CN117326046A - Hypersonic aircraft surface can thermal protection system of dimension shape - Google Patents
Hypersonic aircraft surface can thermal protection system of dimension shape Download PDFInfo
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- CN117326046A CN117326046A CN202311516998.6A CN202311516998A CN117326046A CN 117326046 A CN117326046 A CN 117326046A CN 202311516998 A CN202311516998 A CN 202311516998A CN 117326046 A CN117326046 A CN 117326046A
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- 239000000919 ceramic Substances 0.000 claims abstract description 114
- 239000002826 coolant Substances 0.000 claims abstract description 59
- 229920000049 Carbon (fiber) Polymers 0.000 claims abstract description 55
- 239000004917 carbon fiber Substances 0.000 claims abstract description 55
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims abstract description 54
- 239000011347 resin Substances 0.000 claims abstract description 50
- 229920005989 resin Polymers 0.000 claims abstract description 50
- 239000002131 composite material Substances 0.000 claims abstract description 46
- 239000011159 matrix material Substances 0.000 claims abstract description 45
- 238000013461 design Methods 0.000 claims abstract description 14
- MCMNRKCIXSYSNV-UHFFFAOYSA-N Zirconium dioxide Chemical compound O=[Zr]=O MCMNRKCIXSYSNV-UHFFFAOYSA-N 0.000 claims description 6
- 239000000805 composite resin Substances 0.000 claims description 6
- 229910021426 porous silicon Inorganic materials 0.000 claims description 6
- HBMJWWWQQXIZIP-UHFFFAOYSA-N silicon carbide Chemical compound [Si+]#[C-] HBMJWWWQQXIZIP-UHFFFAOYSA-N 0.000 claims description 6
- BPQQTUXANYXVAA-UHFFFAOYSA-N Orthosilicate Chemical compound [O-][Si]([O-])([O-])[O-] BPQQTUXANYXVAA-UHFFFAOYSA-N 0.000 claims description 5
- PNEYBMLMFCGWSK-UHFFFAOYSA-N aluminium oxide Inorganic materials [O-2].[O-2].[O-2].[Al+3].[Al+3] PNEYBMLMFCGWSK-UHFFFAOYSA-N 0.000 claims description 5
- 229910052727 yttrium Inorganic materials 0.000 claims description 5
- VWQVUPCCIRVNHF-UHFFFAOYSA-N yttrium atom Chemical compound [Y] VWQVUPCCIRVNHF-UHFFFAOYSA-N 0.000 claims description 5
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N Silicium dioxide Chemical compound O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 claims description 4
- KZHJGOXRZJKJNY-UHFFFAOYSA-N dioxosilane;oxo(oxoalumanyloxy)alumane Chemical compound O=[Si]=O.O=[Si]=O.O=[Al]O[Al]=O.O=[Al]O[Al]=O.O=[Al]O[Al]=O KZHJGOXRZJKJNY-UHFFFAOYSA-N 0.000 claims description 3
- 229910052863 mullite Inorganic materials 0.000 claims description 3
- 229910010271 silicon carbide Inorganic materials 0.000 claims description 3
- HQVNEWCFYHHQES-UHFFFAOYSA-N silicon nitride Chemical compound N12[Si]34N5[Si]62N3[Si]51N64 HQVNEWCFYHHQES-UHFFFAOYSA-N 0.000 claims description 3
- GFQYVLUOOAAOGM-UHFFFAOYSA-N zirconium(iv) silicate Chemical compound [Zr+4].[O-][Si]([O-])([O-])[O-] GFQYVLUOOAAOGM-UHFFFAOYSA-N 0.000 claims description 3
- 239000000377 silicon dioxide Substances 0.000 claims 1
- 239000000463 material Substances 0.000 abstract description 18
- 238000000034 method Methods 0.000 abstract description 15
- 230000000694 effects Effects 0.000 abstract description 5
- 238000000197 pyrolysis Methods 0.000 abstract description 5
- 230000009286 beneficial effect Effects 0.000 abstract description 2
- 238000012360 testing method Methods 0.000 description 45
- 239000000523 sample Substances 0.000 description 13
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 13
- 230000008859 change Effects 0.000 description 11
- 230000008569 process Effects 0.000 description 11
- 230000009467 reduction Effects 0.000 description 11
- 239000000047 product Substances 0.000 description 10
- 230000035900 sweating Effects 0.000 description 10
- 238000001816 cooling Methods 0.000 description 7
- QFXZANXYUCUTQH-UHFFFAOYSA-N ethynol Chemical group OC#C QFXZANXYUCUTQH-UHFFFAOYSA-N 0.000 description 7
- 239000007789 gas Substances 0.000 description 7
- 230000001105 regulatory effect Effects 0.000 description 7
- 238000002679 ablation Methods 0.000 description 6
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- 238000010438 heat treatment Methods 0.000 description 6
- 239000012466 permeate Substances 0.000 description 5
- 230000005540 biological transmission Effects 0.000 description 4
- 239000003575 carbonaceous material Substances 0.000 description 3
- 229910010293 ceramic material Inorganic materials 0.000 description 3
- 238000010586 diagram Methods 0.000 description 3
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- 230000002787 reinforcement Effects 0.000 description 3
- XKRFYHLGVUSROY-UHFFFAOYSA-N Argon Chemical compound [Ar] XKRFYHLGVUSROY-UHFFFAOYSA-N 0.000 description 2
- IJGRMHOSHXDMSA-UHFFFAOYSA-N Atomic nitrogen Chemical compound N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 description 2
- CURLTUGMZLYLDI-UHFFFAOYSA-N Carbon dioxide Chemical compound O=C=O CURLTUGMZLYLDI-UHFFFAOYSA-N 0.000 description 2
- 150000001721 carbon Chemical class 0.000 description 2
- 239000005539 carbonized material Substances 0.000 description 2
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- 239000011148 porous material Substances 0.000 description 2
- 229910052814 silicon oxide Inorganic materials 0.000 description 2
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 238000010521 absorption reaction Methods 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 230000004913 activation Effects 0.000 description 1
- 239000004964 aerogel Substances 0.000 description 1
- 239000003570 air Substances 0.000 description 1
- 229910052786 argon Inorganic materials 0.000 description 1
- 230000037237 body shape Effects 0.000 description 1
- 229910052799 carbon Inorganic materials 0.000 description 1
- 239000001569 carbon dioxide Substances 0.000 description 1
- 229910002092 carbon dioxide Inorganic materials 0.000 description 1
- 239000003054 catalyst Substances 0.000 description 1
- 238000005524 ceramic coating Methods 0.000 description 1
- 238000002468 ceramisation Methods 0.000 description 1
- 239000011248 coating agent Substances 0.000 description 1
- 238000000576 coating method Methods 0.000 description 1
- 230000001276 controlling effect Effects 0.000 description 1
- 230000003247 decreasing effect Effects 0.000 description 1
- 239000008367 deionised water Substances 0.000 description 1
- 229910021641 deionized water Inorganic materials 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000001035 drying Methods 0.000 description 1
- 238000002474 experimental method Methods 0.000 description 1
- 239000003292 glue Substances 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
- 238000012544 monitoring process Methods 0.000 description 1
- 229910052757 nitrogen Inorganic materials 0.000 description 1
- 230000001590 oxidative effect Effects 0.000 description 1
- 230000035699 permeability Effects 0.000 description 1
- ISWSIDIOOBJBQZ-UHFFFAOYSA-N phenol group Chemical group C1(=CC=CC=C1)O ISWSIDIOOBJBQZ-UHFFFAOYSA-N 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
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- 238000012546 transfer Methods 0.000 description 1
- 238000011041 water permeability test Methods 0.000 description 1
Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/38—Constructions adapted to reduce effects of aerodynamic or other external heating
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- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Mechanical Engineering (AREA)
- Aviation & Aerospace Engineering (AREA)
Abstract
The invention discloses a hypersonic aircraft surface dimension-capable heat protection system, which comprises a multi-through-hole ceramic body, a feedback control conveying system and a ceramic-coated carbon fiber reinforced resin matrix composite body, wherein the ceramic body is positioned at the head or the front edge of an aircraft, the resin matrix composite body is positioned in a large-area of the aircraft, the ceramic body and the front edge are connected, a coolant channel is arranged between the ceramic body and the resin matrix composite body, the conveying system is positioned in the aircraft, and the coolant is conveyed to an inner cavity of the ceramic body through a pump; the invention has the beneficial effects that: the method comprises the steps of (1) intelligently adjusting the flow of a coolant through a pump according to the surface temperature by a feedback control conveying system, so as to realize the dimension and heat protection of the surface of the head or front material of the aircraft, (2) absorbing heat in a large-area of the aircraft by utilizing the pyrolysis of resin and the thermal choking effect, and meanwhile, realizing the dimension of the large-area by utilizing ceramic-coated carbon fibers, (3) realizing the light weight and dimension of the heat protection structure of the aircraft by adopting a combination matching design according to the pneumatic heat flow distribution characteristics of the aircraft.
Description
Technical Field
The invention relates to the technical field of heat protection, in particular to a heat protection system capable of maintaining the shape of the surface of a hypersonic aircraft.
Background
The heat flow of the standing point of the aircraft is inversely proportional to the 1/2 th power of the radius of the head, and the larger the radius of the head is, the smaller the heat flow is, so that the return cabin of the aircraft adopts the shape of a blunt body to reduce the surface heat flow. However, hypersonic aircrafts in the near space generally adopt wave-supporting bodies or lifting body shapes, have sharp front ends, have severe aerodynamic heat environments, are difficult to dimension on the outer surfaces of aircrafts, and have poor landing accuracy. In the future, hypersonic aircrafts fly faster, the temperature of the head and the front edge of the hypersonic aircrafts is higher than 3000 ℃, the ablation problem generated by passive thermal protection is more and more prominent beyond the temperature-resistant limit of all the current heat-resistant materials, and the thermal protection technology that the standing point and the front edge area of the aircrafts can be in a dimensional shape is needed to be solved. The heat flow of the large-area of the aircraft is low, the surface temperature of the material is generally lower than 1500 ℃, and the material is required to be light enough to reduce the heat-proof structure quality so as to improve the effective heat load of the aircraft, and the dimensional shape is required to be realized at 1500 ℃. Therefore, ensuring that the heat-proof structure is light and can maintain the dimension when used in an extreme heat environment is a neck-blocking technology which restricts the development of hypersonic aircrafts in the future, and is also a worldwide problem.
The aerodynamic heat distribution in the service environment of the hypersonic aircraft in the near space has large difference, the standing point and the heat flow in the front edge area are large, and the heat flow in the large-area is small, so that the combined matching design is a key technology for realizing the light weight and dimension of the thermal protection structure of the aircraft according to the aerodynamic heat flow distribution characteristics of the aircraft. For high heat flow areas such as standing points and front edges of the aircraft, a sweating and heat reducing scheme is a potential solution for realizing the functions of dimension and heat protection. Sweat and heat are reduced by absorbing or taking away heat through a coolant (liquid or gas) to prevent or reduce heat transfer to the inside, so that heat protection is realized in a high heat flow environment; the surface temperature is reduced by cooling, so that the material is kept below the heat-resistant temperature, the appearance of the aircraft is not changed, and therefore, the aircraft can realize the shape maintenance effect in the service process.
In the prior art, the patent with the application number of 2018102059811 discloses a hypersonic leading edge heat protection method capable of realizing self-adaptive local activation, sweating and heat reduction, and the patent with the application number of 2021103759447 discloses a layered gradient porous material sweating and heat reduction structure and an aircraft, but the two patents have the following problems in the use process: firstly, the coolant is arranged in the sweating material, no driving pressure exists in the body in the sweating and heat reducing process, and according to Darcy's law, the coolant is difficult to permeate to the surface of the aircraft to realize cooling, so that the surface of the aircraft is ablated; secondly, no sensor is arranged, and intelligent adjustment of the coolant quantity cannot be realized, so that the coolant is wasted, and the effective load of the aircraft is reduced; thirdly, no dimension method of the large area of the aircraft is proposed, and if the method is used for the large area, the coolant consumption is excessive, and the payload of the aircraft is still reduced.
For the low heat flow characteristic of a large area of a hypersonic aircraft, the materials which can be selected for the heat protection materials at present are modified carbon/carbon materials and heat protection tiles. The heat-proof tile is light, but has lower use temperature and can be shaped in a maintenance way under 1300 ℃; the modified carbon/carbon material has high heat resistance temperature up to 2700 ℃, but has higher density, and the mass ratio of the heat protection system is increased by using the high-density material in a large area, so that the effective load of the aircraft is reduced. The ceramic-coated carbon fiber reinforced resin matrix composite material has a higher use temperature than the heat-resistant tile and a lower density than the carbon/carbon material, so that in a lower heat flow area, the selection of the ceramic-coated carbon fiber reinforced resin matrix composite material is a key for realizing the shape maintenance and light weight of the heat protection of a large area of a hypersonic aircraft in the future.
Disclosure of Invention
The invention aims to provide a heat protection system with a dimension-able hypersonic aircraft surface according to the characteristics of large heat flow at the head or the front edge of the hypersonic aircraft and small heat flow in a large-area region, and the heat protection, light weight and dimension-shaped requirements of the hypersonic aircraft in a hypersonic flight state are met by matching the combined design of heat environments.
In order to achieve the above purpose, the present invention adopts the following technical scheme:
a hypersonic aircraft surface dimension-capable thermal protection system comprises a multi-through-hole ceramic body, a feedback control conveying system and a ceramic-coated carbon fiber reinforced resin matrix composite body;
the multi-through hole ceramic body is positioned at the head or the front edge of the aircraft;
the ceramic-coated carbon fiber reinforced resin matrix composite body is positioned in a large-area of the aircraft;
the multi-through-hole ceramic body is connected with the ceramic-coated carbon fiber reinforced resin matrix composite body, and a coolant channel is arranged between the multi-through-hole ceramic body and the ceramic-coated carbon fiber reinforced resin matrix composite body;
the feedback control delivery system includes a pump located within the aircraft interior, a coolant reservoir, the pump delivering coolant from the coolant reservoir along the coolant channel to the interior cavity of the porous ceramic body.
Preferably, the feedback control transport system further comprises a temperature sensor located at the surface of the multi-via ceramic body and a controller located inside the aircraft;
the controller is connected with the pump and the temperature sensor, and the controller intelligently adjusts the flow of the pump according to the relation between the highest temperature fed back by the sensor and the design allowable temperature.
Preferably, the side wall of the ceramic-coated carbon fiber reinforced resin matrix composite body is provided with a groove, and the side wall of the multi-through-hole ceramic body is provided with a boss which is matched with the groove.
Preferably, the porous ceramic is porous alumina, porous mullite, porous silicon carbide, porous silicon nitride or porous zirconia.
Preferably, the ceramic-coated carbon fiber reinforced resin matrix composite is a silicon oxide-coated carbon fiber reinforced resin matrix composite, a silicon carbide-coated carbon fiber reinforced resin matrix composite, a yttrium silicate-coated carbon fiber reinforced resin matrix composite or a zirconium silicate-coated carbon fiber reinforced resin matrix composite.
The invention has the beneficial effects that:
1. the invention adopts the ceramic body with multiple through holes in the high heat flow area such as the standing point and the front edge, and the coolant is transported by the pump to reach the inner wall of the ceramic body with multiple through holes and permeates to the outer surface to realize the heat-resistant shape-maintaining function.
2. The sensor is buried in the ceramic body with multiple through holes and transmits signals to the control system, the control system adjusts the delivery capacity of the pump by judging the temperature signals of the sensor so as to adjust the sweating amount, and a temperature-sweating real-time feedback control system is formed, so that the purposes of intelligent sweating, heat reduction and coolant consumption reduction are achieved.
3. The ceramic-coated carbon fiber reinforced resin matrix composite material is adopted in a large-area of a main body of the aircraft, has the characteristics of light weight and high heat protection, is different from other carbon fiber reinforced materials, is subjected to ceramic treatment on the surface of the fiber, is a novel carbonized material, can be pyrolyzed in the heat protection process, is not ablated on the surface, and ensures that the pneumatic appearance is maintained.
4. The invention comprises a multi-through-hole ceramic body, a ceramic-coated carbon fiber reinforced resin matrix composite body and a feedback control conveying system, wherein the multi-through-hole ceramic body is connected with the feedback control conveying system, and the perspiration (liquid or gas is used as a coolant) quantity is regulated by a pump according to the perceived temperature change, so that the functions of maintaining shape and heat protection of the material of the head and the front edge surface of the aircraft under the extremely high heat flow environment are realized; the ceramic-coated carbon fiber reinforced resin matrix composite is positioned in a large area of the surface of the aircraft, absorbs heat through thermal choking effect of resin pyrolysis and pyrolysis gas, and utilizes the ceramic-coated carbon fiber to isolate oxidizing gas so as to realize the shape maintaining function of the surface of the thermal protection material. The combined matching design scheme can realize the multifunctional integration of light structure, dimension, heat protection and the like according to the characteristics of the pneumatic heat flow distribution of the aircraft.
5. The feedback control transport system intelligently adjusts the flow of the coolant through the pump according to the relation between the surface temperature of the aircraft and the temperature allowed by design, so that the consumption of the coolant can be saved, and the functions of maintaining the shape and protecting heat of the head and the front edge surface materials of the aircraft are realized; the large-area of the aircraft absorbs heat by utilizing the pyrolysis of resin and the thermal choking effect, and meanwhile, the dimension of the large-area is realized by utilizing the ceramic-coated carbon fiber; according to the pneumatic heat flow distribution characteristics of the aircraft, the light weight and dimension of the thermal protection structure of the aircraft are realized by adopting a combined matching design.
Drawings
FIG. 1 is a schematic diagram of the structure of the present invention;
FIG. 2 is a schematic illustration of a schematic test and intelligent cooling test of the present invention;
FIG. 3 is a diagram of the product dimensions for the first principle test;
FIG. 4 is a graph of product temperature for a first principle test;
FIG. 5 is a diagram of the product dimensions for a second principle test;
FIG. 6 is a graph of product temperature for a second principle test;
FIG. 7 is a graph showing heat flux density over time in a smart cooling test;
FIG. 8 is a graph showing the change of product surface temperature with time in a smart cooling test;
FIG. 9 is a schematic of the change in water flow over time in an intelligent cooling test.
The drawings are for illustrative purposes only and are not to be construed as limiting the present patent; for the purpose of better illustrating the embodiments, certain elements of the drawings may be omitted, enlarged or reduced and do not represent the actual product dimensions; it will be appreciated by those skilled in the art that certain well-known structures in the drawings and descriptions thereof may be omitted.
Detailed Description
The invention is further described below with reference to the accompanying drawings.
Examples
As shown in fig. 1, according to the heat protection system with the shape-maintaining hypersonic aircraft surface, according to the characteristics of large heat flow at the head or the front edge of the hypersonic aircraft and small heat flow in a large-area region, a combined structure system matched and adapted to a thermal environment is provided, and the combined structure system comprises a multi-through-hole ceramic body 1, a ceramic-coated carbon fiber reinforced resin matrix composite body 2 and a feedback control conveying system 3;
the multi-through hole ceramic body 1 is positioned at the head or the front edge of the aircraft, namely the head stagnation point area of the aircraft, such as the head of the windward side and the front edge and the like;
the ceramic-coated carbon fiber reinforced resin matrix composite body 2 is positioned in a large area of the aircraft; and the resin pyrolysis and the thermal choking effect are utilized to absorb heat, and meanwhile, the ceramic is used for coating the carbon fiber to realize the dimension of a large area.
According to the pneumatic heat flow distribution characteristics of the aircraft, the combination matching design of the multi-through-hole ceramic body 1 and the ceramic-coated carbon fiber reinforced resin matrix composite body 2 is adopted to realize the light weight and dimensional shape of the thermal protection structure of the aircraft.
The multi-through-hole ceramic body 1 and the ceramic-coated carbon fiber reinforced resin matrix composite body 2 can be connected by high-temperature glue or threads.
The multi-through hole ceramic body 1 is connected with the ceramic-coated carbon fiber reinforced resin matrix composite body 2, and a coolant channel 31 is arranged between the ceramic-coated carbon fiber reinforced resin matrix composite body and the ceramic-coated carbon fiber reinforced resin matrix composite body;
the feedback controlled transport system 3 comprises a controller 35 located inside the aircraft, a pump 34, a coolant tank 32, the pump delivering coolant from the coolant tank 32 along the coolant channel 31 to the inner cavity of the porous ceramic body 1.
The feedback control transportation system 3 further comprises a temperature sensor 36 positioned on the surface of the multi-through-hole ceramic body 1 and a controller 35 positioned inside the aircraft; the temperature sensor 36 is inserted by the inner chamber of the multi-hole ceramic body 1, and the temperature probe at top is located the surface of the multi-hole ceramic body 1, and the controller 35 passes through transmission line 33 and connects temperature sensor 36, and transmission line 33 penetrates in the coolant channel 31 and the contact department of transmission line 33 and coolant channel 31 is provided with sealedly.
The controller 35 is connected with the pump 34 and the temperature sensor 36, and the controller 35 intelligently adjusts the flow of the pump 34 according to the relation between the highest temperature fed back by the temperature sensor 36 and the design permission temperature.
When the temperature born by the multi-hole ceramic body 1 is higher, the pump 34 pumps out the coolant, the coolant is conveyed into the inner cavity of the multi-hole ceramic body 1 along the coolant channel 31, and the coolant is permeated to the outer surface (i.e. the pneumatic heating surface) of the multi-hole ceramic body by the internal and external pressure difference, so that the surface heat reducing function is realized. The coolant may be water, gas (e.g., nitrogen, argon, air, carbon dioxide), or the like. The coolant flow is regulated by the pump 34 to achieve the function of maintaining shape and thermal protection of the material on the aircraft nose and leading edge surfaces.
In the process of sweating and heat reduction, the pump 34 conveys the coolant into the inner cavity of the multi-through hole ceramic body 1, and then permeates the outer surface of the multi-through hole ceramic material from the inner cavity, when the coolant contacts the inner wall of the multi-through hole ceramic body 1, the pressure generated by the pump 34 on the inner wall is larger than the air pressure on the outer side of the multi-through hole ceramic body 1 under the pressure action of the pump 34, so that the coolant can smoothly permeate the outer surface of the multi-through hole ceramic material, and the heat reduction is effectively carried out on the multi-through hole ceramic body 1.
As shown in fig. 1, the thickness of the ceramic wall surface of the multi-through hole is uneven, the thickness of the wall surface of the head part is minimum, and the thickness gradually increases towards the rear part so as to match the characteristics of heat flow distribution of the head part of the aircraft: the greater the heat flow in the stagnation area, the more coolant is supplied.
The pump 34 is connected with the controller 35, the controller 35 is also connected with five temperature sensors 36 (the number of the connected sensors is increased or decreased according to the size of the multi-through-hole ceramic body 1), the temperature sensors 36 are embedded in the multi-through-hole ceramic body 1 in advance, the controller 35 can monitor the surface temperature of the multi-through-hole ceramic body 1 in real time through the transmission line 33, and the delivery capacity of the pump 34 is regulated according to the surface temperature, so that the sweating amount is regulated, a temperature-sweating real-time feedback system is formed, and the purposes of intelligent sweating heat reduction and coolant consumption reduction are achieved.
When the flying speed is increased, the temperature of the multi-through hole ceramic body 1 is increased, the temperature sensor 36 sends a signal to the pump 34 through the controller 35 to increase the flow rate, and the pressure generated by the coolant contacting the inner wall of the multi-through hole ceramic body 1 is correspondingly increased when the flow rate is increased, so that the pressure of the outside of the multi-through hole ceramic body 1 and the air can be overcome, the perspiration rate is regulated, and the reliable heat reduction is carried out; when the flying speed is reduced, the temperature of the multi-through hole ceramic body 1 is reduced, at the moment, the pressure of the air and the outside of the multi-through hole ceramic body 1 is correspondingly reduced, the temperature sensor 36 sends a signal to the pump 34 through the controller 35, so that the flow rate of the coolant is reduced, and the pressure generated by the coolant contacting the inner wall of the multi-through hole ceramic body 1 is correspondingly reduced while the flow rate is reduced, so that the pressure of the air and the outside of the multi-through hole ceramic body 1 can be overcome, and the heat reduction and the coolant consumption are realized; through the foregoing intelligent control, the pump 34 is adjusted so that the temperature of the porous ceramic body 1 is always within a safe temperature range, and the mass flow rate of the coolant is adjusted according to the change of the heat flow in the time domain, so as to reduce the consumption of the coolant, thereby reducing the dead load of the aircraft and enabling the aircraft to fly faster.
The arrangement of the feedback control conveying system enables the heat protection system to be applied to hypersonic aircraft surfaces under any Mach condition.
The side wall of the ceramic-coated carbon fiber reinforced resin matrix composite body 2 is provided with a groove 21, and the side wall of the multi-through-hole ceramic body 1 is provided with a boss 11 which is matched with the groove 21.
The porous ceramic is porous alumina, porous mullite, porous silicon carbide, porous silicon nitride or porous zirconia, and the coolant permeates to the pneumatic heating surface through the porous ceramic, so that the heat reduction can be realized.
The ceramic-coated carbon fiber reinforced resin matrix composite is a silicon oxide-coated carbon fiber reinforced resin matrix composite, a silicon carbide-coated carbon fiber reinforced resin matrix composite, a yttrium silicate-coated carbon fiber reinforced resin matrix composite or a zirconium silicate-coated carbon fiber reinforced resin matrix composite.
The performance parameters of both the multi-through-hole ceramic body 1 and the ceramic-coated carbon fiber reinforced resin matrix composite body 2 are as follows:
| type(s) | Material name | Density g/cm 3 | Porosity% | Compressive strength MPa | Coefficient of water permeability ×10 -14 m 2 |
| Multi-through hole ceramic body | Porous alumina ceramic | 1.42 | 50.27 | 27 | 16.9 |
| Ceramic-coated carbon fiber reinforcement Resin-based composite material body | Yttrium silicate coated carbon fiber reinforcement Phenolic aerogel composite material | 0.540 | 68.76 | 1.29 | - |
When in use, the multi-through hole ceramic body 1 is positioned in the standing point and front edge area of the aircraft, belongs to a pneumatic heating severe part, and is used for conveying coolant to the arc area of the inner wall of the multi-through hole ceramic body through the pump 34, so that the overhigh temperature during supersonic speed flight is avoided; the ceramic-coated carbon fiber reinforced resin matrix composite body 2 coats a large-area of the aircraft, is a main body part of the aircraft, is subjected to fiber surface ceramization treatment, is a novel carbonized material, can be pyrolyzed in a heat-resistant process, and is free from ablation on the surface, so that the aerodynamic shape is ensured to be maintained.
In order to ensure the reliability of the invention in practical use, the invention is subjected to principle examination test, and a test scheme is shown in fig. 2, and mainly comprises a principle prototype and a heating test system. The principle sample machine consists of a principle sample piece and a feedback control conveying system, wherein the principle sample piece consists of a ceramic body and a resin matrix composite material body, and the feedback control conveying system consists of a temperature sensor, a PLC (programmable logic controller) which is arranged on a main board in a computer, a metering pump, a water tank, a water flow pipeline and the like; the heating test system comprises a gas storage bottle, a flowmeter, a gas channel, a flame nozzle, a test bed and the like, wherein the data acquisition system comprises an infrared thermometer, a thermocouple temperature measuring device, a signal transmission line, a camera, an A/D converter, a computer and the like, the range of the low-temperature infrared thermometer is 50-800 ℃, the range of the high-temperature infrared thermometer is 800-3000 ℃, the thermocouple is a K-type thermocouple, and the test bed is used for fixing a principle sample and a Gordon heat flow meter.
Wherein the oxyacetylene flame heat flux density is 6 MW/m 2 The ablation time is 66 s and 120 s respectively, the front end of the structural member is away from an oxyacetylene flame nozzle 10 mm, deionized water is adopted as a coolant of porous alumina ceramic, the flow rate of water in the porous material is set to be 15 g/min through a metering pump, a surface temperature signal is fed back by a built-in thermocouple in the central standing point area of the head of the multi-through-hole ceramic in the ablation process, and an infrared thermometer is used for monitoring the surface temperature at the middle and lower surface positions of the round table part of the ceramic-coated carbon fiber reinforced resin matrix composite material. The specific test steps are as follows:
(1) Processing a sample of a principle to be tested according to requirements to finish the water permeability test;
(2) Placing the sample to be tested in a muffle furnace at 80 ℃ for drying for standby;
(3) Opening a control system switch, placing a principle sample to be tested on a test bed, adjusting the distance between the principle sample and a flame nozzle, communicating with a temperature measuring sensor, and confirming that all temperature measuring devices are normally used;
(4) Calibrating oxyacetylene flame heat flow by using a Gordon heat flow meter, starting a metering water pump after the heat flow value reaches a preset value and the flow is stable, and moving a principle sample to be tested into a flame range to start a principle test;
(5) Ending the test after the test time is reached. When the instrument is closed, firstly flameout is needed, the oxyacetylene control system is closed, and after the examination and arrangement are correct, the water pump, the temperature measuring device, the panel control system and other devices are closed;
(6) Checking whether the appearance of the principle sample piece can keep the dimension after the test.
The results of two of these experiments were extracted as follows:
first test: the product size is shown in figure 3, the size unit is mm, and the temperature change curve is shown in figure 4:
the product is 6 MW/m 2 The principle test of 66 s is completed under the condition of heat flow, the temperature of the surface of the multi-through hole ceramic (No. 1 temperature measuring point) is 173 ℃ in a dynamic balance state, and the coolant generates phase change heat absorption in the porous ceramic body and then flows out from the surface of the ceramic body in the form of water vapor; the temperature of the middle position (No. 2 temperature measuring point) of the ceramic-coated carbon fiber reinforced resin matrix composite material is close to 1000 ℃ and does not reach the temperature point at which the yttrium silicate ceramic coating is melted. The surfaces of the two materials are not ablated before and after ablation, and the ceramic-coated carbon fiber reinforcement body is taken as a framework, so that the structural appearance of the resin-based heat-resistant material is well maintained. The connection position of the active-passive heat-proof structure has good heat matching under the heat flow condition, and meanwhile, the multi-through-hole ceramic and ceramic-coated carbon fiber reinforced resin matrix composite material also has excellent dimensional capacity. The edge position (No. 3 temperature measuring point) of the back of the round table is far away from the heating source, and the temperature rising rate of the measuring point is slow and the temperature is low.
Second test: the product size is shown in fig. 5, the size unit is mm, and the temperature change curve is shown in fig. 6:
the product is 6 MW/m 2 The principle test of 120 s is completed under the condition of heat flow, and the temperature-time curve recorded by the temperature measuring device is shown in fig. 6. The temperature of the porous ceramic standing point is rapidly increased to 401.02 ℃ and then reaches a dynamic balance state. As the height of the semi-ball head in the first test is higher than the height of the ball top in the test, the distance between the ceramic-coated carbon fiber reinforced resin matrix composite material and the flame nozzle in the test is closer, and the temperatures of the No. 2 thermocouple and the No. 3 thermocouple are higher than the temperature of the same measuring point in the first test and are 1118.11 ℃ and 815.28 ℃ respectively. The structure shows no ablation before and after testing, and has a dimensional capability under the heat flow.
In the principle test process, only water vapor overflows from the surface of the porous ceramic, the ceramic-coated carbon fiber reinforced resin matrix composite material is heated by air flow and is heated from the joint of the two materials, and the surface is changed from black to bright yellow until the whole ceramic-coated carbon fiber reinforced resin matrix composite material is in a bright yellow high-temperature state. The shape of the multi-through-hole ceramic material can be maintained by comparing the shapes before and after the principle test.
The test result shows that the structure can adapt to high temperature and can meet the requirements of dimension and heat protection of the ultra-high speed aircraft.
On the premise of carrying out a principle examination test on the invention, in order to save the consumption of the coolant, an intelligent cooling test is carried out, namely, the temperature of the surface is stably controlled by continuously changing the heat flux density of the oxyacetylene flame and regulating and controlling the supply quantity of the coolant through a feedback control conveying system in the test process. The test platform is shown in fig. 2, and the test process is identical to the steps of the principle check test, except that the oxyacetylene flame heat flux density continuously changes along with time in the test process.
The results of one of the tests were extracted as follows:
the test object is a principle sample as shown in figure 3,
the change of oxyacetylene flame heat flux density with time is shown in figure 7,
the change of the temperature of the standing point position of the surface of the sample piece along with time is shown in figure 8,
the change in the regulated water flow over time is shown in fig. 9.
In fig. 8 and 9, it means that even if the heat flow is changed during the test (shown in fig. 7), the feedback control transportation system can stabilize the temperature of the multi-via ceramic standing point position (temperature measuring point) at the design value (shown in fig. 8, the design value is 300 ℃, 400 ℃, 500 ℃, 600 ℃, 700 ℃) for a long time by adjusting the supply amount of the coolant (shown in fig. 9).
According to the test results, the molecular weight of the catalyst is between 0 and 4 MW/m 2 Under hot flow conditions, maintaining the surface temperature at 300 ℃, the coolant flow only needs to be 0.2g/s, and the time 1400s only consumes about 280g of coolant (water); as the design value increases, the coolant flow is less.
According to the test result, under the condition of the change of the heat flux density, the feedback control conveying system can intelligently adjust the mass flow rate of the coolant through the metering pump according to the relation between the temperature signal fed back by the thermocouple in the multi-through-hole ceramic and the design value of the surface temperature, so that the surface temperature of the multi-through-hole ceramic body is kept constant, the surface is not ablated, and the requirements of thermal protection and dimension of the ultra-high-speed aircraft under any Mach condition can be met.
The present embodiment is not limited in any way by the shape, material, structure, etc. of the present invention, and any simple modification, equivalent variation and modification made to the above embodiments according to the technical substance of the present invention are all included in the scope of protection of the technical solution of the present invention.
The following shall be described: the above embodiments are only for illustrating the technical solution of the present invention, and are not limiting; although the invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may be modified or some technical features may be replaced with others, which may not depart from the spirit and scope of the technical solutions of the embodiments of the present invention.
Claims (5)
1. The hypersonic aircraft surface dimension-capable heat protection system is characterized by comprising a multi-through-hole ceramic body, a feedback control conveying system and a ceramic-coated carbon fiber reinforced resin matrix composite body;
the multi-through hole ceramic body is positioned at the head or the front edge of the aircraft;
the ceramic-coated carbon fiber reinforced resin matrix composite body is positioned in a large-area of the aircraft;
the multi-through-hole ceramic body is connected with the ceramic-coated carbon fiber reinforced resin matrix composite body, and a coolant channel is arranged between the multi-through-hole ceramic body and the ceramic-coated carbon fiber reinforced resin matrix composite body;
the feedback control delivery system includes a pump located within the aircraft interior, a coolant reservoir, the pump delivering coolant from the coolant reservoir along the coolant channel to the interior cavity of the porous ceramic body.
2. The hypersonic aircraft surface capable dimensional thermal protection system of claim 1 wherein the feedback control transport system further comprises a temperature sensor located at the surface of the multi-via ceramic body and a controller located inside the aircraft;
the controller is connected with the pump and the temperature sensor, and the controller intelligently adjusts the flow of the pump according to the relation between the highest temperature fed back by the sensor and the design allowable temperature.
3. The hypersonic aircraft surface dimension-capable thermal protection system of claim 1, wherein grooves are formed in the side walls of the ceramic-coated carbon fiber reinforced resin matrix composite body, and bosses matched with the grooves are formed in the side walls of the multi-through-hole ceramic body.
4. The hypersonic aircraft surface capable dimensional thermal protection system of claim 1 wherein the multi-via ceramic is porous alumina, porous mullite, porous silicon carbide, porous silicon nitride or porous zirconia.
5. The hypersonic vehicle surface dimension capable thermal protection system of claim 1 wherein the ceramic coated carbon fiber reinforced resin based composite is a silica coated carbon fiber reinforced resin based composite, a silicon carbide coated carbon fiber reinforced resin based composite, a yttrium silicate coated carbon fiber reinforced resin based composite, or a zirconium silicate coated carbon fiber reinforced resin based composite.
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