CN108657442B - Aircraft and thermal protection system - Google Patents
Aircraft and thermal protection system Download PDFInfo
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- CN108657442B CN108657442B CN201810468944.XA CN201810468944A CN108657442B CN 108657442 B CN108657442 B CN 108657442B CN 201810468944 A CN201810468944 A CN 201810468944A CN 108657442 B CN108657442 B CN 108657442B
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- 239000012530 fluid Substances 0.000 claims abstract description 17
- CURLTUGMZLYLDI-UHFFFAOYSA-N Carbon dioxide Chemical compound O=C=O CURLTUGMZLYLDI-UHFFFAOYSA-N 0.000 claims description 64
- 229910002092 carbon dioxide Inorganic materials 0.000 claims description 32
- 238000005192 partition Methods 0.000 claims description 11
- 239000001569 carbon dioxide Substances 0.000 claims description 2
- 230000000694 effects Effects 0.000 abstract description 5
- 238000001816 cooling Methods 0.000 description 8
- 238000010586 diagram Methods 0.000 description 7
- 239000011229 interlayer Substances 0.000 description 7
- 239000002826 coolant Substances 0.000 description 4
- 238000010438 heat treatment Methods 0.000 description 4
- 239000000446 fuel Substances 0.000 description 3
- 239000007788 liquid Substances 0.000 description 3
- 230000004048 modification Effects 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 238000002679 ablation Methods 0.000 description 2
- 238000010521 absorption reaction Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 2
- 230000033228 biological regulation Effects 0.000 description 2
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 238000009413 insulation Methods 0.000 description 2
- 239000010410 layer Substances 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 230000005068 transpiration Effects 0.000 description 2
- 230000009471 action Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 239000011248 coating agent Substances 0.000 description 1
- 238000000576 coating method Methods 0.000 description 1
- 239000012809 cooling fluid Substances 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 238000009792 diffusion process Methods 0.000 description 1
- 230000017525 heat dissipation Effects 0.000 description 1
- 239000007791 liquid phase Substances 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 230000008092 positive effect Effects 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 230000005855 radiation Effects 0.000 description 1
- 230000001105 regulatory effect Effects 0.000 description 1
- 238000007789 sealing Methods 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
- 238000003466 welding Methods 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D13/00—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft
- B64D13/06—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft the air being conditioned
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D13/00—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft
- B64D13/06—Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space, or structural parts of the aircraft the air being conditioned
- B64D2013/0603—Environmental Control Systems
- B64D2013/0607—Environmental Control Systems providing hot air or liquid for deicing aircraft parts, e.g. aerodynamic surfaces or windows
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- Health & Medical Sciences (AREA)
- General Health & Medical Sciences (AREA)
- Pulmonology (AREA)
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The utility model provides an aircraft and thermal protection system belongs to aviation aircraft technical field. The thermal protection system is applied to an aircraft, and the aircraft comprises a nose, a fuselage and wings. The thermal protection system comprises a first loop, a second loop and a power supply device. The first loop comprises a first heat source pipeline and a first cold source pipeline, the first heat source pipeline is arranged on the front edge of the machine head or the front edge of the wing, and the first cold source pipeline is arranged on the lee side of the machine body. The second loop comprises a second heat source pipeline and a second cold source pipeline, the second heat source pipeline is arranged on the windward side of the machine body, and the second cold source pipeline is arranged on the leeward side of the machine body. The power supply device is connected with the first circuit and the second circuit and is used for driving the fluid medium to circularly flow in the first circuit and the second circuit. The heat protection carries out heat exchange between the high-temperature area and the low-temperature area of the aircraft in a closed circulation mode through the fluid medium, so that the effect of reducing the temperature of the high-temperature area of the aircraft is achieved.
Description
Technical Field
The utility model belongs to the technical field of the aviation aircraft, particularly, relate to a thermal protection system and have this thermal protection system's aircraft.
Background
When the aircraft flies at a high supersonic speed in the adjacent space, local areas of the aircraft are subjected to severe aerodynamic heating due to shock wave compression and severe friction of airflow with the outer surface of the aircraft. In order to ensure that the aircraft fuselage and its internal environment operate normally within the permitted temperature range, effective structural thermal protection designs are required. Typical hypersonic aircraft thermal protection systems can be divided into three categories: passive thermal protection, semi-passive thermal protection, and active thermal protection.
The basic principle of passive thermal protection is to radiate heat away from the surface of the aircraft fuselage structure or to store it therein without using a cooling fluid medium to absorb the heat, and typical solutions thereof are: heat sink structures, thermal structures, and thermal insulation structures. The heat sink structure relies on the heat capacity of self to absorb heat, and after the heat of absorption reached a definite value, the hot safeguard function of this structure will lose efficacy, if will increase the heat absorption volume then need increase the volume and the weight of heat sink structure, the heat protection efficiency of this structure is lower, is not applicable to the hypersonic aircraft that works for a long time near the space. The thermal structure and the heat insulation structure mainly rely on radiation heat dissipation, have special requirements on factors such as the smooth finish of the outer surface of an aircraft fuselage, a coating and the like, and have poor applicability to the flight environment.
The basic principle of semi-passive thermal protection is to take away heat by using working media or fluid media, and the typical scheme is as follows: ablation structures and heat pipe structures. The ablation structure takes away incident heat by continuously burning and gasifying the outer surface material of the aircraft body, so that the appearance of the aircraft body is changed, the aerodynamic characteristics of the aircraft are changed, and meanwhile, the thermal protection structure is not suitable for the hypersonic aircraft working in the near space for a long time. The driving force for the heat exchange medium of the heat pipe structure to flow comes from the capillary action, the internal flow and heat transfer mechanism of the heat pipe structure are very complex, the problems of complex gas-liquid coupling flow and gas-liquid phase change heat transfer are included, and meanwhile, the thermal protection effect of the heat pipe structure is seriously influenced by the contact thermal resistance between the wall surface of the heat pipe and the wall surface of the aircraft fuselage.
The basic principle of the active thermal protection structure is that all or most of heat is taken away by working medium or cooling flow, and the rest is reflected, and the typical scheme is as follows: liquid film cooling, transpiration cooling, and convection cooling. Liquid film cooling and transpiration cooling require constant consumption of cooling medium, and this thermal protection scheme is not suitable for hypersonic aircraft operating in close proximity for long periods of time. Convection cooling requires an external driving force to drive the coolant to flow in the channels or pipes, and this protection scheme also requires constant consumption of the coolant or a solution to the problem of heat release after heating of the coolant.
It is to be noted that the information invented in the background section above is only for enhancement of understanding of the background of the present disclosure, and thus may include information that does not constitute prior art known to those of ordinary skill in the art.
Disclosure of Invention
The present disclosure provides an aircraft capable of regulating temperature and a thermal protection system.
According to one aspect of the present disclosure, there is provided a thermal protection system for an aircraft, the aircraft including a nose, a fuselage and a wing, the thermal protection system comprising:
the first loop comprises a first heat source pipeline and a first cold source pipeline, the first heat source pipeline is arranged at the front edge of the machine head or the wing, and the first cold source pipeline is arranged on the lee side of the machine body;
the second loop comprises a second heat source pipeline and a second cold source pipeline, the second heat source pipeline is arranged on the windward side of the machine body, and the second cold source pipeline is arranged on the leeward side of the machine body;
and the power supply device is connected with the first circuit and the second circuit and is used for driving the fluid medium to circularly flow in the first circuit and the second circuit.
In an exemplary embodiment of the present disclosure, the power supply apparatus includes a compressor, a turbine, and a connecting shaft connecting the compressor and the turbine, the turbine having an inlet and an outlet, the second heat source pipe being connected to the inlet of the turbine, and the second heat sink pipe being connected to the outlet of the turbine.
In an exemplary embodiment of the present disclosure, the first circuit further includes a first connection pipe connected to the first cold source pipe and the power supply device, the second circuit further includes a second connection pipe connected to the second cold source pipe and the power supply device, and the first connection pipe and the second connection pipe are separately disposed or disposed on the same connection pipe.
In an exemplary embodiment of the present disclosure, the first connection line and the second connection line are provided with a cool source and a valve.
In an exemplary embodiment of the present disclosure, the thermal protection system further includes a bypass line, one end of the bypass line is connected to the first cold source line and the second cold source line, the other end of the bypass line is connected to the power supply device, and the bypass line is provided with a valve.
In an exemplary embodiment of the present disclosure, the second heat source pipeline and the second cold source pipeline are a passage between an outer wall surface and an inner wall surface of the body, and a side of the outer wall surface facing the passage is provided with a rib.
In an exemplary embodiment of the present disclosure, the first heat source pipeline is a passage between an outer wall surface and an inner wall surface of the aircraft nose or the aircraft wing, the first cold source pipeline is a passage between an outer wall surface and an inner wall surface of the aircraft body, partition plates are respectively arranged in the first heat source pipeline and the first cold source pipeline, the partition plates respectively divide the first heat source pipeline and the first cold source pipeline into two chambers, each partition plate is provided with an impact hole, and the impact holes are communicated with the chambers.
In an exemplary embodiment of the present disclosure, the first and second connection pipes are plate heat exchanger structures.
In an exemplary embodiment of the present disclosure, the fluid medium is supercritical carbon dioxide.
According to an aspect of the present disclosure, there is provided an aircraft comprising the thermal protection system of any one of the above.
According to the technical scheme, the aircraft and the thermal protection system have the advantages and positive effects that:
the heat protection system adopts a closed circulation structure, and heat exchange is carried out between the high-temperature area and the low-temperature area of the aircraft in a closed circulation mode through the fluid medium, so that the effect of reducing the temperature of the high-temperature area of the aircraft is achieved.
Drawings
Various objects, features and advantages of the present disclosure will become more apparent from the following detailed description of preferred embodiments thereof, when considered in conjunction with the accompanying drawings. The drawings are merely exemplary illustrations of the disclosure and are not necessarily drawn to scale. In the drawings, like reference characters designate the same or similar parts throughout the different views. Wherein:
fig. 1 shows a schematic structural diagram of a thermal protection system according to an embodiment of the disclosure.
FIG. 2 illustrates supercritical CO in accordance with an embodiment of the disclosure2And (3) schematic diagram of heat exchange in the second heat source pipeline.
FIG. 3 illustrates supercritical CO in accordance with an embodiment of the disclosure2And the schematic diagram of heat exchange in the second cold source pipeline.
Fig. 4 is a schematic diagram illustrating a heat exchange enhancement structure of a second heat source pipeline and a second cold source pipeline according to an embodiment of the disclosure.
FIG. 5 illustrates supercritical CO in accordance with an embodiment of the disclosure2Schematic diagram of heat exchange in the first heat source pipeline.
Fig. 6 shows a schematic view of an impact heat exchange structure of a first heat source pipeline and a first cold source pipeline according to an embodiment of the disclosure.
FIG. 7 illustrates supercritical CO in accordance with an embodiment of the disclosure2Schematic diagram of heat exchange through plate heat exchanger.
Fig. 8 shows a schematic structural view of the plate heat exchanger of fig. 7.
Wherein the reference numerals are as follows:
100. a thermal protection system;
11. a first circuit;
111. a first heat source pipeline;
1111. a partition plate;
11111. an impingement hole;
112. a first cold source pipeline;
113. a first connecting line;
12. a second loop;
121. a second heat source pipeline;
122. a second cold source pipeline;
123. a second connecting line;
1231. a cold source;
1232. a valve;
124. a bypass line;
13. a power supply device;
131. a compressor;
132. a turbine;
1321. an inlet;
1322. an outlet;
133. a connecting shaft;
200. an aircraft;
21. a machine head;
22. a body;
221. the leeward side of the fuselage;
222. the windward side of the fuselage;
223. an outer wall surface;
2231. a rib;
224. an inner wall surface;
23. an airfoil;
231. the leading edge of the wing;
300. a fluid medium.
Detailed Description
Exemplary embodiments that embody features and advantages of the present disclosure are described in detail below in the specification. It is to be understood that the disclosure is capable of various modifications in various embodiments without departing from the scope of the disclosure, and that the description and drawings are to be regarded as illustrative in nature, and not as restrictive.
In the following description of various exemplary embodiments of the disclosure, reference is made to the accompanying drawings, which form a part hereof, and in which are shown by way of illustration various exemplary structures, systems, and steps in which aspects of the disclosure may be practiced. It is to be understood that other specific arrangements of parts, structures, example devices, systems, and steps may be utilized, and structural and functional modifications may be made without departing from the scope of the present disclosure. Nothing in this specification should be construed as requiring a specific three dimensional orientation of structures in order to fall within the scope of this disclosure.
When the Mach number of the space shuttle is 25, the nose, the front edge of the wing and the lower surface of the wing are all high-temperature regions, and the highest temperature can reach 2000K; the temperature in the other zones is relatively low, with a minimum temperature of about 470K. To ensure that the aircraft fuselage and its internal environment operate properly within the allowable temperature range, the present disclosure provides the following embodiments:
example one
The present embodiment provides a thermal protection system 100. Fig. 1 is a schematic structural diagram of a thermal protection system 100 according to an embodiment of the disclosure.
As shown in fig. 1, in the present embodiment, heat shield system 100 is applied to aircraft 200, and aircraft 200 mainly includes a nose 21, a fuselage 22, and wings 23. The thermal protection system 100 basically includes a first circuit 11, a second circuit 12 and a power supply 13. The first circuit 11 may include a first heat source pipe 111 and a first cold source pipe 112, the first heat source pipe 111 may be disposed at the nose 21 or the wing leading edge 231, and the first cold source pipe 112 may be disposed at the fuselage lee side 221. The second circuit 12 may include a second heat source pipe 121 and a second cool source pipe 122, the second heat source pipe 121 may be disposed on the windward side 222 of the body, and the second cool source pipe 122 may be disposed on the leeward side 221 of the body. A power supply 13 is connected to the first and second circuits 11, 12 for driving a circulating flow of the fluid medium 300 within the first and second circuits 11, 12. The thermal protection system 100 may include one or more first circuits 11. The first circuit 11 may be provided with a plurality of sections of first heat source pipes 111 and first cold source pipes 112, and the first heat source pipes 111 and the first cold source pipes 112 are alternately arranged.
The heat protection system adopts a closed circulation structure, and heat exchange is carried out between the high-temperature area and the low-temperature area of the aircraft in a closed circulation mode through the fluid medium, so that the effect of reducing the temperature of the high-temperature area of the aircraft is achieved.
As shown in fig. 1, in the present embodiment, the power supply device 13 may include a compressor 131, a turbine 132, and a connecting shaft 133 connecting the compressor 131 and the turbine 132, the turbine 132 having an inlet 1321 and an outlet 1322, the second heat source pipe 121 being connected to the inlet 1321 of the turbine 132, and the second cool source pipe 122 being connected to the outlet 1322 of the turbine 132.
Further, in the present embodiment, the first circuit 11 may further include a first connection pipe 113, and the first connection pipe 113 is connected to the first cool source pipe 112 and the power supplier 13. The second circuit 12 may further include a second connection line 123, and the second connection line 123 is connected to the second cold source line 122 and the power supplier 13, and particularly, the second connection line 123 is connected to the second cold source line 122 and the compressor 131. The first connection line 113 and the second connection line 123 may be separately disposed or disposed at the same connection line, and a cold source 1231 and a valve 1232 may be disposed at the first connection line 113 and the second connection line 123. As shown in fig. 1, the first connecting pipe 113 and the second connecting pipe 123 are the same connecting pipe in this embodiment.
Further, in this embodiment, the thermal protection system 100 may further include a bypass line 124, one end of the bypass line 124 is connected to the first heat sink line 112 and the second heat sink line 122, the other end of the bypass line 124 is connected to the power supply 13, and the bypass line 124 is provided with a valve 1232.
Further, in the present embodiment, the fluid medium 300 may be supercritical CO2, and the supercritical CO2 is used as the fluid medium, which has characteristics of large fluid density, small viscosity, large heat capacity, and the like, and can significantly reduce the occupied space of the thermal protection system 100. The power supply 13 can drive the supercritical CO2 to flow in the above-mentioned pipeline, and the specific flow path is as follows:
the supercritical CO2 is divided into two parts after being pressurized by the compressor 131, one part of the supercritical CO2 sequentially flows through the second heat source pipeline 121, the turbine 132, the second cold source pipeline 122, the second connecting pipeline 123 and the compressor 131 to form a second loop 12, the second loop 12 is a brayton cycle engine system, and the other part of the supercritical CO2 sequentially flows through the first heat source pipeline, the first cold source pipeline 112, the first connecting pipeline 113 and the compressor 131 to form a first loop 11. In the embodiment, the brayton cycle engine system provides the pressurized supercritical CO2, so as to drive the supercritical CO2 to circularly flow between the heat source pipeline (high temperature region) and the cold source pipeline (low temperature region and normal temperature fuel), and realize the heat exchange between the high temperature region and the low temperature region. The thermal protection system can adopt a multi-flow path design, one flow path of the thermal protection system forms a Brayton cycle engine system, and the other flow paths are used for temperature regulation of the aircraft. I.e. the thermal protection system may comprise a plurality of first circuits 11.
Further, in a brayton cycle engine system, an overall structural layout of a stand-alone centrifugal impeller and a stand-alone centripetal turbine may be employed. The engine has the advantages of simple structure, small volume, light weight, wide working range and the like. Centrifugal impellers for pressurized supercritical CO2 may have a rim diameter of about 50mm and a flow capacity of up to about 3.5 kg/s. The parameters for the components of the brayton cycle engine system may be selected as follows: the supercritical CO2 parameters at the inlet of the compressor 131 can be respectively 8 MPa-10 MPa and 310K-400K, the parameters are close to the critical state (the state parameters of the CO2 critical point are 7.4MPa and 304.4K), and the density of CO2 under the parameter state is higher, so that the compression work of the compressor 131 can be reduced. The pressure ratio of the compressor 131 is 1.5-3.0, the pressure ratio is small, the temperature of the turbine can be 600-923K, and the temperature is low, so that the pneumatic design and structural design difficulty of the compressor 131 and the turbine are reduced, and the design difficulty of the high-pressure resistant equipment of the thermal protection system 100 is reduced. After the supercritical CO2 performs work through turbo expansion, the pressure and the temperature are both reduced, but the temperature is still higher, and the supercritical CO2 needs to enter the second cold source pipeline 122 and the third connecting pipeline for cooling and then can be discharged into the inlet of the compressor 131. The full power produced by the work of turboexpansion drives the compressor 131 via the connecting shaft 133 to compress the supercritical CO 2.
Example two
The present embodiment provides a thermal protection system 100. The thermal protection system 100 is of an integrated design with the aircraft 200.
As shown in fig. 2, 3 and 4, in the present embodiment, the second heat source pipeline 121 and the second heat sink pipeline 122 are a passage between an outer wall surface 223 of the body 22 and an inner wall surface 224 of the body 22, and a rib 2231 is provided on a side of the body outer wall surface 223 facing the passage. In order to ensure the thermal efficiency of the brayton cycle engine, the flow loss of the supercritical CO2 needs to be limited, and the supercritical CO2 can adopt a rib turbulent flow enhanced heat exchange mode when the heat exchange is performed between the second heat source pipeline 121 and the second cold source pipeline 122, which can comprehensively consider the problems of enhanced heat exchange and flow loss. The heat exchange structure adopts an integrated design, thereby being beneficial to saving the occupied space of the heat exchange structure and reducing the weight of the heat exchange structure. As can be seen from fig. 4, the heat exchange structure adopts a double-wall structure, and the supercritical CO2 performs flow heat exchange in an interlayer between the double walls. The double-walled structure may be sized as: the value range of the interlayer thickness s is 1-2.5 mm, the rib thickness t is 0.3-0.5 s, the rib width p is 0.6-1.0 s, and the rib spacing n is 3-6 s.
As shown in fig. 5 and 6, in the present embodiment, the first heat source pipeline 111 is a passage between an outer wall surface 223 and an inner wall surface 224 of the nose 21 or the wing 23, the first cold source pipeline 112 is a passage between the outer wall surface 223 and the inner wall surface 224 of the body 22, a partition 1111 is disposed in each of the first heat source pipeline 111 and the first cold source pipeline 112, the partition 1111 divides the first heat source pipeline 111 and the first cold source pipeline 112 into two chambers, the partition 1111 has an impact hole 11111, and the impact hole 11111 communicates the two chambers. In the first circuit 11, the supercritical CO2 allows a large pressure loss (the driving pressure difference provided can be as high as more than 4 MPa) when flowing through the first heat source pipeline 111 and the first heat sink pipeline 112, so that the supercritical CO2 can adopt an impact heat exchange method with a very strong convection heat exchange capability when performing heat exchange between the first heat source pipeline 111 and the first heat sink pipeline 112. The heat transfer mode has extremely strong heat transfer convection capacity and correspondingly needs to pay larger pressure loss cost. As can be seen from fig. 6, the impingement heat exchange structure has a double sandwich structure, and the size of the double sandwich structure may be: the value range of the diameter d of the impact holes 11111 is 0.3-1.5 mm, the thickness of the interlayer 1 is 1.5-3 d, the thickness of the interlayer 2 is 1-2 d, the thickness of the partition plate 1111 is 0.5-1.5 d, the transverse distance of the impact holes 11111 is 3-10 d, and the longitudinal distance of the impact holes 11111 is 3-10 d.
Further, in the present embodiment, since the supercritical CO2 of the first circuit 11 allows a large pressure loss, the first heat source pipeline 111 and the first heat sink pipeline 112 may adopt an impact heat exchange manner with a very strong convective heat transfer capability, and therefore, the first heat source pipeline 111 in the first circuit 11 is preferably disposed at a position where the pneumatic heating environment of the aircraft 200 is very harsh. The brayton cycle engine system needs to comprehensively consider the cycle efficiency of the engine and limit the pressure loss of the supercritical CO2 in the flow heat exchange process, and then the second heat source pipeline 121 and the second cold source pipeline 122 can adopt a rib turbulent flow enhanced heat exchange mode, and the convective heat exchange capacity of the heat exchange mode is weaker than that of an impact heat exchange mode, so that the second loop 12, namely the second heat source pipeline 121 of the brayton cycle engine system, is preferably arranged at a position where the pneumatic heating environment of the aircraft 200 is severe.
As shown in fig. 7 and 8, in the present embodiment, when the supercritical CO2 exchanges heat between the first connection pipe 113 and the second connection pipe 123, the flow loss of the supercritical CO2 needs to be limited to ensure the thermal efficiency of the brayton cycle engine, and the heat exchange structure of the first connection pipe 113 and the second connection pipe 123 may be a plate heat exchanger. The plate heat exchanger is provided with a rib turbulent flow enhanced heat exchange structure between the laminates, and the laminates can be combined together through diffusion welding. The heat exchange structure has the advantages of high energy transfer density, small volume, light weight, good sealing property and the like. As can be seen from fig. 8, the plate heat exchanger may have the following structural dimensions: the value range of the interlayer thickness s5 is 1.5-3 mm, the rib thickness t is 0.15s 5-0.25 s5, the rib width p is 0.3s 5-0.5 s5, and the rib spacing n is 3s 5-6 s 5.
Further, in this embodiment, a cold source 1231 (not shown in fig. 8) may be disposed inside the plate heat exchanger, and the cold source 1231 may be an onboard normal temperature fuel as an auxiliary cold source. After the supercritical CO2 releases heat through the first cold source pipeline 112 and the second cold source pipeline 122, if the temperature of the supercritical CO2 is still higher than the required value of the inlet of the compressor 131, the heat needs to be further released to the cold source 1231, so that the temperature reaches the required value of the inlet of the compressor 131. If the temperature of the supercritical CO2 meets the requirement of the inlet of the compressor 131, the valve 1232 of the pipeline where the heat sink 1231 is located (e.g., the first connecting pipeline 113 and the second connecting pipeline 123) is closed, and the valve 1232 of the bypass pipeline 124 is opened, so that the supercritical CO2 enters the inlet of the compressor 131 from the bypass pipeline 124.
The thermal protection system provided by the embodiment has the following advantages:
1. the heat protection system adopts a closed circulation structure, and heat exchange is carried out between the high-temperature area and the low-temperature area of the aircraft in a closed circulation mode through the fluid medium, so that the effect of reducing the temperature of the high-temperature area of the aircraft is achieved.
2. The fluid medium can be selected from supercritical CO2, and the supercritical CO2 has the characteristics of high fluid density, low viscosity, high heat capacity and the like, so that the occupied space of a thermal protection system can be obviously reduced.
3. A multiple flow path design can be adopted, one flow path is formed into a Brayton cycle engine system, and the other flow paths are used for temperature regulation of the aircraft. The Brayton cycle engine can adopt the overall structural scheme of a single-stage centrifugal impeller and a single-stage centripetal turbine, and the structure has the advantages of simple structure, small volume, light weight, wide working range and the like.
4. On-board fuel may be used as an auxiliary heat sink to help cool the supercritical CO2 to demand conditions to ensure thermal efficiency of the brayton cycle engine.
5. The thermal protection of the aircraft can be designed integrally, and the aircraft can comprise a rib turbulent flow enhanced heat exchange structure with a double-layer wall and a single interlayer and an impact enhanced heat exchange structure with a three-layer wall and a double interlayer. The integrated structure is beneficial to saving the occupied space of the heat exchange structure and reducing the weight of the heat exchange structure while effectively exchanging heat.
It should be noted herein that the thermal protection system illustrated in the drawings and described in the present specification is but one example of a wide variety of thermal protection systems that can employ the principles of the present disclosure. It should be clearly understood that the principles of this disclosure are in no way limited to any of the details of the thermal protection system or any of the components of the thermal protection system shown in the drawings or described in this specification.
EXAMPLE III
The present embodiments provide an aircraft. The aircraft is provided with the thermal protection system, and the position relationship between the aircraft and the thermal protection system can be seen from the above embodiment, and is not described in detail here.
Exemplary embodiments of an aircraft and thermal protection system as set forth in the present disclosure are described and/or illustrated in detail above. Embodiments of the disclosure are not limited to the specific embodiments described herein, but rather, components and/or steps of each embodiment may be utilized independently and separately from other components and/or steps described herein. Each component and/or step of one embodiment can also be used in combination with other components and/or steps of other embodiments. When introducing elements/components/etc. described and/or illustrated herein, the articles "a," "an," and "the" are intended to mean that there are one or more of the elements/components/etc. The terms "comprising" and "having" are intended to be inclusive and mean that there may be additional elements/components/etc. other than the listed elements/components/etc. Furthermore, the terms "first" and "second" and the like in the claims and the description are used merely as labels, and are not numerical limitations of their objects.
While the aircraft and thermal protection system presented in this disclosure have been described in terms of various specific embodiments, those skilled in the art will recognize that the practice of this disclosure can be practiced with modification within the spirit and scope of the claims.
Claims (9)
1. A thermal protection system for an aircraft comprising a nose, a fuselage and a wing, characterized in that it comprises:
the first loop comprises a first heat source pipeline and a first cold source pipeline, the first heat source pipeline is arranged at the front edge of the machine head or the wing, and the first cold source pipeline is arranged on the lee side of the machine body;
the second loop comprises a second heat source pipeline and a second cold source pipeline, the second heat source pipeline is arranged on the windward side of the machine body, and the second cold source pipeline is arranged on the leeward side of the machine body;
the power supply device is connected with the first circuit and the second circuit and is used for driving the fluid medium to circularly flow in the first circuit and the second circuit;
the first loop further comprises a first connecting pipeline, the first connecting pipeline is connected with the first cold source pipeline and the power supply device, the second loop further comprises a second connecting pipeline, the second connecting pipeline is connected with the second cold source pipeline and the power supply device, and the first connecting pipeline and the second connecting pipeline are separately arranged or arranged on the same connecting pipeline.
2. The thermal protection system of claim 1, wherein said power supply includes a compressor, a turbine and a connecting shaft connecting said compressor and said turbine, said turbine having an inlet and an outlet, said second heat source line being connected to said turbine inlet, said second heat sink line being connected to said turbine outlet.
3. The thermal protection system of claim 1, wherein said first connecting line and said second connecting line are provided with a cold source and a valve.
4. The thermal protection system of claim 1, further comprising a bypass line, wherein one end of the bypass line is connected to the first cold source line and the second cold source line, the other end of the bypass line is connected to the power supply device, and the bypass line is provided with a valve.
5. The thermal protection system according to any one of claims 1 to 4, wherein the second heat source pipeline and the second heat sink pipeline are a passage between an outer wall surface and an inner wall surface of the body, and a side of the outer wall surface facing the passage is provided with a rib.
6. The thermal protection system according to any one of claims 1 to 4, wherein the first heat source pipeline is a passage between an outer wall surface and an inner wall surface of the aircraft nose or the aircraft wing, the first cold source pipeline is a passage between an outer wall surface and an inner wall surface of the aircraft body, partition plates are arranged in the first heat source pipeline and the first cold source pipeline, the partition plates divide the first heat source pipeline and the first cold source pipeline into two chambers respectively, and each partition plate is provided with an impact hole which is communicated with the chambers.
7. The thermal protection system according to claim 1 or 3, wherein said first and second connection lines are plate heat exchanger structures.
8. The thermal protection system of claim 1, wherein said fluid medium is supercritical carbon dioxide.
9. An aircraft comprising a thermal protection system according to any one of claims 1 to 8.
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| CN201810468944.XA CN108657442B (en) | 2018-05-16 | 2018-05-16 | Aircraft and thermal protection system |
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| CN109774981B (en) * | 2019-01-28 | 2020-07-03 | 清华大学 | Thermal protection system of auxiliary rocket of hypersonic vehicle |
| US11465766B2 (en) * | 2019-06-28 | 2022-10-11 | The Boeing Company | Systems and methods for cooling and generating power on high speed flight vehicles |
| EP4001599B8 (en) * | 2020-11-23 | 2023-04-12 | The Boeing Company | Methods and systems for generating power and thermal management having combined cycle architecture |
| CN113335534B (en) * | 2021-06-10 | 2024-03-12 | 西安交通大学 | Phase-change heat-absorbing type air charging and discharging device of passive air charging and discharging system of near space vehicle |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1484599A (en) * | 2000-11-30 | 2004-03-24 | ����ά��ŵķ����-�����ؿع�����˾ | Cooling apparatus |
| CN103562067A (en) * | 2011-05-30 | 2014-02-05 | 涡轮梅坎公司 | Air conditioning system for an aircraft passenger compartment |
| CN106516072A (en) * | 2016-11-10 | 2017-03-22 | 清华大学 | Thermal protection structure for leading edge of hypersonic vehicle |
| CN106640242A (en) * | 2016-09-19 | 2017-05-10 | 清华大学 | Hypersonic flight vehicle engine heat recovery power generation system and control method thereof |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10011361B2 (en) * | 2015-11-16 | 2018-07-03 | Textron Innovations Inc. | Aircraft environmental control system providing improved performance and backup temperature control |
-
2018
- 2018-05-16 CN CN201810468944.XA patent/CN108657442B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1484599A (en) * | 2000-11-30 | 2004-03-24 | ����ά��ŵķ����-�����ؿع�����˾ | Cooling apparatus |
| CN103562067A (en) * | 2011-05-30 | 2014-02-05 | 涡轮梅坎公司 | Air conditioning system for an aircraft passenger compartment |
| CN106640242A (en) * | 2016-09-19 | 2017-05-10 | 清华大学 | Hypersonic flight vehicle engine heat recovery power generation system and control method thereof |
| CN106516072A (en) * | 2016-11-10 | 2017-03-22 | 清华大学 | Thermal protection structure for leading edge of hypersonic vehicle |
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