CN113623086B - Shock wave/boundary layer interference controller - Google Patents
Shock wave/boundary layer interference controller Download PDFInfo
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- CN113623086B CN113623086B CN202110812192.6A CN202110812192A CN113623086B CN 113623086 B CN113623086 B CN 113623086B CN 202110812192 A CN202110812192 A CN 202110812192A CN 113623086 B CN113623086 B CN 113623086B
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- 230000035939 shock Effects 0.000 title claims abstract description 55
- 238000000926 separation method Methods 0.000 claims abstract description 38
- 239000012530 fluid Substances 0.000 claims abstract description 8
- 230000003068 static effect Effects 0.000 claims abstract description 8
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 5
- 238000005452 bending Methods 0.000 claims description 6
- 230000004323 axial length Effects 0.000 claims description 3
- 230000003631 expected effect Effects 0.000 abstract 1
- 238000012795 verification Methods 0.000 abstract 1
- 230000006835 compression Effects 0.000 description 3
- 238000007906 compression Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 3
- 230000000694 effects Effects 0.000 description 3
- 238000011084 recovery Methods 0.000 description 3
- 238000007664 blowing Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 230000008602 contraction Effects 0.000 description 1
- 230000006866 deterioration Effects 0.000 description 1
- 238000007599 discharging Methods 0.000 description 1
- 230000002706 hydrostatic effect Effects 0.000 description 1
- 238000003359 percent control normalization Methods 0.000 description 1
- 239000002243 precursor Substances 0.000 description 1
- 238000004080 punching Methods 0.000 description 1
- 238000001228 spectrum Methods 0.000 description 1
- 230000008719 thickening Effects 0.000 description 1
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
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- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention provides a shock wave/boundary layer interference controller, which is characterized in that high static pressure fluid at the downstream of a turbulence separation packet is introduced into a return channel, is guided by a fin-shaped vortex generator and then is sprayed out from the upstream of the turbulence separation packet to generate a vortex, and the vortex structure is utilized to convey high-energy fluid to the turbulence separation packet so as to reduce flow loss caused by separation. The invention has simple structure and easy realization, and can obtain expected effect after example verification.
Description
Technical Field
The invention relates to the technical field of supersonic flow control, in particular to control of a separation flow field of a turbulent boundary layer induced by incident shock waves.
Background
The hypersonic air inlet is one of the important components of a scramjet engine, and needs to be capable of performing a stamping action on high-speed incoming flow within an extremely limited time and space scale and providing air with the highest possible total pressure and the most uniform possible air flow distribution for a subsequent combustion chamber. The punching effect of the air inlet channel on the incoming flow is mainly completed through a shock wave structure.
Boundary layers are inevitably generated on the precursor body and the compression surface, once flow separation is induced by shock wave/boundary layer interference, the thickness of the boundary layer is increased rapidly, and meanwhile, the main flow is obviously influenced, and an extremely complex wave system structure is formed, which is often accompanied with separation of shock waves, a shear layer, expansion waves and reattachment of a compression wave system structure. The rapid thickening of the boundary layer and the additional shock wave compression greatly reduce the aerodynamic performance of the air inlet channel, even cause the flow of an inner flow channel to be blocked, further fall into an un-starting state, and finally cause the performance deterioration of the propulsion system of the aircraft.
In order to realize the control of the shock wave/boundary layer interference flow, the most mainstream scheme at present is to add a boundary layer leakage flow or an active blowing device on the wall surface of the separation zone. The boundary layer leakage flow is to reduce the size of a separation packet by discharging low-energy fluid near the wall to improve the capability of the boundary layer to resist the inverse pressure gradient and reduce the pressure ratio after the shock wave, but certain flow loss and total pressure loss are caused. The active blowing device injects external high-energy fluid into the boundary layer, changes the velocity distribution in the boundary layer so as to improve the anti-separation capability of the boundary layer, but is difficult to find a stable external high-pressure air source.
Disclosure of Invention
The invention provides a shock wave/boundary layer interference controller, aiming at reducing the flow loss caused by the strong-incidence shock wave/boundary layer interference induced separation packet.
In order to achieve the purpose, the invention provides the following scheme:
a shock wave/boundary layer interference controller comprises a shock wave generator wedge surface and a flat plate; an inner channel is formed between the wedge surface of the laser generator and the flat plate; a backflow channel is arranged in the flat plate and comprises a main channel extending along the extension direction of the inner channel, a front connecting channel extending from the front end of the main channel in a bending way and communicated with the inner channel, and a rear connecting channel extending from the rear end of the main channel in a bending way and communicated with the inner channel; the front connecting runner is bent, inclined and extended from the front end of the main runner until being communicated with the inner channel; the rear connecting flow passage is bent from the rear end of the main flow passage, is inclined forwards and extends until being communicated with the inner passage; a vortex generator is arranged in the front connecting flow passage; and a drainage seam is arranged at the joint of the rear connecting flow passage and the inner channel.
Further, the vortex generator comprises a plurality of fins which are arranged in pairs, and the fins in the pairs are symmetrically arranged in a splayed shape so as to generate a vortex structure when the high-pressure fluid is sprayed out.
Furthermore, when the incoming flow generates an incident shock wave in the inner channel under the action of the wedge surface of the shock wave generator, a turbulence separation packet and a separation shock wave surface extending from the turbulence separation packet to the downstream of the wedge surface of the shock wave generator are generated at the position, close to the flat plate, of the inner channel, and the drainage slit is arranged at the downstream of the reattachment station of the turbulence separation packet and used for capturing the high-pressure airflow in the area. The turbulent boundary layer is separated from the flat plate by the inverse pressure gradient, and the flow is reattached to the near wall under the influence of an external wave system structure, and the reattachment station refers to the flow direction position when the flow is reattached. Generally, the pressure here is significantly greater than the incoming hydrostatic pressure, accompanied by a strong counter pressure gradient. It has been verified that the placement of the bleed slot downstream of the reattachment station provides the best control.
Further, let outOne or more flow slits arranged downstream L of the reattachment station of the turbulent separation package 2x Satisfy L 2x ≤0.2L ref Length of flow direction L of the run-off slit 2 The flow direction length is equivalent to that of a local high static pressure area and meets 0.05L ref ≤L 2 ≤0.15L ref Wherein L is ref Is the length of the original split packet without control.
Furthermore, the inclined extending direction of the front connecting flow passage forms an included angle alpha with the surface of the flat plate 1 (ii) a The inclined extending direction of the rear connecting flow passage forms an included angle alpha with the surface of the flat plate 2 ；35°≤α 1 ≤55°；35°≤α 2 ≤55°。
Further, a fin-shaped vortex generator intersecting the plate wall upstream of the separation shock wave surface; the distance between the position where the separated shock wave surface is generated and the front connecting flow passage is L 1x Satisfy 0.5L ref ≤L 1x ≤L ref 。
Furthermore, the maximum width of the backflow channel is d and meets 0.6L 2 ≤d≤2L 2 。
Further, the bottom surface of the vortex generator is parallel to the flat plate; the vortex generator occupies a total axial length L 1 Satisfy 0.3L 2 ≤L 1 ≤1.2L 2 Each fin forms an angle beta with the horizontal direction 1 Satisfies the condition that the beta is not less than 20 degrees 1 Not more than 40 degrees, and the width of each channel is D 1 The minimum distance between adjacent fins is L 3 Satisfy D 1 ≤0.1L ref ，2D 1 ≤L 3 ≤0.3L ref 。
Further, the device is applied to an air inlet channel of an aircraft, and the wedge surface of the shock wave generator is an air inlet channel lip cover; the flat plate is the inner wall of the air inlet channel.
The technical scheme of the invention has the following beneficial effects:
because of the existence of the incident shock wave, the pressure intensity at the downstream of the turbulence separation packet is obviously higher than the pressure intensity at the upstream of the separation shock wave surface, namely the incoming static pressure, so that the static pressure difference in the backflow channel is increased, the high-pressure fluid entering the backflow channel from the discharge slit is sprayed out from the vortex generator, and a vortex structure is generated under the action of the profile of the vortex generator, and the vortex structure develops along the flow direction and simultaneously draws high-energy main flow into the turbulence separation packet to carry out work. On one hand, the low-energy airflow in part of the boundary layer flows out from the leakage seams, so that the speed distribution in the boundary layer is improved, and the total pressure loss in the main flow is reduced; on the other hand, the swirl flow generated by the vortex generator can convey the energy in the main flow into the turbulence separation packet, inhibit separation and reduce flow loss. These two factors combine to significantly reduce the total pressure loss through the separation package.
Drawings
FIG. 1 is a three-dimensional isometric view of a shock/boundary layer disturbance controller provided by the present invention.
Fig. 2 is a schematic two-dimensional size diagram of a shock wave/boundary layer interference controller provided by the invention, wherein the upper diagram is a front view, and the lower diagram is a top view.
FIG. 3 is a flow spectrum of a reference scheme of a specific application example of the shock wave/boundary layer interference controller provided by the invention.
FIG. 4 is a flow chart of a control scheme of a specific application example of the shock wave/boundary layer interference controller provided by the invention.
Detailed Description
As shown in fig. 1, the shock wave/boundary layer interference controller provided by the invention comprises a shock wave generator wedge surface 1 and a flat plate 2; an inner channel 3 is formed between the wedge surface 1 of the laser generator and the flat plate 2. When the shock wave/boundary layer interference controller is applied to an aircraft air inlet structure, the wedge surface 1 of the shock wave generator is an air inlet lip cover; the flat plate 2 is the inner wall of the air inlet channel. When the incoming flow is acted by the wedge surface 1 of the laser generator, an incident shock wave 7 is generated in the inner channel 3, a turbulent flow separation packet 8 is generated at the position, close to the flat plate 2, of the inner channel 3, and a separation shock wave surface 9 extends from the turbulent flow separation packet 8 to the downstream of the wedge surface 1 of the laser generator.
The flat plate 2 is provided with a return channel 4 inside, and the return channel 4 comprises a main channel 41 extending along the extension direction of the inner channel, a front connecting channel 42 extending from the front end of the main channel 41 in a bending way and communicating with the inner channel 3, and a rear connecting channel 43 extending from the rear end of the main channel 41 in a bending way and communicating with the inner channel 3. What is needed isThe front connecting flow passage 42 is bent, inclined and extended from the front end of the main flow passage 41 until being communicated with the inner passage 3; then, the connecting duct 43 is bent from the rear end of the main duct 41 and extends forward until communicating with the inner duct 3. In the present embodiment, as shown in fig. 2, the front connecting flow channel extends obliquely at an angle α with respect to the surface of the flat plate 1 (ii) a The inclined extending direction of the rear connecting flow passage forms an included angle alpha with the surface of the flat plate 2 ；35°≤α 1 ≤55°；35°≤α 2 Is less than or equal to 55 degrees. The maximum width of the return channel is d and satisfies 0.6L 2 ≤d≤2L 2 。
As shown in fig. 1 and 2, a vortex generator 5 is provided in the front connection flow passage 42; and a drainage slit 6 is arranged at the joint of the rear connecting flow passage 43 and the inner channel 3. One or more discharge slits arranged downstream L of the reattachment station of the turbulent separation package 2x Satisfy L 2x ≤0.2L ref Length of flow direction L of the run-off slit 2 The flow direction length is equivalent to that of a local high static pressure area and meets 0.05L ref ≤L 2 ≤0.15L ref Wherein L is ref Is the length of the original split packet without control.
As shown in fig. 2, the vortex generator includes a plurality of fins in pairs, and the fins in pairs are symmetrically arranged in a shape of a Chinese character 'ba' to generate a vortex structure when the high-pressure fluid is ejected. The finned vortex generator intersects the flat wall upstream of the separation shock wave surface; the distance between the position where the separated shock wave surface is generated and the front connecting flow passage is L 1x Satisfy 0.5L ref ≤L 1x ≤L ref . The bottom surface of the vortex generator is parallel to the flat plate; the vortex generator occupies a total axial length L 1 Satisfy 0.3L 2 ≤L 1 ≤1.2L 2 Each fin forms an angle beta with the horizontal direction 1 Satisfies the condition that the beta is not less than 20 degrees 1 Not more than 40 degrees, and the width of each channel is D 1 The minimum distance between adjacent fins is L 3 Satisfy D 1 ≤0.1L ref ，2D 1 ≤L 3 ≤0.3L ref 。
The shock/boundary layer disturbance controller effect is demonstrated below in conjunction with a specific configuration:
mach number Ma of incoming flow 0 Static pressure p ═ 3 0 6998Pa, static temperature t 0 Reynolds number based on boundary layer thickness of 107K is about 1.83 × 10 5 The deflection angle of the shock wave generator was 13 °, and the internal contraction ratio (inlet area/outlet area) of the flow channel was about 1.47. The flow patterns without control scheme and control scheme are shown in fig. 3 and fig. 4, respectively, the key geometric parameters of the control scheme are shown in table 1, and the total pressure recovery coefficient of the control scheme and the same station downstream of the separation packet of the reference scheme is compared with table 2. It can be seen that after the controller is adopted to control the incident shock wave/boundary layer interference, the downstream of the vortex generator has an obvious vortex structure and the form of the separation packet is also changed as can be seen from the spanwise section. And the overall pressure recovery coefficient on the same flow direction station section downstream of the separation packet is improved by 1.9%, which shows that the shock wave/boundary layer interference controller is effective in reducing the separation flow loss.
TABLE 1
| L 1 | L 2 | L 3 | α 1 | α 2 | β 1 | H 1 | D 1 |
| 5mm | 10mm | 20mm | 45 | 45 | 30 | 9mm | 2.5mm |
TABLE 2
| Total pressure recovery coefficient | |
| Prototype | 76.1％ |
| Control scheme | 78％ |
The implementation form of the present invention is explained by applying specific examples, which are intended to illustrate the core idea of the present invention and the specific control effect that can be achieved by the core idea of the present invention, and for a person skilled in the art, on the premise of applying the idea of the present invention, the specific implementation manner and the application range may be changed, and these should also be regarded as the protection scope of the present invention.
Claims (8)
1. A shock wave/boundary layer interference controller comprises a shock wave generator wedge surface and a flat plate; an inner channel is formed between the wedge surface of the laser generator and the flat plate; the device is characterized in that a backflow channel is arranged in the flat plate, and the backflow channel comprises a main channel extending along the extension direction of the inner channel, a front connecting channel extending from the front end of the main channel in a bending way and communicated with the inner channel, and a rear connecting channel extending from the rear end of the main channel in a bending way and communicated with the inner channel; the front connecting runner is bent, inclined and extended from the front end of the main runner until being communicated with the inner channel; the rear connecting flow passage is bent from the rear end of the main flow passage, is inclined forwards and extends until being communicated with the inner passage; a vortex generator is arranged in the front connecting flow passage; a drainage seam is arranged at the joint of the rear connecting flow passage and the inner channel;
the vortex generator comprises a plurality of fins which are arranged in pairs, and the fins in pairs are symmetrically arranged in a splayed mode so that a vortex structure is generated when high-pressure fluid is sprayed out.
2. The shock/boundary layer disturbance controller according to claim 1, wherein when the incoming flow generates an incident shock in the inner channel under the action of the wedge surface of the shock generator, and generates a turbulent separation packet at the position close to the flat plate in the inner channel and a separation shock surface extending from the turbulent separation packet to the downstream of the wedge surface of the shock generator, the bleed slot is arranged at the downstream of the reattachment station of the turbulent separation packet to capture the high pressure air flow in the area.
3. The shock/boundary layer disturbance controller according to claim 2, wherein the bleed slot is one or more, and is disposed downstream L of the turbulent separation packet reattachment station 2x Satisfy L 2x ≤0.2L ref Length of flow direction L of the run-off slit 2 The flow direction length is equivalent to that of a local high static pressure area and meets 0.05L ref ≤L 2 ≤0.15L ref Wherein L is ref Is the length of the original split packet without control.
4. The shock/boundary layer disturbance controller according to claim 3, wherein the inclined extension direction of the front connecting flow channel forms an included angle α with the surface of the flat plate 1 (ii) a Rear connecting flow passageThe inclined extending direction forms an included angle alpha with the surface of the flat plate 2 ；
；
。
5. The shock/boundary layer disturbance controller of claim 4, wherein the fin-shaped vortex generators intersect the planar wall upstream of the separation shock surface; the distance between the position where the separated shock wave surface is generated and the front connecting flow passage is L 1x Satisfy 0.5L ref ≤L 1x ≤L ref 。
6. The shock/boundary layer interference controller of claim 5, wherein the maximum width of the return channel is d, satisfying 0.6L 2 ≤d≤2L 2 。
7. The shock/boundary layer interference controller of claim 5 wherein the vortex generator bottom surface is parallel to the flat plate; the vortex generator occupies a total axial length L 1 Satisfy 0.3L 2 ≤L 1 ≤1.2L 2 Each fin forms an angle beta with the horizontal direction 1 Satisfy the following requirements
Each channel having a width D 1 The minimum distance between adjacent fins is L 3 Satisfy D 1 ≤0.1L ref ，2D 1 ≤L 3 ≤0.3L ref 。
8. The shock/boundary layer interference controller according to any of claims 1 to 7, applied to an aircraft inlet, wherein the shock generator wedge surface is an inlet lip shroud; the flat plate is the inner wall of the air inlet channel.
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5424824A (en) * | 1993-05-12 | 1995-06-13 | The Boeing Company | Method and apparatus for normal shock sensing within the focal region of a laser beam |
| 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 |
| CN103953448A (en) * | 2014-04-15 | 2014-07-30 | 南京航空航天大学 | Hypersonic air inlet channel |
| CN107013367A (en) * | 2017-02-17 | 2017-08-04 | 北京空天技术研究所 | Turbine base double combustion chamber's punching press combined cycle engine |
| 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 |
| 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 |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9429071B2 (en) * | 2011-06-23 | 2016-08-30 | Continuum Dynamics, Inc. | Supersonic engine inlet diffuser with deployable vortex generators |
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2021
- 2021-07-19 CN CN202110812192.6A patent/CN113623086B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5424824A (en) * | 1993-05-12 | 1995-06-13 | The Boeing Company | Method and apparatus for normal shock sensing within the focal region of a laser beam |
| 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 |
| CN103953448A (en) * | 2014-04-15 | 2014-07-30 | 南京航空航天大学 | Hypersonic air inlet channel |
| CN107013367A (en) * | 2017-02-17 | 2017-08-04 | 北京空天技术研究所 | Turbine base double combustion chamber's punching press combined cycle engine |
| 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 |
| 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 |
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