CN117465653A - Shock rod based on supercritical fluid coolant, heat-proof device and method - Google Patents
Shock rod based on supercritical fluid coolant, heat-proof device and method Download PDFInfo
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- CN117465653A CN117465653A CN202310839661.2A CN202310839661A CN117465653A CN 117465653 A CN117465653 A CN 117465653A CN 202310839661 A CN202310839661 A CN 202310839661A CN 117465653 A CN117465653 A CN 117465653A
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- 230000035939 shock Effects 0.000 title claims abstract description 83
- 239000012530 fluid Substances 0.000 title claims abstract description 68
- 239000002826 coolant Substances 0.000 title claims abstract description 53
- 238000000034 method Methods 0.000 title claims abstract description 23
- 238000001816 cooling Methods 0.000 claims abstract description 86
- 230000008602 contraction Effects 0.000 claims abstract description 14
- 230000008569 process Effects 0.000 claims description 13
- 230000007704 transition Effects 0.000 claims description 4
- 238000004891 communication Methods 0.000 claims description 2
- 230000000694 effects Effects 0.000 abstract description 24
- 230000001965 increasing effect Effects 0.000 description 10
- 230000005284 excitation Effects 0.000 description 6
- 238000010521 absorption reaction Methods 0.000 description 3
- 230000009467 reduction Effects 0.000 description 3
- 238000002679 ablation Methods 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 230000010355 oscillation Effects 0.000 description 2
- 238000000926 separation method Methods 0.000 description 2
- 239000007787 solid Substances 0.000 description 2
- 230000002411 adverse Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008094 contradictory effect Effects 0.000 description 1
- 238000009413 insulation Methods 0.000 description 1
- 230000003993 interaction Effects 0.000 description 1
- 238000002955 isolation Methods 0.000 description 1
- 230000002265 prevention Effects 0.000 description 1
- 238000010992 reflux Methods 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 238000005507 spraying 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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- 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
- B64C1/0009—Aerodynamic aspects
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C30/00—Supersonic type aircraft
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- Aviation & Aerospace Engineering (AREA)
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Abstract
The invention belongs to the technical field of heat protection of high-speed aircrafts, and particularly relates to a shock rod based on a supercritical fluid coolant, a heat protection device and a method, wherein the shock rod comprises a rod body and a pneumatic disc; one end of the rod body is connected with the pneumatic disc; the pneumatic disc is provided with a cooling channel I along the axial direction, the pneumatic disc is provided with a cooling channel II along the radial direction, the cooling channel II comprises a contraction section, a roar and an expansion section which are sequentially arranged from the axis to the outside of the disc, and the other end of the contraction section is communicated with the cooling channel I. The invention combines the structure of the shock wave rod with the characteristic of the supercritical fluid used, so that the supercritical fluid is in the optimal state at a specific position, and on the basis of reducing the quality of a supercritical coolant storage tank, the invention can not only improve the cooling effect of the shock wave rod, ensure the structural strength of the shock wave rod, but also improve the thermal environment of the head part and the body part of the aircraft, thereby achieving the thermal protection effect. But also can reduce the flight resistance and improve the aerodynamic performance of the aircraft.
Description
Technical Field
The invention belongs to the technical field of high-speed aircraft heat protection, and particularly relates to a shock rod based on a supercritical fluid coolant, a heat protection device and a heat protection method.
Background
To meet the thermal protection requirements of high speed aircraft and to increase their payload, blunt body head configurations are commonly employed. However, during high speed flight, the aircraft nose can generate intense bow shock waves, thereby causing the pressure and heat flow in the area to rise sharply, causing the resistance to rise rapidly, and the temperature to rise rapidly, which brings serious challenges to the power and heat insulation system of the whole aircraft. How to realize the effects of drag reduction and heat protection is a concern of many students, and drag reduction and heat protection are generally difficult to be simultaneously considered.
In order to solve the problem of overlarge resistance and heat flow of an aircraft in the high-speed flight process, researchers at home and abroad propose a plurality of flow control methods, such as shock wave rods, reverse jet flow, windward grooves, energy delivery, pneumatic disk additionally arranged and the like, and respectively carry out more intensive researches. The active cooling method mainly realizes the aims of drag reduction and heat prevention by spraying cooling working medium or designing some specific mechanical structures.
The results of the study show that shock beam is one of the simplest and most effective ways. The introduction of shock beam structures also presents other problems, such as the harsher heat flow environment to which the shock beam tip is subjected, and the incidence of shock beam induced oblique shock waves to the bulb causes localized "hot spots" that cause localized ablation of the aircraft head. In order to weaken the adverse effect of the shock rod, a reverse jet hole is formed at the top end of the shock rod, taking Chinese patent CN 200610169683.9-a hypersonic aircraft ablation-free self-adaptive heat-protection and drag-reduction system as an example, a shock rod combined jet mode is adopted to push out shock waves attached to the top end of the shock rod on one hand, so that a separation area covers a larger aircraft head area, and on the other hand, a layer of cold air isolation is added between the top end of the shock rod and high-speed incoming flow, so that the effect of reducing the heat flow at the top end of the shock rod is achieved. However, the container in this way requires a large weight, i.e. the cooling structure occupies a large space and is of a high load, which is not conducive to the lightweight design of high-speed aircraft, and the thermal protection of the shock rod itself and of the high-speed aircraft itself remains to be improved, and the head of the aircraft is still susceptible to ablation.
Disclosure of Invention
The invention aims to solve the technical problem of providing the shock rod, the heat-proof device and the method based on the supercritical fluid coolant, which can improve the cooling effect of the shock rod and the thermal environment of a high-speed aircraft on the basis of reducing the quality of a supercritical coolant storage tank.
The invention provides a shock rod based on supercritical fluid coolant, which comprises a rod body and a pneumatic disc;
one end of the rod body is connected with the pneumatic disc, and the other end of the rod body is arranged on the head of the aircraft;
the pneumatic disc is provided with a cooling channel I along the axial direction, the pneumatic disc is provided with a cooling channel II along the radial direction, the cooling channel II comprises a contraction section, a roar channel and an expansion section which are sequentially arranged from the axis to the outside of the disc, and the other end of the contraction section is communicated with the cooling channel I;
the supercritical fluid exchanges heat and cools the rod body and the pneumatic disc through the cooling channel I and the cooling channel II, and is converted into a gas state after being accelerated, decompressed and cooled in the process of passing through the cooling channel II, and the low-temperature supersonic gas film is sprayed out of the expansion section.
Further, the cooling channel II is an annular channel.
Further, the cooling channel II is perpendicular to the cooling channel I.
Still further, the convergent section transitions in a circular arc toward one side of the aircraft nose.
Furthermore, a windward concave cavity is arranged at the axial center position of one end of the pneumatic disc, which is away from the rod body.
Still further, the channel wall that the roar and expansion section deviate from body of rod one end is straight structure, the channel wall that roar and expansion section are close to body of rod one end is the arc structure of variable cross section.
The invention also provides a heat protection device based on the supercritical fluid coolant, which comprises a supercritical coolant storage tank and a shock rod, wherein the end part of the rod body of the shock rod is arranged at the head of the aircraft, the supercritical coolant storage tank is arranged inside the head of the aircraft, and the supercritical coolant storage tank is communicated with the cooling channel I of the rod body.
Further, the radius of the aircraft head is 20-30 mm, and the half cone angle of the aircraft body 2 is 20-30 °.
The invention also provides a heat-proof method based on the supercritical fluid coolant, which uses the heat-proof device and comprises the following steps:
the supercritical coolant storage tank releases supercritical fluid, the supercritical fluid exchanges heat and cools the rod body and the pneumatic disc through the cooling channel I and the cooling channel II, the supercritical fluid is accelerated, depressurized and cooled in the process of passing through the cooling channel II and then is converted into a gas state, a low-temperature supersonic gas film is sprayed out of the expansion section, the low-temperature supersonic gas film pushes out bow-shaped laser generated by the pneumatic disc, so that bow-shaped laser waves are far away from the high-speed aircraft, and the low-temperature supersonic gas film and high-temperature gas flow after bow-shaped laser are mixed for absorbing heat, so that the total air flow temperature after bow-shaped laser is reduced.
The shock rod provided by the invention has the beneficial effects that the supercritical fluid is used as a coolant, the supercritical fluid is always in a supercritical state when in the cooling channel I and part of the cooling channel II, the rod body and the pneumatic disc can be subjected to heat exchange and cooling, the structural strength of the shock rod is ensured, the temperature of the supercritical fluid can be slightly increased in the heat exchange and cooling process of the rod body and the pneumatic disc, the temperature and the pressure can be ensured to be still higher than the critical point, the supercritical state is still ensured, and the heat exchange and cooling effects are met. The supercritical fluid has the highest specific heat capacity near the critical point and has huge heat absorption potential, so that the heat exchange effect can be improved by using the coolant, the volume of a supercritical coolant storage tank can be reduced, and the space of a heat protection system can be saved.
In the process of passing through the cooling channel II, the supercritical fluid can be accelerated, depressurized and cooled when passing through the contraction section, the roar and the expansion section, and can be converted into a gaseous state, and finally the gaseous state is sprayed out from an outlet of the expansion section, so that a low-temperature supersonic gas film perpendicular to the high-speed gas flow is formed by the gaseous state coolant, and on one hand, the low-temperature supersonic gas film pushes the bow-shaped shock wave outwards to be far away from the high-speed aircraft, and the bow-shaped shock wave is prevented from being incident on the wall surface of the body of the aircraft to generate local high heat flow; on the other hand, the low-temperature supersonic air film and the high-temperature air flow after the bow-shaped excitation are mixed for absorbing heat, so that the total air flow temperature after the bow-shaped excitation is reduced, the thermal environment of the head part and the body part of the aircraft is improved, and the thermal protection effect is achieved. In addition, the supersonic air film increases the low speed reflux zone between the shock rod and the aircraft head, further reducing drag.
The invention combines the structure of the shock wave rod and the characteristic of the supercritical fluid used, so that the supercritical fluid is in the optimal state at a specific position, and on the basis of reducing the quality of a supercritical coolant storage tank, the invention can not only improve the cooling effect of the shock wave rod, ensure the structural strength of the shock wave rod, but also improve the thermal environment of the head and the body of the aircraft, thereby achieving the thermal protection effect. But also can reduce the flight resistance and improve the aerodynamic performance of the aircraft.
Drawings
FIG. 1 is a schematic diagram of the structure of the present invention;
FIG. 2 is a schematic diagram of the local structure of the shock beam in the present invention (wherein the solid thick arrows represent high velocity gas flow, the open thick arrows represent low temperature supersonic gas film, and the solid thin arrows represent supercritical fluid).
In the figures, 1-aircraft head, 2-aircraft body, 3-supercritical coolant reservoir, 4-flow valve, 5-rod, 51-cooling channel I, 6-pneumatic disk, 61-cooling channel II, 611-convergent section, 612-throat, 613-divergent section, 614-arcuate tip, 62-windward cavity.
Detailed Description
The following description of the embodiments of the present invention will be made clearly and fully with reference to the accompanying drawings, in which it is evident that the embodiments described are only some, but not all embodiments of the invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
It should be noted that all directional indicators (such as up, down, left, right, front, and rear … …) in the embodiments of the present invention are merely used to explain the relative positional relationship, movement, etc. between the components in a particular posture (as shown in the drawings), and if the particular posture is changed, the directional indicator is changed accordingly.
Furthermore, descriptions such as those referred to as "first," "second," and the like, are provided for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implying an order of magnitude of the indicated technical features in the present disclosure. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include at least one such feature. In the description of the present invention, the meaning of "plurality" means at least two, for example, two, three, etc., unless specifically defined otherwise.
In the present invention, unless specifically stated and limited otherwise, the terms "connected," "affixed," and the like are to be construed broadly, and for example, "affixed" may be a fixed connection, a removable connection, or an integral body; the device can be mechanically connected, electrically connected, physically connected or wirelessly connected; either directly or indirectly, through intermediaries, or both, may be in communication with each other or in interaction with each other, unless expressly defined otherwise. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
In addition, the technical solutions of the embodiments of the present invention may be combined with each other, but it is necessary to be based on the fact that those skilled in the art can implement the technical solutions, and when the technical solutions are contradictory or cannot be implemented, the combination of the technical solutions should be considered as not existing, and not falling within the scope of protection claimed by the present invention.
As shown in fig. 1-2, the present invention provides a shock rod based on supercritical fluid coolant, comprising a rod body 5 and a pneumatic disc 6;
one end of the rod body 5 is connected with the pneumatic disc 6, and the other end of the rod body is arranged on the head 1 of the aircraft;
the pneumatic disc 6 is provided with a cooling channel I51 along the axial direction, the pneumatic disc 6 is provided with a cooling channel II 61 along the radial direction, the cooling channel II 61 comprises a contraction section 611, a roar 612 and an expansion section 613 which are sequentially arranged from the axis to the outside of the disc, and the other end of the contraction section 611 is communicated with the cooling channel I51. The area of the constriction 611 in the direction from the cooling channel i 51 to the throat 612 is sequentially reduced, that is, the channel area is gradually reduced when the supercritical fluid passes through, which can lead to the increase of the speed of the fluid, realize the speed increasing effect, but the flow is still at subsonic speed; in throat 612, the passage area is minimal, where the supercritical fluid flow velocity reaches a maximum, i.e., the speed of sound; the area of the expansion section 613 in the radial direction from the throat 612 to the air disk 6 increases in order, that is, the passage area is gradually widened when the supercritical fluid passes therethrough, and the supercritical fluid continues to accelerate from the sound velocity to the supersonic velocity. The constriction 611, the throat 612 and the expansion 613 adopt specific configurations to form a laval nozzle configuration, so that the effect of increasing, reducing and reducing the temperature after the supercritical fluid flows is finally achieved, specifically, the supercritical fluid in the supercritical coolant storage tank 3 is in a larger pressure state, the supercritical fluid can be regarded as a "dense gas", and supersonic flow is formed after the supercritical fluid flows through the cooling channel ii 61 of the laval nozzle configuration.
The supercritical fluid exchanges heat and cools the rod body 5 and the pneumatic disc 6 through the cooling channel I51 and the cooling channel II 61, and is converted into a gas state after being accelerated, decompressed and cooled in the process of passing through the cooling channel II 61, and the low-temperature supersonic gas film is sprayed out from the expansion section 613.
The shock rod provided by the invention uses the supercritical fluid as the coolant, the supercritical fluid is always in a supercritical state when in the cooling channel I51 and part of the cooling channel II 61, the rod body 5 and the pneumatic disc 6 can be subjected to heat exchange and cooling, the structural strength of the shock rod is ensured, the temperature of the supercritical fluid is slightly increased in the heat exchange and cooling process of the rod body 5 and the pneumatic disc 6, but the temperature and the pressure can be ensured to be still higher than the critical point and still in the supercritical state, and the heat exchange and cooling effects are met. The supercritical fluid has the highest specific heat capacity near the critical point and has huge heat absorption potential, so that the heat exchange effect can be improved by using the coolant, the volume of the supercritical coolant storage tank 3 can be reduced, and the space of a heat protection system can be saved.
In the process of passing through the cooling channel II 61, the supercritical fluid can be accelerated, depressurized and cooled when passing through the contraction section 611, the roar 612 and the expansion section 613, and can be converted into a gaseous state, and finally the gaseous state is sprayed out from the outlet of the expansion section 613 to form a low-temperature supersonic air film perpendicular to the high-speed air flow, and the low-temperature supersonic air film pushes the bow-shaped shock wave outwards on one hand to enable the bow-shaped shock wave to be far away from the high-speed aircraft, so that the bow-shaped shock wave is prevented from being incident on the wall surface of the aircraft body 2 to generate local high heat flow; on the other hand, the low-temperature supersonic air film and the high-temperature air flow after the bow-shaped excitation are mixed for absorbing heat, so that the total air flow temperature after the bow-shaped excitation is reduced, the thermal environment of the aircraft head part 1 and the aircraft body part 2 is improved, and the thermal protection effect is achieved. In addition, the supersonic gas film increases the low speed recirculation zone between the shock rod and the aircraft nose 1, further reducing drag.
The invention combines the structure of the shock wave rod and the characteristic of the supercritical fluid used, so that the supercritical fluid is in the optimal state at a specific position, and on the basis of reducing the quality of the supercritical coolant storage tank 3, the invention can not only improve the cooling effect of the shock wave rod, ensure the structural strength of the shock wave rod, but also improve the thermal environment of the head part 1 and the body part 2 of the aircraft, thereby achieving the thermal protection effect. But also can reduce the flight resistance and improve the aerodynamic performance of the aircraft.
In one embodiment, the cooling channel ii 61 is an annular channel, that is, the contraction section 611, the roar 612 and the expansion section 613 are all annular channels, so that the supercritical fluid sprayed from the expansion section 613 can be ensured to be a uniformly distributed low-temperature supersonic air film with a disc structure, and the heat protection effect is ensured. In this embodiment, the pneumatic disc 6 is divided into two parts separated from each other, the first part is fixedly connected with the rod body 5, the second part is parallel to the first part, and when the pneumatic disc is specifically used, the first part and the second part are connected with each other through the rib plate, so that the second part is fixed, the rib plate can be a plate arranged between the first part and the second part, can be a connecting column, and can reduce flow resistance as much as possible under the premise of ensuring the connection strength.
In one embodiment, the cooling channel ii 61 is perpendicular to the cooling channel i 51, so that the bow shock wave formed by the pneumatic disc 6 and the high-speed air flow and the low-temperature supersonic air film of the cooling channel ii 61 can increase the area of thermal protection without increasing the pneumatic resistance, thereby increasing the thermal environment of the high-speed aircraft.
In one embodiment, the shrinkage section 611 is in arc transition towards one side of the aircraft head 1, that is, the shrinkage section 611 is in an arc shape as a whole, so that flow separation occurring when the supercritical fluid is diverted from the cooling channel i 51 to the cooling channel ii 61 is avoided, pressure loss is reduced, flow field quality of the supersonic air film is improved, and in specific use, an arc tip 614 is convexly arranged towards the cooling channel i 51 from the center of the second part of the pneumatic disc 6, and the shrinkage section 611 is formed by the arc corner in the first part.
In this embodiment, a windward cavity 62 is disposed at an axial position of the end of the pneumatic disc 6 facing away from the rod body 5, preferably, a surface of the second portion of the pneumatic disc 6 facing away from the rod body 5 is spherical, the windward cavity 62 is disposed in a central residence area of the pneumatic disc 6, on one hand, energy of high-speed airflow is consumed by gas oscillation in the windward cavity 62, and the air cooling function is provided for the residence area of the pneumatic disc 6. On the other hand, because the contraction section 611 adopts diversion arc transition, the thickness (arc tip 614) of the central position of the pneumatic disc 6 is increased, the cooling effect of the supercritical fluid on the central standing point area of the pneumatic disc 6 is weakened, the thickness of the central standing point area of the windward concave cavity 62 is reduced, the heat conduction process of the supercritical fluid and the pneumatic disc 6 is improved, and the cooling effect is enhanced. Namely, the windward concave cavity 62 structure not only can cool the residence area of the pneumatic disc 6, but also can ensure the uniform thickness of the pneumatic disc 6 and ensure the heat exchange effect of the supercritical fluid on the pneumatic disc 6. In addition, the provision of the windward cavity 62 reduces the overall weight of the shock rod, avoiding excessive increases in the weight of the aircraft.
In one embodiment, the axial center position of the end of the pneumatic disc 6 away from the rod body 5 is provided with a windward concave cavity 62, the windward concave cavity 62 is arranged in the central residence area of the pneumatic disc 6, the air oscillation in the windward concave cavity 62 consumes the energy of high-speed air flow, and the air flow has a cooling effect on the residence area of the pneumatic disc 6, and in addition, the windward concave cavity 62 is arranged on and lightens the overall weight of the shock rod, so that the weight of the aircraft is prevented from being excessively increased.
In one embodiment, the channel walls of the throat 612 and the expansion section 613 facing away from the end of the rod body 5 are in a straight structure, and the channel walls of the throat 612 and the expansion section 613 near the end of the rod body 5 are in a variable-section arc-shaped structure, that is, the contraction section 611, the throat 612 and the expansion section 613 integrally form a binary half-Laval nozzle structure. The binary half Laval nozzle structure reduces half of the Laval nozzle volume while realizing the supersonic velocity air film, namely reduces the whole weight of the pneumatic disc 6, and avoids excessively increasing the weight of the aircraft.
The invention also provides a heat protection device based on the supercritical fluid coolant, which comprises a supercritical coolant storage tank 3 and a shock rod, wherein the end part of a rod body 5 of the shock rod is arranged at the head 1 of the aircraft, the supercritical coolant storage tank 3 is arranged inside the head 1 of the aircraft, the supercritical coolant storage tank 3 is communicated with a cooling channel I51 of the rod body 5, and the supercritical coolant storage tank 3 is used for providing the supercritical fluid for the cooling channel I51 and a cooling channel II 61 to be used as the coolant. In addition, a flow valve 4 is preferably provided between the supercritical coolant reservoir 3 and the cooling passage i 51 to facilitate control of the flow rate of the supercritical fluid into the cooling passage i 51.
In one embodiment, the radius of the aircraft nose 1 is 20-30 mm, the half cone angle of the aircraft body 2 is 20-30 degrees, and compared with the conventional high-speed aircraft, especially the high-speed missile, the aircraft nose 1 has smaller radius and smaller windward area, thus the heat-proof pressure is small and the shock resistance is small. One of the reasons why conventional high-speed aircraft cannot be made into a tip is that the heat flow of the tip part is extremely high and is extremely easy to burn out. Based on the improved shock rod and the heat-proof device, the invention can achieve smaller radius of the head 1 of the aircraft, and greatly reduces shock resistance while solving the heat protection problem.
The invention also provides a heat-proof method based on the supercritical fluid coolant, which uses the heat-proof device and comprises the following steps:
the supercritical coolant storage tank 3 releases supercritical fluid, the supercritical fluid exchanges heat and cools the rod body 5 and the pneumatic disc 6 through the cooling channel I51 and the cooling channel II 61, the supercritical fluid is accelerated, depressurized and cooled in the process of passing through the cooling channel II 61 and then is converted into a gas state, a low-temperature supersonic gas film is sprayed out from the expansion section 613, the low-temperature supersonic gas film pushes outwards bow-shaped shock waves generated by the pneumatic disc 6, so that bow-shaped shock waves are far away from the high-speed aircraft, and the low-temperature supersonic gas film and the bow-shaped shock waves are mixed and absorbed by high-temperature airflow, so that the total air flow temperature after bow-shaped shock is reduced.
The present invention also provides a specific embodiment, in which the supercritical fluid coolant in the supercritical coolant storage tank 3 is at a pressure near the critical point, typically having a pressure of several megapascals to tens of megapascals, and the flow valve 4 is a pressure reducing valve that is subjected to a pressure of 10 megapascals or more. The cooling channel II 61 adopts a design method of a shortened Laval nozzle structure, and the total length is 10mm (namely, the diameter of the pneumatic disc 6 is 10 mm). The shortened Laval nozzle structure design has the advantages that the windward area of the pneumatic disc 6 is reduced, the normal shock wave area of the pneumatic disc 6 is further reduced, shock wave resistance is further reduced, and meanwhile, the heat protection area and heat protection pressure are further reduced.
The radius of the aircraft head 1 is 20-30 mm and the half cone angle of the aircraft body 2 is 20-30 deg.. The aircraft head 1 has smaller radius and smaller windward area, thus reducing the heat protection area and the heat load, leading to small heat-proof pressure and small shock resistance.
During high-speed flight of the high-speed aircraft, the high-speed airflow oscillates in the windward concave cavity 62 of the pneumatic disc 6, part of energy is consumed, and the first heat protection of the pneumatic disc 6 is realized. The pneumatic disk 6 pre-compresses the high-speed air flow, and generates a bow shock wave in front, and compared with the bow shock wave generated by the head 1 of the original aircraft, the strength of the bow shock wave is greatly reduced, so that the effects of reducing heat and drag are realized. At this time, the flow valve 4 is opened, the supercritical fluid in the supercritical coolant storage tank 3 enters the cooling channel I51 to absorb heat and cool down the rod body 5, so as to ensure the structural strength of the rod body 5, then the supercritical coolant enters the cooling channel II 61, and when the pneumatic disc 6 is subjected to heat absorption and cooling down to realize the second heat protection, the process of speed increasing, pressure reducing and cooling down is completed through the contraction section 611, the roar 612 and the expansion section 613, supersonic flow is achieved at the outlet, and the low-temperature supersonic air film is sprayed. The low-temperature supersonic air film pushes the bow-shaped shock wave outwards on one hand to enable the bow-shaped shock wave to be far away from the high-speed aircraft, so that the bow-shaped shock wave is prevented from being incident on the wall surface of the aircraft body 2 to generate local high heat flow; on the other hand, the ultrasonic air film and the high-temperature air flow after the bow-shaped excitation are mixed for absorbing heat, so that the total air flow temperature after the bow-shaped excitation is reduced, the thermal environment of the head and the body of the aircraft is improved, and the thermal protection effect is achieved. In addition, the supersonic gas film increases the low speed recirculation zone between the shock rod and the aircraft nose 1, further reducing drag.
What is not described in detail in this specification is prior art known to those skilled in the art.
Claims (10)
1. A shock rod based on a supercritical fluid coolant, which is characterized by comprising a rod body (5) and a pneumatic disc (6);
one end of the rod body (5) is connected with the pneumatic disc (6), and the other end of the rod body is arranged on the head (1) of the aircraft;
a cooling channel I (51) is arranged on the rod body (5) along the axial direction, a cooling channel II (61) is arranged on the pneumatic disc (6) along the radial direction, the cooling channel II (61) comprises a contraction section (611), a roar (612) and an expansion section (613) which are sequentially arranged from the axis to the outside of the disc, and the other end of the contraction section (611) is communicated with the cooling channel I (51);
the supercritical fluid exchanges heat and cools the rod body (5) and the pneumatic disc (6) through the cooling channel I (51) and the cooling channel II (61), and is converted into a gas state after being accelerated, decompressed and cooled in the process of passing through the cooling channel II (61), and the low-temperature supersonic gas film is sprayed out of the expansion section (613).
2. The shock rod based on supercritical fluid coolant according to claim 1, characterized in that the cooling channel ii (61) is an annular channel.
3. A shock rod based on supercritical fluid coolant according to claim 1 or 2, characterized in that the cooling channel ii (61) is perpendicular to the cooling channel i (51).
4. A shock rod based on supercritical fluid coolant according to claim 3, characterized in that the constriction (611) has a circular arc transition towards one side of the aircraft head (1).
5. The shock rod based on supercritical fluid coolant according to claim 4, characterized in that the pneumatic disc (6) is provided with a windward cavity (62) at the axial position of the end facing away from the rod body (5).
6. The shock rod based on supercritical fluid coolant according to claim 1, characterized in that the pneumatic disc (6) is provided with a windward cavity (62) at the axial position of the end facing away from the rod body (5).
7. A shock rod based on supercritical fluid coolant according to claim 1, characterized in that the channel walls of the throat (612) and the expansion section (613) at the end facing away from the rod body (5) are in a straight structure, and the channel walls of the throat (612) and the expansion section (613) at the end near the rod body (5) are in a variable cross-section arc structure.
8. A heat protection device based on supercritical fluid coolant, characterized by comprising a supercritical coolant storage tank (3) and a shock rod according to any of claims 1-7, the end of the rod body (5) of the shock rod being arranged at the aircraft head (1), the supercritical coolant storage tank (3) being arranged inside the aircraft head (1), the supercritical coolant storage tank (3) being in communication with the cooling channel i (51) of the rod body (5).
9. A shock rod based on supercritical fluid coolant according to claim 8, characterized in that the radius of the aircraft nose (1) is 20-30 mm and the half cone angle of the aircraft body 2 is 20-30 °.
10. A heat protection method based on supercritical fluid coolant, characterized in that a heat protection device according to claim 8 or 9 is used, comprising the steps of:
the supercritical coolant storage tank (3) releases supercritical fluid, the supercritical fluid exchanges heat and cools the rod body (5) and the pneumatic disc (6) through the cooling channel I (51) and the cooling channel II (61), the supercritical fluid is accelerated, depressurized and cooled in the process of passing through the cooling channel II (61) and then is converted into a gas state, a low-temperature supersonic gas film is sprayed out of the expansion section (613), the low-temperature supersonic gas film pushes outwards an arch-shaped shock wave generated by the pneumatic disc (6), so that the arch-shaped shock wave is far away from the high-speed aircraft, the low-temperature supersonic gas film and the arch-shaped shock wave are mixed and absorbed by high-temperature gas flow, and the total air flow temperature after arch-shaped shock is reduced.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310839661.2A CN117465653A (en) | 2023-07-10 | 2023-07-10 | Shock rod based on supercritical fluid coolant, heat-proof device and method |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310839661.2A CN117465653A (en) | 2023-07-10 | 2023-07-10 | Shock rod based on supercritical fluid coolant, heat-proof device and method |
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| Publication Number | Publication Date |
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| CN202310839661.2A Pending CN117465653A (en) | 2023-07-10 | 2023-07-10 | Shock rod based on supercritical fluid coolant, heat-proof device and method |
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| Country | Link |
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| CN (1) | CN117465653A (en) |
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