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 PDFInfo
- Publication number
- 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
- Authority
- CN
- China
- Prior art keywords
- compression
- flow field
- air inlet
- lip
- forebody
- 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.)
- Granted
Links
- 238000000034 method Methods 0.000 title claims abstract description 13
- 230000008602 contraction Effects 0.000 claims abstract description 17
- 230000004048 modification Effects 0.000 claims abstract description 12
- 238000012986 modification Methods 0.000 claims abstract description 12
- 239000002243 precursor Substances 0.000 claims abstract description 9
- 230000006835 compression Effects 0.000 claims description 138
- 238000007906 compression Methods 0.000 claims description 138
- 230000035939 shock Effects 0.000 claims description 29
- 230000007704 transition Effects 0.000 claims description 6
- 238000004364 calculation method Methods 0.000 claims description 3
- 238000009792 diffusion process Methods 0.000 claims description 3
- 238000004088 simulation Methods 0.000 claims description 3
- 230000007547 defect Effects 0.000 abstract description 3
- 238000004519 manufacturing process Methods 0.000 abstract description 2
- 238000010586 diagram Methods 0.000 description 2
- 230000010354 integration Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000007599 discharging Methods 0.000 description 1
- 238000009527 percussion Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Images
Classifications
-
- 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—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/02—Arrangements 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—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
- B64D2033/0253—Arrangements 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—Arrangements 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
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Manufacturing & Machinery (AREA)
- Transportation (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention relates to a ramjet, in particular to a design method of an inner contraction hypersonic inlet passage integrated with a forebody, which is used for solving the defect that the prior hypersonic aerocraft with a longer forebody and a thicker boundary layer is not provided with a perfect solution due to the integrated design and manufacture of the forebody and the inner contraction hypersonic inlet passage. The design method of the inner contraction hypersonic inlet channel integrated with the precursor comprises the following steps: step (1): designing an aircraft precursor; step (2): designing a reference flow field; and (3): determining the inlet profile of the air inlet channel; and (4): determining a three-dimensional pneumatic profile of an 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.
Description
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. A method for designing an inner contraction hypersonic inlet channel integrated with a precursor is characterized by comprising the following steps:
step (1): designing an aircraft forebody (1)
The aircraft forebody (1) is a revolution body or a revolution body-like structure, a generatrix of the aircraft forebody (1) 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 standard flow field (2)
The reference flow field (2) is of an axisymmetric structure and comprises 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) comprises an outer compression conical surface (211) and an outer compression shock wave surface (212) positioned on 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 located 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 trailing compression wave (225) in the inner compression area (223); the lip-reflecting 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 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 further determine an 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 a lip mask compression surface (222), and determining the lip mask compression surface (222), thereby determining a lip reflection shock wave (23), an initial compression wave (224) and a final compression wave (225);
(2.4) determining an inner compression conical surface (221) according to the incoming flow mass conservation principle;
and (3): determining inlet profile of inlet duct
The reference flow field (2) is partially overlapped with the flow field of the aircraft forebody (1), the symmetry axes of the reference flow field and the flow field of the aircraft forebody are parallel to each other, and the inlet molded line of the air inlet comprises an intersecting line (3) of the aircraft forebody (1) and an outer compression shock wave surface (212), a reference flow field lip molded line (4) located at a lip point B, and two side straight lines forming 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
Dispersing inlet molded lines of an air inlet, and tracking flow lines in a reference flow field (2) 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 (5) of the air inlet and a lip cover compression surface (6) of the air inlet;
and (5): and according to the matching relation 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. The method of claim 1, wherein the method further comprises the step of:
in step (4), the specific step of determining the three-dimensional aerodynamic profile of the air inlet channel includes: dispersing inlet molded lines of the air inlet channel, and slicing sectors formed by the inlet molded lines of the air inlet channel by taking a symmetry axis of the reference flow field (2) as a 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) to obtain an external compression surface (5) of the air inlet; and tracking the flow line in the reference flow field (2) by taking the intersection point of each slice and the lip molded line (4) of the reference flow field as a starting point to obtain a compression surface (6) of the lip cover of the air inlet channel.
3. A method of designing an internal-contraction hypersonic inlet duct integrated with a precursor according to claim 1 or 2, characterized in that:
in the step (5), the concrete steps of completing the integrated design of the air inlet and the aircraft forebody (1) through geometric modification according to the matching relation of the profile of the air inlet and the aircraft forebody (1) comprise: the smooth transition of the outer compression face (5) both sides of intake duct and the forebody (1) of aircraft is realized through the geometric model modification, intake duct lip cover compression face (6) and intake duct export (8) form intake duct throat diffusion section through the smooth transition of geometric model modification to realize intake duct and forebody (1) of aircraft integrated design.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111164255.8A CN113800001B (en) | 2021-09-30 | 2021-09-30 | Design method of inner-shrinkage hypersonic air inlet channel integrated with precursor |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111164255.8A CN113800001B (en) | 2021-09-30 | 2021-09-30 | Design method of inner-shrinkage hypersonic air inlet channel integrated with precursor |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113800001A true CN113800001A (en) | 2021-12-17 |
| CN113800001B CN113800001B (en) | 2024-02-27 |
Family
ID=78897299
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111164255.8A Active CN113800001B (en) | 2021-09-30 | 2021-09-30 | Design method of inner-shrinkage hypersonic air inlet channel integrated with precursor |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113800001B (en) |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115371933A (en) * | 2022-10-24 | 2022-11-22 | 中国航发四川燃气涡轮研究院 | Method for testing aerodynamic coupling between air inlet channel and aircraft forebody |
| CN116341106A (en) * | 2023-03-14 | 2023-06-27 | 南京航空航天大学 | Strong-expansion-direction pressure gradient compression surface design method based on flow field similarity transformation |
| CN117048839A (en) * | 2023-07-21 | 2023-11-14 | 南京理工大学 | Hypersonic integrated air inlet channel design method for giving shape of precursor |
| CN118289204A (en) * | 2024-03-29 | 2024-07-05 | 南京航空航天大学 | Plane and conical shock wave based multiplication-wave front/compression surface integrated configuration and design method thereof |
Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2008045108A2 (en) * | 2005-12-15 | 2008-04-17 | Gulfstream Aerospace Corporation | Isentropic compression inlet for supersonic aircraft |
| CN102996253A (en) * | 2012-12-31 | 2013-03-27 | 中国人民解放军国防科学技术大学 | Supersonic air intake duct and wall face determination method of supersonic air intake duct |
| CN104908975A (en) * | 2015-05-04 | 2015-09-16 | 厦门大学 | Aircraft fore-body and internal waverider-derived hypersonic inlet integrated design method |
| CN104975950A (en) * | 2015-06-16 | 2015-10-14 | 南京航空航天大学 | Method for determining binary hypersonic inlet passage based on appointed wall pressure distribution |
| CN105947230A (en) * | 2016-05-24 | 2016-09-21 | 中国人民解放军63820部队吸气式高超声速技术研究中心 | Design method for wave rider and air inlet duct integrated configuration |
| CN110210096A (en) * | 2019-05-24 | 2019-09-06 | 南昌航空大学 | The variable cross-section three-dimensional contract Design of Inlet method of the bent cone bomb body of matching |
| CN110304267A (en) * | 2019-07-19 | 2019-10-08 | 中国人民解放军国防科技大学 | Hypersonic aircraft design method and system |
| CN110450963A (en) * | 2019-08-28 | 2019-11-15 | 中国人民解放军国防科技大学 | Hypersonic aircraft body and inward turning type air inlet channel integrated design method and system |
| CN111348169A (en) * | 2020-04-27 | 2020-06-30 | 南昌航空大学 | Integrated design method for circumferential four-inlet-channel layout of conical aircraft forebody |
| DE102020117768A1 (en) * | 2019-07-23 | 2021-01-28 | Deutsches Zentrum für Luft- und Raumfahrt e.V. | Supersonic inlet with contour bulge |
| US20210200917A1 (en) * | 2019-12-31 | 2021-07-01 | Southwest University Of Science And Technology | Basic flow-field of double straight conical shock waves with controllable downstream flow-field parameters and design method thereof |
-
2021
- 2021-09-30 CN CN202111164255.8A patent/CN113800001B/en active Active
Patent Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2008045108A2 (en) * | 2005-12-15 | 2008-04-17 | Gulfstream Aerospace Corporation | Isentropic compression inlet for supersonic aircraft |
| CN102996253A (en) * | 2012-12-31 | 2013-03-27 | 中国人民解放军国防科学技术大学 | Supersonic air intake duct and wall face determination method of supersonic air intake duct |
| CN104908975A (en) * | 2015-05-04 | 2015-09-16 | 厦门大学 | Aircraft fore-body and internal waverider-derived hypersonic inlet integrated design method |
| CN104975950A (en) * | 2015-06-16 | 2015-10-14 | 南京航空航天大学 | Method for determining binary hypersonic inlet passage based on appointed wall pressure distribution |
| CN105947230A (en) * | 2016-05-24 | 2016-09-21 | 中国人民解放军63820部队吸气式高超声速技术研究中心 | Design method for wave rider and air inlet duct integrated configuration |
| CN110210096A (en) * | 2019-05-24 | 2019-09-06 | 南昌航空大学 | The variable cross-section three-dimensional contract Design of Inlet method of the bent cone bomb body of matching |
| CN110304267A (en) * | 2019-07-19 | 2019-10-08 | 中国人民解放军国防科技大学 | Hypersonic aircraft design method and system |
| DE102020117768A1 (en) * | 2019-07-23 | 2021-01-28 | Deutsches Zentrum für Luft- und Raumfahrt e.V. | Supersonic inlet with contour bulge |
| CN110450963A (en) * | 2019-08-28 | 2019-11-15 | 中国人民解放军国防科技大学 | Hypersonic aircraft body and inward turning type air inlet channel integrated design method and system |
| US20210200917A1 (en) * | 2019-12-31 | 2021-07-01 | Southwest University Of Science And Technology | Basic flow-field of double straight conical shock waves with controllable downstream flow-field parameters and design method thereof |
| CN111348169A (en) * | 2020-04-27 | 2020-06-30 | 南昌航空大学 | Integrated design method for circumferential four-inlet-channel layout of conical aircraft forebody |
Non-Patent Citations (1)
| Title |
|---|
| 李怡庆，周驯黄，朱呈祥，尤延铖: "《曲锥前体／三维内转进气道一体化设计与分析》", 航空动力学报, vol. 33, no. 1, pages 87 - 96 * |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115371933A (en) * | 2022-10-24 | 2022-11-22 | 中国航发四川燃气涡轮研究院 | Method for testing aerodynamic coupling between air inlet channel and aircraft forebody |
| CN115371933B (en) * | 2022-10-24 | 2023-03-24 | 中国航发四川燃气涡轮研究院 | Method for testing aerodynamic coupling between air inlet channel and aircraft forebody |
| CN116341106A (en) * | 2023-03-14 | 2023-06-27 | 南京航空航天大学 | Strong-expansion-direction pressure gradient compression surface design method based on flow field similarity transformation |
| CN116341106B (en) * | 2023-03-14 | 2024-06-07 | 南京航空航天大学 | Strong-expansion-direction pressure gradient compression surface design method based on flow field similarity transformation |
| CN117048839A (en) * | 2023-07-21 | 2023-11-14 | 南京理工大学 | Hypersonic integrated air inlet channel design method for giving shape of precursor |
| CN118289204A (en) * | 2024-03-29 | 2024-07-05 | 南京航空航天大学 | Plane and conical shock wave based multiplication-wave front/compression surface integrated configuration and design method thereof |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113800001B (en) | 2024-02-27 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN113800001A (en) | Design method of internal shrinkage hypersonic inlet channel integrated with forebody | |
| CN101392685B (en) | Internal waverider hypersonic inlet and design method based on random shock form | |
| CN109927917B (en) | Integrated design method for internal rotation type wave-rider forebody air inlet channel of supersonic aircraft | |
| CN104210672B (en) | Hypersonic rider fuselage and inlet channel integrated design method | |
| CN112340014B (en) | Inner-outer flow decoupling double-waverider high-speed air suction type aircraft and generation method thereof | |
| CN101813027B (en) | Bump air inlet method for realizing integration of unequal-strength wave system with forebody | |
| CN104908975B (en) | Aircraft fore-body and internal waverider-derived hypersonic inlet integrated design method | |
| CN113153529B (en) | Wide-speed-range air inlet channel design method based on double-incidence bending shock waves | |
| CN107089341B (en) | With aircraft integrated hypersonic inlet external compression face design method | |
| CN106741976B (en) | A kind of mimetic design method of waverider forebody derived air intake duct integration configuration | |
| CN111767613B (en) | Design method for inlet section of fixed-geometry high-speed air inlet channel with offset compression surface | |
| CN108301926B (en) | Hypersonic convex-to-circular inner contraction air inlet channel and design method thereof | |
| CN212337458U (en) | Three-dimensional inner rotation damping expansion wave high-performance double-channel TBCC air inlet | |
| CN201301751Y (en) | Inner wave rider type hypersonic speed air inlet channel based on arbitrary shaped shock wave | |
| CN111348169A (en) | Integrated design method for circumferential four-inlet-channel layout of conical aircraft forebody | |
| CN105539863B (en) | Hypersonic aircraft precursor, air intake duct and support plate integrated pneumatic layout method | |
| CN105151307A (en) | Method for cutting Mach surface of hypersonic aircraft with forebody/air inlet pipeline in integrated design | |
| CN104912667A (en) | Design method of hypersonic speed internal-contraction air inlet channel carried out in steps | |
| CN113279860A (en) | Inner bulge S-shaped air inlet channel with middle control section and method | |
| CN101392686A (en) | Internal waverider hypersonic inlet and design method for giving attention to internal and external flow performance | |
| CN110210096B (en) | Design method of variable-section three-dimensional internal-contraction air inlet channel matched with curved cone bullet body | |
| CN113104198B (en) | Wave-rider precursor self-adaptive boundary layer displacement inward-transfer air inlet channel integrated design method | |
| CN201301753Y (en) | Inner wave rider type air inlet channel taking internal and external flow performance into consideration | |
| CN109488459A (en) | Rotatable air intake duct and aircraft in the hypersonic three-dimensional of one kind | |
| CN115659705B (en) | Fully-parameterized high-stealth air inlet channel design method and high-stealth air inlet channel |
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 |