CN113800001A - Design method of internal shrinkage hypersonic inlet channel integrated with forebody - Google Patents

Design method of internal shrinkage hypersonic inlet channel integrated with forebody Download PDF

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CN113800001A
CN113800001A CN202111164255.8A CN202111164255A CN113800001A CN 113800001 A CN113800001 A CN 113800001A CN 202111164255 A CN202111164255 A CN 202111164255A CN 113800001 A CN113800001 A CN 113800001A
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compression
flow field
lip
aircraft
intake port
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CN113800001B (en
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莫建伟
王玉峰
梁俊龙
李光熙
南向军
呼延霄
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Xian Aerospace Propulsion Institute
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    • 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
    • B64C30/00—Supersonic type aircraft
    • 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
    • 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
    • B64D2033/0253—Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes specially adapted for particular type of aircraft
    • B64D2033/026—Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes specially adapted for particular type of aircraft for supersonic or hypersonic aircraft

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Abstract

本发明涉及冲压发动机，具体涉及一种与前体一体化的内收缩高超声速进气道设计方法，用于解决目前针对前体较长、附面层较厚的高超声速飞行器，其前体与内收缩进气道的一体化设计制作，仍然未有完善解决方案的不足之处。该与前体一体化的内收缩高超声速进气道设计方法，包括如下步骤：步骤(1)：设计飞行器前体；步骤(2)：设计基准流场；步骤(3)：确定进气道进口型线；步骤(4)：确定进气道三维气动型面；步骤(5)：根据进气道型面与飞行器前体吻合关系，通过几何修型完成进气道与飞行器前体一体化设计。

Figure 202111164255

The invention relates to a ramjet, in particular to a design method for an inner-shrinking hypersonic air inlet integrated with a precursor, which is used to solve the problem that the current hypersonic aircraft with a longer precursor and a thicker boundary layer, the precursor and the The integrated design and production of the inwardly retracted air intake still has not perfected the insufficiency of the solution. The method for designing an inner-shrinking hypersonic air inlet integrated with a precursor includes the following steps: step (1): designing the aircraft precursor; step (2): designing a reference flow field; step (3): determining the air inlet Inlet profile; Step (4): Determine the three-dimensional aerodynamic profile of the intake port; Step (5): According to the matching relationship between the profile of the intake port and the aircraft precursor, complete the integration of the intake port and the aircraft precursor through geometric modification design.

Figure 202111164255

Description

Design method of internal shrinkage hypersonic inlet channel integrated with forebody
Technical Field
The invention relates to a ramjet engine, in particular to a design method of an inner contraction hypersonic inlet channel integrated with a forebody.
Background
In the hypersonic flight, because the aircraft forebody is strongly coupled with the air inlet flow field, the shape, the flow field uniformity and the boundary layer characteristics of the aircraft forebody have direct influence on the performance and the starting performance of the air inlet, and thus, the hypersonic aircraft mostly adopts the integrated design of an inner contraction air inlet and the forebody.
At present, the integrated design of an inner contraction air inlet and a front body is mainly divided into two types, one type is the inner contraction air inlet as a wave-multiplying front body, for example, a three-module and four-module Busemann air inlet is directly adopted by a 20 th century 60-year Johns Hopkins university SCRAM missile to form a circular section configuration, and a four-module inner contraction air inlet is also adopted by a scramjet engine developed in England in the United kingdom in the Hyshot plan; the other is that the forebody and the inner contraction inlet are directly integrated, such as the SR-72 hypersonic scout concept issued by Rockschid Martin company in the United states and the newly-introduced high-speed percussion weapon (HSSW) concept, which adopt the integration design of the projectile body and the inner contraction inlet, but detailed design details are not published yet; for the hypersonic flight vehicle with a long forebody and a thick boundary layer, the integral design of the forebody and the inner contraction air inlet channel is not disclosed, and a perfect solution is not provided because the former cannot be effectively applied and the latter is not disclosed.
Disclosure of Invention
The invention aims to solve the problem that the prior hypersonic aerocraft with a longer forebody and a thicker boundary layer has the defects that the integral design and manufacture of the forebody and an inner contraction air inlet channel still has no perfect solution, and provides a design method of the inner contraction hypersonic air inlet channel integrated with the forebody.
In order to solve the defects of the prior art, the invention provides the following technical solutions:
a design method of an inner contraction hypersonic inlet channel integrated with a forebody is characterized by comprising the following steps:
step (1): designing an aircraft forebody
The aircraft forebody is a revolution body or a revolution body-like structure, a generatrix of the aircraft forebody consists of a cubic curve and a flat straight line, and coefficients of the cubic curve are determined according to preset starting point coordinates, preset end point coordinates, a starting cone angle and a preset end point slope;
step (2): design of reference flow field
The reference flow field is of an axisymmetric structure and comprises an external compression flow field, an internal compression flow field and a lip reflection shock wave; the outer compression flow field comprises an outer compression conical surface and an outer compression shock wave surface positioned on the periphery of the outer compression conical surface, and an outer compression area is formed between the outer compression conical surface and the outer compression shock wave surface; the inner compression flow field comprises an inner compression conical surface and a lip cover compression surface, the lip cover compression surface is positioned at the periphery of the inner compression conical surface, the intersection point of the lip cover compression surface and the outer compression shock wave surface is a lip opening point B, an inner compression area is formed between the inner compression conical surface and the lip cover compression surface, and the lip cover compression surface generates an initial compression wave and a tail compression wave in the inner compression area; the lip reflected shock wave is positioned between the inner compression area and the outer compression area;
(2.1) obtaining flow field parameters of the aircraft forebody through numerical simulation calculation according to the inflow parameters of the aircraft forebody, and extracting the non-uniform flow field parameters at the inlet section of the air inlet channel as inflow parameters of a reference flow field;
(2.2) giving pressure distribution on the outer compression conical surface according to the inflow parameters of the reference flow field, so as to determine the outer compression conical surface and determine an outer compression shock wave surface;
(2.3) interpolating and determining a lip point B on the outer compression shock wave surface according to the capture radius of the reference flow field, and giving out pressure distribution on the lip cover compression surface so as to determine the lip cover compression surface, thereby determining a lip reflection shock wave, an initial compression wave and an end compression wave;
(2.4) determining an internal compression conical surface according to the incoming flow mass conservation principle;
and (3): determining inlet profile of inlet duct
The reference flow field and the aircraft forebody flow field are partially overlapped, the symmetry axes of the reference flow field and the aircraft forebody flow field are mutually parallel, and the inlet molded line of the air inlet channel comprises an intersecting line of the aircraft forebody and an external compression shock wave surface, a reference flow field lip molded line located at a lip point B, and two side straight lines forming a sector with the intersecting line and the reference flow field lip molded line;
and (4): determining three-dimensional aerodynamic profile of air inlet duct
Dispersing inlet molded lines of an air inlet, and tracking flow lines in a reference flow field to obtain a three-dimensional pneumatic molded surface of the air inlet, wherein the three-dimensional pneumatic molded surface of the air inlet comprises an outer compression surface of the air inlet and a lip cover compression surface of the air inlet;
and (5): and according to the matching relation between the profile of the air inlet and the front body of the aircraft, the integrated design of the air inlet and the front body of the aircraft is completed through geometric modification.
Further, in step (4), the specific step of determining the three-dimensional aerodynamic profile of the air inlet channel includes: dispersing the inlet profile of the air inlet, and slicing the sector formed by the inlet profile of the air inlet by taking the symmetry axis of the reference flow field as the center; taking the intersection point of each slice and the intersecting line as a starting point, and carrying out streamline tracing in a reference flow field to obtain an external compression surface of the air inlet; and tracking the flow line in the reference flow field by taking the intersection point of each slice and the lip-shaped line of the reference flow field as a starting point to obtain the compression surface of the inlet lip cover.
Further, in the step (5), the concrete step of completing the integrated design of the air inlet and the aircraft forebody through geometric modification according to the matching relationship between the profile of the air inlet and the aircraft forebody comprises: the smooth transition between the outer compression surface of the air inlet channel and the aircraft precursor is realized through geometric model modification on two sides, and the compression surface of the lip cover of the air inlet channel and the outlet of the air inlet channel form an air inlet channel throat diffusion section through geometric model modification smooth transition, so that the integrated design of the air inlet channel and the aircraft precursor is realized.
Compared with the prior art, the invention has the beneficial effects that:
the invention provides a design method of an internal and external cone mixed compression reference flow field based on a Bump compression surface, realizes the integrated design of a three-dimensional internal contraction air inlet channel and an aircraft forebody, realizes the automatic drainage overflow of a forebody incoming flow boundary layer on the external compression surface of the air inlet channel through a transverse pressure gradient, improves the starting capacity of the air inlet channel, widens the working range of the air inlet channel, does not need additional systems such as boundary layer channels, air discharging devices and the like, reduces the aerodynamic resistance of the aircraft, and has the remarkable characteristics of high integration degree, good performance, high volume ratio, simple structure and the like.
Drawings
FIG. 1 is a schematic structural diagram of an aircraft precursor;
FIG. 2 is a schematic diagram of a reference flow field;
FIG. 3 is a schematic view of a configuration in which the aircraft forebody partially coincides with the reference flow field;
FIG. 4 is a schematic view of inlet profile of the inlet duct;
FIG. 5 is a schematic view of an integrated internal contracting hypersonic inlet with a forebody.
The reference numerals are explained below: 1-an aircraft precursor; 2-reference flow field, 21-outer compression flow field, 211-outer compression conical surface, 212-outer compression shock wave surface, 213-outer compression area, 22-inner compression flow field, 221-inner compression conical surface, 222-lip cover compression surface, 223-inner compression area, 224-initial compression wave, 225-end compression wave, 23-lip reflection shock wave; 3-intersecting line; 4-a reference flow field lip molded line; 5-an external compression surface of the air inlet; 6-inlet lip cover compression surface; 7-boundary layer; 8-outlet of air inlet channel.
Detailed Description
The invention will be further described with reference to the drawings and exemplary embodiments.
A design method of an inner contraction hypersonic inlet channel integrated with a precursor comprises the following steps:
step (1): designing an aircraft forebody 1
Referring to fig. 1, the aircraft forebody 1 is a revolution body or revolution body-like structure, the head half-cone angle of the aircraft forebody is 15 degrees, the generatrix of the aircraft forebody is composed of a cubic curve and a flat straight line, and the coefficient of the cubic curve is determined according to the preset coordinates of a starting point and an end point, the slope of the starting cone angle and the slope of the end point;
step (2): design reference flow field 2
Referring to fig. 2, the reference flow field 2 is an axisymmetric structure, and includes an external compression flow field 21, an internal compression flow field 22, and a lip reflection shock wave 23; the outer compression flow field 21 comprises an outer compression conical surface 211 and an outer compression shock wave surface 212 positioned at the periphery of the outer compression conical surface 211, and an outer compression area 213 is formed between the outer compression conical surface 211 and the outer compression shock wave surface 212; the inner compression flow field 22 comprises an inner compression conical surface 221 and a lip cover compression surface 222, the lip cover compression surface 222 is positioned on the periphery of the inner compression conical surface 221, the intersection point of the lip cover compression surface 222 and the outer compression shock wave surface 212 is a lip opening point B, an inner compression area 223 is formed between the inner compression conical surface 221 and the lip cover compression surface 222, and the lip cover compression surface 222 generates an initial compression wave 224 and a tail compression wave 225 in the inner compression area 223; the lip reflected shock wave 23 is located between the inner compression zone 223 and the outer compression zone 213;
(2.1) obtaining flow field parameters of the aircraft forebody 1 through numerical simulation calculation according to the inflow parameters of the aircraft forebody 1, and extracting the non-uniform flow field parameters at the inlet section of the air inlet channel as inflow parameters of the reference flow field 2;
(2.2) according to the inflow parameters of the reference flow field 2, giving pressure distribution on the outer compression conical surface 211, so as to determine the outer compression conical surface 211, and thus determine the outer compression shock wave surface 212;
(2.3) interpolating and determining a lip point B on the outer compression shock wave surface 212 according to the capture radius of the reference flow field 2, giving a pressure distribution on the lip cover compression surface 222, thereby determining the lip cover compression surface 222, and thereby determining the lip reflected shock wave 23, the initial compression wave 224 and the trailing compression wave 225;
(2.4) determining the inner compression conical surface 221 according to the incoming flow mass conservation principle;
and (3): determining inlet profile of inlet duct
Referring to fig. 3 and 4, the reference flow field 2 and the flow field of the aircraft forebody 1 are partially overlapped, the symmetry axes of the reference flow field and the flow field of the aircraft forebody 1 are parallel to each other, and the inlet molded line of the air inlet channel comprises an intersecting line 3 of the aircraft forebody 1 and the external compression shock wave surface 212, a reference flow field lip molded line 4 located at a lip point B, and two side straight lines ac and nh which form a sector with the intersecting line 3 and the reference flow field lip molded line 4;
and (4): determining three-dimensional aerodynamic profile of air inlet duct
With reference to fig. 4 and 5, dispersing the inlet profile of the air inlet, and slicing the sector formed by the inlet profile of the air inlet by taking the symmetry axis of the reference flow field 2 as the center; taking the intersection point of each slice and the intersecting line 3 as a starting point, carrying out streamline tracing in the reference flow field 2, and circumferentially distributing the streamlines to obtain an external compression surface 5 of the air inlet; taking the intersection point of each slice and the reference flow field lip molded line 4 as a starting point, tracking flow lines in the reference flow field 2, and circumferentially distributing the flow lines to obtain an air inlet lip cover compression surface 6; after supersonic air flow passes through the outer compression surface 5 of the air inlet, the incoming flow boundary layer 7 is discharged to two sides on the outer compression surface 5 of the air inlet due to the action of transverse pressure gradient, the low-energy flow ratio entering the air inlet is reduced, and the performance of the air inlet is improved;
and (5): referring to fig. 5, smooth transition between the two sides of the external compression surface 5 of the inlet channel and the aircraft forebody 1 is realized through geometric modification, and the inlet channel lip compression surface 6 and the inlet channel outlet 8 form an inlet channel throat diffusion section through geometric modification smooth transition, so that the integrated design of the inlet channel and the aircraft forebody 1 is realized.
The above embodiments are only used for illustrating the technical solutions of the present invention, and not for limiting the same, and it is obvious for a person skilled in the art to modify the specific technical solutions described in the foregoing embodiments or to substitute part of the technical features, and these modifications or substitutions do not make the essence of the corresponding technical solutions depart from the scope of the technical solutions protected by the present invention.

Claims (3)

1.一种与前体一体化的内收缩高超声速进气道设计方法，其特征在于，包括如下步骤：1. an in-contraction hypersonic air inlet design method integrated with a precursor, is characterized in that, comprises the steps: 步骤(1)：设计飞行器前体(1)Step (1): Design the Aircraft Front Body (1) 所述飞行器前体(1)为旋成体或类旋成体结构，其母线由三次曲线和平直线组成，所述三次曲线系数依据预设的起点、终点坐标，起始锥角和终点斜率确定；The aircraft precursor (1) is a gyratory or quasi-rotational structure, and its generatrix is composed of a cubic curve and a straight line, and the cubic curve coefficient is determined according to the preset coordinates of the starting point, the ending point, the starting cone angle and the ending point slope; 步骤(2)：设计基准流场(2)Step (2): Design basis flow field (2) 所述基准流场(2)为轴对称结构，包括外压缩流场(21)、内压缩流场(22)和唇口反射激波(23)；所述外压缩流场(21)包括外压缩锥面(211)，以及位于外压缩锥面(211)外围的外压缩激波面(212)，所述外压缩锥面(211)与外压缩激波面(212)之间构成外压缩区(213)；所述内压缩流场(22)包括内压缩锥面(221)和唇罩压缩面(222)，所述唇罩压缩面(222)位于内压缩锥面(221)外围，唇罩压缩面(222)与外压缩激波面(212)的交点为唇口点B，所述内压缩锥面(221)与唇罩压缩面(222)之间构成内压缩区(223)，所述唇罩压缩面(222)在内压缩区(223)产生初始压缩波(224)和结尾压缩波(225)；所述唇口反射激波(23)位于内压缩区(223)与外压缩区(213)之间；The reference flow field (2) is an axisymmetric structure, including an outer compression flow field (21), an inner compression flow field (22) and a lip reflection shock wave (23); the outer compression flow field (21) includes an outer compression flow field (21); A compression cone (211), and an outer compression shock surface (212) located at the periphery of the outer compression cone (211), an outer compression zone (212) is formed between the outer compression cone (211) and the outer compression shock surface (212). 213); the inner compression flow field (22) includes an inner compression cone surface (221) and a lip cover compression surface (222), the lip cover compression surface (222) is located at the periphery of the inner compression cone surface (221), and the lip cover compression surface (222) The intersection point of the compression surface (222) and the outer compression shock surface (212) is the lip point B, and an inner compression zone (223) is formed between the inner compression cone surface (221) and the lip cover compression surface (222). The lip mask compression surface (222) generates an initial compression wave (224) and an ending compression wave (225) in the inner compression zone (223); the lip reflected shock wave (23) is located in the inner compression zone (223) and the outer compression zone (213); (2.1)根据飞行器前体(1)来流参数，通过数值仿真计算得到飞行器前体(1)流场参数，提取进气道入口截面处非均匀流场参数作为基准流场(2)来流参数；(2.1) According to the inflow parameters of the aircraft precursor (1), the flow field parameters of the aircraft precursor (1) are obtained by numerical simulation, and the non-uniform flow field parameters at the inlet section of the intake port are extracted as the reference flow field (2) Incoming flow parameter; (2.2)根据基准流场(2)来流参数，给出外压缩锥面(211)上的压力分布，从而确定外压缩锥面(211)，由此确定外压缩激波面(212)；(2.2) According to the flow parameters of the reference flow field (2), the pressure distribution on the outer compression cone (211) is given, thereby determining the outer compression cone (211), thereby determining the outer compression shock surface (212); (2.3)根据基准流场(2)的捕获半径在外压缩激波面(212)上插值确定唇口点B，给出唇罩压缩面(222)上的压力分布，从而确定唇罩压缩面(222)，由此确定唇口反射激波(23)、初始压缩波(224)和结尾压缩波(225)；(2.3) According to the capture radius of the reference flow field (2), the lip point B is determined by interpolation on the outer compression shock surface (212), and the pressure distribution on the lip compression surface (222) is given, thereby determining the lip compression surface (222) ), thereby determining the lip reflection shock wave (23), the initial compression wave (224) and the ending compression wave (225); (2.4)根据来流质量守恒原理确定内压缩锥面(221)；(2.4) Determine the inner compression cone surface (221) according to the principle of mass conservation of incoming flow; 步骤(3)：确定进气道进口型线Step (3): Determine the inlet profile of the intake port 将基准流场(2)与飞行器前体(1)流场部分重合，并且二者对称轴相互平行，所述进气道进口型线包括飞行器前体(1)与外压缩激波面(212)的相贯线(3)、位于唇口点B的基准流场唇口型线(4)，以及与相贯线(3)和基准流场唇口型线(4)形成扇区的两侧直线；The reference flow field (2) is partially coincident with the flow field of the aircraft precursor (1), and the axes of symmetry of the two are parallel to each other, and the inlet profile of the air inlet includes the aircraft precursor (1) and the outer compression shock surface (212) The intersection line (3) of , the reference flow field lip shape line (4) at the lip point B, and the two sides of the sector formed with the intersection line (3) and the reference flow field lip shape line (4) straight line; 步骤(4)：确定进气道三维气动型面Step (4): Determine the three-dimensional aerodynamic profile of the intake port 对进气道进口型线离散，在基准流场(2)中进行流线追踪，得到进气道三维气动型面，所述进气道三维气动型面包括进气道外压缩面(5)和进气道唇罩压缩面(6)；The inlet profile of the intake port is discretized, and streamline tracing is performed in the reference flow field (2) to obtain a three-dimensional aerodynamic profile of the intake port. The three-dimensional aerodynamic profile of the intake port includes the outer compression surface of the intake port (5) and the Intake port lip cover compression surface (6); 步骤(5)：根据进气道型面与飞行器前体(1)吻合关系，通过几何修型完成进气道与飞行器前体(1)一体化设计。Step (5): According to the matching relationship between the profile of the air inlet and the front body (1) of the aircraft, the integrated design of the air inlet and the front body (1) of the aircraft is completed through geometric modification. 2.根据权利要求1所述的一种与前体一体化的内收缩高超声速进气道设计方法，其特征在于：2. a kind of inner shrinking hypersonic air inlet design method integrated with precursor according to claim 1, is characterized in that: 步骤(4)中，所述确定进气道三维气动型面的具体步骤包括：将进气道进口型线进行离散，以基准流场(2)对称轴为中心，对进气道进口型线形成的扇区进行切片；以每个切片与相贯线(3)的交点为起点，在基准流场(2)中进行流线追踪，得到进气道外压缩面(5)；以每个切片与基准流场唇口型线(4)的交点为起点，在基准流场(2)中进行流线追踪，得到进气道唇罩压缩面(6)。In step (4), the specific step of determining the three-dimensional aerodynamic profile of the intake port includes: discretizing the inlet profile line of the intake port, and taking the symmetry axis of the reference flow field (2) as the center, to the inlet port profile line. Slice the formed sector; take the intersection of each slice and the intersecting line (3) as the starting point, carry out streamline tracing in the reference flow field (2), and obtain the outer compression surface of the intake port (5); take each slice as the starting point. The intersection with the reference flow field lip profile line (4) is taken as the starting point, and streamline tracing is performed in the reference flow field (2) to obtain the intake port lip cover compression surface (6). 3.根据权利要求1或2所述的一种与前体一体化的内收缩高超声速进气道设计方法，其特征在于：3. a kind of inner shrinking hypersonic air inlet design method integrated with precursor according to claim 1 and 2, is characterized in that: 步骤(5)中，所述根据进气道型面与飞行器前体(1)吻合关系，通过几何修型完成进气道与飞行器前体(1)一体化设计的具体步骤包括：所述进气道外压缩面(5)两侧通过几何修型实现和飞行器前体(1)的光滑过渡，所述进气道唇罩压缩面(6)与进气道出口(8)通过几何修型光滑过渡形成进气道喉部扩压段，从而实现进气道与飞行器前体(1)一体化设计。In step (5), the specific steps of completing the integrated design of the air inlet and the aircraft precursor (1) through geometric modification according to the matching relationship between the profile of the air intake and the aircraft precursor (1) include: The two sides of the outer compression surface (5) of the airway are geometrically modified to achieve a smooth transition with the aircraft front body (1), and the compression surface (6) of the air intake lip cover and the air intake outlet (8) are smoothed by geometrical modification The diffuser section at the throat of the intake port is formed by transition, thereby realizing the integrated design of the intake port and the aircraft front body (1).
CN202111164255.8A 2021-09-30 2021-09-30 Design method of inner-shrinkage hypersonic air inlet channel integrated with precursor Active CN113800001B (en)

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