CN100423989C - Ablation-free self-adaptive heat-resistant and damping system for high supersonic aerocraft - Google Patents
Ablation-free self-adaptive heat-resistant and damping system for high supersonic aerocraft Download PDFInfo
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- CN100423989C CN100423989C CNB2006101696839A CN200610169683A CN100423989C CN 100423989 C CN100423989 C CN 100423989C CN B2006101696839 A CNB2006101696839 A CN B2006101696839A CN 200610169683 A CN200610169683 A CN 200610169683A CN 100423989 C CN100423989 C CN 100423989C
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Abstract
The invention relates to a non-erosion self-adaptive flameproof damping system of hypersonic speed airplane, wherein the front end of airplane is arranged with one pressure container with better thermal conductivity; the container has some liquid; the head of airplane has one support rod; the top of support rod has baffle plate; when airplane is at hypersonic speed, via the support rod, the head flow field can be rectified to change the wave activation structure; since the wave can heat air and via the friction resistance, the surface temperature of airplane will increase quickly; the pressure container transmits energy to the liquid; and the liquid will be evaporated to be ejected at the sealing open at the top of support rod, to cool the baffle plate, and hold the head flow field.
Description
Technical field:
The present invention is a kind of new and high technology of space flight and aviation engineering, relates to a kind of control hypersonic aircraft header stream field structure that passes through, and reduces thermal environment, reduces the aircraft aerodynamic drag, forms a cover aircraft surface ablation-free self-adaptive solar heat protection and a drag reduction system.
Background technology:
For the space shuttle that can reentry and re-entry vehicle and the class hypersonic aircrafts such as hypersonic aircraft, interception guided missile and mobile missile that can in atmospheric envelope, fly, in with hypersonic flight course, because the effect and the friction drag of head bow shock, its ambient air is by pneumatic heating, temperature rises to thousands of rapidly even up to ten thousand degrees centigrade, and material and structure of aircraft itself all proposed high thermal protection requirement.Several passive thermal protection structure of present more employing has all increased the weight of aircraft in various degree, also makes the surperficial pneumatic structure of aircraft complicated simultaneously, just might cause catastrophic consequence and a certain local structure of these thermal protection systems are damaged.
Summary of the invention:
The problems referred to above at hypersonic aircraft exists the object of the present invention is to provide a kind of ablation-free self-adaptive control system, carry out hypersonic aircraft cephalic protection, reduce pneumatic heating and reduce flight resistance.
For achieving the above object, the invention provides solar heat protection of a kind of hypersonic aircraft ablation-free self-adaptive and drag reduction system, its front end at aircraft is provided with heat-resisting and a pressure container that heat transfer character is good, and pole is set at the head of aircraft, utilize pole reformation Vehicle nose flow field, bow shock is changed into cone shock, alleviate the pneumatic heating of Vehicle nose and reduce drag due to shock wave; The pole top is provided with baffle plate, forms steam behind the heated liquid by the ejection of pole top, cooling pole top baffle plate, and protection pole is not burnt, and keeps the function in its reformation flow field, and can reduce aircraft ambient gas temperature.
Further, also be provided with in the end of this pole and seal, be heated when arriving liquid evaporation in uniform temperature or the described high-temperature resistant container and reaching certain pressure when this seals, the steam of described liquid evaporation can seal ejection by this.
Further, described baffle plate is a hemisphere face shape, and perhaps other are suitable for the shape of gas flow and heat exchange.
Further, described container is to be suitable for the hemispherical of gas flow or semielliptical shape.
Further, described liquid is water or liquid with the evaporation properties of being heated.
System of the present invention is that the head at aircraft is provided with a pole and a pressure container heat-resisting and heat transfer character is good.This pole Vehicle nose flow field of can reforming, the change bow shock is a cone shock, reduces the aircraft ambient gas heating effect and the aircraft drag due to shock wave that are caused by the normal shock wave compression.The a certain amount of evaporated liquid of being heated of packing in the pressure container is set, the pneumatic heating of Vehicle nose causes the heated liquid evaporation in the container in the flight course, and phase transition process absorbs a large amount of pneumatic heat, has improved the pressure of vaporization back gas, and pneumatic to add heat big more, and pressure for vaporization is high more.
According to heat transfer theory, the pneumatic environment on Vehicle nose pole top is complicated more, and pneumatic heating is more severe.In order to protect a club head, guiding liquids evaporation back gas is by the ejection of pole top, and a cooling club head is kept the function in pole reformation flow field.Further, after the ejection of gas after the vaporization, can absorb air-flow heat behind the cone shock ripple, alleviate the harsh thermal environment of Vehicle nose.Pneumatic to add heat big more, and ejection gas is many more, and cooling performance is strong more, has the positive feed back adaptation function.
Comprehensively, the present invention is by being provided with a pole and a pressure container heat-resisting and heat transfer character is good in the aircraft end, and the evaporated liquid of being heated of packing in this container.Under hypersonic flying condition, the pole Vehicle nose flow field of can reforming becomes cone shock with the head bow shock; The air of shock wave heating and aircraft and windage resistance make the vessel surface temperature rise rapidly, and the liquid of pressure container is by absorbing heat of vaporization, cooled containers wall; Gas after the evaporation is from the ejection of pole top, and cooling pole is kept the function in pole reformation flow field; The gas of ejection can reduce the temperature of aircraft ambient gas, further reduces the surface temperature of aircraft; Simultaneously, can reduce aloft drag due to shock wave greatly by this profile structure.
Description of drawings
Fig. 1 specifically uses scheme drawing for the present invention;
Fig. 3 is the enlarged diagram of D among Fig. 1;
Constitution diagram when Fig. 2 is an aircraft flight of the present invention.
The specific embodiment
As shown in Figure 1, the present invention is that the front end at aircraft 1 is provided with heat-resisting and a pressure container 2 that heat transfer character is good, container 2 can be designed as and is suitable for the hemispherical of gas flow, also can be other suitable shapes such as semielliptical shape, makes gas flow relatively good just passable as long as satisfy.The a certain amount of evaporated liquid 5 of being heated of in this container 2, packing into, and away from the end face of aircraft 1 pole 3 and pole top baffle plate 4 are set at this container 2.
When hypersonic flight, can form head cone shock 8 around the aircraft 1, as shown in Figure 2.High Temperature Gas behind the shock wave 8 is known from experience pressure container 2 heating to Vehicle nose, and the heat that liquid 5 absorbs wall of container evaporates, and by the ejection of pole 3 tops, and cooling pole top baffle plate 4, keep the function in pole 3 reformation flow fields.The steam 9 of ejection absorbs the heat of the air behind the shock wave 8, further reduces the temperature on aircraft 1 surface.Pole top baffle plate 4 can be designed as hemisphere face shape, and its concave surface also can be designed as other shape that is suitable for gas flow and heat exchange and size towards nozzle 3.
Can from pole 3, not flow out before use for the ease of the liquid 5 that is contained in the heat-resisting pressure container 2, can be provided with in the end of this pole 3 and seal 7, as shown in Figure 1.Be heated when arriving liquid 5 evaporations in uniform temperature or the container 2 and reaching certain pressure when this seals 7, under the effect of the temperature and pressure that liquid 5 evaporations produce, steam can seal 7 by this and spray.Sealing 7 and can select low-melting-point metal or non-metallic material, also can be that the material that plastic film etc. can be washed open by steam under certain pressure is made.In addition, liquid 5 can be water, also can be other suitable liquid with the evaporation properties of being heated.
In addition, the present invention also can be designed as other similar structures, and the Vessel Design that for example will place liquid is the shell of aircraft; Perhaps container 2 is designed to the evaporator of sandwich-like hemisphere face shape, pole 3 communicates with the interlayer of evaporator, when aircraft 1 is in state of flight, by pump the liquid in the container is supplied with evaporator, after reaching uniform temperature or pressure, steam just sprays in the nozzle of pole 3.
The foregoing description just is used for explanation of the invention, and can not be as limitation of the present invention, and therefore the embodiment that design philosophy every and of the present invention is identical is all in protection scope of the present invention.
Claims (5)
1. hypersonic aircraft ablation-free self-adaptive solar heat protection and drag reduction system, it is characterized in that, at the front end of aircraft one heat-resisting and pressure container that heat transfer character is good is set, and pole is set at the head of aircraft, utilize pole reformation Vehicle nose flow field, bow shock is changed into cone shock, alleviate the pneumatic heating of Vehicle nose and reduce drag due to shock wave; The pole top is provided with baffle plate, forms steam behind the heated liquid by the ejection of pole top, cooling pole top baffle plate, and protection pole is not burnt, and keeps the function in its reformation flow field, and can reduce aircraft ambient gas temperature.
2. a kind of hypersonic aircraft ablation-free self-adaptive according to claim 1 solar heat protection and drag reduction system, it is characterized in that, also be provided with in the end of this pole and seal, when the liquid evaporation in this seals be heated arrival uniform temperature or described high-temperature resistant container reached certain pressure, the steam of described liquid evaporation can seal ejection by this.
3. a kind of hypersonic aircraft ablation-free self-adaptive according to claim 2 solar heat protection and drag reduction system is characterized in that, described baffle plate is a hemisphere face shape, and perhaps other are suitable for the shape of gas flow and heat exchange.
4. a kind of hypersonic aircraft ablation-free self-adaptive according to claim 3 solar heat protection and drag reduction system is characterized in that, described container is to be suitable for the hemispherical of gas flow or semielliptical shape.
5. a kind of hypersonic aircraft ablation-free self-adaptive according to claim 4 solar heat protection and drag reduction system is characterized in that, described liquid is water or liquid with the evaporation properties of being heated.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNB2006101696839A CN100423989C (en) | 2006-12-27 | 2006-12-27 | Ablation-free self-adaptive heat-resistant and damping system for high supersonic aerocraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNB2006101696839A CN100423989C (en) | 2006-12-27 | 2006-12-27 | Ablation-free self-adaptive heat-resistant and damping system for high supersonic aerocraft |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN1994823A CN1994823A (en) | 2007-07-11 |
| CN100423989C true CN100423989C (en) | 2008-10-08 |
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| Application Number | Title | Priority Date | Filing Date |
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| CNB2006101696839A Expired - Fee Related CN100423989C (en) | 2006-12-27 | 2006-12-27 | Ablation-free self-adaptive heat-resistant and damping system for high supersonic aerocraft |
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Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102928189A (en) * | 2012-05-25 | 2013-02-13 | 中国科学院力学研究所 | Experimental device for reducing heat flow rate by applying local reverse overflow of aircraft |
| CN104326079B (en) * | 2014-10-14 | 2016-07-06 | 中国科学院力学研究所 | Self adaptation active thermal preventer and aircraft |
| CN106828887B (en) * | 2016-11-14 | 2018-11-23 | 中国航天空气动力技术研究院 | A kind of adaptive thermal protection method of Vehicle nose |
| CN106628111B (en) * | 2016-12-06 | 2018-05-11 | 清华大学 | A kind of supersonic speed air film cooling structure of adaptive Shock Wave |
| CN112193401B (en) * | 2020-04-07 | 2022-05-20 | 北京空天技术研究所 | Thermal protection method for front edge of hypersonic aircraft |
| CN111559492A (en) * | 2020-04-26 | 2020-08-21 | 南京航空航天大学 | High-efficiency shock wave resistance reduction system of hypersonic aircraft |
| CN113306697A (en) * | 2021-07-08 | 2021-08-27 | 南京航空航天大学 | Novel hypersonic aircraft wing |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4014485A (en) * | 1975-04-14 | 1977-03-29 | Martin Marietta Corporation | Gas cooling system for hypersonic vehicle nosetip |
| 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 |
| US6581870B1 (en) * | 1999-11-08 | 2003-06-24 | Lfk Lenkflugkoerpersysteme Gmbh | Method and apparatus for reducing pressure and temperature on the front of a missile at ultrasonic speed |
-
2006
- 2006-12-27 CN CNB2006101696839A patent/CN100423989C/en not_active Expired - Fee Related
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4014485A (en) * | 1975-04-14 | 1977-03-29 | Martin Marietta Corporation | Gas cooling system for hypersonic vehicle nosetip |
| 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 |
| US6581870B1 (en) * | 1999-11-08 | 2003-06-24 | Lfk Lenkflugkoerpersysteme Gmbh | Method and apparatus for reducing pressure and temperature on the front of a missile at ultrasonic speed |
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| CN1994823A (en) | 2007-07-11 |
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