CN115072012A - Thermal examination test system for thermal protection structure of hypersonic aircraft - Google Patents
Thermal examination test system for thermal protection structure of hypersonic aircraft Download PDFInfo
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- CN115072012A CN115072012A CN202210457645.2A CN202210457645A CN115072012A CN 115072012 A CN115072012 A CN 115072012A CN 202210457645 A CN202210457645 A CN 202210457645A CN 115072012 A CN115072012 A CN 115072012A
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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
-
- 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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- Engineering & Computer Science (AREA)
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- General Physics & Mathematics (AREA)
- Immunology (AREA)
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- Health & Medical Sciences (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Testing Resistance To Weather, Investigating Materials By Mechanical Methods (AREA)
Abstract
The invention relates to a thermal assessment test system for a thermal protection structure of a hypersonic aircraft. The test system comprises a closed test cabin, a fuel storage tank, an air inlet pipeline and a fuel gas generator, wherein 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 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 on the cabin body. 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 examination tests of the aerospace craft.
Description
Technical Field
The invention relates to a thermal assessment test system for a thermal protection structure 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 thermal assessment test system of a hypersonic aircraft thermal protection structure, which is energy-saving, low in consumption, small in thermal inertia, high in heating speed and stable in temperature rise, and aims to solve the problems in the prior art.
The technical scheme of the invention is as follows: a thermal examination test system of a thermal protection structure of a hypersonic aircraft comprises a closed test cabin, 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 transmission 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 on the cabin body.
Further, the cabin body is a square cabin body, and the cabin body is composed of a top plate, a bottom plate and four side walls;
the observation window subassembly be the multiunit observation window, the multiunit observation window set up respectively on roof and four side walls, the observation window include big glass disk and the little glass disk that transparent material made, little glass disk set up on the outer face of side wall, big glass disk set up on the interior face of side wall, little glass disk and big glass disk between be equipped with the inner chamber, the inner chamber on be equipped with the air conditioning import.
Furthermore, the inner cavity is horn-shaped, one end facing the small glass wafer is small, the other end facing the large glass wafer is large, and the combination parts of the small glass wafer and the large glass wafer and the inner cavity are respectively provided with a first high-temperature-resistant sealing element.
Furthermore, the plurality of direct flame burners are respectively arranged on the four side walls, high-temperature-resistant sealing pieces II are arranged at the joint parts of the direct flame burners and the side walls, and the number of the direct flame burners on each side wall is the same.
Furthermore, a plurality of direct flame burners on each side wall are arranged from top to bottom, the positions of the plurality of groups of direct flame burners on each side wall are consistent, and the direct flame burners on the side walls on four sides are arranged in the clockwise circumferential direction.
Furthermore, 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 communicated to the outside of the cabin body through a test cable channel, and the detachable cover plate is used for fixedly placing a test platform for the test piece.
Further, the detachable cover plate comprises a cylindrical mounting seat and a square mounting seat, the cylindrical mounting seat and the square mounting seat are of an integrated structure, a chamfer is arranged on the outer ring of the top of the cylindrical mounting seat, and a sealing groove is formed between the cylindrical mounting seat and the square mounting seat and used for placing a high-temperature-resistant sealing element III.
Furthermore, the test piece test platform be used for placing the spacecraft subassembly test piece, the test piece test platform set up the sealed cowling outward, the sealed cowling place and can dismantle the apron on.
Furthermore, a waste heat recovery pipeline is arranged on the side wall of the cabin body.
Furthermore, the test system also comprises a PLC intelligent controller and a plurality of infrared temperature measuring recorders, the infrared temperature measuring recorders are randomly arranged outside the cabin body, and the gas pipeline is provided with a pressure regulating valve.
The invention has the beneficial effects that:
(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 the 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) most aerospace craft need to be connected with cables during thermal examination, and the cables are more in general conditions, so that a test cable channel needs to be arranged in a test cabin and communicated to the outside of the cabin body, the test cable channel can be passed by workers, and in the invention, the test cable channel is arranged on a bottom plate, so that the workers can be connected with the cables from bottom to top after entering the test cable channel, 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 used for heating an aviation aircraft test piece in a closed test cabin made of refractory materials in a full field, 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 examination 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
A thermal examination test system of a thermal protection structure of a 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 transmission pipeline, and the fuel gas generator transmits 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 and 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 arranged in the clockwise circumferential direction.
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 communicated to the outside of the cabin body 11 for a test cable channel, and the detachable cover plate 4 is used for fixedly placing a test platform 6 of a test piece.
The test piece test platform 6 is used for placing the spacecraft component test piece, the seal cover 7 is arranged outside the test piece test platform 6, and the seal 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 cabin body 11 is provided with a waste heat recovery pipeline 8, so that high-temperature gas in the test cabin 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 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, and may be communication between two elements or interaction relationship between two elements, unless otherwise specifically limited, and the specific meaning of the terms in the present invention will 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 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 (10)
1. The utility model provides a hot examination test system of hot protective structure of hypersonic aircraft which characterized in that: the test system comprises a closed test cabin (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 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 on the cabin body (11).
2. The thermal assessment test system for the thermal protection structure of the hypersonic aircraft according to claim 1, 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 and advance mouthful (34).
3. The thermal assessment test system for the thermal protection structure of the hypersonic aircraft according to claim 2, characterized in that: inner chamber (33) be the loudspeaker form, it is little towards the one end of little glass disk (32), and the one end towards big glass disk (31) is big, little glass disk (32) and big glass disk (31) and the joint part of inner chamber (33) set up high temperature resistant sealing member one respectively.
4. The thermal assessment test system for the thermal protection structure of the hypersonic aircraft according to claim 1, characterized in that: the plurality of the direct flame burners (2) are respectively arranged on the side walls (113) on four sides, the joint parts of the direct flame burners (2) and the side walls (113) are provided with high-temperature-resistant sealing pieces II, and the number of the direct flame burners (2) on each side wall (113) is the same.
5. The thermal assessment test system of the thermal protection structure of the hypersonic aircraft according to claim 4, characterized in that: a plurality of direct flame burners (2) on every side wall (113) top-down set up, the position of multiunit direct flame burner (2) on every side wall (113) is unanimous, direct flame burner (2) on the side wall (113) of four sides are clockwise circumference and arrange.
6. The thermal assessment test system for the thermal protection structure of the hypersonic flight vehicle of claim 2, characterized in that: the center is equipped with and can dismantle apron (4) on bottom plate (112) of the cabin body (11), can dismantle apron (4) center and communicate outside the cabin body (11) for test cable passageway (5), can dismantle apron (4) be used for the fixed test platform (6) of placing of test piece.
7. The thermal assessment test system of the thermal protection structure of the hypersonic aircraft according to claim 6, characterized in that: detachable cover board (4) include cylindrical mount pad (41) and square mount pad (42), cylindrical mount pad (41) and square mount pad (42) formula structure as an organic whole, cylindrical mount pad (41) top outer lane be equipped with the chamfer, cylindrical mount pad (41) and square mount pad (42) between be equipped with the seal groove and be used for placing high temperature resistant sealing member three.
8. The thermal assessment test system of the thermal protection structure of the hypersonic aircraft according to claim 6, characterized in that: the test piece testing platform (6) is used for placing the spacecraft component test pieces, 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).
9. The thermal assessment test system for the thermal protection structure of the hypersonic aircraft according to claim 2, characterized in that: and a waste heat recovery pipeline (8) is arranged on the side wall (113) of the cabin body (11).
10. The thermal assessment test system of the thermal protection structure of the hypersonic aircraft according to claim 6, 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 randomly arranged outside the cabin body (11), and the gas pipeline is provided with a pressure regulating valve.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202210457645.2A CN115072012A (en) | 2022-04-28 | 2022-04-28 | Thermal examination test system for thermal protection structure of hypersonic aircraft |
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| CN202210457645.2A CN115072012A (en) | 2022-04-28 | 2022-04-28 | Thermal examination test system for thermal protection structure of hypersonic aircraft |
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