CN102183312B - Surface high-temperature measurement device for nonmetallic heat resistant material plane test piece of hypersonic speed aircraft - Google Patents
Surface high-temperature measurement device for nonmetallic heat resistant material plane test piece of hypersonic speed aircraft Download PDFInfo
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- CN102183312B CN102183312B CN201110063310.4A CN201110063310A CN102183312B CN 102183312 B CN102183312 B CN 102183312B CN 201110063310 A CN201110063310 A CN 201110063310A CN 102183312 B CN102183312 B CN 102183312B
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Abstract
The invention discloses a surface high-temperature measurement device for a nonmetallic heat resistant material plane test piece of a hypersonic speed aircraft. The device comprises a high-temperature ceramic framework, a temperature measurement thermocouple, a nonmetallic heat resistant material flat test piece, a silicon molybdenum infrared radiation heating tube and a computer. The high-temperature ceramic framework is provided with a circular hole; the temperature measurement thermocouple sleeved with a ceramic insulating sleeve is inserted into the circular hole of the high-temperature ceramic framework; and the front part of the temperature measurement thermocouple is bent into an upward arch; the nonmetallic heat resistant material flat test piece is pressed on the high-temperature ceramic framework by means of self weight to press the forefront of the temperature measurement thermocouple to form a ball-shaped temperature sensing part by welding. Due to interaction between the downward pressure of the nonmetallic heat resistant material flat test piece and the front arch upward elasticity of the temperature measurement thermocouple and adoption of double technology measures of applying pre-deformation and increasing pre-stress, tight contact of the nonmetallic heat resistant material flat test piece and the temperature sensing part of the front end of the thermocouple under the high-temperature environment of between 1,000 and 1,400 DEG C is ensured, so that the surface high-temperature test result of the nonmetallic heat resistant material plane test piece of the hypersonic speed aircraft such as a missile, a space shuttle and the like is accurate and reliable.
Description
Technical field
The present invention relates to the nonmetal heat insulation material of hypersonic aircraft plane testpieces surface high-temp measurement mechanism, in the hot test environment of the hypersonic aircrafts such as simulated missile, space shuttle with nonmetal plane light heat-insulating material, the non-metal heat-insulating material surface is measured and record in real time up to 1000 ℃-1400 ℃ high temperature dynamic change.
Background technology
Along with the development of world's space flight and aviation technology, the design rate of long-range maneuvering-vehicle is increasing substantially, the development work that the developed countries such as present the United States, Russia, European Union just competitively carry out hypersonic aircraft.China is also making great efforts to carry out the research of the long-range maneuvering-vehicle of hypersonic speed of new generation, the research of pointing out the long-range maneuvering-vehicle of hypersonic speed in the project of national nature science fund project guide " relates to national security and peaceful use space; being that one of focus of space technology is competitively fought in the world at present, is the embodiment of overall national strength ".When the high speed aircraft such as guided missile flew with High Mach number, " thermal boundary " problem that is caused by Aerodynamic Heating was very serious.When flight Mach number near 4 the time, guided missile front end stagnation temperature can reach 700 ℃, with the hypersonic aircraft of 6 Mach numbers flights, the stagnation temperature in its nose of wing and antenna house tapering will be above 1200 ℃.Therefore the design of the research of high-temperature structural material and anti-heat insulation structural is gordian technique and the core technology of the reliability design of hypersonic aircraft general safety.
When temperature will be above 1000 ℃, even if adopt the refractory metal materials such as nickel base superalloy or titanium alloy, also can occur being out of shape, soften, the phenomenon of rigidity decline, have a strong impact on aerodynamic configuration and the safe flight of high-speed aircraft.The interior temperature of instrument room that sophisticated electronics is installed does not in addition generally allow above 80 ℃, necessary design and installation thermal protection struc ture or laying heat-barrier material, to reduce the conduction of velocity of the guided missile surface inside section of heat, therefore, the hypersonic vehicle of a new generation generally all adopts the nonmetal light materials such as carbon fibre composite, porous foam type refractory ceramics watt, resurrection glass fibre to manufacture the shell of aircraft or as anti-heat-barrier material.
When carrying out the security and Reliability Design of the hypersonic aircrafts such as guided missile, need in advance the surface temperature of employed nonmetal light material to be measured, to check its heat resistance, ablation property and heat-proof quality.Because nonmetallic materials can not can be with temperature thermocouple silk direct spot welding on material surface, so generally adopt the method meter surface temperature of bonding or crimping as metal material.If employing bonding way, since the temperature probe of metal material and nonmetallic materials difference of thermal expansion coefficients very big, be higher than under 1000 ℃ the thermal environment, metal temperature probe and the non-metal material surface segregation phenomenon of coming unstuck often occurring, causing the inaccurate situation of surface temperature measurement.
If can guarantee that metal temperature probe and non-metal material surface are in intimate-association state all the time under hot environment, can realize the temperature survey of non-metal material surface under the hot environment on principle.But, make the temperature thermocouple front end in process of the test, remain close contact to non-metal material surface, must make between near the part of thermopair front end and the non-metal material surface large relative displacement can not be arranged, otherwise inevitable related thermopair front end temperature-sensitive section's warpage and dislocation, particularly it will can also keep steady state (SS) under greater than 1000 ℃ hot environment, and this is key and the technological difficulties that are related to the success or failure of crimping thermometric mode.
In addition, because under greater than 1000 ℃ hot environment, the testpieces surface can thermal ablation occur and heat waste is ruined, generally can not repeat the multiple high temp test to same testpieces, and the solar heat protection test specimen of high-speed aircraft is all very expensive, just surpass 50,000 U.S. dollars such as U.S.'s Glenn than the heat-proof tile monolithic cost of inferior space shuttle, so the test data that each test obtains is all very valuable.Therefore, must exploitation can at the nonmetal heat insulation material of the high-speed aircraft of reliably working under hot environment plane test specimen surface high-temp measurement mechanism, record in the high temperature heat test process situation of change of heat insulation material surface temperature.This work has very important practical significance for thermal protection and the security and Reliability Design of the hypersonic aircrafts such as guided missile, space shuttle.
Summary of the invention
Technology of the present invention is dealt with problems and is: overcome the deficiencies in the prior art, the nonmetal heat insulation material of a kind of hypersonic aircraft plane testpieces surface high-temp measurement mechanism is provided, this device can be accurately and is measured reliably and record in the hypersonic aircraft heat test process, the dynamic high temperature of nonmetal heat insulation material plane testpieces surface temperature field up to 1000 ℃-1400 ℃ changes, and simple in structure, easy to use, checking with safety and Protection for the hot strength of the hypersonic aircrafts such as guided missile, space shuttle provides reliable test basis.
Technical solution of the present invention is: the nonmetal heat insulation material of hypersonic aircraft plane testpieces surface high-temp measurement mechanism comprises: refractory ceramics framework, temperature thermocouple, nonmetal heat insulation material treadmill test part, adiabatic bracing frame, heating source and computing machine; The refractory ceramics framework lies on the adiabatic bracing frame, one circular hole is arranged on the refractory ceramics framework, temperature thermocouple is inserted in the circular hole, and the front portion that makes temperature thermocouple curves upwards arc, nonmetal heat insulation material treadmill test part is pressed on the refractory ceramics framework, make the refractory ceramics framework push down the thermopair front end temperature-sensitive section that is welded into ball-point pen type of temperature thermocouple forefront, interaction between the elastic force that makes progress by the downward gravity of nonmetal heat insulation material treadmill test part and temperature thermocouple front arcuate position, so that nonmetal heat insulation material treadmill test part and thermopair front end temperature-sensitive section can remain close contact under 1000 ℃-1400 ℃ hot environment under hot environment, guarantee accuracy and the reliability of temperature-measuring results; Can produce the below that is placed on nonmetal heat insulation material treadmill test part up to the silicon chrome vermillion external radiation heating source of 1600 ℃ of high temperature, heating power carries out radiation heating to the lower surface of nonmetal heat insulation material treadmill test part; The temperature thermocouple that is crimped on nonmetal heat insulation material treadmill test part lower surface is sent to computing machine with output signal by wire, and computer recording also calculates the high temperature delta data on the nonmetal heat insulation material treadmill test of guided missile, space shuttle and so on hypersonic aircraft part surface.
Refractory ceramics framework and temperature thermocouple do not adopt traditional adhering fixed mode, but the interaction between the elastic force that makes progress by the downward pressure of nonmetal heat insulation material treadmill test part and temperature thermocouple front arcuate position makes that nonmetal heat insulation material treadmill test part and thermopair front end temperature-sensitive section remain close contact under the hot environment.
Bad phenomenon can not appear being in contact with one another in order to ensure thermopair front end temperature-sensitive section in the high temperature heat test process and nonmetal heat insulation material treadmill test part, before nonmetal heat insulation material treadmill test part is installed, it is larger that the arc displacement that first front portion of temperature thermocouple is made progress is done, make thermopair front end temperature-sensitive section be higher than plane 5mm-6mm on the refractory ceramics framework, form an initial predeformation that makes progress, when nonmetal heat insulation material treadmill test part is pressed on the refractory ceramics framework, the displacement of 5mm-6mm will appear descending mutually in thermopair front end temperature-sensitive section, after this 5mm-6mm distortion of thermopair front portion is forced to push back, can produce the larger elastic force that makes progress, so that thermopair front end temperature-sensitive section and nonmetal heat insulation material treadmill test part surface can keep close contact, adopt this predeformation reinforcing method to guarantee the reliability and stability of thermometric work under 1000 ℃ of-1400 ℃ of hot environments.
Principle of work of the present invention is: make a refractory ceramics framed structure that equates with plane test specimen boundary dimensions, this ceramic frame adopts the alumina content of 1800 ℃ of abilities to fire up to 99% corundum high-temperature ceramic materials and forms, therefore, this ceramic frame can be guaranteed steady operation under 1000 ℃ to 1400 ℃ hot environments.One circular hole is arranged in the refractory ceramics framed structure, temperature thermocouple is installed in the mesopore of refractory ceramics framed structure, the anterior upwards bow shape of temperature thermocouple.When carrying out the test specimen surface high-temp measurement of the nonmetal heat insulation material of hypersonic aircraft plane, nonmetal heat insulation material plane test specimen is pressed on the refractory ceramics framed structure, produce the bowed shape interaction that downward pressure and thermopair front portion upwards curve by the deadweight of plane test specimen, form the close contact between test specimen surface, nonmetal heat insulation material plane and the thermocouple temperature sensitive section.Make thermopair front end thermometric section at high temperature also can remain contact force upwards.When heating for test specimen surface, nonmetal heat insulation material plane according to heat flow curve or temperature curve, because the rigidity that the bowed shape that warm galvanic couple front end makes progress and thermocouple wire itself have, and corundum refractory ceramics framed structure has under 1800 ℃ of high temperature not yielding advantages, even when temperature up to 1000 ℃ during to 1400 ℃, also can close contact between test specimen surface and the thermocouple temperature sensitive section, thermopair front end temperature-sensitive section, the temperature variation on perception test specimen surface is passed through computer recording rapidly, and calculate the dynamic temperature change curve on nonmetal heat insulation material test specimen surface in the heat test process.
The present invention's beneficial effect compared with prior art is: because the present invention has taked temperature thermocouple is installed in the mesopore of refractory ceramics framed structure, the anterior upwards bow shape of temperature thermocouple, deadweight by the plane test specimen produces the interaction that upwards pressure that bowed shape that downward pressure and thermopair front portion upwards curve produces, and has guaranteed the close contact of test specimen surface, nonmetal heat insulation material plane and thermocouple temperature sensitive section.Because the temperature thermocouple support zone do not use bonding agent, avoided traditional bonding support fixing means under greater than 1000 ℃ of hot environments, the test failure that comes unglued and cause very easily appears in the temperature thermocouple support zone.In addition, because thermopair front end temperature-sensitive section does not have adhesive linkage yet, it is depended on pressure rather than by bonding contact between test specimen surface and the thermocouple temperature sensitive section, under 1000 ℃-1400 ℃ hot environment, even because high temperature thermal deformation produces a small amount of transversal displacement, because larger interactional prestress is arranged, can guarantee between test specimen surface and the thermocouple temperature sensitive section it is the state that is in close contact all the time between the two.Make measurement result more accurately, reliably.Above invention has solved under 1000 ℃-1400 ℃ excessive temperature environment, is related to the gordian technique difficult point of crimping thermometric mode success or failure.In addition, need to be to the lot of experiments spare thermometric of different effect of heat insulation, different-thickness the time, owing to there not being adhesive linkage, do not need to clear up adhesive spots, only need change new nonmetal heat insulation material plane test specimen, so the present invention also has simple and direct, the easy to use advantage of installation.
Description of drawings
Fig. 1 is structure side view of the present invention;
Fig. 2 is the perspective view of refractory ceramics framework of the present invention;
Fig. 3 is the temperature changing curve diagram on certain testpieces testpieces surface under 1000 ℃ of high temperature, guided missile carbon fibre composite plane of using the present invention and recording;
Fig. 4 is the temperature changing curve diagram on certain testpieces testpieces surface under 1400 ℃ of high temperature, guided missile carbon fibre composite plane of using the present invention and recording.
Embodiment
As depicted in figs. 1 and 2, the present invention is comprised of refractory ceramics framework 1, temperature thermocouple 2, nonmetal heat insulation material treadmill test part 6, adiabatic bracing frame 7, heating source 8 (the present embodiment adopts the silicon chrome vermillion external radiation heating tube that can form 1600 ℃ of hot environments), wire 9 and computing machine 10.Refractory ceramics framework 1 adopts the corundum high-temperature ceramic materials of the alumina content 99% of 1800 ℃ of abilities to fire, and temperature thermocouple 2 adopts the two platinum Rhodium thermopairs of measurement range up to 1800 ℃.Refractory ceramics framework 1 is placed on the adiabatic bracing frame 7, one circular hole 5 is arranged on the refractory ceramics framework 1, there is the temperature thermocouple 2 of ceramic insulation sleeve pipe 4 to insert in the circular hole 5 of refractory ceramics framework 1 in cover, temperature thermocouple 2 front portions curve upwards arc, nonmetal heat insulation material treadmill test part 6 is pressed on the refractory ceramics framework 1, make it push down the thermopair front end temperature-sensitive section 3 that temperature thermocouple 2 forefronts are welded into round bead shape, interaction between the elastic force that makes progress by the downward gravity of nonmetal heat insulation material treadmill test part 6 and temperature thermocouple 2 front arcuate positions is so that nonmetal heat insulation material treadmill test part 6 and thermopair front end temperature-sensitive section 3 can remain close contact.
The diameter of the circular hole 5 among the present invention is 5mm; The front portion of temperature thermocouple 2 is to up-bow, and makes thermopair front end temperature-sensitive section 3 be higher than the upper plane 5mm-6mm of refractory ceramics framework 1, forms the upwards initial predeformation of 5mm-6mm.When nonmetal heat insulation material treadmill test part 6 is pressed on the refractory ceramics framework 1, the thermopair front end temperature-sensitive section 3 that is fixed on temperature thermocouple 2 forefronts on the refractory ceramics framework 1 is forced to push back 5mm-6mm downwards by nonmetal heat insulation material treadmill test part 6, be that thermopair front end temperature-sensitive section 3 is forced to depress and get back to original initial position, because the rigidity that the metal thermoelectric thermo wires has, can produce the very large elastic force that makes progress when pressure is depressed, this apply prestressed technical method guaranteed between thermopair front end temperature-sensitive section 3 and the nonmetal heat insulation material treadmill test part 6 contact more closely and reliable.
In the pneumatic heat test process of high temperature up to 1000 ℃-1400 ℃ of the hypersonic aircrafts such as simulated missile, when heating source 8 carries out radiation heating for the surface of nonmetal heat insulation material treadmill test part 6 according to the heating-up temperature curve, temperature thermocouple front end temperature-sensitive section 3 with nonmetal heat insulation material treadmill test part 6 surperficial close contacts, detect the temperature variation on nonmetal heat insulation material treadmill test part 6 surfaces, and change temperature variation into electric signal, send into computing machine 10 through wire 9 and store and calculate, draw the dynamic changing curve of nonmetal heat insulation material treadmill test part 6 surface temperatures shown in Fig. 3 (1000 ℃ of steady temperatures) and Fig. 4 (1400 ℃ of steady temperatures).
Can see from Fig. 3 and Fig. 4, the surface temperature of the nonmetal heat insulation material treadmill test of hypersonic aircraft part 6 rises to respectively 1000 ℃ and 1400 ℃ of high temperature within 150 seconds, and heating process finishes when remaining to for 600 second afterwards.Can also be seen by Fig. 3 and Fig. 4, in whole test overall process, design temperature curve and observing and controlling temperature curve repeatability are good, and tracking error is very little, and the observing and controlling temperature curve is level and smooth, does not have jitter phenomenon.Proved the present invention under 1000 ℃ of-1400 ℃ of very high temperature environments, also can measure in real time and record the dynamic change of the nonmetal heat insulation material treadmill test of hypersonic aircraft part 6 surface temperatures accurately and effectively.
The content that is not described in detail in the instructions of the present invention belongs to the known prior art of this area professional and technical personnel.
Claims (9)
1. the nonmetal heat insulation material of hypersonic aircraft plane testpieces surface high-temp measurement mechanism is characterized in that comprising: refractory ceramics framework (1), temperature thermocouple (2), nonmetal heat insulation material treadmill test part (6), adiabatic bracing frame (7), heating source (8) and computing machine (10); Refractory ceramics framework (1) lies on the adiabatic bracing frame (7), one circular hole (5) is arranged on the refractory ceramics framework (1), temperature thermocouple (2) is inserted in the circular hole (5), and the front portion that makes temperature thermocouple (2) curves upwards arc, nonmetal heat insulation material treadmill test part (6) is pressed on the refractory ceramics framework (1), make refractory ceramics framework (1) push down the thermopair front end temperature-sensitive section (3) that is welded into ball-point pen type of temperature thermocouple (2) forefront, interaction between the elastic force that makes progress by the downward gravity of nonmetal heat insulation material treadmill test part (6) and temperature thermocouple (2) front arcuate position, so that nonmetal heat insulation material treadmill test part (6) and thermopair front end temperature-sensitive section (3) can remain close contact under 1000 ℃-1400 ℃ hot environment under hot environment, guarantee accuracy and the reliability of temperature-measuring results; Heating source (8) is placed on the below of nonmetal heat insulation material treadmill test part (6), and heating power carries out radiation heating to the lower surface of nonmetal heat insulation material treadmill test part (6); The output signal that is crimped on the temperature thermocouple (2) of nonmetal heat insulation material treadmill test part (6) lower surface is sent to computing machine (10) by wire (9), and computing machine (10) records and calculate the high temperature delta data on the nonmetal heat insulation material treadmill test of guided missile, space shuttle and so on hypersonic aircraft part (6) surface.
2. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism, it is characterized in that: the diameter of described circular hole (5) is 5mm.
3. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism, it is characterized in that: described nonmetal heat insulation material treadmill test part (6) and thermopair front end temperature-sensitive section (3) remain close contact.
4. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism, it is characterized in that: the front portion of described temperature thermocouple (2) is to up-bow, make thermopair front end temperature-sensitive section (3) be higher than the upper plane 5mm-6mm of refractory ceramics framework (1), form initial predeformation upwards.
5. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism is characterized in that: the thermopair front end temperature-sensitive section (3) that is welded into ball-point pen type of described temperature thermocouple (2) forefront is forced to depress 5mm-6mm by refractory ceramics framework (1).
6. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism is characterized in that: described refractory ceramics framework (1) adopts the corundum high-temperature ceramic materials of the alumina content 99% of 1800 ℃ of abilities to fire.
7. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism is characterized in that: the upper cover of described temperature thermocouple (2) has ceramic insulation sleeve pipe (4).
8. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism is characterized in that: described heating source (8) adopts the silicon chrome vermillion external radiation heating tube that can form 1600 ℃ of hot environments.
9. the nonmetal heat insulation material of hypersonic aircraft according to claim 1 plane testpieces surface high-temp measurement mechanism is characterized in that: the double platinum rhodium thermopair of described temperature thermocouple (2) employing measurement range up to 1800 ℃.
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| CN201110063310.4A CN102183312B (en) | 2011-03-16 | 2011-03-16 | Surface high-temperature measurement device for nonmetallic heat resistant material plane test piece of hypersonic speed aircraft |
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| CN105509917B (en) * | 2015-11-27 | 2018-12-11 | 中国航空工业集团公司沈阳飞机设计研究所 | A kind of aircraft tail fairing thermometric experimental rig |
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| CN109655484B (en) * | 2018-12-25 | 2019-11-08 | 中国空气动力研究与发展中心超高速空气动力研究所 | A kind of experimental rig for heat insulation material Performance Assessment |
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| CN112179522A (en) * | 2020-09-14 | 2021-01-05 | 中国空气动力研究与发展中心超高速空气动力研究所 | Aircraft surface temperature sensor testing system and testing method |
| CN113310655B (en) * | 2021-07-29 | 2022-04-26 | 中国空气动力研究与发展中心高速空气动力研究所 | Quick-response temperature measuring device and temperature measuring method for hypersonic wind tunnel |
| CN114018436B (en) * | 2021-11-08 | 2023-11-10 | 北京卫星环境工程研究所 | Spacecraft material space strong electromagnetic environment effect test system |
| CN114722543B (en) * | 2022-06-09 | 2022-08-12 | 中国飞机强度研究所 | Design method for heat reflecting screen in structural heat strength test of hypersonic aircraft |
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