CN203441627U - Supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device - Google Patents
Supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device Download PDFInfo
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- CN203441627U CN203441627U CN201320281995.4U CN201320281995U CN203441627U CN 203441627 U CN203441627 U CN 203441627U CN 201320281995 U CN201320281995 U CN 201320281995U CN 203441627 U CN203441627 U CN 203441627U
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- 238000005457 optimization Methods 0.000 description 2
- 238000007789 sealing Methods 0.000 description 2
- 239000000243 solution Substances 0.000 description 2
- 108091053398 TRIM/RBCC family Proteins 0.000 description 1
- 102000011408 Tripartite Motif Proteins Human genes 0.000 description 1
- 102100040255 Tubulin-specific chaperone C Human genes 0.000 description 1
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Abstract
The utility model discloses a supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device. The device comprises a nozzle contracting segment, a nozzle expanding segment, a gas-guiding bypass and a control device for controlling the opening and closing of the gas-guiding bypass, the gas-guiding bypass is an internal flowing passage through which the nozzle contracting segment and the nozzle expanding segment are communicated, and an internal flow field is established by the aid of the natural pressure difference between a gas inlet of the nozzle contracting segment and an gas outlet of the nozzle expanding segment. When an aerocraft works is in a transonic and over-expanding state, the bypass is opened, high-pressure gas of the nozzle contracting segment is guided to the nozzle expanding segment to form a secondary flow injecting effect, and a thrust coefficient of the nozzle is increased; and when the aerocraft is in a normal working state, the gas-guiding bypass passage is closed, and the nozzle design point performances cannot be affected. The supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device is simple in structure, needs no extra gas guiding sources and is free of additional weight, and the thrust performance of the nozzle is improved.
Description
Technical field
The new open form of this practicality relates to a kind of ultrasound velocity/hypersonic aircraft jet pipe inner flow passage design.
Background technique
For meeting the combined cycle engine across the requirement, particularly similar TBCC/RBCC of large range of Mach numbers flight, require vent systems in broad scope, all to there is good aeroperformance, comprise thrust, lift, pitch moment etc.There is the problems such as the mechanical constraint of expanding noodles and sealing in traditional axisymmetric nozzle, makes jet pipe to be issued to the thrust performance that all reaches higher in the difference operating mode of flying, and this makes the living space of ultrasound velocity/hypersonic aircraft be subject to serious compression.Therefore, expand its flight Mach number, improving off-design performance is an important research direction.
Can find out with reference to figure 1 and Fig. 2, Fig. 1 shows when aircraft is scope transonic speed (be Mach number 0.8 ~ 1.2 near), particularly near M=1, afterbody drag (Afterbody Drag) increases severely, and in Fig. 2, jet pipe net thrust (Thrust-Minus-Drag Performance) is more violent compared with hour decline in flight Mach number (Flight Mach Number) simultaneously.Due to the vitals of jet pipe as pressed engine, its performance quality directly affects the propulsive efficiency of whole aircraft propulsion.Therefore, the present invention is intended to solve hypersonic aircraft when wide range of Mach numbers flight, due to jet pipe off-design point, even causes serious overexpansion, the problem that causes thus nozzle performance sharply to worsen.
In general solve this technical problem and conventionally have two schemes: the first, the jet pipe of use variable-geometry, jet pipe is adjustable jet, can adjust nozzle section according to flight condition long-pending.But due to technical problems such as mechanical realization is complicated, impost is large, anti-heat sealings, realize difficulty large.The second, Secondary Flow bleed scheme, this scheme is similar to bleed scheme of the present invention, and maximum difference is that the present invention has utilized the pressurized gas of jet pipe contraction section dexterously, and do not need extra bleed air source, solved the technical barrier that Secondary Flow bleed scheme needs additional gas sources.
Summary of the invention
The utility model needs the deficiencies such as optional equipment Secondary Flow source of the gas, the limited impost of jet pipe thrust performance be large for prior art, a kind of ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type control gear has been proposed, the present invention is by causing nozzle divergence cone by the pressurized gas of jet pipe contraction section, form the effect of Secondary Flow injection, improved the thrust coefficient of jet pipe, be that jet pipe thrust performance has improved, reduced negative lift, the pitch moment that is conducive to like this to reduce aircraft is poor.By the high pressure draught of jet pipe contraction section is caused to extending section injection, and produce a series of shock wave reflection at extending section, extending section wall pressure is increased, nozzle thrust coefficient raises, can effectively improve the off-design performance of motor, expand the living space of ultrasound velocity/hypersonic aircraft.
For realizing above technical purpose, the utility model will be taked following technological scheme:
A kind of ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device, described bypass type device comprises overexpansion jet pipe contraction section and extending section, characterized by further comprising be communicated with the bleed bypass of jet pipe contraction section and extending section, for the catch of opening and closing bleed bypass channel, bleed bypass utilizes the natural pressure reduction between the suction port of jet pipe contraction section and the air outlet of nozzle divergence cone to set up interior flow field; Described bleed bypass has comprised the suction port 1, first passage 2, the first corner 3, second channel 4, the second corner 5, third channel 6, the air outlet 7 that are sequentially communicated with, and described suction port 1 is communicated with jet pipe contraction section, and air outlet 7 is communicated with nozzle divergence cone.
As a preferred version of the present utility model, bleed bypass channel is for waiting width design, channel width be nozzle divergence cone and contraction section joint height 5% ~ 20%; The angle [alpha] of described bleed bypass channel suction port 1 is 90 °, the angle ranging from the angle between first passage and the tangent line of jet pipe contraction section; The angle of air outlet 7 is 90 ° ~ 160 °, the angle ranging from third channel the initial segment to the line of outlet side and the angle between horizontal line; Described suction port 1 and the position, design space of air outlet 7 are arranged within the scope of jet pipe contraction section and nozzle divergence cone 0 ~ 50% place; Third channel 6 incident angle β and horizontal line are counterclockwise 90 ° ~ 160 °.
As further preferred version of the present utility model, described suction port 1 is arranged on jet pipe contraction section 30.53%, described air outlet 7 is arranged on nozzle divergence cone 30.0%, described channel width be nozzle divergence cone and contraction section joint height 5.2%, it is 150.4 ° counterclockwise that described third channel 6 incident angle β and level are.
On the basis of technique scheme, the radius of turn arc length of the first corner 3, the second corner 5 is at least 4 times of channel width.
For realizing the control gear of the technological scheme of above-mentioned ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device, it is characterized in that, when ultrasound velocity/hypersonic aircraft shunting device is worked on design point, now aircraft transonic speed, while working under overexpansion state, control gear sends instruction, opens the catch of tap air channel; When the work of ultrasound velocity/hypersonic aircraft shunting device off-design point, now aircraft is when normal working, and control gear sends instruction, closes the catch of bleed bypass channel.
Catch can lay respectively at bleed bypass and import and export two ends, and is connected with flight control system actuation mechanism.
Catch also can be positioned at bleed bypass channel, and is connected with flight control system actuation mechanism.
Compared with prior art, adopt the effect of technique scheme to be: simple in structure, because the utility model has utilized the high pressure draught of jet pipe contraction section, only need bleed air line that pressurized gas are caused to nozzle divergence cone injection, avoid traditional air machinery because of the required interpolation of secondary gas drainage, can under the condition that does not increase impost, improve the thrust performance of jet pipe.
Bypass type bleed control gear described in the utility model to open with closing be to be controlled by flight control system, when aircraft is operated in transonic speed or serious overexpansion state lower time, the bleed bypass that internal flow channel is set between jet pipe contraction section and extending section is opened, bypass channel, when opening, causes extending section injection by the high pressure draught in jet pipe contraction section.Air-flow enters bypass channel from suction port 1, from air outlet 7, penetrate, this strand of Secondary Flow is from air outlet 7 is penetrated, interact with the main flow in jet pipe, the main flow in jet pipe is lifted, make main flow, at upper expanding noodles, a series of shock wave reflection occur, expanding noodles wall pressure integration is increased, and lower expanding noodles is because main flow is lifted, by environmental gas, to be filled, jet pipe thrust performance obtains gain.Now, nozzle thrust coefficient increases, and negative lift reduces.When ultrasound velocity/hypersonic aircraft is worked on design point, control gear sends instruction, closes bleed bypass channel, and bleed bypass is closed, and can not affect Nozzle Design point performance.
Internal flow channel being provided with, comprise suction port 1, first passage 2, the first corner 3, second channel 4, the second corner 5, third channel 6, the air outlet 7 that order is communicated with, wherein suction port 1 is arranged on jet pipe contraction section, and air outlet 7 is arranged on nozzle divergence cone.In the technical solution of the utility model, the pressure reduction of bleed bypass utilization between the suction port of jet pipe contraction section and the air outlet of extending section is set up interior flow field, and shunting device work pressure ratio scope can be controlled and need to choose according to aircraft, and determined by each design parameter size of bypass channel.The utility model is the design that is optimized to the position of internal flow channel, suction port, air outlet under the condition of considering coupling, can reduce the loss of air-flow in tap air channel, guarantee to obtain nozzle performance maximum gain with minimum bleed cost, further improve thrust performance.
The width of bypass channel, is the size of passage, and it has determined the size of bleed flow, and the size of bleed bypass channel width need to be optimized design according to the needs of aircraft.The width design that waits to bleed bypass channel, wherein, passage 3 width adopt jet pipe contraction section and extending section joint height to carry out nondimensionalization; The angle [alpha] of bleed bypass suction port 1, the incident angle β of third channel 6, the position of suction port 1 and the position of air outlet 7, the definite of these four parameters is the optimum results of having considered coupling.Optimization aim, for can obtain good thrust performance under the condition of bleed flow-rate ratio minimum, further improves thrust performance.
The angle [alpha] of bleed bypass channel suction port 1 is 90 °, and the angle between first passage 2 and the tangent line of jet pipe contraction section is 90 °, and it is 90 ° ~ 160 ° that third channel 6 incident angle β and level are counterclockwise angle; The angle of air outlet 7 is 90 ° ~ 160 °, and this angle is that third channel the initial segment is to the line of outlet side and the angle between horizontal line.Because airspeed in contraction section is larger, suction port 1 adopts the design perpendicular with jet pipe ground profile, is in order to guarantee that drainage is entering the larger flow losses of bypass Shi Buhui inlet generation; The angle design of air outlet 7 is in order to obtain best Secondary Flow jet injection effect; Before the position, design space of suction port 1 and air outlet 7 is arranged in jet pipe contraction section and within the scope of nozzle divergence cone 0 ~ 50% place, do not considering under the condition of coupling, the design space of suction port 1 is in jet pipe contraction section 0 ~ 50% scope, and more forward nozzle performance promotes larger; The design space of air outlet 7 is also within the scope of nozzle divergence cone 0 ~ 50% place, and more forward nozzle performance promotes larger; Therefore after suction port 1 and air outlet 7 can not too be leaned on, otherwise will affect the thrust performance of jet pipe; Third channel 6 incident angles and level are counterclockwise 90 ° ~ 160 °, are not considering that, under the condition of coupling, the angle of passage 6 is larger, and the performance boost of jet pipe thrust is larger.But because the related scheme work of the utility model pressure ratio scope need to be determined according to the needs of aircraft, therefore need to be according to the selected reasonable number range of effect of concrete scheme.In the angle [alpha] of considering bleed bypass suction port 1, be 90 °, the angle of air outlet 7, the coupling of the position of the angle beta of third channel 6, suction port 1 and these key parameters of position of air outlet 7, chosen above number range.
But because each factor exists certain coupling, the parameters combination of best performance can not make nozzle performance reach optimum.Having considered, bleed flow is smaller and after restriction that performance gain is higher, according to optimum results, a preferred embodiment of the present utility model is: when bypass channel width be jet pipe contraction section and extending section joint height 5.2%, suction port 1 be positioned at jet pipe import 30.53%, air outlet 7 to be positioned at nozzle divergence cone before 30.0%, and the angle beta of third channel 6 and horizontal sextant angle are while being 150.4 °, under the condition that is 3.22% in bleed flow-rate ratio, nozzle thrust coefficient raises 8.57%, and negative lift declines 95.1%.
From accompanying drawing 6, can find out, the line segment that triangle forms is that the working effect of aircraft jet pipe after shunting device has been installed, and the line segment that quadrangle forms is that the service behaviour of aircraft jet pipe after shunting device has been installed.After shunting device has been installed, the work pressure ratio scope maximum value of bypass channel is pressure ratio 10.5 ~ 11 o'clock, and nozzle performance is more superior than the benchmark nozzle performance that bypass is not installed.
On the basis of above technological scheme, the radius of turn arc length of the first corner 3, the second corner 5 is at least 4 times of channel width, and airflow breakaway does not occur when guaranteeing that air-flow passes through corner.If radius of turn is too small, during air-flow process corner, can produce flow separation, produce larger line loss.
It is symmetrical jet pipe or offset nozzle that technical solutions according to the invention are applicable to jet pipe simultaneously, has expanded the scope of the utility model application.Because no matter be symmetrical jet pipe or offset nozzle, the principle that they use is all identical, is all by Secondary Flow, to produce shock wave to make wall pressure raise to obtain the gain of thrust performance.
Accompanying drawing explanation
Body resistance characteristic after aircraft under the different flight Mach numbers of Fig. 1.
Jet pipe net thrust under the different flight Mach numbers of Fig. 2.
Fig. 3 structural representation of the present utility model.
Fig. 4 schematic perspective view of the present utility model.
The structural representation of Fig. 5 the utility model preferred embodiment.
Fig. 6 adopts after auto-adaptive control scheme, and jet pipe is performance chart in Design compression ratio.
In Fig. 1 and Fig. 2: Fig. 1 shows when aircraft is scope transonic speed (be Mach number 0.8 ~ 1.2 near), particularly near M=1, afterbody drag (Afterbody Drag) increases severely, and in Fig. 2, jet pipe net thrust (Thrust-Minus-Drag Performance) is more violent compared with hour decline in flight Mach number (Flight Mach Number) simultaneously.
In Fig. 3: 1, suction port, 2, first passage, the 3, first corner, 4, second channel, the 5, second corner, 6, third channel, 7, air outlet.
In Fig. 5: 8, air outlet is positioned at the position of nozzle divergence cone front 30.0%, 9, nozzle throat height, 10, suction port is positioned at the position of jet pipe import front 30.53%, 11, bypass channel width.
In Fig. 6: the line segment that triangle forms is that the working effect of aircraft jet pipe after shunting device has been installed, the line segment that quadrangle forms is not for installing the service behaviour of the aircraft jet pipe of shunting device.
Embodiment
Below in conjunction with drawings and Examples, the utility model is elaborated:
As shown in Fig. 3, Fig. 4 and Fig. 5, ultrasound velocity described in the utility model/hypersonic bypass type bleed control gear, comprise overexpansion jet pipe contraction section and extending section, and the bleed bypass channel between jet pipe contraction section and extending section is set, bypass channel connecting spray nozzle contraction section and extending section, described bypass channel comprises suction port 1, first passage 2, the first corner 3, second channel 4, the second corner 5, third channel 6, air outlet 7.Suction port 1 is arranged on jet pipe contraction section, and the design space of suction port 1 is preferably within the scope of 0 ~ 50% place, jet pipe import front portion; Air outlet 7 is arranged on nozzle divergence cone, and the design space of air outlet 7 is preferably within the scope of 0 ~ 50% place, nozzle divergence cone front portion.Bleed bypass suction port 1 α angle is 90 °, and exporting 7 angles is 90 ° ~ 160 °, and bypass width is the height 5% ~ 20% of nozzle divergence cone and contraction section joint.First passage 2 is orthogonal with the tangent line of jet pipe contraction section; The radius of turn arc length of the first corner 3 is at least 4 times of channel width, guarantees that air-flow is not separated by making; Second channel 4 is a straight line; The design of the second corner 5 is identical with the first corner 3; Third channel 6 reference angle β and level are 90 ° ~ 160 ° (counterclockwise) angles.
The given bypass channel structure of the utility model is optimum results, the lower getable maximum thrust performance gain of condition that optimization aim is bleed flow-rate ratio minimum.
At this, provide an embodiment, when bypass channel width is the height 5.2% of nozzle divergence cone and contraction section joint, third channel 6 incident angle β are 150.4 ° (angles counterclockwise), suction port 1 is positioned at jet pipe import front 30.53%, air outlet 7 is positioned at nozzle divergence cone front 30.0%, under the condition that is 3.22% in bleed flow-rate ratio, the work pressure ratio scope maximum value of bypass channel is pressure ratio 10.5 ~ 11 o'clock, as can be seen from Figure 6, the line segment that triangle forms is that the working effect of aircraft jet pipe after shunting device has been installed, the line segment that quadrangle forms is not for installing the service behaviour of the aircraft jet pipe of shunting device, the nozzle performance that shunting device has been installed is obviously more superior than the benchmark nozzle performance that bypass is not installed.After shunting device has been installed, when ultrasound velocity/hypersonic aircraft shunting device is worked on design point, now aircraft transonic speed, while working under overexpansion state, control gear sends instruction, open bleed bypass channel, when bypass channel is when opening, the high pressure draught in jet pipe contraction section is caused to extending section injection.This strand of Secondary Flow is from air outlet 7 is penetrated, main flow in jet pipe is lifted, make main flow, at upper expanding noodles, a series of shock wave reflection occur, expanding noodles wall pressure integration is increased, and lower expanding noodles is because main flow is lifted, by environmental gas, filled, therefore, nozzle thrust coefficient increases, and negative lift reduces.Experiment showed, that nozzle thrust coefficient has raise 8.57%, negative lift declines 95.1%.
If certainly think further to improve jet pipe thrust performance, on the basis of given technological scheme, can realize by expanding the width of bypass channel in the above, also can by expanding the width of bypass channel, realize separately.
The shape of body/pipeline described in the utility model and section area can design according to the concrete condition of ultrasound velocity/hypersonic aircraft.
Claims (6)
1. ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device, described bypass type device comprises overexpansion jet pipe contraction section and extending section, characterized by further comprising be communicated with the bleed bypass of jet pipe contraction section and extending section, for the catch of opening and closing bleed bypass channel, bleed bypass utilizes the natural pressure reduction between the suction port of jet pipe contraction section and the air outlet of nozzle divergence cone to set up interior flow field; Described bleed bypass has comprised the suction port (1), first passage (2), the first corner (3), second channel (4), the second corner (5), third channel (6), air outlet (7) that are sequentially communicated with, described suction port (1) is communicated with jet pipe contraction section, and air outlet (7) are communicated with nozzle divergence cone.
2. according to ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device claimed in claim 1, it is characterized in that, described bleed bypass channel is for waiting width design, channel width be nozzle divergence cone and contraction section joint height 5%～20%; The angle of described bleed bypass channel suction port (1) is 90 °, the angle ranging from the angle between first passage and the tangent line of jet pipe contraction section; The angle of air outlet (7) is 90 °～160 °, the angle ranging from third channel (6) the initial segment to the line of outlet side and the angle between horizontal line; The position, design space of described suction port (1) and air outlet (7) is arranged within the scope of jet pipe contraction section and nozzle divergence cone 0～50% place, and third channel (6) incident angle and level are counterclockwise 90 °～160 °.
3. according to ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device claimed in claim 2, it is characterized in that, described suction port (1) is arranged on jet pipe contraction section 30.53%, described air outlet (7) is arranged on nozzle divergence cone 30.0%, described channel width be nozzle divergence cone and contraction section joint height 5%, it is 150.4 ° counterclockwise that described third channel (6) incident angle and level are.
4. according to ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device claimed in claim 3, it is characterized in that, the radius of turn arc length of the first corner (3), the second corner (5) is at least 4 times of channel width.
5. according to the arbitrary ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device described in claim 1-4, it is characterized in that, described catch lays respectively at bleed bypass and imports and exports two ends.
6. according to the arbitrary ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device described in claim 1-4, it is characterized in that, described catch is positioned at bleed bypass channel.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201320281995.4U CN203441627U (en) | 2013-05-21 | 2013-05-21 | Supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201320281995.4U CN203441627U (en) | 2013-05-21 | 2013-05-21 | Supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device |
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| CN203441627U true CN203441627U (en) | 2014-02-19 |
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| CN201320281995.4U Expired - Lifetime CN203441627U (en) | 2013-05-21 | 2013-05-21 | Supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device |
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Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103291495A (en) * | 2013-05-21 | 2013-09-11 | 南京航空航天大学 | Supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device |
| CN106014684A (en) * | 2016-05-30 | 2016-10-12 | 西北工业大学 | Combined flow control method and structure for improving SERN for TBCC |
| CN107120210A (en) * | 2017-06-25 | 2017-09-01 | 北京航天三发高科技有限公司 | A kind of supersonic nozzle |
| CN107655694A (en) * | 2017-08-24 | 2018-02-02 | 南京理工大学 | A kind of supersonic nozzle jet blends experimental provision |
| CN110450964A (en) * | 2018-05-07 | 2019-11-15 | 南京普国科技有限公司 | Class axial symmetry tilt outlet, which is received, expands jet pipe and its design method |
| CN115075981A (en) * | 2021-03-15 | 2022-09-20 | 中国科学院沈阳自动化研究所 | Thrust vectoring nozzle adopting flow control |
-
2013
- 2013-05-21 CN CN201320281995.4U patent/CN203441627U/en not_active Expired - Lifetime
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103291495A (en) * | 2013-05-21 | 2013-09-11 | 南京航空航天大学 | Supersonic/hypersonic aerocraft engine overexpansion nozzle bypass type device |
| CN103291495B (en) * | 2013-05-21 | 2016-04-06 | 南京航空航天大学 | Ultrasound velocity/hypersonic aircraft motor overexpansion jet pipe bypass type device |
| CN106014684A (en) * | 2016-05-30 | 2016-10-12 | 西北工业大学 | Combined flow control method and structure for improving SERN for TBCC |
| CN107120210A (en) * | 2017-06-25 | 2017-09-01 | 北京航天三发高科技有限公司 | A kind of supersonic nozzle |
| CN107120210B (en) * | 2017-06-25 | 2023-05-23 | 北京航天三发高科技有限公司 | Supersonic jet pipe |
| CN107655694A (en) * | 2017-08-24 | 2018-02-02 | 南京理工大学 | A kind of supersonic nozzle jet blends experimental provision |
| CN110450964A (en) * | 2018-05-07 | 2019-11-15 | 南京普国科技有限公司 | Class axial symmetry tilt outlet, which is received, expands jet pipe and its design method |
| CN110450964B (en) * | 2018-05-07 | 2020-11-24 | 南京普国科技有限公司 | Axisymmetric inclined outlet convergent-divergent nozzle and design method thereof |
| CN115075981A (en) * | 2021-03-15 | 2022-09-20 | 中国科学院沈阳自动化研究所 | Thrust vectoring nozzle adopting flow control |
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| C14 | Grant of patent or utility model | ||
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| AV01 | Patent right actively abandoned |
Granted publication date: 20140219 Effective date of abandoning: 20160406 |
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| C25 | Abandonment of patent right or utility model to avoid double patenting |