CN103953448A - Hypersonic air inlet channel - Google Patents
Hypersonic air inlet channel Download PDFInfo
- Publication number
- CN103953448A CN103953448A CN201410151860.5A CN201410151860A CN103953448A CN 103953448 A CN103953448 A CN 103953448A CN 201410151860 A CN201410151860 A CN 201410151860A CN 103953448 A CN103953448 A CN 103953448A
- Authority
- CN
- China
- Prior art keywords
- return flow
- flow line
- intake duct
- air inlet
- channel
- 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
- 238000000926 separation method Methods 0.000 claims abstract description 21
- 230000008676 import Effects 0.000 claims description 23
- 230000007704 transition Effects 0.000 claims description 6
- 230000035939 shock Effects 0.000 abstract description 6
- 230000006698 induction Effects 0.000 abstract description 5
- 230000003068 static effect Effects 0.000 abstract description 5
- 230000003247 decreasing effect Effects 0.000 abstract description 2
- 238000000034 method Methods 0.000 description 16
- 230000008859 change Effects 0.000 description 7
- 238000004088 simulation Methods 0.000 description 6
- 238000010992 reflux Methods 0.000 description 5
- 238000013461 design Methods 0.000 description 4
- 238000013459 approach Methods 0.000 description 3
- 230000006835 compression Effects 0.000 description 3
- 238000007906 compression Methods 0.000 description 3
- 230000008602 contraction Effects 0.000 description 3
- 230000007423 decrease Effects 0.000 description 3
- 238000010586 diagram Methods 0.000 description 3
- 230000006872 improvement Effects 0.000 description 3
- 239000002243 precursor Substances 0.000 description 3
- 230000008569 process Effects 0.000 description 3
- 238000011084 recovery Methods 0.000 description 3
- 230000001105 regulatory effect Effects 0.000 description 3
- 238000011160 research Methods 0.000 description 3
- 208000036366 Sensation of pressure Diseases 0.000 description 2
- 230000008034 disappearance Effects 0.000 description 2
- 230000002411 adverse Effects 0.000 description 1
- 230000003321 amplification Effects 0.000 description 1
- 238000005452 bending Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000001276 controlling effect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- YGANSGVIUGARFR-UHFFFAOYSA-N dipotassium dioxosilane oxo(oxoalumanyloxy)alumane oxygen(2-) Chemical compound [O--].[K+].[K+].O=[Si]=O.O=[Al]O[Al]=O YGANSGVIUGARFR-UHFFFAOYSA-N 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 230000010534 mechanism of action Effects 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 229910052627 muscovite Inorganic materials 0.000 description 1
- 238000003199 nucleic acid amplification method Methods 0.000 description 1
- 238000002161 passivation Methods 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 230000002829 reductive effect Effects 0.000 description 1
- 230000004044 response Effects 0.000 description 1
- 230000002441 reversible effect Effects 0.000 description 1
- 239000000523 sample Substances 0.000 description 1
- 238000012546 transfer Methods 0.000 description 1
- 238000002604 ultrasonography Methods 0.000 description 1
Landscapes
- Characterised By The Charging Evacuation (AREA)
Abstract
The invention discloses a hypersonic air inlet channel. The hypersonic air inlet channel comprises an air inlet channel main body, an air inlet channel lip cover and a backflow channel, wherein the backflow channel comprises a backflow channel inlet, a backflow channel outlet and a uniform-section drainage pipeline; the uniform-section drainage pipeline is connected with the backflow channel inlet and the backflow channel outlet. According to the working principle of the hypersonic air inlet channel, low-energy flow in separation bubbles at an air inlet channel inlet is drained through the backflow channel inlet by utilizing the difference, generated by induction of large separation bubbles in front of the an air inlet channel inlet when the air inlet channel is not started, between front and back static pressures of a shock wave, passes through the drainage pipeline and then is re-filled into a front air inlet channel body from the backflow channel outlet to form closed circulation flow. By virtue of a backflow channel structure, the self-starting Mach number of the air inlet channel can be remarkably decreased, and meanwhile, an external flow field of the air inlet channel is hardly disturbed. Moreover, the performance of the air inlet channel is hardly affected by the backflow channel under a high Mach number so as to be guaranteed. The hypersonic air inlet channel is simple in structure and easy to implement.
Description
Technical field
The invention belongs to Ramjet Technique field, particularly a kind of hypersonic inlet.
Background technique
Hypersonic flight refers to the flight that Mach number is greater than 5.The research of hypersonic long-range maneuvering-vehicle becomes world today's hot research problem of competitively carrying out of making the country prosperous because of its important strategic importance.Intake duct, as the critical piece in hypersonic propulsion system, is one of key technology of the hypersonic push technological development of air-inlet type, and the quality of its performance often produces vital impact to whole propulsion system performance.But under present stage hypersonic inlet low mach, self-starting problem has often limited the operating range of aircraft, and then directly assist system and flight cost are exerted a decisive influence.Therefore, probe into and how effectively reduce intake duct self-starting Mach number and there is outstanding realistic meaning.
Conventionally the technological approaches of, widening intake duct work range of Mach numbers mainly contains two classes: become how much regulating methods and fixed how much profile Fluid field controlling methods.How much intake duct schemes of the change of taking at present are mainly divided into following a few class: rotary type, translational, vari ramp etc.The intake duct that the hypersonic missiles of mechanism's research such as France state aviation aerospace flight technology research center (ONERA) and the X-43A aircraft of the U.S. adopt is lip and rotates change geometrical solution.With respect to lip scheme of rotation, it is relatively low that contraction lip formula becomes how much intake duct control difficulty, and F.Falempin and the Muscovite M.Goldfeld etc. of France is studied the starting process of how much intake ducts of flexible lip formula change.How much intake ducts of change change mechanically object plane parameter and throatpiston is long-pending, and then oral area wave system and contraction ratio are regulated, therefore can effectively widen intake duct work range of Mach numbers and ensure key state or even different operating state under intake duct approach optimum performance work.But its shortcoming is also very outstanding: weight increases, complex structure, reliability decrease, and obturage, thermal protection problem is comparatively outstanding.
Flow field control method under fixed how much profiles is to reach the object that reduces intake duct self-starting Mach number, as initiatively aspirated, offer passive bypass channel etc. in intake duct mostly by the overflow under low mach.This type of regulating method is actually throatpiston long-pending " amplification ", has alleviated the problem that under low mach, inlet throat sectional area seems too small, therefore can effectively improve self-starting performance under intake duct low mach.But also can be with and serve adverse effect by the flow field control method under these type of fixed how much profiles, after intake duct starting, the overflow meeting that continues to occur causes the loss of intake duct flow, causes motor power loss.Overflow also can produce to disturb to External airflow field and cause the resistance of intake duct and even whole aircraft to increase.
Summary of the invention
The problem to be solved in the present invention is to provide a kind of hypersonic inlet with return flow line, this intake duct is based on closed flow field control technique, by easy drainage system, intake duct self-starting performance is obviously improved, intake duct work range of Mach numbers is significantly widened.This drainage system neither can significantly increase the weight of former intake duct, almost produces without refluxing under High Mach number simultaneously, and inlet characteristic under High Mach number is exerted an influence hardly.
A kind of hypersonic inlet disclosed by the invention, comprise intake duct main body, intake duct lip cover, the intake duct internal channel forming between intake duct main body, intake duct lip cover, intake duct internal channel top is Fighter Inlet, intake duct compressing surface next-door neighbour Fighter Inlet, there is separation zone at Fighter Inlet place; This hypersonic inlet also comprises return flow line, and described return flow line comprises return flow line import, return flow line outlet and connects return flow line import and the drainage pipeline of return flow line outlet; Return flow line import is opened in intake duct internal channel, and it is latter half of to be positioned at separation zone, and return flow line outlet is opened in the residing same grading air flue compressing surface of return flow line import.
As the further improvement of technique scheme, the import of described return flow line and return flow line outlet wall are all perpendicular to intake duct compressing surface.
As the further improvement of technique scheme, the kernel of section line of described return flow line import and the distance L of separation zone originating point
1meet:
0.65L
B≤L
1≤0.95
B，
Wherein, L
bfor separation zone is along the width flowing to;
The kernel of section line of return flow line outlet and the distance L of intake duct compressing surface initial point
2meet:
0.5b≤L
2≤1.0L
S，
Wherein, L
sfor the distance between initial point and the separation zone originating point of the residing compressing surface in return flow line;
As further improving again of technique scheme, described drainage pipeline is constant section duct.
As further improving again of technique scheme, arc transition is all passed through in connecting between drainage pipeline and return flow line import, return flow line outlet.
As further improving again of technique scheme, the cross-sectional width b of drainage pipeline meets:
0.2W≤b≤0.5W，
Wherein, W is the entrance width of intake duct internal channel;
Drainage pipeline is that R, this internal face are L apart from the perpendicular distance of compressing surface near the transition arc radius of the internal face of compressing surface one side
d, and meet:
1.0b≤R≤2.0b，
1.5R≤L
D≤2.0R。
Beneficial effect of the present invention:
Only, by drainage system simple in structure, significantly reduce the self-starting Mach number of hypersonic inlet, widened the work range of Mach numbers of intake duct, and almost noiseless to intake duct Flow Field outside.And under High Mach number, this structure exerts an influence hardly to inlet characteristic, has ensured inlet characteristic under High Mach number.And working stability, reliable, is easy to realize.
Brief description of the drawings
Fig. 1 is hypersonic inlet structural representation of the present invention;
Fig. 2 is the each parts of hypersonic inlet of the present invention and relative position schematic diagram;
Fig. 3-1 and Fig. 3-2nd, hypersonic inlet fundamental diagram of the present invention;
Fig. 4 is hypersonic inlet mechanism of action schematic diagram of the present invention;
Fig. 5-1,5-2 and 5-3 are the Mach number isopleth maps under typicalness in hypersonic inlet self-starting process of the present invention;
Fig. 6 is the Mach number isopleth map under typicalness in former profile intake duct (not with return flow line) self-starting process;
Embodiment
Below in conjunction with accompanying drawing, a kind of hypersonic inlet that the present invention is proposed is elaborated.
As depicted in figs. 1 and 2, a kind of hypersonic inlet comprises intake duct main body 5, intake duct lip cover 4, the intake duct internal channel 6 forming between intake duct main body 5, intake duct lip cover 4, intake duct internal channel 6 tops are Fighter Inlet 8, intake duct compressing surface 7 is close to Fighter Inlet 8.
When under low mach, intake duct is inoperative, Fighter Inlet 8 places often form large separation zone 9, and separation zone 9 is L along the width flowing to
b.This hypersonic inlet also comprises return flow line, and described return flow line comprises return flow line import 1, return flow line outlet 2 and connects return flow line import 1 and prismatic drainage pipeline 3 of return flow line outlet 2.
Return flow line import 1 is opened in intake duct internal channel 6, and is positioned in the intake duct main body 5 that separation zone 9 is latter half of covered.The distance L of the kernel of section line of return flow line import 1 and separation zone 9 originating points
1meet: 0.65L
b≤ L
1≤ 0.95L
b.Return flow line outlet 2 is opened in the residing same grading air flue compressing surface 7 of return flow line import 1, and is positioned at the front of separation zone 9.Return flow line import 1 and return flow line export 2 walls all perpendicular to intake duct compressing surface 7.The kernel of section line of return flow line outlet 2 and the distance L of intake duct compressing surface 7 initial points
2meet: 0.5b≤L
2≤ 1.0L
s, wherein, L
sfor the distance between initial point and separation zone 9 originating points of compressing surface 7.The cross-sectional width b of drainage pipeline 3 meets: 0.2W≤b≤0.5W, wherein, W is the entrance width of intake duct internal channel.Arc transition is all passed through in connecting between drainage pipeline 3 and return flow line import 1, return flow line outlet 2.Drainage pipeline 3 is that R, this internal face are L apart from the perpendicular distance of compressing surface 7 near the transition arc radius of a side internal face of compressing surface 7
d, and meet: 1.0b≤R≤2.0b, 1.5R≤L
d≤ 2.0R.
As shown in Fig. 1 and Fig. 3-1, Fig. 3-2, for hypersonic inlet, under low mach, when intake duct is caught flow can not be all by venturi time, often form large separate bubble at Fighter Inlet 8 places, and then generation induction shock wave 10, air-flow static pressure after induction shock wave 10 significantly raises, now utilize induction shock wave 10 front and back differential static pressures as power source, " ordering about " in separation zone low energy flows in return flow line import 1 place and draws, and drawn back into air flue precursor by drainage pipeline 3, Secondary Flow 11 forms.Refluxing exports 2 places and refills intake duct in return flow line, and forms " bulge shape " pneumatic wall 12 in herein.The pneumatic wall 12 of this convex of flowing through at a high speed, bring out produce a series of weak compressions, expansion wave system carries out " modification " to wedge surface compressional wave 13, make wedge surface compressional wave 13 have obvious bending skew outward, cause transfer window 14 to become large, the overflow of intake duct precursor ultrasound velocity increases, intake duct is caught flow and is declined, and this is obviously conducive to the starting under intake duct low mach.Simultaneously, as shown in Figure 4, under low mach, along with the beginning of drainage, in separation zone, low energy flows in return flow line inlet and draws, separate bubble reduces gradually, and induction shock strength weakens thereupon, causes differential static pressure between the inlet/outlet of return flow line to decline, drive pressure reduction reduce make capacity of reflux decline, the suffered impact of intake duct precursor compression wave system weakens, and so forms " reverse feedback response " until separate bubble and induce shockwave cancellation, and intake duct starts immediately.
Application example 1:
(1) technical order:
Work range of Mach numbers: 5.0～7.0, design work state is Mach 6.0
(2) scheme introduction:
Designed a binary hypersonic inlet with three grades of compressing surfaces, three roads compression wedge surface angles are respectively 5 °, and 5.4 ° and 5.9 °, venturi height A
t=18.7mm, intake duct distance piece length is 7 times of venturi width, contraction ratio CR=1.6 in venturi considers for thermal protection, and compressing surface leading edge and lip cover leading edge have all been carried out to passivation.In this former profile intake duct, offer return flow line, and return flow line design parameter is: L
1≈ 0.70L
b, b=0.2W, L
2=0.5b, R=1.0b, L
d=1.5R.By numerical simulation, two-dimensional analog is carried out in band inlet duct flow field, return flow line, from the simulation result shown in Fig. 5-1, can find out that intake duct realizes self-starting in Mach 4.2.
Application example 2:
In the former profile intake duct described in application example 1, offer return flow line, change the parameter of return flow line in embodiment 1, the present embodiment return flow line design parameter is: L
1≈ 090L
b, b=0.4W, L
2=1.0L
s, R=1.5b, L
d=2.0R.By numerical simulation, two-dimensional analog is carried out in band inlet duct flow field, return flow line, from the simulation result shown in Fig. 5-2, can find out that intake duct realizes self-starting in Mach 4.2.
Application example 3:
In the former profile intake duct described in application example 1, offer return flow line, change the parameter of return flow line in embodiment 1, the present embodiment return flow line design parameter is: L
1≈ 084L
b, b=0.27W, L
2=0.33L
s, R=1.25b, L
d=1.5R.By numerical simulation, former profile intake duct and band inlet duct flow field, return flow line are carried out to two-dimensional analog, and simulation result is analyzed to contrast.
(1) self-starting Character Comparison:
Definition separation zone is intake duct starting state while disappearance completely.For former profile intake duct, as shown in Figure 6, under low mach, Fighter Inlet place forms large separate bubble, along with the increase of incoming flow Mach number, separate bubble exists until Mach 5.4 just disappears always, intake duct is realized self-starting, and this self-starting Mach number has significantly exceeded the lower limit of normal range of operation, has dwindled the normal work range of Mach numbers of intake duct.And in intake duct, offer return flow line, and by easy drainage system, can find out from Fig. 5-3, Fighter Inlet place separate bubble is decreased to rapidly disappearance, and intake duct realizes self-starting in Mach 3.7.Visible, return flow line makes intake duct self-starting Mach number be reduced to Mach 3.7 by Mach 5.4, and intake duct self-starting performance is obviously improved, and intake duct work range of Mach numbers is significantly widened.
(2) inlet characteristic contrast in full range of Mach numbers:
Table 1 has contrasted former profile intake duct and intake duct of the present invention under typicalness and has been with return flow line intake duct distance piece outlet performance, and wherein σ is inlet total pres sure recovery coefficient,
for Flow coefficient of inlet.As can be seen from the table, Mach 3.5 o'clock, the two equal unstart, the flow coefficient with return flow line intake duct is lower than former profile intake duct.Along with incoming flow Mach number is increased to Mach 3.7, the unstart of former profile intake duct, and band return flow line intake duct starting, now inlet total pres sure recovery coefficient, flow coefficient all significantly raise.And after under High Mach number, former profile intake duct also starts, can find out from accompanying drawing 5-3, now because Fighter Inlet oblique shock wave incidence point is positioned at import rear, return flow line, between passage inlet/outlet, differential static pressure is very little, and capacity of reflux is little.During as Mach 5.5, capacity of reflux is about 0.004kg/s, accounts for 0.1% of intake duct flow, and Flow coefficient of inlet almost remains unchanged, and total pressure recovery coefficient also slightly raises.Visible, return flow line structure exerts an influence hardly to inlet characteristic under High Mach number, has ensured the performance under intake duct High Mach number.
Intake duct distance piece outlet performance comparison in the full range of Mach numbers of table 1
The concrete application approach of the present invention is a lot, and the above is only the preferred embodiment of the present invention, should be understood that; for those skilled in the art; under the premise without departing from the principles of the invention, can also make some improvement, these improve and also should be considered as protection scope of the present invention.
Claims (6)
1. a hypersonic inlet, comprise intake duct main body (5), intake duct lip cover (4), the intake duct internal channel (6) forming between intake duct main body (5), intake duct lip cover (4), intake duct internal channel (6) top is Fighter Inlet (8), intake duct compressing surface (7) next-door neighbour's Fighter Inlet (8), Fighter Inlet (8) has been located separation zone (9); It is characterized in that: this hypersonic inlet also comprises return flow line, described return flow line comprises return flow line import (1), return flow line outlet (2) and connects return flow line import (1) and the drainage pipeline (3) of return flow line outlet (2); Return flow line import (1) is opened in intake duct internal channel (6), and it is latter half of to be positioned at separation zone (9), and return flow line outlet (2) is opened in the residing same grading air flue compressing surface of return flow line import (1) (7).
2. hypersonic inlet according to claim 1, is characterized in that: described return flow line import (1) and return flow line outlet (2) wall are all perpendicular to intake duct compressing surface (7).
3. hypersonic inlet according to claim 2, is characterized in that: the distance L of the kernel of section line of described return flow line import (1) and separation zone (9) originating point
1meet:
0.65L
B≤L
1≤0.95
B，
Wherein, L
bfor separation zone (9) are along the width flowing to;
The return flow line outlet kernel of section line of (2) and the distance L of intake duct compressing surface (7) initial point
2meet:
0.5b≤L
2≤1.0
S，
Wherein, L
sfor the distance between initial point and separation zone (9) originating point of compressing surface (7).
4. hypersonic inlet according to claim 3, is characterized in that: described drainage pipeline (3) is constant section duct.
5. hypersonic inlet according to claim 4, is characterized in that: arc transition is all passed through in connecting between drainage pipeline (3) and return flow line import (1), return flow line outlet (2).
6. hypersonic inlet according to claim 5, is characterized in that: the cross-sectional width b of drainage pipeline (3) meets:
0.2W≤b≤0.5W，
Wherein, W is the entrance width of intake duct internal channel;
Drainage pipeline (3) is that R, this internal face are L apart from the perpendicular distance of compressing surface (7) near the transition arc radius of the internal face of compressing surface (7) one sides
d, and meet:
1.0b≤R≤2.0b，
1.5R≤L
D≤2.0R。
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410151860.5A CN103953448B (en) | 2014-04-15 | 2014-04-15 | A kind of hypersonic inlet |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410151860.5A CN103953448B (en) | 2014-04-15 | 2014-04-15 | A kind of hypersonic inlet |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN103953448A true CN103953448A (en) | 2014-07-30 |
| CN103953448B CN103953448B (en) | 2016-05-18 |
Family
ID=51330772
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201410151860.5A Expired - Fee Related CN103953448B (en) | 2014-04-15 | 2014-04-15 | A kind of hypersonic inlet |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN103953448B (en) |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109026396A (en) * | 2018-08-07 | 2018-12-18 | 北京空间技术研制试验中心 | Supersonic Three Dimensional air intake duct pneumatic control method |
| CN109184952A (en) * | 2018-08-21 | 2019-01-11 | 西安理工大学 | A kind of hypersonic inlet not self-holding ability quantitative analysis method in starting state Disengagement zone |
| CN109204849A (en) * | 2018-08-07 | 2019-01-15 | 北京空间技术研制试验中心 | For the anti-self-oscillation control method of high speed near space vehicle air intake duct |
| CN109296473A (en) * | 2018-08-10 | 2019-02-01 | 西安理工大学 | A kind of magnetic control pulsed discharge hypersonic inlet assistant starting flow control method |
| CN109356723A (en) * | 2018-11-27 | 2019-02-19 | 北京空间技术研制试验中心 | Closed return flow line flow field control method |
| CN109436293A (en) * | 2018-11-21 | 2019-03-08 | 南京航空航天大学 | A kind of shock wave control device |
| CN113148192A (en) * | 2021-04-08 | 2021-07-23 | 南京航空航天大学 | Binary adjustable air inlet channel and aircraft forebody integrated assembly and design method |
| CN113623086A (en) * | 2021-07-19 | 2021-11-09 | 南京航空航天大学 | Shock wave/boundary layer interference controller |
| CN114320661A (en) * | 2021-12-21 | 2022-04-12 | 哈尔滨工业大学 | Backflow injection pressurization system based on detonation combustion excitation and pressurization method thereof |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20020124758A1 (en) * | 2001-03-07 | 2002-09-12 | Cartland Harry E. | Stagnation pressure activated fuel release mechanism for hypersonic projectiles |
| CN101029597A (en) * | 2007-03-22 | 2007-09-05 | 南京航空航天大学 | Fixed geometrical supersonic-speed and high supersonic-speed adjusting air inlet |
| EP2180164A1 (en) * | 2008-10-23 | 2010-04-28 | Mbda Uk Limited | Method and system for altering engine air intake geometry |
| CN102953825A (en) * | 2012-11-22 | 2013-03-06 | 南京航空航天大学 | Pneumatic supersonic velocity/hypersonic velocity adjustable air inlet passage for self-circulation of forebody |
| CN203962164U (en) * | 2014-04-15 | 2014-11-26 | 南京航空航天大学 | A kind of hypersonic inlet |
-
2014
- 2014-04-15 CN CN201410151860.5A patent/CN103953448B/en not_active Expired - Fee Related
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20020124758A1 (en) * | 2001-03-07 | 2002-09-12 | Cartland Harry E. | Stagnation pressure activated fuel release mechanism for hypersonic projectiles |
| CN101029597A (en) * | 2007-03-22 | 2007-09-05 | 南京航空航天大学 | Fixed geometrical supersonic-speed and high supersonic-speed adjusting air inlet |
| EP2180164A1 (en) * | 2008-10-23 | 2010-04-28 | Mbda Uk Limited | Method and system for altering engine air intake geometry |
| CN102953825A (en) * | 2012-11-22 | 2013-03-06 | 南京航空航天大学 | Pneumatic supersonic velocity/hypersonic velocity adjustable air inlet passage for self-circulation of forebody |
| CN203962164U (en) * | 2014-04-15 | 2014-11-26 | 南京航空航天大学 | A kind of hypersonic inlet |
Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109026396A (en) * | 2018-08-07 | 2018-12-18 | 北京空间技术研制试验中心 | Supersonic Three Dimensional air intake duct pneumatic control method |
| CN109204849B (en) * | 2018-08-07 | 2020-07-14 | 北京空间技术研制试验中心 | Anti-self-oscillation control method for air inlet channel of high-speed near space aircraft |
| CN109204849A (en) * | 2018-08-07 | 2019-01-15 | 北京空间技术研制试验中心 | For the anti-self-oscillation control method of high speed near space vehicle air intake duct |
| CN109296473A (en) * | 2018-08-10 | 2019-02-01 | 西安理工大学 | A kind of magnetic control pulsed discharge hypersonic inlet assistant starting flow control method |
| CN109184952B (en) * | 2018-08-21 | 2019-06-18 | 西安理工大学 | A kind of hypersonic inlet not self-holding ability quantitative analysis method in starting state Disengagement zone |
| CN109184952A (en) * | 2018-08-21 | 2019-01-11 | 西安理工大学 | A kind of hypersonic inlet not self-holding ability quantitative analysis method in starting state Disengagement zone |
| CN109436293A (en) * | 2018-11-21 | 2019-03-08 | 南京航空航天大学 | A kind of shock wave control device |
| CN109356723A (en) * | 2018-11-27 | 2019-02-19 | 北京空间技术研制试验中心 | Closed return flow line flow field control method |
| CN113148192A (en) * | 2021-04-08 | 2021-07-23 | 南京航空航天大学 | Binary adjustable air inlet channel and aircraft forebody integrated assembly and design method |
| CN113148192B (en) * | 2021-04-08 | 2022-04-05 | 南京航空航天大学 | Binary adjustable air inlet channel and aircraft forebody integrated assembly and design method |
| CN113623086A (en) * | 2021-07-19 | 2021-11-09 | 南京航空航天大学 | Shock wave/boundary layer interference controller |
| CN113623086B (en) * | 2021-07-19 | 2022-08-02 | 南京航空航天大学 | Shock wave/boundary layer interference controller |
| CN114320661A (en) * | 2021-12-21 | 2022-04-12 | 哈尔滨工业大学 | Backflow injection pressurization system based on detonation combustion excitation and pressurization method thereof |
Also Published As
| Publication number | Publication date |
|---|---|
| CN103953448B (en) | 2016-05-18 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN103953448A (en) | Hypersonic air inlet channel | |
| CN203962164U (en) | A kind of hypersonic inlet | |
| CN105840551B (en) | The pneumatic implementation method of multi-state point high load capacity compressor blade | |
| CN102953825B (en) | Pneumatic supersonic velocity/hypersonic velocity adjustable air inlet passage for self-circulation of forebody | |
| CN102953826B (en) | Pneumatic supersonic velocity/hypersonic velocity adjustable air inlet passage for forebody-inner passage circulation | |
| Guleren et al. | Numerical simulation of the stalled flow within a vaned centrifugal pump | |
| CN106050469A (en) | Nozzle structure for realizing throat area adjustment and thrust vector and adjusting method | |
| CN109896027A (en) | A kind of bump inlet and Boundary layer flow method based on plasma synthesis jet stream | |
| CN108425887A (en) | Ultra-wide string trigonometric function waveform blade | |
| Sun et al. | Design, modification and optimization of an ultra-high-load transonic low-reaction aspirated compressor | |
| CN108412618A (en) | Hypersonic/supersonic axisymmetric inlet lip and design method thereof | |
| CN109353527A (en) | Using the BLI air intake duct of mixed flow control method | |
| CN105221479A (en) | Centrifugal blower fan blade wheel, centrifugal blower and air-conditioning | |
| CN103029830B (en) | A kind of two Waverider is to spelling air suction type hypersonic vehicle precursor and method of designing thereof | |
| CN107489651A (en) | A kind of blade profile optimization method for suppressing fan shock wave noise based on quadratic function | |
| CN100580258C (en) | Method for improving air compressor blade load by using pumping and sucking | |
| CN109488459A (en) | Rotatable air intake duct and aircraft in the hypersonic three-dimensional of one kind | |
| CN100567082C (en) | A kind of method that is used to construct intake oblique cut inlet | |
| CN106122107A (en) | Complex bend stator blade for multi stage axial flow compressor | |
| CN103077317B (en) | A kind of method of calculating flux of the leakage grooves for stable shock wave | |
| CN206017269U (en) | Complex bend stator blade for multi stage axial flow compressor | |
| CN109063407A (en) | A kind of modeling method of scramjet engine steady-state model | |
| CN105485743B (en) | A kind of range hood with noise reducing mechanism | |
| CN108533406A (en) | A kind of sliding block is moved forward and backward variable geometry inlet under the jaw adjusted | |
| CN208138201U (en) | Ultra-wide string trigonometric function waveform blade |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| TR01 | Transfer of patent right | ||
| TR01 | Transfer of patent right |
Effective date of registration: 20180511 Address after: No. 29, Qinhuai District, Qinhuai District, Nanjing, Jiangsu Co-patentee after: Beijing Institute of Power Machinery Patentee after: Nanjing University of Aeronautics and Astronautics Address before: No. 29, Qinhuai District, Qinhuai District, Nanjing, Jiangsu Patentee before: Nanjing University of Aeronautics and Astronautics |
|
| CF01 | Termination of patent right due to non-payment of annual fee | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20160518 Termination date: 20210415 |