CN115416880A - Application method of thermal assessment test system for hypersonic aircraft - Google Patents
Application method of thermal assessment test system for hypersonic aircraft Download PDFInfo
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- CN115416880A CN115416880A CN202210462605.7A CN202210462605A CN115416880A CN 115416880 A CN115416880 A CN 115416880A CN 202210462605 A CN202210462605 A CN 202210462605A CN 115416880 A CN115416880 A CN 115416880A
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G7/00—Simulating cosmonautic conditions, e.g. for conditioning crews
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N25/00—Investigating or analyzing materials by the use of thermal means
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Abstract
The invention relates to a use method of a thermal assessment test system of a hypersonic aircraft. A worker enters the closed test chamber, a test piece testing platform is placed at the center of the detachable cover plate, an aircraft component test piece is placed on the test piece testing platform, a sealing cover is placed on the outer side of the test piece testing platform, and the closed test chamber is closed; a worker enters a test cable channel from the outside to reach the bottom of the closed test cabin, and cable connection is carried out on an aircraft assembly test piece placed on a test piece test platform; igniting a plurality of direct flame burners simultaneously, heating a test piece of an aircraft component, collecting data by a test data collecting module, closing the direct flame burners after the temperature in a closed test cabin reaches a set temperature, and entering a heat preservation stage; and when the temperature in the closed test chamber is lower than 200 ℃, opening the closed test chamber to finish the hot test. The method has the advantages of energy conservation, low consumption, small thermal inertia, high heating speed, stable temperature rise and convenient operation.
Description
Technical Field
The invention relates to a use method of thermal assessment equipment, in particular to a use method of a thermal assessment test system of a hypersonic aircraft.
Background
When the hypersonic aircraft flies at a high speed in the atmospheric layer, the thermal protection system is subjected to multiple thermal-force coupling actions such as high-enthalpy pneumatic heating, mechanical load, pneumatic shear force, thermal stress generated by temperature gradient and the like, the thermal protection material can be ablated and peeled off under the action of high temperature, the mechanical load, the pneumatic shear force and the thermal stress can damage the material substrate to generate microcracks, and the ablation and peeling processes of the thermal protection material are accelerated, so that the material structure deformation and the strength change of the aircraft can be caused, the accurate control of the aircraft is influenced, or the safety of the aircraft is endangered.
At present, the high-temperature test field for testing aerospace devices at home mostly adopts an electric heating mode, and the electric heating mode is mostly a plurality of static heating modes such as quartz lamp heating, graphite heating or electric arc heating. The quartz lamp or graphite heating mode has the characteristics of small thermal inertia, wide applicability, convenience in control and the like, but the direct use of electric energy generates heat, so that the power consumption is huge and the use cost is high; the electric arc heating mode has the advantages of high heating enthalpy value and simple components of the airflow medium, but the heating time is short, the size of a product which can be used for testing is small, and the application range is greatly limited. Therefore, a new heating mode is needed for high-temperature tests of hypersonic aircrafts, especially for high-temperature thermal examination of thermal protection structures.
Disclosure of Invention
The invention aims to provide a using method of a thermal assessment test system of a hypersonic aircraft, which is energy-saving, consumption-reducing, small in thermal inertia, high in heating speed, stable in temperature rise and convenient to operate, and aims to solve the problems in the prior art.
The technical scheme of the invention is as follows: the use method of the test system for the thermal assessment of the hypersonic aircraft comprises a closed test cabin, a fuel storage tank, an air inlet pipeline, a fuel gas generator and a test data acquisition module, wherein the fuel storage tank is communicated to the fuel gas generator through an air conveying pipeline, and the fuel gas generator conveys high-temperature fuel gas into the closed test cabin through the fuel gas pipeline; the closed test chamber comprises a chamber body, a chamber door, a plurality of direct flame burners and a plurality of observation window assemblies; the plurality of direct flame burners are arranged in the cabin body and connected with an external gas pipeline, and the plurality of observation window assemblies are respectively arranged in the cabin body; the center of the bottom plate of the cabin body is provided with a detachable cover plate, the center of the detachable cover plate is a test cable channel communicated to the outside of the cabin body, and the detachable cover plate is used for fixedly placing a test platform of a test piece; the test piece testing platform is used for placing a test piece of the aerospace craft component, a sealing cover is arranged outside the test piece testing platform, and the sealing cover is placed on the detachable cover plate;
the using method comprises the following steps: a worker enters the closed test chamber, a test piece testing platform is placed at the center of the detachable cover plate, an aircraft component test piece is placed on the test piece testing platform, a sealing cover is placed on the outer side of the test piece testing platform, the lower end of the sealing cover is sealed with the detachable cover plate, and after the setting is finished, the worker exits from the closed test chamber and seals the closed test chamber;
a worker enters the test cable channel from the outside to reach the bottom of the closed test cabin, cable connection is carried out on an aircraft assembly test piece placed on the test piece test platform, after the arrangement is completed, the worker leaves from the test cable channel, and the external inlet of the test cable channel is closed;
simultaneously igniting a plurality of direct flame burners, heating the closed test chamber, heating the aircraft component test piece at high temperature after the heating is finished, acquiring data by a test data acquisition module, observing the flame state of the direct flame burners through an observation window component to ensure that the flame is in a light blue and yellowish state, and closing the direct flame burners to enter a heat preservation stage after the temperature in the closed test chamber reaches a set temperature;
and when the temperature in the closed test chamber is lower than 200 ℃, opening the closed test chamber to finish the hot test.
Further, the cabin body is a square cabin body, and the cabin body consists of a top plate, a bottom plate and four side walls;
the observation window assembly is a plurality of groups of observation windows, the plurality of groups of observation windows are respectively arranged on the top plate and the four-side wall, each observation window comprises a large glass wafer and a small glass wafer which are made of transparent materials, the small glass wafer is arranged on the outer plate surface of the side wall, the large glass wafer is arranged on the inner plate surface of the side wall, an inner cavity is formed between the small glass wafer and the large glass wafer, and a cold air inlet is formed in the inner cavity.
Furthermore, the plurality of direct flame burners are respectively arranged on the four side walls, the joint parts of the direct flame burners and the side walls are provided with second high-temperature-resistant sealing parts, and the number of the direct flame burners on each side wall is the same.
Furthermore, the test system also comprises a combustion fan, and the combustion fan is interlocked with the direct flame burner.
Furthermore, the direct flame burner fails to ignite continuously for three times, and can ignite again after the direct flame nozzle is purged.
Furthermore, the test system also comprises a PLC intelligent controller and a plurality of infrared temperature measurement recorders, the infrared temperature measurement recorders are randomly arranged outside the cabin body, a pressure regulating valve is arranged on the gas pipeline, and the pressure regulating valve and the infrared temperature measurement recorders are interlocked in an overtemperature way.
The beneficial effects of the invention are:
(1) Although the prior heating cabin is also provided with the observation window, the prior heating cabin is mostly provided with a single-sheet glass window, the glass window with the structure has the defects that the observation visual field is limited, the state in the heating cabin cannot be clearly observed, and the observation window is easy to damage in a high-temperature environment due to higher temperature in the heating cabin;
(2) The direct flame burners are arranged at the same position of each side wall, namely the direct flame burners on the side walls on four sides are arranged in the clockwise circumferential direction, when a plurality of burners are ignited simultaneously, rotary airflow can be generated rapidly, the flowability of high-temperature gas in the test chamber is enhanced, meanwhile, the heating effect in the test chamber is more uniform, the heating speed is increased, the test working efficiency is improved, and the damage of a test piece caused by the uneven rapid heating can be prevented;
(3) The test cable channel is arranged on the bottom plate, so that after a worker enters the test cable channel, the worker can be connected with the cable from bottom to top, and the operation is more convenient.
The test system takes natural gas as a heating source, high-temperature gas generated by the direct flame burner is utilized to carry out full-field heating on an aviation aircraft test piece in a closed test cabin made of refractory materials, the highest heating temperature can reach 1500K, the test system has the advantages of low test cost, large test structure size, wide applicability, short test preparation period and the like, and can meet the requirements of high-frequency and low-cost thermal check tests on aerospace and aviation equipment.
Drawings
Fig. 1 is a schematic structural view of the present invention.
Fig. 2 is a schematic structural diagram of the present invention.
Fig. 3 is a schematic structural view of a side wall.
Fig. 4 is a cross-sectional schematic view of a viewing window assembly.
In the figure, 1 is a closed test chamber, 11 is a chamber body, 111 is a top plate, 112 is a bottom plate, and 113 is a side wall;
2 is a direct flame burner;
3 is an observation window component, 31 is a large glass wafer, 32 is a small glass wafer, 33 is an inner cavity, and 34 is a cold air inlet;
4 is a detachable cover plate, 41 is a cylindrical mounting seat, and 42 is a square mounting seat;
5 is a test cable channel;
6 is a test piece testing platform;
7 is a sealing cover;
and 8 is a waste heat recovery pipeline.
Detailed Description
The use method of the test system for the thermal examination of the hypersonic aircraft comprises a closed test chamber 1, a fuel storage tank, an air inlet pipeline and a fuel gas generator, wherein the fuel storage tank is communicated to the fuel gas generator through an air conveying pipeline, and the fuel gas generator conveys high-temperature fuel gas into the closed test chamber 1 through a fuel gas pipeline;
the closed test chamber 1 comprises a chamber body 11, a chamber door, a plurality of direct flame burners 2 and a plurality of observation window assemblies 3;
the plurality of direct flame burners 2 are arranged in the cabin body 11 and connected with a gas pipeline, and the plurality of observation window assemblies 3 are respectively arranged on the cabin body 11.
The cabin body 11 is a square cabin body made of refractory materials, and is composed of a top plate 111, a bottom plate 112 and four side walls 113;
The inner cavity 33 is trumpet-shaped, one end facing the small glass wafer 32 is small, the other end facing the large glass wafer 31 is large, and the combination parts of the small glass wafer 32, the large glass wafer 31 and the inner cavity 33 are respectively provided with a first high-temperature-resistant sealing element.
The observation window is composed of two pieces of glass and a horn-shaped inner cavity 33 between the two pieces of glass, and can be arranged on the top plate 111 and the four side walls 113, so that the observation visual field is widened, the state of a test piece in the test chamber can be clearly observed, and meanwhile, cooling gas is injected into the inner cavity 33 through the cold air inlet 34, the temperature of the observation window is reduced, and the observation window is prevented from being damaged due to high temperature.
The plurality of the direct flame burners 2 are respectively arranged on the four side walls 113, the joint parts of the direct flame burners 2 and the side walls 113 are provided with second high-temperature-resistant sealing parts, and the number of the direct flame burners 2 on each side wall 113 is the same.
The plurality of direct flame burners 2 on each side wall 113 are arranged from top to bottom, the positions of the plurality of groups of direct flame burners 2 on each side wall 113 are consistent, and the direct flame burners 2 on the four side walls 113 are circumferentially arranged clockwise.
This test system uses the natural gas as the heating heat source, in airtight test chamber 1 by refractory material constitution, utilize the high temperature gas that gas generator produced, carry out the hot examination test to the test piece of placing in test chamber 1 through direct flame nozzle 2, direct flame nozzle 2 sets up the same position at every side wall 113, also be exactly that totally 12 direct flame nozzles 2 on the four sides wall 113 are clockwise circumference and arrange, after 12 direct flame nozzles ignite simultaneously, can produce rotation type high temperature air current rapidly, the mobility of high temperature gas in test chamber has not only been strengthened, and simultaneously, make the effect of test chamber heating intensification also more even, rate of heating also more accelerates, improve experimental work efficiency, can also prevent to cause the damage of test piece because of rapid heating is inhomogeneous.
The invention is provided with a set of gas main pipelines and twelve sets of gas branch pipelines, various valves, switches, pressure gauges and other elements are arranged on the main pipelines and the branch pipelines, and every three burners 2 share one set of electric actuating mechanism, thereby having the functions of automatic ignition, sequential control, flame monitoring, automatic temperature regulation and the like.
The center of the bottom plate 112 of the cabin body 11 is provided with a detachable cover plate 4, the center of the detachable cover plate 4 is a test cable channel communicated to the outside of the cabin body 11, and the detachable cover plate 4 is used for fixedly placing a test piece test platform 6.
The test piece testing platform 6 is used for placing the spacecraft component test piece, the sealing cover 7 is arranged outside the test piece testing platform 6, and the sealing cover 7 is placed on the detachable cover plate 4.
The test system is mainly used for thermal examination of aerospace equipment, most aerospace equipment needs to be connected with cables during thermal examination, and the cables are generally more, so that a test cable channel 5 needs to be arranged in a test cabin and communicated to the outside of a cabin body 11, the test cable channel 5 can be passed by workers, and in the test system, the test cable channel 5 is arranged on a bottom plate 112, so that the workers can be connected with the cables from bottom to top after entering the test cable channel 5, and the operation is more convenient.
Because the size and the size of the test piece are different, the detachable cover plate 4 comprises a cylindrical mounting seat 41 and a square mounting seat 42, and is suitable for test piece test platforms 6 with different assembly requirements, so that the detachable cover plate can adapt to test pieces with different sizes and sizes, is wide in application range and is firm in connection.
The test system still include PLC intelligent control ware and a plurality of infrared temperature measurement record appearance, a plurality of infrared temperature measurement record appearance set up outside the cabin body 11 wantonly, can be with the accurate record of the temperature in a plurality of different regions in the cabin body 11, carry out the temperature measurement to the test piece from a plurality of directions, note the temperature variation of test piece in the cabin body 11, the test result is more accurate, the gas pipeline on be equipped with the air-vent valve to connect the pressure in the PLC intelligent control ware automatic regulation gas pipeline, improve experimental stability.
The cabin body 11 further comprises a cabin shell, a cabin lining, a cabin opening and a cabin door, wherein the cabin shell is of a rectangular frame structure, a frame is formed by combining and welding angle steel and channel steel, and a steel plate is adhered and welded on the inner surface of the frame; the cabin body 11 side part is provided with a ladder stand for facilitating the loading and unloading of the cabin top, the cabin shell is firmly welded, the appearance is attractive, the whole cabin shell is simple and large, polishing and derusting are carried out after the manufacturing is finished, the antirust primer is coated, the outer surface is sprayed with paint, and the paint surface is smooth and clean.
The cabin lining sets up in 11 bottoms in the cabin body, side wall and hatch department, the cabin lining include flame retardant coating, heat preservation and insulating layer, the flame retardant coating is built by laying bricks or stones for high alumina standard brick, the heat preservation is built by laying bricks or stones for light firebrick, the insulating layer is built by laying bricks or stones for the diatomaceous earth brick, compromise good heat preservation and thermal insulation performance when guaranteeing intensity, it is durable, impact resistance is strong, the cabin top adopts the preparation of zirconium-containing fiber module to form and adopts the aluminium silicate fibre blanket to lay, there is certain inclined plane when the hatch is built by laying bricks or stones, can reach good sealed when guaranteeing that the hatch door closes and compresses tightly, reduce the calorific loss of hatch.
The hatch door adopts the mode of singly opening the door by hand, and the hatch door adopts two sets of bearings to fix on runing rest, and the hatch door sets up all around and compresses tightly the hand wheel, and when the hatch door was closed to needs, the hatch door was pressed into to artifical rotatory hatch door, adopts the hand wheel to compress tightly the hatch door. The cabin door is formed by welding profile steel and plates, the cabin door is internally manufactured by adopting a zirconium-containing mixed fiber module, and an aluminum silicate fiber blanket is adopted for laying and insulating heat. The hatch has certain inclined planes around, and can form a flexible sealing surface with the hatch, so as to achieve the best sealing effect.
The side wall 113 of the chamber body 11 is provided with a waste heat recovery pipeline 8, so that high-temperature gas in the test chamber can be recycled after the test is finished.
The working principle of the invention is as follows: the natural gas is used as a heating source, high-temperature gas generated by the direct flame burner 2 is used for heating the structural member of the aerospace craft in a full field in the closed test chamber 1 made of refractory materials, and the maximum heating temperature can reach 1500K.
The specific using method comprises the following steps: a worker enters the closed test chamber 1, a test piece testing platform 6 is placed at the center of the detachable cover plate 4, an aircraft component test piece is placed on the test piece testing platform 6, a sealing cover 7 is placed on the outer side of the test piece testing platform 6, the lower end of the sealing cover 7 is sealed with the detachable cover plate 4, and after the setting is completed, the worker exits from the closed test chamber 1 and closes the closed test chamber 1;
a worker enters the test cable channel 5 from the outside to reach the bottom of the closed test cabin 1, cable connection is carried out on an aircraft assembly test piece placed on the test piece test platform 6, after the arrangement is completed, the worker leaves from the test cable channel 5, and an external inlet of the test cable channel 5 is closed;
simultaneously igniting the plurality of direct flame burners 2, and heating the closed test chamber, wherein the heating technical parameters are as follows:
after the heating is finished, heating the aircraft component test piece at a high temperature, acquiring data by a test data acquisition module, and simultaneously observing the flame state of the direct flame burner 2 through an observation window component 3 to ensure that the flame is in a light blue and yellowish state, and after the temperature in the closed test chamber 1 reaches a set temperature, closing the direct flame burner 2 to enter a heat preservation stage;
and when the temperature in the closed test chamber 1 is lower than 200 ℃, opening the closed test chamber 1 to finish the thermal examination test.
The standard parts used by the invention can be purchased from the market, and the special-shaped parts can be customized according to the description and the description of the attached drawings.
In the present invention, unless otherwise explicitly specified or limited, for example, it may be fixedly attached, detachably attached, or integrated; can be mechanically or electrically connected; the terms may be directly connected or indirectly connected through an intermediate agent, and may be used for communicating the inside of two elements or interacting relation of two elements, unless otherwise specifically defined, and the specific meaning of the terms in the present invention can be understood by those skilled in the art according to specific situations.
Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions and alterations can be made in these embodiments without departing from the principles and spirit of the invention, the scope of which is defined in the appended claims and their equivalents.
Claims (6)
1. A use method of a thermal assessment test system of a hypersonic aircraft is characterized by comprising the following steps: the test system comprises a closed test cabin (1), a fuel storage tank, an air inlet pipeline, a fuel gas generator and a test data acquisition module, wherein the fuel storage tank is communicated to the fuel gas generator through an air transmission pipeline, and the fuel gas generator transmits high-temperature fuel gas into the closed test cabin (1) through a fuel gas pipeline; the closed test chamber (1) comprises a chamber body (11), a chamber door, a plurality of direct flame burners (2) and a plurality of observation window assemblies (3); the plurality of direct flame burners (2) are arranged in the cabin body (11) and connected with an external gas pipeline, and the plurality of observation window assemblies (3) are respectively arranged in the cabin body (11); a detachable cover plate (4) is arranged in the center of a bottom plate (112) of the cabin body (11), a test cable channel (5) is arranged in the center of the detachable cover plate (4) and communicated to the outside of the cabin body (11), and the detachable cover plate (4) is used for fixedly placing a test platform (6) of a test piece; the test piece testing platform (6) is used for placing a test piece of the aerospace craft component, a sealing cover (7) is arranged outside the test piece testing platform (6), and the sealing cover (7) is placed on the detachable cover plate (4);
the using method comprises the following steps: a worker enters the closed test cabin (1), a test piece testing platform (6) is placed at the center position of the detachable cover plate (4), an aircraft component test piece is placed on the test piece testing platform (6), a sealing cover (7) is placed on the outer side of the test piece testing platform (6), the lower end of the sealing cover (7) is sealed with the detachable cover plate (4), and after the setting is completed, the worker exits the closed test cabin (1) and closes the closed test cabin (1);
a worker enters the test cable channel (5) from the outside to reach the bottom of the closed test cabin (1), cable connection is carried out on an aircraft assembly test piece placed on the test piece test platform (6), after the arrangement is completed, the worker leaves from the test cable channel (5), and an external inlet of the test cable channel (5) is closed;
simultaneously igniting a plurality of direct flame burners (2), heating the closed test chamber, heating the aircraft component test piece at high temperature after the heating is finished, acquiring data by a test data acquisition module, observing the flame state of the direct flame burners (2) through an observation window component (3) to ensure that the flame is in a light blue and yellowish state, and closing the direct flame burners (2) to enter a heat preservation stage after the temperature in the closed test chamber (1) reaches a set temperature;
and when the temperature in the closed test chamber (1) is lower than 200 ℃, opening the closed test chamber (1) to finish the thermal examination test.
2. The use method of the hypersonic aircraft thermal assessment test system according to claim 1, is characterized in that: the cabin body (11) is a square cabin body, and the cabin body is composed of a top plate (111), a bottom plate (112) and four side walls (113);
the observation window subassembly be the multiunit observation window, the multiunit observation window set up respectively on roof (111) and four sides side wall (113), the observation window include big glass disk (31) and little glass disk (32) that transparent material made, little glass disk (32) set up on the outer face of side wall (113), big glass disk (31) set up on the interior face of side wall (113), little glass disk (32) and big glass disk (31) between be equipped with inner chamber (33), inner chamber (33) on be equipped with air conditioning import (34).
3. The use method of the hypersonic aircraft thermal assessment test system according to claim 1, characterized in that: the plurality of direct flame burners (2) are respectively arranged on the four side walls (113), the joint parts of the direct flame burners (2) and the side walls (113) are provided with high-temperature-resistant sealing parts II, and the number of the direct flame burners (2) on each side wall (113) is the same.
4. The use method of the hypersonic aircraft thermal assessment test system according to claim 1, characterized in that: the test system also comprises a combustion fan, and the combustion fan is interlocked with the direct flame burner (2).
5. The use method of the hypersonic aircraft thermal assessment test system according to claim 1, is characterized in that: the straight flame burner (2) fails to ignite continuously for three times, and can ignite again after the straight flame burner (2) is purged.
6. The use method of the hypersonic aircraft thermal assessment test system according to claim 1, is characterized in that: the test system further comprises a PLC intelligent controller and a plurality of infrared temperature measuring recorders, the infrared temperature measuring recorders are arranged outside the cabin body (11) at will, a pressure regulating valve is arranged on the gas pipeline, and the pressure regulating valve and the infrared temperature measuring recorders are in over-temperature interlocking.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210462605.7A CN115416880A (en) | 2022-04-28 | 2022-04-28 | Application method of thermal assessment test system for hypersonic aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210462605.7A CN115416880A (en) | 2022-04-28 | 2022-04-28 | Application method of thermal assessment test system for hypersonic aircraft |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN115416880A true CN115416880A (en) | 2022-12-02 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117284508A (en) * | 2023-09-25 | 2023-12-26 | 北京交通大学 | Pneumatic relative pressure difference distribution and thermal simulation test system and test method thereof |
Citations (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH07260370A (en) * | 1994-03-25 | 1995-10-13 | Sumitomo Metal Ind Ltd | Window for viewing thermal-spray repair inside kiln/ furnace |
| JP2001165878A (en) * | 1999-12-14 | 2001-06-22 | Tostech:Kk | Infrared ray high temperature heating furnace |
| KR101018010B1 (en) * | 2010-11-18 | 2011-03-02 | 한국건설기술연구원 | Test furnace having heating device for circulating fire flame, and heating method for the same |
| CN202145199U (en) * | 2011-06-10 | 2012-02-15 | 上海化工研究院 | Safety barrier device for improving heat accumulation experiment state |
| WO2012035041A2 (en) * | 2010-09-16 | 2012-03-22 | Eads Deutschland Gmbh | Testing device for testing test objects made of high-temperature-resistant material |
| KR101264246B1 (en) * | 2012-12-12 | 2013-05-27 | 한국건설기술연구원 | Multipurpose vertical test furnace having flat-type burner for evaluating fire resistance capacity of vertical structure member, and heating method using the same |
| CN103982115A (en) * | 2014-05-28 | 2014-08-13 | 上海交通大学 | Vacuum glass high-low-temperature observation window device having temperature compensation function |
| CN104596781A (en) * | 2013-10-30 | 2015-05-06 | 北京强度环境研究所 | Flame heating type thermal noise combined environment testing device |
| US20170059527A1 (en) * | 2015-09-01 | 2017-03-02 | Corestar International Corporation | Hand-Held Eddy Current Test Instrument and Sensor |
| CN109632886A (en) * | 2018-12-12 | 2019-04-16 | 北京空天技术研究所 | Fine hot certification test system and method in a kind of high-speed aircraft cabin |
| CN111077855A (en) * | 2019-11-26 | 2020-04-28 | 上海空间电源研究所 | Be applied to spacecraft unit heat balance control system |
| WO2020096477A1 (en) * | 2018-11-09 | 2020-05-14 | Дмитрий Вячеславович ФЕДОТОВ | Thermodynamic test bench for simulating aerodynamic heating |
| CN111929016A (en) * | 2020-07-27 | 2020-11-13 | 中国航发湖南动力机械研究所 | Heating furnace and high-temperature high-cycle fatigue test device |
| CN112129483A (en) * | 2020-08-21 | 2020-12-25 | 中国航天空气动力技术研究院 | Ablation ground simulation test device and method based on ablation backset compensation |
| US20210080181A1 (en) * | 2019-09-16 | 2021-03-18 | Nanoracks, Llc | Space Oven |
| CN212806531U (en) * | 2020-07-27 | 2021-03-26 | 昊洋科技(天津)有限公司 | Modular building fire resistance test furnace |
| CN214846721U (en) * | 2021-02-03 | 2021-11-23 | 北京空间飞行器总体设计部 | Ground cable wiring structure for thermal test of large spacecraft |
-
2022
- 2022-04-28 CN CN202210462605.7A patent/CN115416880A/en active Pending
Patent Citations (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH07260370A (en) * | 1994-03-25 | 1995-10-13 | Sumitomo Metal Ind Ltd | Window for viewing thermal-spray repair inside kiln/ furnace |
| JP2001165878A (en) * | 1999-12-14 | 2001-06-22 | Tostech:Kk | Infrared ray high temperature heating furnace |
| WO2012035041A2 (en) * | 2010-09-16 | 2012-03-22 | Eads Deutschland Gmbh | Testing device for testing test objects made of high-temperature-resistant material |
| KR101018010B1 (en) * | 2010-11-18 | 2011-03-02 | 한국건설기술연구원 | Test furnace having heating device for circulating fire flame, and heating method for the same |
| CN202145199U (en) * | 2011-06-10 | 2012-02-15 | 上海化工研究院 | Safety barrier device for improving heat accumulation experiment state |
| KR101264246B1 (en) * | 2012-12-12 | 2013-05-27 | 한국건설기술연구원 | Multipurpose vertical test furnace having flat-type burner for evaluating fire resistance capacity of vertical structure member, and heating method using the same |
| CN104596781A (en) * | 2013-10-30 | 2015-05-06 | 北京强度环境研究所 | Flame heating type thermal noise combined environment testing device |
| CN103982115A (en) * | 2014-05-28 | 2014-08-13 | 上海交通大学 | Vacuum glass high-low-temperature observation window device having temperature compensation function |
| US20170059527A1 (en) * | 2015-09-01 | 2017-03-02 | Corestar International Corporation | Hand-Held Eddy Current Test Instrument and Sensor |
| WO2020096477A1 (en) * | 2018-11-09 | 2020-05-14 | Дмитрий Вячеславович ФЕДОТОВ | Thermodynamic test bench for simulating aerodynamic heating |
| CN109632886A (en) * | 2018-12-12 | 2019-04-16 | 北京空天技术研究所 | Fine hot certification test system and method in a kind of high-speed aircraft cabin |
| US20210080181A1 (en) * | 2019-09-16 | 2021-03-18 | Nanoracks, Llc | Space Oven |
| CN111077855A (en) * | 2019-11-26 | 2020-04-28 | 上海空间电源研究所 | Be applied to spacecraft unit heat balance control system |
| CN111929016A (en) * | 2020-07-27 | 2020-11-13 | 中国航发湖南动力机械研究所 | Heating furnace and high-temperature high-cycle fatigue test device |
| CN212806531U (en) * | 2020-07-27 | 2021-03-26 | 昊洋科技(天津)有限公司 | Modular building fire resistance test furnace |
| CN112129483A (en) * | 2020-08-21 | 2020-12-25 | 中国航天空气动力技术研究院 | Ablation ground simulation test device and method based on ablation backset compensation |
| CN214846721U (en) * | 2021-02-03 | 2021-11-23 | 北京空间飞行器总体设计部 | Ground cable wiring structure for thermal test of large spacecraft |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117284508A (en) * | 2023-09-25 | 2023-12-26 | 北京交通大学 | Pneumatic relative pressure difference distribution and thermal simulation test system and test method thereof |
| CN117284508B (en) * | 2023-09-25 | 2024-03-22 | 北京交通大学 | Pneumatic relative pressure difference distribution and thermal simulation test system and test method thereof |
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