CN102152848A - Super-speed aircraft frontal edge impact and small snakelike channel cooling structure - Google Patents
Super-speed aircraft frontal edge impact and small snakelike channel cooling structure Download PDFInfo
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- CN102152848A CN102152848A CN2011100693092A CN201110069309A CN102152848A CN 102152848 A CN102152848 A CN 102152848A CN 2011100693092 A CN2011100693092 A CN 2011100693092A CN 201110069309 A CN201110069309 A CN 201110069309A CN 102152848 A CN102152848 A CN 102152848A
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Abstract
The invention discloses a super-speed aircraft frontal edge impact and small snakelike channel cooling structure, comprising an impact cavity, impact holes and a small snakelike channel, wherein the impact cavity is arranged at the head part of the super-speed aircraft frontal edge, one side of the impact cavity close to a machine body is provided with a row of impact holes at a middle axis position, the two ends of the impact hole are communicated with the impact cavity and an air supply cavity, two small snakelike channels are arranged at the positions corresponding to the impact holes on the upper surface and the lower surface of the impact cavity close to the surface of a super-speed aircraft wedge, the two ends of the small snakelike channel are communicated with the impact cavity and atmosphere at the tail part, the small snakelike channel is a continuous U-shaped curve channel, the vertical distance before the U-shaped channel is curved is equal to the distance between the two impact holes, and an inlet position, an outlet position and the impact holes are arranged in the same vertical plane. In the invention, according to the characteristic of heat exchange of the super-speed aircraft frontal edge, an impact cooling structure with high heat exchange capacity is adopted in a high heat-flow density region, and the small snakelike channels are adopted on the surface of the wedge with low heat-flow density and large heat exchange area, thus the heat exchange efficiency is greatly improved.
Description
Technical field
The present invention relates to a kind of leading edge of hypersonic vehicle impact+small serpentine channel cooling structure, belong to the heat exchange fields that local high heat flux is higher, other large area region density of heat flow rate are not too high such as Aeronautics and Astronautics, power machine.
Background technology
The proposition of pneumatic heating problems is because hypersonic aircraft development and the needs that develop.Aircraft is during with supersonic speed or hypersonic flight, and air is subjected to strong compression and violent rubbing effect, and most of kinetic energy is converted into heat energy, causes aircraft ambient air temperature sharply to raise.Have the very big temperature difference between this high-temperature gas and the aircraft surface, partial heat energy is rapidly to the object plane transmission, and this thermal energy transfer mode is called pneumatic heating.The serious high temperature that pneumatic heating produced can reduce the strength limit of material and the load-carrying capacity of Flight Vehicle Structure, makes structure produce thermal deformation, destroys the aerodynamic configuration of parts and influences the safe flight of aircraft.
The thermal protection problem in high heat flux zones such as leading edge stationary point is one of hypersonic aircraft design-calculated key issue, has become critical restraining factors and technical bottleneck in the hypersonic vehicle development process.Near hypersonic aircraft density of heat flow rate the leading edge stationary point when flight is very big (up to 10
6W/m
2More than), and produce solid wall surface localized hyperthermia (more than the 3000K), and might cause the change of aircraft profile, structural strength and rigidity, have a strong impact on the safety performance and the life-span of supersonic aircraft.Therefore, especially important for thermal protection research status in the hypersonic aircraft thermal protective system at position, leading edge stationary point.The pneumatic heating characteristics of hypersonic aircraft are: 1, and the flight time is longer, reaches dozens of minutes to several hrs; 2, the leading edge density of heat flow rate distributes and is bell-shaped profile, near the density of heat flow rate maximum stationary point, and longshore current sharply reduces to density of heat flow rate.Traditional passive cooling, cool off as radiation, reach high radiant heat flux density, then need very high surface temperature, therefore can't satisfy strength of material and requirement in service life, though and the ablation layer thermal protection structure can satisfy the requirement of high heat flux, the flight time is reached the hypersonic aircraft of dozens of minutes even several hours, then can't use.Therefore develop novel active cooling method, become the gordian technique of hypersonic vehicle design and development.
Summary of the invention
The objective of the invention is is in order to address the above problem, to propose a kind of leading edge of hypersonic vehicle impact+small serpentine channel cooling structure.
A kind of leading edge of hypersonic vehicle impact+small serpentine channel cooling structure of the present invention comprises and impacts chamber, impact opening and small serpentine channel.
Impact the chamber and be positioned at high-speed aircraft leading edge head, impact the chamber and offer row's impact opening near body one middle side part axial location,, the impact opening two ends are communicated with impacts chamber and air feed chamber, press close to the position of the corresponding impact opening of upper and lower surface on high-speed aircraft sphenoid surface in the impact chamber and offer two small straight channel, small serpentine channel two ends are communicated with impacts chamber and afterbody atmosphere, small serpentine channel is continuous U-shaped bending channel, width is two distances between impact opening, and entrance location, exit position and impact opening are positioned at same perpendicular.
The invention has the advantages that:
(1) the present invention is directed to the leading edge of hypersonic vehicle heat exchange characteristic, adopt the impact cooling structure of high exchange capability of heat in the high heat flux zone, at density of heat flow rate is not very big but small serpentine channel heat exchange is adopted on sphenoid that heat interchanging area is bigger surface, and heat exchange efficiency significantly improves;
(2) adopt millimeter cooling structure of level yardstick, do not change the pneumatic external form of hypersonic vehicle.
Description of drawings
Fig. 1 is an overall schematic of the present invention;
Fig. 2 is model inner structure of the present invention (the B-B cross section of a Fig. 3) scheme drawing;
Fig. 3 is an A-A schematic cross-section among Fig. 2;
Among the figure:
1-impacts the chamber, 2-impact opening, the small serpentine channel of 3-, 4-high-speed aircraft sphenoid surface, 5-leading edge, 6-air feed chamber, 7-rib
The specific embodiment
The present invention is described in further detail below in conjunction with drawings and Examples.
For reducing aerodynamic drag in the flight course, the surperficial external form of hypersonic aircraft is generally glut shape, and head generally is designed to the circular arc of millimeter magnitude.This has determined that its structure of initiatively cooling off should be millimetre-sized miniature scale structure.The pneumatic heating characteristics of hypersonic aircraft are: near density of heat flow rate maximum leading edge 5 stationary points, and longshore current is to sharply reducing, in sphenoid surface heat flow distribution uniform.
Based on the pneumatic heating characteristics and the profile characteristics of hypersonic vehicle above-mentioned, the present invention is a kind of leading edge of hypersonic vehicle impact+small serpentine channel cooling structure, comprises impacting chamber 1, impact opening 2 and small serpentine channel 3.High-speed aircraft leading edge inside is opened and is impacted chamber 1, impacts chamber 1 and offers row's impact opening 2 near body one middle side part axial location, and the diameter of impact opening 2 is 0.5mm～1.0mm, and spacing is 5.0mm～20.0mm, and impact opening 2 two ends are communicated with impacts chamber 1 and air feed chamber 6.Press close to the position of the corresponding impact opening 2 of upper and lower surface on high-speed aircraft sphenoid surface 4 in impact chamber 1 and offer two small serpentine channels, small serpentine channel 3 with impact that the chamber links to each other the import sectional dimension for being (0.5mm～1.0mm) * (0.5mm～1.0mm), entrance location, exit position and impact opening 2 are positioned at same perpendicular, small serpentine channel 3 is continuous U type bending channel, the vertical distance of each U type passage before revolution is two distances between the impact opening, and the width of the rib 7 between the U-shaped bending channel (fin is exactly the solid portion between the passage) is 0.5mm～1.0mm.Small serpentine channel 3 two ends are communicated with impacts chamber 1 and afterbody atmosphere.
Cooling media enters from impact opening with certain speed and impacts chamber 1, carry out heat exchange with leading edge of hypersonic vehicle 5 inside, then from the afterbody diffluence of the small serpentine channel 3 of both sides to aircraft, the leading edge 5 and the afterbody of aircraft are cooled off, reduce the temperature on hypersonic vehicle surface, drain into atmosphere by afterbody.
The present invention not only improved single-piece heat exchange effect, and bulk temperature is evenly distributed from thermodynamic (al) angle.
Embodiment:
The present invention opens in high-speed aircraft leading edge inside and impacts chamber 1, the impact opening 2 that diameter is 0.5mm～1.0mm is opened at the middle part, impact opening 2 spacings are 5.0mm～20.0mm, open the small serpentine channel 3 identical with impact opening 2 quantity in aircraft upper and lower surface 4 inside, small serpentine channel 3 with impact import cross section that the chamber links to each other and be of a size of to (0.5mm～1.0mm) * (0.5mm～1.0mm), the width of the fin between the small serpentine channel 3 is 0.5mm～1.0mm.Cooling media enters from circular impact hole 2 and impacts chamber 1, form large-area impact cooled region at leading edge 5 inside faces, impact jet flow has the characteristics of the high coefficient of heat transfer, therefore cooling media can be in impacting chamber 1 carries out good heat exchange with leading edge 5 inwalls of hypersonic vehicle, reduces near the temperature in leading edge 5 stationary points.Carry out cooling media behind the thermal exchange along of the afterbody diffluence of small serpentine channel 3 with leading edge 5 inside faces, further the sphenoid surface of hypersonic vehicle is cooled off, the effect of cooling media is fully played to hypersonic vehicle.
As shown in Figure 2, the diameter of impact opening 2 is 0.5mm～1.0mm, impact opening 2 spacings are 5.0mm～20.0mm, small serpentine channel 3 with impact import cross section that the chamber links to each other and be of a size of to (0.5mm～1.0mm) * (0.5mm～1.0mm), the width of the fin between the small serpentine channel 3 is 0.5mm～1.0mm.The density of heat flow rate on the sphenoid surface of hypersonic vehicle is not very high, but heat interchanging area is bigger, adopts small serpentine channel 3 heat exchange just in time can satisfy this heat exchange characteristic.The existence of small serpentine channel 3 not only can increase the heat interchanging area between cooling media and the high-temp solid, and can strengthen the disturbance of cooling media, thereby can make heat exchange more abundant, thereby makes the drop in temperature on sphenoid surface.
Claims (4)
1. a leading edge of hypersonic vehicle impact+small serpentine channel cooling structure is characterized in that, comprises impacting chamber, impact opening and small serpentine channel;
Impact the chamber and be positioned at high-speed aircraft leading edge head, impact the chamber and offer row's impact opening near body one middle side part axial location, the impact opening two ends are communicated with impacts chamber and air feed chamber, press close to the position of the corresponding impact opening of upper and lower surface on high-speed aircraft sphenoid surface in the impact chamber and offer two small serpentine channels, small serpentine channel two ends are communicated with impacts chamber and afterbody atmosphere, small serpentine channel is continuous U-shaped bending channel, the U-shaped passage is two distances between impact opening in the vertical distance before the bending, entrance location, exit position and impact opening are positioned at same perpendicular.
2. a kind of leading edge of hypersonic vehicle impact+small serpentine channel cooling structure according to claim 1 is characterized in that the diameter of impact opening is 0.5mm～1.0mm, and spacing is 5.0mm～20.0mm.
3. a kind of leading edge of hypersonic vehicle impact+small serpentine channel cooling structure according to claim 1 is characterized in that, small serpentine channel with impact that the chamber links to each other the import cross section be of a size of (0.5mm～1.0mm) * (0.5mm～1.0mm).
4. a kind of leading edge of hypersonic vehicle impact+small serpentine channel cooling structure according to claim 1 is characterized in that the width of fin is 0.5mm～1.0mm between the U-shaped bending channel of small serpentine channel.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN2011100693092A CN102152848A (en) | 2011-03-22 | 2011-03-22 | Super-speed aircraft frontal edge impact and small snakelike channel cooling structure |
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| Application Number | Priority Date | Filing Date | Title |
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| CN2011100693092A CN102152848A (en) | 2011-03-22 | 2011-03-22 | Super-speed aircraft frontal edge impact and small snakelike channel cooling structure |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109264030A (en) * | 2018-09-29 | 2019-01-25 | 北京机械设备研究所 | A kind of cooling active thermal protection structure of convection current |
| CN109941424A (en) * | 2019-03-25 | 2019-06-28 | 西北工业大学 | A kind of thermal protection struc ture integration leading edge for Air-breathing hypersonic vehicle |
| CN110979633A (en) * | 2019-12-13 | 2020-04-10 | 西北工业大学 | Cooling enhancement structure for front edge of hypersonic aircraft |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5299762A (en) * | 1991-10-15 | 1994-04-05 | Grumman Aerospace Corporation | Injection-cooled hypersonic leading edge construction and method |
| US5351917A (en) * | 1992-10-05 | 1994-10-04 | Aerojet General Corporation | Transpiration cooling for a vehicle with low radius leading edges |
| JP3096312B2 (en) * | 1991-02-25 | 2000-10-10 | 石川島播磨重工業株式会社 | Manufacturing method of cooling structure |
| CN2744599Y (en) * | 2004-07-27 | 2005-12-07 | 南京师范大学 | Heat protection device for pneumatic heating heated-surface of superhigh speed aircraft |
| CN1994824A (en) * | 2006-12-27 | 2007-07-11 | 中国科学院力学研究所 | Reverse pulse explosion heat-resistant and damping method for high supersonic aerocraft |
-
2011
- 2011-03-22 CN CN2011100693092A patent/CN102152848A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP3096312B2 (en) * | 1991-02-25 | 2000-10-10 | 石川島播磨重工業株式会社 | Manufacturing method of cooling structure |
| US5299762A (en) * | 1991-10-15 | 1994-04-05 | Grumman Aerospace Corporation | Injection-cooled hypersonic leading edge construction and method |
| US5351917A (en) * | 1992-10-05 | 1994-10-04 | Aerojet General Corporation | Transpiration cooling for a vehicle with low radius leading edges |
| CN2744599Y (en) * | 2004-07-27 | 2005-12-07 | 南京师范大学 | Heat protection device for pneumatic heating heated-surface of superhigh speed aircraft |
| CN1994824A (en) * | 2006-12-27 | 2007-07-11 | 中国科学院力学研究所 | Reverse pulse explosion heat-resistant and damping method for high supersonic aerocraft |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109264030A (en) * | 2018-09-29 | 2019-01-25 | 北京机械设备研究所 | A kind of cooling active thermal protection structure of convection current |
| CN109941424A (en) * | 2019-03-25 | 2019-06-28 | 西北工业大学 | A kind of thermal protection struc ture integration leading edge for Air-breathing hypersonic vehicle |
| CN110979633A (en) * | 2019-12-13 | 2020-04-10 | 西北工业大学 | Cooling enhancement structure for front edge of hypersonic aircraft |
| CN110979633B (en) * | 2019-12-13 | 2022-04-26 | 西北工业大学 | Cooling enhancement structure for front edge of hypersonic aircraft |
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Application publication date: 20110817 |