CN113800001B - Design method of inner-shrinkage hypersonic air inlet channel integrated with precursor - Google Patents
Design method of inner-shrinkage hypersonic air inlet channel integrated with precursor Download PDFInfo
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- CN113800001B CN113800001B CN202111164255.8A CN202111164255A CN113800001B CN 113800001 B CN113800001 B CN 113800001B CN 202111164255 A CN202111164255 A CN 202111164255A CN 113800001 B CN113800001 B CN 113800001B
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- 239000002243 precursor Substances 0.000 title claims abstract description 63
- 238000000034 method Methods 0.000 title claims abstract description 12
- 230000004048 modification Effects 0.000 claims abstract description 12
- 238000012986 modification Methods 0.000 claims abstract description 12
- 230000006835 compression Effects 0.000 claims description 136
- 238000007906 compression Methods 0.000 claims description 136
- 230000035939 shock Effects 0.000 claims description 12
- 238000009826 distribution Methods 0.000 claims description 6
- 230000007704 transition Effects 0.000 claims description 6
- 238000004364 calculation method Methods 0.000 claims description 3
- 230000005284 excitation Effects 0.000 claims description 3
- 238000004088 simulation Methods 0.000 claims description 3
- 230000015572 biosynthetic process Effects 0.000 claims 1
- 230000008602 contraction Effects 0.000 abstract description 8
- 230000007547 defect Effects 0.000 abstract description 3
- 239000010410 layer Substances 0.000 description 6
- 238000009987 spinning Methods 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 238000009792 diffusion process Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 238000005086 pumping Methods 0.000 description 1
- 238000000926 separation method Methods 0.000 description 1
- 239000002344 surface layer Substances 0.000 description 1
Classifications
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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
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- 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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- 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 relates to a ramjet engine, in particular to a design method of an internal contraction hypersonic air inlet channel integrated with a precursor, which is used for solving the defect that the prior hypersonic air vehicle with a longer precursor and a thicker boundary layer is designed and manufactured by the integrated design of the precursor and the internal contraction air inlet channel, and still has no perfect solution. The design method of the internal shrinkage hypersonic inlet channel integrated with the precursor comprises the following steps: step (1): designing an aircraft precursor; step (2): designing a reference flow field; step (3): determining an inlet molded line of the air inlet channel; step (4): determining a three-dimensional pneumatic profile of an air inlet channel; step (5): and according to the coincidence relation between the molded surface of the air inlet channel and the precursor of the aircraft, the integrated design of the air inlet channel and the precursor 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 internal contraction hypersonic air inlet channel integrated with a precursor.
Background
In hypersonic flight, because the shape, the flow field uniformity and the boundary layer characteristics of the aircraft precursor have direct influence on the performance and the starting performance of the air inlet channel due to strong coupling between the aircraft precursor and the flow field of the air inlet channel, the hypersonic aircraft mostly adopts an inner contraction air inlet channel and precursor integrated design.
At present, the integrated design of an internal shrinkage air inlet channel and a precursor is mainly divided into two types, wherein one type is that the internal shrinkage air inlet channel is used as a passenger front body; the other is that the precursor and the internal contraction air inlet channel are directly integrated, but the detailed design details are not published yet; for hypersonic aircrafts with longer precursors and thicker boundary layers, the precursors and the inner contraction air inlet channels are integrally designed, and the former is not effectively applicable, while the latter is not disclosed, so that a perfect solution is still not available.
Disclosure of Invention
The invention aims to solve the defect that the prior hypersonic air vehicle with a longer precursor and a thicker boundary layer has no perfect solution for the integrated design and manufacture of the precursor and the inner contraction air inlet channel, and provides a method for designing the inner contraction hypersonic air inlet channel integrated with the precursor.
In order to solve the defects existing in the prior art, the invention provides the following technical solutions:
the design method of the inner-contraction hypersonic air inlet channel integrated with the precursor is characterized by comprising the following steps of:
step (1): design of aircraft precursors
The aircraft precursor is in a spinning body or spinning-like body structure, a bus of the aircraft precursor consists of a cubic curve and a flat straight line, and the coefficient of the cubic curve is determined according to preset starting point and ending point coordinates, and a starting cone angle and an ending point slope;
step (2): design reference flow field
The reference flow field is of an axisymmetric structure and comprises an outer compression flow field, an inner compression flow field and lip reflection shock waves; the outer compression flow field comprises an outer compression conical surface and an outer compression laser surface positioned at 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 laser 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 on the periphery of the inner compression conical surface, the intersection point of the lip cover compression surface and the outer compression excitation surface is a lip 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 initial compression waves and ending compression waves in the inner compression area; the lip reflection shock wave is positioned between the inner compression area and the outer compression area;
(2.1) according to the incoming flow parameters of the precursor of the aircraft, obtaining flow field parameters of the precursor of the aircraft through numerical simulation calculation, and extracting non-uniform flow field parameters at the inlet section of the air inlet channel as reference flow field incoming flow parameters;
(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, thereby determining the outer compression laser surface;
(2.3) interpolating to determine lip point B on the external compression laser 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 lip reflection shock waves, initial compression waves and final compression waves;
(2.4) determining an internal compression conical surface according to an incoming flow mass conservation principle;
step (3): determining inlet profile of air inlet channel
The reference flow field is partially overlapped with the aircraft precursor flow field, symmetry axes of the reference flow field and the aircraft precursor flow field are parallel to each other, and the inlet molded line of the air inlet channel comprises an intersecting line of the aircraft precursor and an external compression laser surface, a reference flow field lip molded line positioned 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;
step (4): determining three-dimensional pneumatic profile of air inlet channel
Dispersing the inlet molded line of the air inlet channel, and tracking a streamline in a reference flow field to obtain a three-dimensional pneumatic molded surface of the air inlet channel, wherein the three-dimensional pneumatic molded surface of the air inlet channel comprises an outer compression surface of the air inlet channel and a lip cover compression surface of the air inlet channel;
step (5): and according to the coincidence relation between the molded surface of the air inlet channel and the precursor of the aircraft, the integrated design of the air inlet channel and the precursor of the aircraft is completed through geometric modification.
Further, in the 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 channels, and slicing a sector formed by the inlet molded lines of the air inlet channels by taking a symmetrical axis of a reference flow field as a center; taking the intersection point of each slice and the intersecting line as a starting point, and carrying out streamline tracking in a reference flow field to obtain an outer compression surface of the air inlet channel; and (3) taking the intersection point of each slice and the lip profile of the reference flow field as a starting point, and tracking a streamline in the reference flow field to obtain the compression surface of the lip cover of the air inlet channel.
Further, in the step (5), the specific step of completing the integrated design of the air inlet and the aircraft precursor through geometric modification according to the matching relation between the air inlet profile and the aircraft precursor comprises the following steps: the two sides of the outer compression surface of the air inlet channel are in smooth transition with the front body of the aircraft through geometric modification, 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 modification smooth transition, so that the integrated design of the air inlet channel and the front body of the aircraft 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 shrinkage air inlet channel and an aircraft precursor, realizes the automatic overflow of a precursor inflow boundary layer through a transverse pressure gradient on an air inlet channel external compression surface, improves the starting capability of the air inlet channel, widens the working range of the air inlet channel, does not need additional systems such as boundary layer channel separation, air release 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 view of an aircraft precursor;
FIG. 2 is a schematic structural view of a reference flow field;
FIG. 3 is a schematic view of a configuration of an aircraft precursor partially coincident with a reference flow field;
FIG. 4 is a schematic view of the inlet profile of the inlet duct;
fig. 5 is a schematic structural view of an internal contracting hypersonic air intake duct integrated with a precursor.
The reference numerals are explained as follows: 1-an aircraft precursor; 2-reference flow field, 21-external compression flow field, 211-external compression conical surface, 212-external compression laser surface, 213-external compression zone, 22-internal compression flow field, 221-internal compression conical surface, 222-lip cover compression surface, 223-internal compression zone, 224-initial compression wave, 225-ending compression wave and 23-lip reflection shock wave; 3-intersecting lines; 4-a lip profile of a reference flow field; 5-an air inlet channel outer compression surface; 6-an inlet lip shroud compression face; 7-a boundary layer; 8-outlet of air inlet channel.
Detailed Description
The invention is further described below with reference to the drawings and exemplary embodiments.
The design method of the inner-contraction hypersonic air inlet channel integrated with the precursor comprises the following steps:
step (1): designing an aircraft precursor 1
Referring to fig. 1, the aircraft precursor 1 is in a spinning or quasi-spinning structure, the half cone angle of the head is 15 degrees, the generatrix of the aircraft precursor is composed of a cubic curve and a flat straight line, and the coefficient of the cubic curve is determined according to preset starting point and ending point coordinates, and the starting cone angle and the ending point slope;
step (2): design reference flow field 2
Referring to fig. 2, the reference flow field 2 is an axisymmetric structure, and includes an outer compression flow field 21, an inner compression flow field 22 and lip reflection shock waves 23; the outer compression flow field 21 comprises an outer compression conical surface 211, an outer compression laser surface 212 positioned at the periphery of the outer compression conical surface 211, and an outer compression region 213 is formed between the outer compression conical surface 211 and the outer compression laser 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 excitation surface 212 is a lip 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 final compression wave 225 in the inner compression area 223; the lip reflection shock wave 23 is located between the inner compression zone 223 and the outer compression zone 213;
(2.1) according to the incoming flow parameters of the aircraft precursor 1, obtaining the flow field parameters of the aircraft precursor 1 through numerical simulation calculation, and extracting the non-uniform flow field parameters at the inlet section of the air inlet channel as the incoming flow parameters of the reference flow field 2;
(2.2) giving a pressure distribution on the outer compression cone 211 according to the inflow parameter of the reference flow field 2, thereby determining the outer compression cone 211, thereby determining the outer compression pumping surface 212;
(2.3) interpolating the lip point B on the outer compression shock surface 212 according to the capture radius of the reference flow field 2, giving the pressure distribution on the lip shroud compression surface 222, thereby determining the lip reflected shock wave 23, the initial compression wave 224 and the final compression wave 225;
(2.4) determining an internal compression cone 221 according to the principle of conservation of incoming flow mass;
step (3): determining inlet profile of air inlet channel
Referring to fig. 3 and 4, the reference flow field 2 is partially overlapped with the flow field of the aircraft precursor 1, and the symmetry axes of the reference flow field 2 and the flow field of the aircraft precursor 1 are parallel to each other, and the inlet molded line of the air inlet comprises an intersecting line 3 of the aircraft precursor 1 and the external compression laser surface 212, a reference flow field lip molded line 4 positioned at a lip point B, and two side straight lines ac and nh forming a sector with the intersecting line 3 and the reference flow field lip molded line 4;
step (4): determining three-dimensional pneumatic profile of air inlet channel
Referring to fig. 4 and 5, the inlet molded line of the air inlet is discretized, and a sector formed by the inlet molded line of the air inlet is sliced 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, tracking the streamline in the reference flow field 2, and circumferentially arranging the streamline to obtain an air inlet channel outer compression surface 5; taking the intersection point of each slice and the lip profile 4 of the reference flow field as a starting point, tracking the flow line in the reference flow field 2, and circumferentially arranging the flow line to obtain a compression surface 6 of the lip cover of the air inlet channel; after the supersonic air flow passes through the air inlet outer compression surface 5, due to the effect of transverse pressure gradient, the inflow auxiliary surface layer 7 overflows to two sides on the air inlet outer compression surface 5, so that the low energy flow rate of the air entering the air inlet is reduced, and the performance of the air inlet is improved;
step (5): referring to fig. 5, the two sides of the outer compression surface 5 of the air inlet are in smooth transition with the front body 1 of the aircraft through geometric modification, and the compression surface 6 of the lip cover of the air inlet and the outlet 8 of the air inlet form an air inlet throat diffuser through geometric modification smooth transition, so that the integrated design of the air inlet and the front body 1 of the aircraft is realized.
The foregoing embodiments are merely for illustrating the technical solutions of the present invention, and not for limiting the same, and it will be apparent to those skilled in the art that modifications may be made to the specific technical solutions described in the foregoing embodiments, or equivalents may be substituted for some of the technical features thereof, without departing from the spirit of the technical solutions protected by the present invention.
Claims (3)
1. The design method of the inner-contraction hypersonic air inlet channel integrated with the precursor is characterized by comprising the following steps of:
step (1): design of aircraft front body (1)
The aircraft precursor (1) is in a rotary or quasi-rotary structure, a bus of the aircraft precursor is composed of a cubic curve and a flat straight line, and the coefficient of the cubic curve is determined according to preset starting point and ending point coordinates, a starting cone angle and an ending point slope;
step (2): design reference 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 lip reflection shock waves (23); the outer compression flow field (21) comprises an outer compression conical surface (211) and an outer compression laser 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 laser 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 excitation surface (212) is a lip 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 an end compression wave (225) in the inner compression area (223); the lip reflection shock wave (23) is positioned between the inner compression zone (223) and the outer compression zone (213);
(2.1) according to the incoming flow parameters of the aircraft precursor (1), obtaining flow field parameters of the aircraft precursor (1) through numerical simulation calculation, and extracting non-uniform flow field parameters at the inlet section of the air inlet channel as the incoming flow parameters of the reference flow field (2);
(2.2) giving a pressure distribution on the outer compression cone (211) according to the incoming flow parameters of the reference flow field (2) to determine the outer compression cone (211), thereby determining the outer compression laser surface (212);
(2.3) interpolating the lip point B on the outer compression shock surface (212) according to the capture radius of the reference flow field (2), giving a pressure distribution on the lip shroud compression surface (222), thereby determining the lip reflected shock wave (23), the initial compression wave (224) and the final compression wave (225);
(2.4) determining an internal compression cone (221) according to the principle of conservation of incoming flow mass;
step (3): determining inlet profile of air inlet channel
The reference flow field (2) is partially overlapped with the flow field of the aircraft precursor (1) and the symmetry axes of the reference flow field and the flow field are parallel to each other, and the inlet molded line of the air inlet comprises an intersecting line (3) of the aircraft precursor (1) and an external compression laser surface (212), a reference flow field lip molded line (4) positioned 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);
step (4): determining three-dimensional pneumatic profile of air inlet channel
Dispersing the inlet molded line of the air inlet channel, and carrying out streamline tracking in a reference flow field (2) to obtain a three-dimensional pneumatic molded surface of the air inlet channel, wherein the three-dimensional pneumatic molded surface of the air inlet channel comprises an outer compression surface (5) of the air inlet channel and a lip cover compression surface (6) of the air inlet channel;
step (5): and according to the coincidence relation between the molded surface of the air inlet channel and the aircraft precursor (1), the integrated design of the air inlet channel and the aircraft precursor (1) is completed through geometric modification.
2. The method for designing an internal contracting hypersonic air intake duct integrated with a precursor as set forth in claim 1, wherein:
in the step (4), the specific step of determining the three-dimensional aerodynamic profile of the air inlet channel comprises the following steps: dispersing inlet molded lines of the air inlet channels, and slicing a sector formed by the inlet molded lines of the air inlet channels 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, and carrying out streamline tracking in the reference flow field (2) to obtain an air inlet channel outer compression surface (5); and (3) taking the intersection point of each slice and the lip profile (4) of the reference flow field as a starting point, and carrying out streamline tracking in the reference flow field (2) to obtain the compression surface (6) of the lip cover of the air inlet channel.
3. The method for designing an internal contracting hypersonic air intake duct integrated with a precursor according to claim 1 or 2, wherein:
in the step (5), the specific step of completing the integrated design of the air inlet and the aircraft precursor (1) through geometric modification according to the anastomotic relationship between the air inlet molded surface and the aircraft precursor (1) comprises the following steps: smooth transition with aircraft precursor (1) is realized through geometric modification in the outer compression face of intake duct (5) both sides, intake duct lip cover compression face (6) and intake duct export (8) are through geometric modification smooth transition formation intake duct throat diffuser to realize intake duct and aircraft precursor (1) integrated design.
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Families Citing this family (5)
| Publication number | Priority date | Publication date | Assignee | Title |
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| 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 |
| 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 |
| CN117048839B (en) * | 2023-07-21 | 2025-02-14 | 南京理工大学 | Design method of hypersonic integrated inlet with given forebody shape |
| CN118289204B (en) * | 2024-03-29 | 2025-02-14 | 南京航空航天大学 | A waverider precursor/compression surface integrated configuration based on plane and conical shock waves and its design method |
| CN118728556B (en) * | 2024-05-30 | 2025-04-01 | 南京航空航天大学 | A parametric design method for axisymmetric precooling inlet |
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