, obtaining a pressure distribution curve on the osculating plane of each spanwise scale factor according to the functional relation between the pressure on each osculating plane and the independent variable;
step 13, obtaining a flow field corresponding to the pressure distribution curve in each intimate plane based on a characteristic line method;
and step 14, performing curved surface lofting on the wall lines of the flow field in all the intimate planes, and taking the obtained profile as a boundary layer guide profile.
Preferably, the step 11 comprises:
defining the origin of coordinates as the intersection point 4 of the starting line of the boundary layer guide profile and the symmetrical plane, and defining the pressure on each section of the osculating plane by the following functional relation:
P(θ,x)＝kθy(x)+P0 (1)
wherein, P0The pressure value is a given value and is the origin of coordinates; the independent variable x is a horizontal coordinate and takes a value rangeEnclose as 0<x<L, wherein L is the abscissa of the boundary layer guide profile termination line, namely the abscissa of point 5; y (x) is a function of x, satisfying the following condition: monotonically increases with x; the function value is 0 at the origin of coordinates, namely y (0) is 0; the external rotation axis symmetric flow field of theta indicates that the included angle range of the tangent plane where the flow field is positioned and the symmetric plane is-thetac<θ<θcWherein thetacIs a given value; k is a radical ofθIs a spanwise scaling factor, which is a function of θ; because the hypersonic precursor is required to take on a portion of the compression task, the boundary layer guide profile flow field is increasingly pressurized in the flow direction (the boundary layer guide profile acts to compress the flow, so the wall pressure is increasingly pressurized in the flow direction).
The step 12 comprises:
by varying kθControlling each spanwise scaling factor to be kθThe pressure distribution curve in the intimate surface realizes the control of the pressure distribution in the span direction, forms a transverse pressure gradient on the guiding molded surface of the boundary layer and realizes the guiding of the boundary layer.
Preferably, the step 12 comprises:
kθmonotonically decreasing as θ increases, k is given by the second derivativeθThe curve is divided into the following three sections:
theta is located at 0, theta1]In the interval, kθ"(θ)>0, and kθ(0)＝1；
Theta is located at [ theta ]1,θ2]In the interval, kθ"(θ)<0；
Theta is located at [ theta ]2,θc]In the interval, kθ"(θ)>0, and kθ(θc)＝kminWherein k isminFor a given value, the range is 0<kmin<1；
kθThe first derivative of the curve is zero at θ ═ 0, θ ═ θ1And θ ═ θ2The first derivative continues.
Preferably, k is within each intervalθThe specific functional relationship comprises at least one specific analytic expression in the given polynomial, trigonometric, exponential and the like functional relationships.
As a preferred embodiment of the present invention: for example, given y (x) as a function of a parabolic form, where a can be given according to the compression requirement of the actual precursor on the gas flow, and the value range is generally [0.01,1 ].
P(θ,x)＝kθ(ax2)+P0 (2)
kθWhich is a spanwise scaling factor, is a function of theta. The invention changes kθThe control of the pressure distribution in the span direction is realized, so that a transverse pressure gradient is formed on the boundary layer guide profile, and the guidance of the boundary layer is further realized. Lower pair of kθThe functional relationship with theta is described. k is a radical ofθAlong the theta direction, the curve is monotonically decreased and can be divided into three sections according to the positive and negative of the second derivative. The first stage is [0, theta ]1]A section having kθ"(θ)>0, satisfies kθ(0) 1. The second stage is [ theta ]1,θ2]Interval, k in the intervalθ"(θ)<0. The third stage is [ theta ]2,θc]A section having kθ"(θ)>0, satisfies kθ(θc)＝kminWherein k isminFor a given value, the range is 0<kmin<1. Furthermore, kθThe first derivative of the curve is zero at θ ═ 0, θ ═ θ1And θ ═ θ2The first derivative is continuous to ensure the curve is smooth. Specific analytic expressions such as polynomial, trigonometric, exponential and the like can be given to the specific functional form in each interval.
One specific implementation case is as follows: setting a quadratic function distribution form in all three intervals, and taking theta1＝10，θ2＝25，θc＝60；θ1Corresponding kθTake 0.8, kmin0.6. Meanwhile, the curve meets all the requirements, and the number of equations is equal to the number of variables under the condition, so that quadratic function expressions in three intervals can be solved. (solving equations directly or converting into linear algebra can obtain 9 variables and 9 equations in total)
Within each interval kθThe values are dispersed along the theta direction, and the dispersion precision range is [1 degree ], 10 degree]. According to formula 1Solving to obtain each spanwise scaling factor as kθThe pressure profile in the osculating plane of (a).
And solving the pressure-controllable flow field of the osculating plane, and solving the flow field corresponding to the given pressure distribution curve in each osculating plane based on a characteristic line method. The method is a public technology in the field, and the solving process can refer to the application number as follows: 201710784957.3 and Beijing Gao education Press-2012 "gas dynamics [ M ].
In the concrete solving process, based on a spiral characteristic line method, a prediction-correction method is adopted to solve two types of unit processes of a downstream inner point of two known adjacent inner points, a known upstream wall surface point and a downstream wall surface point of the adjacent inner points. Then, the whole flow field and the wall molded line corresponding to the pressure distribution curve are solved according to a space stepping mode.
Thus, the wall surface corresponding to the pressure distribution curve in the tangent plane is obtained. The walls corresponding to the pressure profiles in all the osculating planes are then solved in this way.
And finally, performing curved surface lofting on the wall surface lines of the flow field in all the close surfaces to obtain a profile which is the boundary layer guide profile.
Preferably, the step 2 includes:
and step 21, performing numerical simulation on the boundary layer guide molded surface to obtain the boundary layer thickness distribution on the boundary layer guide molded surface. In the invention kθIs arranged such that the boundary layer is close to theta1<θ<θ2And θ ═ θcThe boundary layer thickness is formed spanwise to correspond to the first protrusion and the second protrusion.
The step 3 comprises the following steps:
step 31, determining a suction distribution zone through a threshold according to the distribution characteristics of the boundary layer on the boundary layer molded surface;
in step 32, suction holes are respectively arranged on each suction distribution belt.
Preferably, said step 31 comprises:
in step 311, when the boundary layer is not of the boundary layer guide type, x is x0The thickness of the boundary layer of the cross section is taken as T0；
Step 312, for the boundary layer first convex region, given the scaling factor r1, when the boundary layer thickness satisfies T>r1×T0While, recognizing the location on the first suction distribution belt;
for the boundary layer second raised region, given a scaling factor r2, when the boundary layer thickness at that location satisfies T>r2×T0While, recognizing the location on the second suction distribution belt;
r1 and r2 are given values, r1 and r2 are both greater than zero, and r1< r 2;
step 313, the identification of the position of the suction profile is performed for each section perpendicular to the x-direction according to step 312, obtaining a first suction profile and a second suction profile over the entire boundary layer guiding profile.
As a specific embodiment of the invention, a series of equidistant spanwise sections are arranged on the boundary layer guide profile along the x direction (axial direction of the air inlet channel), and the section pitch has a value range of [10,100 ].
The suction distribution strip position is identified in each spanwise section, where x is x0The cross section is described as an example. With no boundary layer leading profile, the same x as x under the condition of free development of the boundary layer0The thickness of the boundary layer in the cross section is taken as a reference thickness and is denoted as To。
For the boundary layer first raised area, a scaling factor r is given1When the boundary layer thickness satisfies T>r1·ToThe location is identified on the first suction distribution belt. For the boundary layer second raised area, a scaling factor r is given2When the boundary layer thickness at that position satisfies T>r2·ToThe location is identified on the second suction distribution belt. r is1And r2The given value is the value, and the value ranges of the given value and the given value are [0.5,2 ] according to design experience]Suitably, r is satisfied in general1<r2。
In this way, the suction profile over the entire boundary layer guiding profile can be obtained by identifying the position of the suction profile for each spanwise section.
Thereby, a first suction profile and a second suction profile are obtained on the boundary layer guiding profile.
One specific implementation case is as follows: giving x ═ x0The distribution of the boundary layer in the cross-section in the spanwise direction specifies the process of identifying the suction profile. The boundary layer thickness in the figure forms two distinct protrusions in the spanwise direction. Obtaining T in the section by numerical simulationo0.58. Get r10.9, r2The range of the suction profile of the first raised area is 1.04, which ultimately results in a range of [10.2,28.8 ]]The range of the suction distribution band of the second convex area is [48.0,58.6 ]]。
Preferably, the step 32 specifically includes:
step 321, obtaining the total number of the suction holes according to the given shape, the suction area ratio and the area of a single suction hole; the suction area ratio is the percentage of the total suction hole area to the area of the suction profile, and the total suction area is divided by the area of a single suction hole to obtain the total hole number.
The holes may be distributed in an equidistant array on the suction distribution belt, at step 322, and may of course be arranged according to actual requirements. The suction holes are distributed on each suction distribution belt according to the distribution form of the holes.
The arrangement mode of the suction holes is the common technology in the field, and the arrangement mode of the suction holes can be divided into two parts: determining the shape and the number of the suction holes; given the distribution pattern of the suction holes.
Given the shape of the hole first, the suction hole may be circular, triangular, rectangular, etc. Then, the total hole number is obtained by dividing the total suction area by the area of the single suction hole, given the suction area ratio, which is the percentage of the total suction hole area to the area of the suction profile, and the area of the single suction hole.
The distribution pattern of the holes can be in the form of an equidistant array on the suction distribution belt, and can of course be arranged according to the actual requirements.
The invention introduces the working principle of supersonic Bump into the design of hypersonic precursor, so that the profile of the precursor generates transverse pressure gradient, thereby actively guiding the boundary layer airflow, ensuring that the thickness of the boundary layer is distributed on the profile of the boundary layer guide to form two bulges, and then arranging suction holes at the bulges, thereby realizing the boundary layer flow control method which is suitable for the boundary layer guide matched suction of a hypersonic air inlet channel.
Compared with the traditional suction technology, under the same suction effect, the suction holes and the suction cavity adopted by the invention are fewer, which is beneficial to reducing the damage degree of the structural strength of the fuselage material and saving the space in the aircraft; compared with a Bump air inlet, the invention provides a boundary layer controllable mode of matching boundary layer guiding and suction aiming at the problem that the boundary layer displacement is relatively difficult in hypersonic flow, and the problem is basically solved; compared with the existing control mode of cooperation of Bump and suction, the boundary layer is guided by the special design of the front body profile, and then the suction is performed in a targeted manner, so that the function of the suction hole can be exerted to a greater extent. The scheme is proved to be feasible through numerical simulation.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention, and all modifications and equivalents of the present invention, which are made by the contents of the present specification and the accompanying drawings, or directly/indirectly applied to other related technical fields, are included in the scope of the present invention.

Claims (8)

1.一种边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，包括以下步骤：1. a hypersonic air inlet flow control method combining boundary layer guidance and suction, is characterized in that, comprises the following steps: 步骤1，在高超声速进气道前体上形成边界层引导型面，通过在边界层引导型面上产生的横向压力梯度，使流经边界层引导型面的气流的流动方向发生定向改变；Step 1: A boundary layer guiding profile is formed on the precursor of the hypersonic inlet port, and the flow direction of the airflow flowing through the boundary layer guiding profile is directionally changed by the transverse pressure gradient generated on the boundary layer guiding profile; 步骤2，通过作用在所述边界层引导型面上的横向压力梯度将边界层引导到所述边界层引导型面的局部；Step 2, guiding the boundary layer to a part of the boundary layer guiding profile through the lateral pressure gradient acting on the boundary layer guiding profile; 步骤3，在所述边界层引导型面的局部设置抽气孔；Step 3, setting air extraction holes in the part of the boundary layer guiding profile; 其中，所述步骤1中形成边界层引导型面的步骤包括：Wherein, the step of forming the boundary layer guiding profile in the step 1 includes: 步骤11，以变量x，θ为自变量通过函数关系定义每一个密切面上的压力；x为进气道轴向变量，对于外转轴对称流场，所述密切面指与对称面呈θ角度的流向切面；Step 11, using variables x and θ as independent variables to define the pressure on each close surface through a functional relationship; x is the axial variable of the intake port, and for the outer rotational axis symmetric flow field, the close surface refers to the symmetric plane at an angle of θ. the flow direction section; 步骤12，根据每个密切面上的压力与自变量的函数关系，获得每一个展向比例因子的密切面上的压力分布曲线；Step 12, according to the functional relationship between the pressure and the independent variable on each approximation surface, obtain the pressure distribution curve on the approximation surface of each spanwise scaling factor; 步骤13，在每个密切面内基于特征线法获得压力分布曲线对应的流场；Step 13, obtain the flow field corresponding to the pressure distribution curve based on the characteristic line method in each close surface; 步骤14，将所有密切面内的流场壁面线进行曲面放样，获得的型面作为边界层引导型面。Step 14: Perform surface lofting of the flow field wall lines in all the close planes, and the obtained profile is used as the boundary layer guiding profile. 2.如权利要求1所述的边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，所述步骤11包括：2. The hypersonic air inlet flow control method combining boundary layer guidance and suction according to claim 1, wherein the step 11 comprises: 定义坐标原点为边界层引导型面起始线与对称面的交点，通过下面的函数关系定义每个密切面上的压力：The origin of coordinates is defined as the intersection of the starting line of the boundary layer guiding profile and the symmetry plane, and the pressure on each close surface is defined by the following functional relationship: P(θ，x)＝kθy(x)+P0 (1)P(θ, x)=k θ y(x)+P 0 (1) 其中，P0为坐标原点的压力值，为给定值；自变量x为横坐标，取值范围为0＜x＜L，其中L为边界层引导型面终止线横坐标；y(x)为x的函数，满足以下条件：随x单调增加；在坐标原点函数值为0，即y(0)＝0；θ针对外转轴对称流场表示该流场所处密切面与对称面的夹角，范围为-θc＜θ＜θc，其中θc为给定值；kθ为展向比例因子，是θ的函数；Among them, P 0 is the pressure value of the coordinate origin, which is a given value; the independent variable x is the abscissa, and the value range is 0<x<L, where L is the abscissa of the boundary layer guiding profile termination line; y(x) is a function of x, which satisfies the following conditions: it increases monotonically with x; the function value at the coordinate origin is 0, that is, y(0)=0; , the range is -θ c <θ < θ c , where θ c is a given value; k θ is the spanwise scaling factor, which is a function of θ; 所述步骤12包括：The step 12 includes: 通过改变kθ，每一个展向比例因子为kθ的密切面内的压力分布曲线，实现对压力分布在展向的控制，在边界层引导型面形成横向压力梯度，实现对边界层的引导。By changing k θ , each pressure distribution curve in the close plane with the spanwise scaling factor k θ realizes the control of the pressure distribution in the spanwise direction, and forms a lateral pressure gradient on the boundary layer guiding profile to realize the guidance of the boundary layer. . 3.如权利要求2所述的边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，所述步骤12包括：3. The hypersonic air inlet flow control method combining boundary layer guidance and suction according to claim 2, wherein the step 12 comprises: kθ随着θ增加单调递减，按二阶导数将kθ曲线分为以下三段：k θ decreases monotonically as θ increases, and the k θ curve is divided into the following three segments according to the second derivative: θ位于[0，θ1]区间时，kθ″(θ)＞0，且kθ(0)＝1；When θ is in the interval [0, θ 1 ], k θ "(θ)>0, and k θ (0)=1; θ位于[θ1，θ2]区间时，kθ″(θ)＜0；When θ is in the interval [θ 1 , θ 2 ], k θ "(θ)<0; θ位于[θ2，θc]区间时，kθ″(θ)＞0，且kθ(θc)＝kmin，其中kmin为给定的值，范围为0＜kmin＜1；When θ is in the interval [θ 2 , θ c ], k θ ″(θ)>0, and k θ (θ c )=k min , where k min is a given value, and the range is 0<k min <1; kθ曲线在θ＝0处一阶导数为零，θ＝θ1和θ＝θ2处一阶导数连续。The first derivative of the kθ curve is zero at θ = 0 , and the first derivative is continuous at θ= θ1 and θ=θ2. 4.如权利要求3所述的边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，每个区间内kθ具体函数关系包括给定多项式型、三角函数型、指数型中至少一种具体解析式。4. The hypersonic inlet flow control method combining boundary layer guidance and suction according to claim 3, wherein the specific functional relationship of k θ in each interval includes a given polynomial type, a trigonometric function type, an exponential At least one specific analytical formula in the type. 5.如权利要求4所述的边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，所述步骤2中边界层引导到所述边界层引导型面的局部包括：5 . The hypersonic intake port flow control method combining boundary layer guidance and suction according to claim 4 , wherein in the step 2, the boundary layer is guided to the part of the boundary layer guidance profile comprising: 6 . 所述边界层厚度方向的变化集中在所述边界层引导型面沿展向的中部和底部；The change in the thickness direction of the boundary layer is concentrated in the middle and bottom of the boundary layer guiding profile along the span; 所述步骤3包括：The step 3 includes: 步骤31，根据边界层在边界层型面上的分布特点，通过阈值确定抽吸分布带；Step 31, according to the distribution characteristics of the boundary layer on the boundary layer profile, determine the suction distribution zone through a threshold; 步骤32，在每一抽吸分布带上分别设置抽吸孔。Step 32, respectively setting suction holes on each suction distribution belt. 6.如权利要求5所述的边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，所述步骤2包括：6. The hypersonic air inlet flow control method combining boundary layer guidance and suction according to claim 5, wherein the step 2 comprises: 步骤21，对边界层引导型面进行数值仿真，获得在靠近θ1＜θ＜θ2以及θ＝θc的区域，边界层厚度分布沿展向对应形成第一凸起和第二凸起。Step 21: Numerical simulation is performed on the boundary layer guiding profile, and it is obtained that in the region close to θ 1 <θ < θ 2 and θ=θ c , the thickness distribution of the boundary layer correspondingly forms the first protrusion and the second protrusion along the span. 7.如权利要求6所述的边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，所述步骤31包括：7. The hypersonic air inlet flow control method combining boundary layer guidance and suction according to claim 6, wherein the step 31 comprises: 步骤311，边界层没有加入边界层引导型的情况下，x＝x0时的垂直于x方向的截面的边界层厚度作为T0；Step 311 , in the case where the boundary layer guide type is not added to the boundary layer, the thickness of the boundary layer of the section perpendicular to the x direction when x=x 0 is taken as T 0 ; 步骤312，对于边界层第一凸起区域，给定比例因子r1，当边界层厚度满足T＞r1×T0时，认定该位置在第一抽吸分布带上；Step 312, for the first convex region of the boundary layer, given a scale factor r1, when the thickness of the boundary layer satisfies T>r1×T 0 , it is determined that the position is on the first suction distribution zone; 对于边界层第二凸起区域，给定比例因子r2，当该位置处的边界层厚度满足T＞r2×T0时，认定该位置在第二抽吸分布带上；For the second convex area of the boundary layer, given the scale factor r2, when the thickness of the boundary layer at the position satisfies T>r2×T 0 , it is determined that the position is on the second suction distribution zone; r1和r2为给定值，r1，r2均大于零，且r1＜r2；r1 and r2 are given values, both r1 and r2 are greater than zero, and r1 < r2; 步骤313，按照步骤312对每一个垂直于x方向的截面都进行抽吸分布带位置的确认，获得整个边界层引导型面上的第一抽吸分布带和第二抽吸分布带。Step 313 , according to step 312 , confirm the position of the suction distribution zone for each section perpendicular to the x-direction, and obtain the first suction distribution zone and the second suction distribution zone on the entire boundary layer guiding profile. 8.根据权利要求5～7任一项所述的边界层引导和抽吸结合的高超声速进气道流动控制方法，其特征在于，所述步骤32包括：8 . The hypersonic intake port flow control method combining boundary layer guidance and suction according to claim 5 , wherein the step 32 comprises: 步骤321，根据给定的形状、抽吸面积比和单个抽吸孔的面积，获得抽气孔的总数量；Step 321, according to the given shape, the suction area ratio and the area of a single suction hole, obtain the total number of suction holes; 步骤322，根据孔的分布形式在每一抽吸分布带上分别布设抽吸孔。Step 322, according to the distribution form of the holes, respectively arrange suction holes on each suction distribution belt.
CN201811603785.6A 2018-12-26 2018-12-26 A Hypersonic Inlet Flow Control Method Combining Boundary Layer Guidance and Suction Active CN109649666B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201811603785.6A CN109649666B (en) 2018-12-26 2018-12-26 A Hypersonic Inlet Flow Control Method Combining Boundary Layer Guidance and Suction

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201811603785.6A CN109649666B (en) 2018-12-26 2018-12-26 A Hypersonic Inlet Flow Control Method Combining Boundary Layer Guidance and Suction

Publications (2)

Publication Number Publication Date
CN109649666A CN109649666A (en) 2019-04-19
CN109649666B true CN109649666B (en) 2021-07-02

Family

ID=66116505

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201811603785.6A Active CN109649666B (en) 2018-12-26 2018-12-26 A Hypersonic Inlet Flow Control Method Combining Boundary Layer Guidance and Suction

Country Status (1)

Country Link
CN (1) CN109649666B (en)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110990955B (en) * 2019-12-12 2024-05-28 中国人民解放军国防科技大学 Hypersonic speed Bump air inlet channel design method and hypersonic speed Bump air inlet channel design system

Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4815279A (en) * 1985-09-27 1989-03-28 The United States Of America As Represented By The National Aeronautics And Space Administration Hybrid plume plasma rocket
RU2385820C1 (en) * 2009-05-04 2010-04-10 Государственное образовательное учреждение высшего профессионального образования Томский государственный университет (ТГУ) Cooling method of taper head end of flying machine
CN102562461A (en) * 2010-12-21 2012-07-11 通用电气公司 System and method of operating an active flow control system for manipulating a boundary layer across a rotor blade of a wind turbine
CN105956286A (en) * 2016-05-06 2016-09-21 北京航空航天大学 Prediction method of forecabin thermal protection system whole trajectory temperature boundary of hypersonic velocity aircraft
CN107590330A (en) * 2017-09-04 2018-01-16 中国人民解放军国防科技大学 Design method of two-dimensional pre-compressed precursor with boundary layer displacement
CN107628266A (en) * 2017-09-04 2018-01-26 中国人民解放军国防科技大学 Design method of axisymmetric pre-compression precursor with boundary layer displacement
US10071798B2 (en) * 2012-11-19 2018-09-11 The Regents Of The University Of California Hypersonic laminar flow control
US10118696B1 (en) * 2016-03-31 2018-11-06 Steven M. Hoffberg Steerable rotating projectile
CN108999704A (en) * 2018-08-17 2018-12-14 中国人民解放军国防科技大学 Hypersonic inlet starting method and starting device
CN109033525A (en) * 2018-06-27 2018-12-18 浙江大学 A kind of hypersonic transition prediction method based on simplified three equation transition models

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19820097C2 (en) * 1998-05-06 2003-02-13 Airbus Gmbh Arrangement for boundary layer suction and impact boundary layer control for an aircraft
US7648100B2 (en) * 2000-05-31 2010-01-19 Kevin Kremeyer Shock wave modification method and system
CN104890887B (en) * 2015-04-20 2016-01-13 南京航空航天大学 Supersonic and hypersonic inlets using pneumatic non-start control method
CN104975950B (en) * 2015-06-16 2017-09-29 南京航空航天大学 The binary hypersonic inlet of wall pressure distribution is specified to determine method
US10953979B2 (en) * 2015-11-11 2021-03-23 The Arizona Board Of Regents On Behalf Of The University Of Arizona Control of hypersonic boundary layer transition
US10252790B2 (en) * 2016-08-11 2019-04-09 General Electric Company Inlet assembly for an aircraft aft fan

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4815279A (en) * 1985-09-27 1989-03-28 The United States Of America As Represented By The National Aeronautics And Space Administration Hybrid plume plasma rocket
RU2385820C1 (en) * 2009-05-04 2010-04-10 Государственное образовательное учреждение высшего профессионального образования Томский государственный университет (ТГУ) Cooling method of taper head end of flying machine
CN102562461A (en) * 2010-12-21 2012-07-11 通用电气公司 System and method of operating an active flow control system for manipulating a boundary layer across a rotor blade of a wind turbine
US10071798B2 (en) * 2012-11-19 2018-09-11 The Regents Of The University Of California Hypersonic laminar flow control
US10118696B1 (en) * 2016-03-31 2018-11-06 Steven M. Hoffberg Steerable rotating projectile
CN105956286A (en) * 2016-05-06 2016-09-21 北京航空航天大学 Prediction method of forecabin thermal protection system whole trajectory temperature boundary of hypersonic velocity aircraft
CN107590330A (en) * 2017-09-04 2018-01-16 中国人民解放军国防科技大学 Design method of two-dimensional pre-compressed precursor with boundary layer displacement
CN107628266A (en) * 2017-09-04 2018-01-26 中国人民解放军国防科技大学 Design method of axisymmetric pre-compression precursor with boundary layer displacement
CN109033525A (en) * 2018-06-27 2018-12-18 浙江大学 A kind of hypersonic transition prediction method based on simplified three equation transition models
CN108999704A (en) * 2018-08-17 2018-12-14 中国人民解放军国防科技大学 Hypersonic inlet starting method and starting device

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
高超声速边界层转捩研究现状与发展趋势;陈坚强;《空气动力学学报》;20170630;全文 *

Also Published As

Publication number Publication date
CN109649666A (en) 2019-04-19

Similar Documents

Publication Publication Date Title
CN101813027B (en) Bump air inlet method for realizing integration of unequal-strength wave system with forebody
CN115048753B (en) Continuous transonic wind tunnel aerodynamic shape design method
CN110186688B (en) Suction type transonic plane cascade turbine test bench with hole-slot structure
CN103711753B (en) A kind of vortex generator structure suppressing Shock Wave boundary layer to be separated
CN112576546B (en) Optimization method of non-uniform-thickness airfoil axial flow blade
CN109367795A (en) Fuselage bilateral air inlet high-speed aircraft aerodynamic arrangement
CN106081069A (en) Limit the aerofoil assembly in fluid-actuated hole
CN111120401A (en) Multi-wing centrifugal ventilator blade design method based on NACA wing profile
CN104481700A (en) Combined dynamic axisymmetric variable geometry inlet, engine and air inlet control method
CN114802776B (en) Embedded air inlet based on swept step to displace the forebody boundary layer
CN109649666B (en) A Hypersonic Inlet Flow Control Method Combining Boundary Layer Guidance and Suction
CN110990955B (en) Hypersonic speed Bump air inlet channel design method and hypersonic speed Bump air inlet channel design system
CN108583847A (en) A kind of low reynolds number high power factor aerofoil profile suitable for long endurance unmanned aircraft
CN107933895B (en) Micro-blow structure and method for drag reduction and lift increase of supercritical airfoil
CN104908957A (en) Ridge type sweeping vortex generator and generation method
CN109850128B (en) Multistage blowing annular quantity lift-increasing device and aircraft
CN211650472U (en) Air deflector assembly, air conditioner indoor unit and air conditioner
CN113153815B (en) Supersonic adsorption type compressor blade based on multiple holes
CN205311919U (en) Controllable swell intake duct of lateral pressure gradient
CN107016199B (en) A Design Method of Shock-Free Boundary Layer Displacement Bulge
CN109376385A (en) Boundary layer suction inlet with controllable wall pressure gradient
CN105351248A (en) High-performance airfoil for fan
CN117902051B (en) Air inlet using embedded microchannel array and its design method
CN113062803A (en) Layered air inlet channel for separating boundary layer and modeling method thereof
EP2487371B1 (en) Wing structure and fairing device

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
GR01 Patent grant
GR01 Patent grant