CN103867339A - Ablation-proof structure of solid rocket engine - Google Patents
Ablation-proof structure of solid rocket engine Download PDFInfo
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- CN103867339A CN103867339A CN201210539206.2A CN201210539206A CN103867339A CN 103867339 A CN103867339 A CN 103867339A CN 201210539206 A CN201210539206 A CN 201210539206A CN 103867339 A CN103867339 A CN 103867339A
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- insulation layer
- heat insulation
- proof structure
- antiscour
- protective coating
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Abstract
Disclosed is an ablation-proof structure of a solid rocket engine. The ablation-proof structure of the solid rocket engine comprises a pressure-bearing structure casing, a thermal insulating layer and an anti-scouring low-residue protective layer, wherein the thermal insulating layer is located in the pressure-bearing structure casing, and the anti-scouring low-residue protective layer is embedded into the inner surface of the thermal insulating layer. By means of the technical scheme, solid residues generated during ablation of the thermal insulating layer can be reduced, and moving parts are prevented from being blocked.
Description
Technical field
The present invention relates to Solid Rocket Motor Technology field, particularly a kind of solid propellant rocket burn-out proof structure.
Background technique
The kinetic energy that solid propellant rocket relies on, derives from the combustion gas of the solid propellant high pressure combustion reaction generation being seated in motor body.These combustion gas in jet pipe, expand accelerate to very high-speed, thereby gas internal energy is converted into the kinetic energy of exhaust.Propellant agent is in combustion process, and the heat insulation layer of motor stands after washing away of high-temperature fuel gas, and heat insulation layer can produce ablation.The ablation process of heat insulation layer can produce solid residue and come off, and these solid residues flow with high-temperature fuel gas, and wherein part solid residue flows out through jet pipe, and other can be attached on other component of motor.If these solid residues are attached on moving element, the motion of meeting impede motion parts, even stuck.
The rocket motor that possesses thrust regulatory function adopts gas flow controlling device at present, and the moving element relating to is more.The running of once in motion parts is restricted, or stuck, will bring accident difficult to the appraisal.Therefore how to reduce the solid residue that heat insulation layer produces in ablation process, prevent stuck one of the current problem demanding prompt solution that becomes of moving element.
Summary of the invention
The problem that the present invention solves is how to reduce the solid residue that heat insulation layer produces in ablation process, prevents that moving element is stuck.
For addressing the above problem, technological scheme of the present invention provides a kind of solid propellant rocket burn-out proof structure, comprising:
Bearing structure housing;
Be positioned at the heat insulation layer of described bearing structure enclosure interior;
The low residue protective coating of antiscour embedding at the interior shape face of described heat insulation layer.
Compared with prior art, the present invention has the following advantages:
The features such as that technological scheme of the present invention has is simple in structure, compact, easy to assembly, reliable operation, have reduced the solid residue that heat insulation layer produces in ablation process, prevent that moving element is stuck.Can play effective heat-blocking action, again can antiscour.Not only can reduce the ablation amount of heat insulation layer, and bear higher fuel gas temperature and longer operating time.This design proposal is reliably effective in addition, implements fairly simple convenience.
Accompanying drawing explanation
Fig. 1 is the structural drawing of the solid propellant rocket burn-out proof structure that provides of the embodiment of the present invention.
Embodiment
A lot of details are set forth in the following description so that fully understand the present invention.But the present invention can implement to be much different from alternate manner described here, and those skilled in the art can do similar popularization without prejudice to intension of the present invention in the situation that, and therefore the present invention is not subject to the restriction of following public concrete enforcement.
Secondly, the present invention utilizes schematic diagram to be described in detail, and in the time that the embodiment of the present invention is described in detail in detail, for ease of explanation, described schematic diagram is example, and it should not limit the scope of protection of the invention at this.
At present, novel solid rocket motor adopt coating improve heat insulation layer erosion resistibility, reduce ablating rate, bear higher fuel gas temperature and longer operating time, but there is following shortcoming in this method:
1. manufacturing process more complicated, needs professional manufacturing equipment, experience various technical processes, last moulding.
2. different heat insulation layer material, the technological requirement difference of coating demand, has strengthened difficulty of processing.
3. adopt coating process, comparing does not add coating process and can bear the longer operating time, but not too applicable to the solid propellant rocket of long operating time.
Therefore how to reduce the solid residue that heat insulation layer produces in ablation process, prevent stuck one of the current problem demanding prompt solution that becomes of moving element.
For addressing the above problem, technological scheme of the present invention provides a kind of solid propellant rocket burn-out proof structure, and Fig. 1 is the structural drawing of the solid propellant rocket burn-out proof organization plan that provides of the embodiment of the present invention, describes in detail below in conjunction with Fig. 1.
Bearing structure housing 1;
Be positioned at the heat insulation layer 2 of described bearing structure housing 1 inside;
The low residue protective coating 3 of antiscour embedding at the interior shape face of described heat insulation layer 2.
The outermost surface of burn-out proof structure is bearing structure housing 1, and heat insulation layer 2 is positioned at bearing structure housing 1, and the low residue protective coating 3 of antiscour embeds in heat insulation layer 2 profiles.In the bearing structure housing 1 of burn-out proof structure, adopt the heat insulation layer 2 of low thermal conductivity, play heat-blocking action.Embed the low residue protective coating 3 of antiscour at the interior shape face of heat insulation layer 2, the low residue protective coating 3 of antiscour prevents that high-temperature fuel gas from directly contacting with heat insulation layer 2, can bear washing away of high-temperature high-pressure fuel gas, can effectively reduce heat insulating construction residue and produce.The low residue protective coating 3 of this antiscour is formed by high temperature resistant low ablative material manufacture, such as rhenium, tungsten, molybdenum etc., and alloy, also can use the stupalith of infusibility etc., its Main Function bears washing away of warm high-pressure gas, can effectively reduce heat insulation layer 2 ablation amounts.
In specific implementation process, when solid propellant rocket work, produce high-temperature high-pressure fuel gas, bearing structure housing 1 need bear gaseous-pressure, and bearing structure housing 1 adopts high-strength material, as high duty metal or nonmetallic material etc.Heat insulation layer 2 adopts low thermal conductivity material, such as Ethylene Propylene Terpolymer, mainly plays heat-blocking action.
The region that high-temperature fuel gas is flowed through, under heat conduction, thermoconvection, thermal-radiating effect, the low residue protective coating 3 of high-temperature fuel gas and antiscour, heat insulation layer 2 and bearing structure housing 1 carry out heat transmission, cause the low residue protective coating 3 of antiscour, heat insulation layer 2 and bearing structure housing 1 temperature all can raise, but because heat insulation layer 2 adopts low thermal conductivity material, can obviously weaken heat transfer process, reduce 1 temperature rise of bearing structure housing, protection bearing structure housing 1 can obviously not reduce because excess Temperature causes bearing structure housing 1 intensity in engine working process; Reduce the thermal energy loss of combustion gas simultaneously, improved engine power utilization ratio.
The low residue protective coating 3 of antiscour is formed by high temperature resistant low ablative material manufacture, such as rhenium, tungsten, molybdenum etc., and alloy, also can use the stupalith of infusibility etc.The low residue protective coating 3 of antiscour is mainly used to bear washing away of high-temperature high-pressure fuel gas, intercepts contacting of high-temperature high-pressure fuel gas and heat insulation layer 2 simultaneously.If there is no the low residue protective coating 3 of antiscour, high-temperature high-pressure fuel gas directly contacts with heat insulation layer 2, can produce a large amount of residues.These residues flow with high-temperature fuel gas, and wherein part flows out through jet pipe, and other can be attached on other component of motor.If these residues are attached on moving element, the motion of meeting impede motion parts, even stuck.Adopt the low residue protective coating 3 of antiscour, avoided high-temperature high-pressure fuel gas directly to contact with heat insulation layer 2.Because the low residue protective coating 3 of antiscour is by high temperature resistant low ablative material, can effectively reduce the generation of residue.
Technological scheme of the present invention at least has following beneficial effect:
Technological scheme of the present invention provides a kind of solid propellant rocket burn-out proof organization plan, improve heat insulating construction erosion resistibility, reduce its ablating rate, and can bear higher fuel gas temperature and longer operating time.
Technological scheme of the present invention, not only has the features such as simple in structure, compact, easy to assembly, reliable operation; Can play effective heat-blocking action, again can antiscour, guarantee that heat insulation layer 2 ablation amounts are very little, bear higher fuel gas temperature and longer operating time, this design proposal is reliably effective, implements fairly simple convenience.
Claims (5)
1. a solid propellant rocket burn-out proof structure, is characterized in that, comprising:
Bearing structure housing;
Be positioned at the heat insulation layer of described bearing structure enclosure interior;
The low residue protective coating of antiscour embedding at the interior shape face of described heat insulation layer.
2. solid propellant rocket burn-out proof structure as claimed in claim 1, is characterized in that, described heat insulation layer adopts low thermal conductivity material.
3. solid propellant rocket burn-out proof structure as claimed in claim 1, is characterized in that, the low residue protective coating of described antiscour adopts high temperature resistant low ablative material.
4. solid propellant rocket burn-out proof structure as claimed in claim 3, is characterized in that, the material of the low residue protective coating of described antiscour adopt in rhenium, tungsten, molybdenum one or more etc.
5. solid propellant rocket burn-out proof structure as claimed in claim 3, is characterized in that, the material of the low residue protective coating of described antiscour is pottery.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201210539206.2A CN103867339A (en) | 2012-12-14 | 2012-12-14 | Ablation-proof structure of solid rocket engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201210539206.2A CN103867339A (en) | 2012-12-14 | 2012-12-14 | Ablation-proof structure of solid rocket engine |
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| Publication Number | Publication Date |
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| CN103867339A true CN103867339A (en) | 2014-06-18 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN201210539206.2A Pending CN103867339A (en) | 2012-12-14 | 2012-12-14 | Ablation-proof structure of solid rocket engine |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104833768A (en) * | 2015-03-11 | 2015-08-12 | 西北工业大学 | Simulation device of thermal insulation layer ablation under condition of particle phase deposition in rocket engine |
| CN105736177A (en) * | 2014-12-09 | 2016-07-06 | 上海新力动力设备研究所 | Tailpipe nozzle heat insulation structure for double-layer composite material formed integrally |
| CN105736184A (en) * | 2014-12-09 | 2016-07-06 | 上海新力动力设备研究所 | Large thrust ratio, long-working micro-ablation throat insert and throat structure of expansion section |
| CN106150748A (en) * | 2016-08-29 | 2016-11-23 | 潍柴动力股份有限公司 | A kind of heat insulating coat |
| CN109667682A (en) * | 2018-11-20 | 2019-04-23 | 西安航天动力技术研究所 | A kind of heat-insulation assembly of solid propellant rocket bend pipe |
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| CN1066708A (en) * | 1991-04-30 | 1992-12-02 | 北京理工大学 | The material of flammable storepipe and jet pipe and structure |
| DE4201600A1 (en) * | 1992-01-22 | 1993-07-29 | Dornier Gmbh | PROTECTION SYSTEM FOR COMPONENTS EXPOSED TO HIGH THERMAL LOADS OR THERMAL LOADS AND DAM PRESSURE |
| US20060064984A1 (en) * | 2004-09-27 | 2006-03-30 | Gratton Jason A | Throat retention apparatus for hot gas applications |
| CN1782360A (en) * | 2004-12-03 | 2006-06-07 | 中国北方工业公司 | Rocket engine jet nozzle of composite distance increasing shell |
| CN2908821Y (en) * | 2005-12-30 | 2007-06-06 | 中国北方工业公司 | Nozzle of rocket engine |
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2012
- 2012-12-14 CN CN201210539206.2A patent/CN103867339A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1066708A (en) * | 1991-04-30 | 1992-12-02 | 北京理工大学 | The material of flammable storepipe and jet pipe and structure |
| DE4201600A1 (en) * | 1992-01-22 | 1993-07-29 | Dornier Gmbh | PROTECTION SYSTEM FOR COMPONENTS EXPOSED TO HIGH THERMAL LOADS OR THERMAL LOADS AND DAM PRESSURE |
| US20060064984A1 (en) * | 2004-09-27 | 2006-03-30 | Gratton Jason A | Throat retention apparatus for hot gas applications |
| CN1782360A (en) * | 2004-12-03 | 2006-06-07 | 中国北方工业公司 | Rocket engine jet nozzle of composite distance increasing shell |
| CN2908821Y (en) * | 2005-12-30 | 2007-06-06 | 中国北方工业公司 | Nozzle of rocket engine |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105736177A (en) * | 2014-12-09 | 2016-07-06 | 上海新力动力设备研究所 | Tailpipe nozzle heat insulation structure for double-layer composite material formed integrally |
| CN105736184A (en) * | 2014-12-09 | 2016-07-06 | 上海新力动力设备研究所 | Large thrust ratio, long-working micro-ablation throat insert and throat structure of expansion section |
| CN105736184B (en) * | 2014-12-09 | 2018-04-03 | 上海新力动力设备研究所 | High thrust is than, the micro-ablation that works long hours larynx lining and the laryngeal structure of diffuser |
| CN104833768A (en) * | 2015-03-11 | 2015-08-12 | 西北工业大学 | Simulation device of thermal insulation layer ablation under condition of particle phase deposition in rocket engine |
| CN106150748A (en) * | 2016-08-29 | 2016-11-23 | 潍柴动力股份有限公司 | A kind of heat insulating coat |
| CN109667682A (en) * | 2018-11-20 | 2019-04-23 | 西安航天动力技术研究所 | A kind of heat-insulation assembly of solid propellant rocket bend pipe |
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Application publication date: 20140618 |