A hypersonic or supersonic aircraft heat reduction structure and heat reduction method

Abstract

The application discloses a hypersonic or supersonic aircraft heat reduction structure and a heat reduction method, and by arranging the hypersonic or supersonic aircraft heat reduction structure, the angle of the first shock wave at the upstream shock wave is increased, so that the shock wave-shock wave interference mode at the upstream shock wave and the downstream shock wave of the aircraft is changed from the fourth type shock wave-shock wave interference or the fourth r type shock wave-shock wave interference into the fifth type shock wave-shock wave interference or the sixth type shock wave-shock wave interference. The application is applied to the thermal protection field, and by inhibiting the shock wave-shock wave interference, effective heat reduction is realized. Compared with the method of directly eliminating the interference, the application is more feasible and has practical application potential, can effectively overcome the complexity of the interference, improve the heat reduction effect, and reduce the complexity and cost of the technical implementation, improve the thermal management effect of the hypersonic or supersonic aircraft, and has a positive influence on the performance and safety of the aircraft, and provides a new method for the thermal protection of the aircraft.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of supersonic or hypersonic aircraft, in particular to a supersonic or hypersonic aircraft heat reduction structure and heat reduction method. BACKGROUND

[0002] Supersonic and hypersonic aircraft face severe thermal management challenges during high-speed flight. Due to the aircraft crossing the air at supersonic speed, a large amount of aerodynamic heat load is generated, which leads to rapid rise of the aircraft surface temperature, and has adverse effects on the aircraft structure and performance. Therefore, reducing the aircraft wall temperature and controlling heat transfer are key issues to ensure the safe operation and performance optimization of the aircraft. Traditional heat reduction methods include surface coating and thermal insulation materials, etc. passive means to reduce heat transfer, however, because high-speed flight will cause severe aerodynamic heat load, which may limit the passive heat reduction method under high-speed flight conditions.

[0003] In supersonic and hypersonic aircraft, shock wave-shock wave interference is an important problem in heat reduction. When a shock wave interacts with the surrounding fluid, complex interference effects occur, causing changes in the propagation direction, angle and intensity of the shock wave. This interference effect can cause the temperature of the aircraft wall to rise and increase the aerodynamic drag of the aircraft, which has adverse effects on performance and safety. Existing shock wave control methods mainly rely on adjusting the geometry of the aircraft, optimizing the parameters of the leading edge and trailing edge, or using flow control technology to affect the propagation of the shock wave. However, traditional shock wave control methods have some limitations in practical application. For example, methods that directly eliminate shock wave-shock wave interference often require complex aircraft structure adjustments and flow control devices, and the effect may be limited under different working conditions. SUMMARY

[0004] In view of the deficiencies in the prior art, the present application provides a supersonic or hypersonic aircraft heat reduction structure and heat reduction method, which can adjust the angle of the upstream shock wave and the downstream shock wave according to the change of the wall temperature, and is expected to achieve more effective heat reduction effect, and has a positive impact on the thermal management and performance optimization of supersonic and hypersonic aircraft.

[0005] To achieve the above-mentioned purpose, the present application provides a supersonic or hypersonic aircraft heat reduction structure, comprising a wedge-shaped structure body, the structure body comprising a first plane at the bottom, a second plane and a third plane at the top, a fourth plane at the tail, and a fifth plane and a sixth plane at the side;

[0006] The first plane, the second plane, the third plane and the fourth plane are sequentially connected end to end and form a closed ring structure, and the fifth plane and the sixth plane cover both sides of the ring structure;

[0007] The first plane and the third plane are perpendicular to the fourth plane.

[0008] In one of the embodiments, the hypersonic or supersonic aircraft heat reduction structure further comprises a jet flow structure, the jet flow structure comprising an air inlet channel, a gas collection cavity, a first jet flow channel and a second jet flow channel arranged inside the structure body.

[0009] The gas collection cavity is located inside the structure body and close to the third plane and the fourth plane, the first end of the air inlet channel is located on the second plane, and the second end of the air inlet channel is in communication with the gas collection cavity.

[0010] The first end of the first jet flow channel is in communication with the gas collection cavity, and the second end of the first jet flow channel is located on the third plane.

[0011] The first end of the second jet flow channel is in communication with the gas collection cavity, and the second end of the second jet flow channel is located on the fourth plane.

[0012] In one of the embodiments, the jet flow structure further comprises a third jet flow channel and a fourth jet flow channel arranged inside the structure body.

[0013] The first end of the third jet flow channel is in communication with the gas collection cavity, and the second end of the third jet flow channel is located on the fifth plane.

[0014] The first end of the fourth jet flow channel is in communication with the gas collection cavity, and the second end of the fourth jet flow channel is located on the sixth plane.

[0015] In one of the embodiments, the second end of the third jet flow channel and the second end of the fourth jet flow channel are inclined towards the top and tail of the structure body.

[0016] In one of the embodiments, the jet flow structure further comprises a discharge electrode.

[0017] The discharge electrode is arranged in the gas collection cavity.

[0018] To achieve the above-mentioned purpose, the application further provides a hypersonic or supersonic aircraft heat reduction method, comprising:

[0019] The above-mentioned hypersonic or supersonic aircraft heat reduction structure is arranged at the upstream shock wave of the aircraft, the angle of the first shock wave at the upstream shock wave is increased by active and / or passive flow control mode, so that the shock wave-shock wave interference mode at the upstream shock wave and the downstream shock wave of the aircraft is changed from the fourth type shock wave-shock wave interference or the fourth r type shock wave-shock wave interference to the fifth type shock wave-shock wave interference or the sixth type shock wave-shock wave interference.

[0020] In one of the embodiments, the heat reduction method further comprises: arranging active and / or passive flow control devices at the downstream shock wave of the aircraft to reduce the angle of the second shock wave at the downstream shock wave.

[0021] In one of the embodiments, at the upstream shock wave and the downstream shock wave of the aircraft, respectively, active and / or passive flow control devices are arranged, and when the aircraft wall temperature monitored by the sensor system is 1300K-1500K, the heat reduction method comprises the following steps:

[0022] Step 101: After increasing the attack angle of the aircraft by δθ degrees, it is determined whether the aircraft wall temperature is less than 1300K:

[0023] If yes, the aircraft is controlled to fly at the current attack angle, and after a preset period of time, the attack angle of the aircraft is adjusted to the initial value;

[0024] Otherwise, step 102 is performed;

[0025] Step 102: It is determined whether the current attack angle of the aircraft reaches the attack angle threshold:

[0026] If yes, step 103 is performed;

[0027] Otherwise, return to step 101;

[0028] Step 103: The discharge electrode on the heat reduction structure of the supersonic or hypersonic aircraft at the upstream shock wave is controlled to high-voltage discharge, and the shock-shock interference mode at the upstream shock wave and the downstream shock wave of the aircraft is changed from the fourth type shock-shock interference or the fourth r type shock-shock interference to the fifth type shock-shock interference;

[0029] Step 104: After a preset period of time, it is determined whether the aircraft wall temperature is less than 1300K:

[0030] If yes, the discharge electrode on the heat reduction structure of the supersonic or hypersonic aircraft at the upstream shock wave is turned off, and after a preset period of time, the attack angle of the aircraft is adjusted to the initial value;

[0031] Otherwise, step 105 is performed;

[0032] Step 105: It is determined whether the aircraft wall temperature exceeds 1500K:

[0033] If yes, the active and / or passive flow control devices at the downstream shock are turned on, the shock-shock interference mode at the upstream shock and the downstream shock of the aircraft is changed from the 5th shock-shock interference to the 6th shock-shock interference, and after the aircraft wall temperature is less than 1300K, the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock and the active and / or passive flow control devices at the downstream shock are turned off, and after a preset period of time, the attack angle of the aircraft is adjusted to the initial value.

[0034] Otherwise, return to step 104.

[0035] In one of the embodiments, the active and / or passive flow control devices are arranged at the upstream shock and the downstream shock of the aircraft respectively, and when the aircraft wall temperature monitored by the sensor system is greater than 1500K, the heat reduction method comprises the following steps:

[0036] Step 201, after the attack angle of the aircraft is increased by δθ degrees, it is determined whether the aircraft wall temperature is less than 1300K:

[0037] If yes, the aircraft is controlled to fly at the current attack angle, and after a preset period of time, the attack angle of the aircraft is adjusted to the initial value;

[0038] Otherwise, step 202 is performed.

[0039] Step 202, it is determined whether the current attack angle of the aircraft reaches the attack angle threshold:

[0040] If yes, step 203 is performed.

[0041] Otherwise, return to step 201.

[0042] Step 203, the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock is controlled to discharge at high voltage, and the active and / or passive flow control devices at the downstream shock are turned on, the shock-shock interference mode at the upstream shock and the downstream shock of the aircraft is changed from the 4th shock-shock interference or the 4th r shock-shock interference to the 6th shock-shock interference, and after the aircraft wall temperature is less than 1300K, the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock and the active and / or passive flow control devices at the downstream shock are turned off, and after a preset period of time, the attack angle of the aircraft is adjusted to the initial value.

[0043] In one of the embodiments, the process of increasing the first shock angle by the active and / or passive flow control devices at the upstream shock is as follows:

[0044] The first shock wave angle is increased by forming a high-temperature and high-speed jet flow on the surface of the aircraft at the upstream shock wave through spark discharge or arc discharge, and then inducing a new shock wave to lift the first shock wave.

[0045] The first shock wave angle is increased by forming a high-temperature and high-speed jet flow on the surface of the aircraft at the upstream shock wave through high-temperature and high-pressure carbon dioxide gas generated by the aircraft semi-Brayton cycle active cooling and power generation system, and then inducing a new shock wave to lift the first shock wave.

[0046] The first shock wave angle is increased by arranging a leading edge boundary layer around the surface of the aircraft at the upstream shock wave, thereby generating vortex around the leading edge and affecting the formation and propagation of the first shock wave.

[0047] In one embodiment, the process of reducing the second shock wave angle by the active and / or passive flow control device at the downstream shock wave is:

[0048] The second shock wave angle is reduced by generating a local high temperature through abnormal glow discharge and laser-induced plasma on the surface of the aircraft at the downstream shock wave, and then generating a local virtual solid wall at the slope of the second shock wave under the action of the local high temperature, so that the turning angle of the slope of the second shock wave is reduced.

[0049] The second shock wave angle is reduced by generating a mixture of free electrons and positive ions through normal glow discharge on the surface of the aircraft at the downstream shock wave, and then generating Lorentz force through the coupling of the magnetic field and the mixture, and applying the same volume force as the flow direction to the boundary layer fluid to locally accelerate the fluid.

[0050] The second shock wave angle is reduced by arranging an adjustable surface structure as a shock wave control surface on the surface of the aircraft at the downstream shock wave, and adjusting the shape and position of the shock wave control surface.

[0051] Compared with the prior art, the present application has the following beneficial technical effects:

[0052] 1. The hypersonic or supersonic aircraft heat reduction structure in the present application improves the traditional forward triangular wedge vortex generator by changing the sharp corners at the top and tail into flat surfaces, thereby reducing the resistance and solving the problem of aircraft ablation avoidance.

[0053] 2. The hypersonic or supersonic aircraft heat reduction structure in the present application is provided with a jet structure in the preferred scheme, thereby generating a jet flow at the top and tail of the structure, making up for the sharp corners at the top and tail, and maintaining good flow control effect of the hypersonic or supersonic aircraft heat reduction structure.

[0054] 3. The heat reduction method in this invention is designed for supersonic or hypersonic vehicles. It transforms the shock wave-shock wave interference mode from type IV or type IVr to type V or type VI, thereby achieving effective heat reduction. Compared with the method of directly eliminating interference, it is more feasible and has the potential for practical application. It can effectively overcome the complexity of interference, improve the heat reduction effect, and reduce the complexity and cost of technical implementation.

[0055] 4. This invention improves the thermal management of supersonic or hypersonic vehicles by transforming the shock-shock interference mode at the upstream and downstream shock waves of the supersonic or hypersonic vehicle into Class V or Class VI, and has a positive impact on the vehicle's performance and safety, providing a novel method for vehicle thermal protection. Attached Figure Description

[0056] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the structures shown in these drawings without creative effort.

[0057] Figure 1 This is a schematic diagram of a traditional forward triangular wedge vortex generator.

[0058] Figure 2 This is a first axonometric view of the heat dissipation structure of a supersonic or hypersonic vehicle in Embodiment 1 of the present invention;

[0059] Figure 3 This is a second axonometric view of the heat dissipation structure of the supersonic or hypersonic vehicle in Embodiment 1 of the present invention;

[0060] Figure 4 This is a cross-sectional view of the heat dissipation structure of a supersonic or hypersonic vehicle in Embodiment 1 of the present invention.

[0061] Figure 5 This is a schematic diagram of the internal structure of the heat dissipation structure of the supersonic or hypersonic vehicle in Embodiment 1 of the present invention.

[0062] Figure 6Fig. 1 is a schematic diagram of the working principle of the heat reduction structure of the supersonic or hypersonic aircraft in the embodiment 1 of the present application, wherein (a) is a schematic diagram of the shock wave angle when the aircraft surface is not provided with the heat reduction structure of the supersonic or hypersonic aircraft, (b) is a schematic diagram of the shock wave angle when the aircraft surface is provided with the heat reduction structure of the supersonic or hypersonic aircraft having the first jet flow channel and the second jet flow channel, and (c) is a schematic diagram of the shock wave angle when the aircraft surface is provided with the heat reduction structure of the supersonic or hypersonic aircraft having the first jet flow channel, the second jet flow channel, the third jet flow channel and the fourth jet flow channel;

[0063] Figure 7 Fig. 2 is a flow chart of the unsteady working mode point of the heat reduction method when the aircraft wall surface temperature is 1300K-1500K in the embodiment 2 of the present application;

[0064] Figure 8 Fig. 3 is a flow chart of the unsteady working mode point of the heat reduction method when the aircraft wall surface temperature exceeds 1500K in the embodiment 2 of the present application;

[0065] Figure 9 Fig. 4 is a schematic diagram of the flow field structure change before and after the shock wave control in the application example of the heat reduction method in the embodiment 2 of the present application, wherein (a) is a schematic diagram of the flow field structure before the shock wave control, and (b) is a schematic diagram of the flow field structure after the shock wave control;

[0066] Figure 10 Fig. 5 is a schematic diagram of the wall surface maximum heat flow change before and after the shock wave control in the application example of the heat reduction method in the embodiment 2 of the present application.

[0067] The accompanying drawings are included to provide a further understanding of the present application, and are incorporated herein and constitute a part of the detailed description. The embodiments described in the drawings merely are some of the embodiments of the present application, and not all, and consequently, they should not be used in a restrictive sense. In the drawings:

[0068] The present application is achieved, the functional characteristics and advantages will be further described with reference to the embodiments and the accompanying drawings. DETAILED DESCRIPTION

[0069] The technical solutions in the embodiments of the present application will be clearly and completely described with reference to the accompanying drawings in the embodiments of the present application. Obviously, the described embodiments are only some of the embodiments of the present application, and not all. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor are within the scope of protection of the present application.

[0070] It should be noted that all the direction indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present application are only used to explain the relative position relationship, movement condition, etc. between components in a certain specific posture (as shown in the drawings), and if the specific posture changes, the direction indications will also change accordingly.

[0071] In addition, the descriptions such as "first", "second" and the like in the present application are only for the purpose of description, and cannot be understood as indicating or implying the relative importance of the indicated technical features or implicitly indicating the number of the indicated technical features. Therefore, the features defined with "first", "second" can explicitly or implicitly include at least one of the features. In the description of the present application, the meaning of "a plurality of" is at least two, for example, two, three, etc., unless otherwise specifically limited.

[0072] In the present application, unless otherwise specifically defined and limited, the terms "connection", "fixation" and the like should be understood in a broad sense, for example, "fixation" can be fixed connection, or detachable connection, or integral; can be mechanical connection, or electrical connection, or physical connection, or wireless communication connection; can be directly connected, or indirectly connected through intermediate medium, can be the internal communication of two elements or the interaction relationship between two elements, unless otherwise specifically limited. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0073] In addition, the technical solutions of each embodiment of the present application can be combined with each other, but it must be based on the fact that a person skilled in the art can realize it, and when the combination of technical solutions appears contradictory or unachievable, it should be considered that the combination of technical solutions does not exist, nor is it within the scope of protection required by the present application.

[0074] Embodiment 1

[0075] The hypersonic aircraft heat reduction structure can change the properties of fluid flow by introducing vortex or vortex motion, so as to realize the control and adjustment of flow. Referring to Figure 1The traditional forward triangular wedge vortex generator can generate abundant flow direction counter-rotating vortex pairs, thereby controlling the boundary layer flow separation and further affecting the shape of the shock wave. Moreover, due to the more stable triangular wedge structure, the forward triangular wedge vortex generator is widely used in the thermal protection field of supersonic or hypersonic aircraft. When the forward triangular wedge vortex generator is installed on the surface of the supersonic or hypersonic aircraft, the sharp corner structure at the top of the forward triangular wedge vortex generator inevitably generates relatively large resistance, and the sharp corner structure at the tail of the forward triangular wedge vortex generator also causes the ablation problem of the aircraft. Based on this, the supersonic or hypersonic aircraft heat reduction structure provided in the embodiment improves the sharp corner structure at the top and the tail into a planar structure on the basis of the traditional forward triangular wedge vortex generator, thereby reducing the resistance and solving the ablation problem of the aircraft wall surface.

[0076] Reference Figures 2 to 5 The supersonic or hypersonic aircraft heat reduction structure in the embodiment includes a wedge-shaped structure body 1, which includes a first plane 101 at the bottom, a second plane 102 and a third plane 103 at the top, a fourth plane 104 at the tail, and a fifth plane 105 and a sixth plane 106 at the side. The first plane 101, the second plane 102, the third plane 103, and the fourth plane 104 are sequentially connected end to end and form a closed ring structure, the fifth plane 105 and the sixth plane 106 cover the two sides of the ring structure, and the first plane 101 and the third plane 103 are perpendicular to the fourth plane 104, that is, the first plane 101 and the third plane 103 are parallel to each other. Among them, the third plane 103 replaces the sharp corner structure at the top of the traditional forward triangular wedge vortex generator, and the fourth plane 104 replaces the sharp corner structure at the tail of the traditional forward triangular wedge vortex generator, thereby reducing the resistance and solving the ablation problem of the aircraft.

[0077] Considering that replacing the sharp corner structure at the top and the tail of the vortex generator with a planar structure will have a certain impact on the flow control effect of the vortex generator, in the embodiment, as preferred, a jet structure is arranged on the structure body 1 to compensate for the removed sharp corner structure.

[0078] In this embodiment, the jet flow structure includes an air inlet channel 201, a gas collection cavity 202, a first jet flow channel 203 and a second jet flow channel 204 arranged inside the structure 1. The gas collection cavity 202 is a rectangular cavity arranged inside the structure 1 and close to the tail end of the third plane 103 and the fourth plane 104. The air inlet channel 201 is also rectangular in cross section, with its length direction parallel to the first plane 101. The first end of the air inlet channel 201 is arranged on the second plane 102, and the second end of the air inlet channel 201 is in communication with the gas collection cavity 202. The length direction of the first jet flow channel 203 is perpendicular to the third plane 103. The first end of the first jet flow channel 203 is in communication with the gas collection cavity 202, and the second end of the first jet flow channel 203 is arranged in the central region of the third plane 103. In this way, after the gas flow enters the gas collection cavity 202 through the air inlet channel 201, a jet flow is formed in the central region of the second plane 102 through the first jet flow channel 203, so as to compensate for the sharp corner structure at the top of the hypersonic aircraft heat reduction structure, and the hypersonic aircraft heat reduction structure maintains good flow control effect. The length direction of the second jet flow channel 204 is perpendicular to the fourth plane 104. The first end of the second jet flow channel 204 is in communication with the gas collection cavity 202, and the second end of the second jet flow channel 204 is arranged in the central region of the fourth plane 104. In this way, after the gas flow enters the gas collection cavity 202 through the air inlet channel 201, a jet flow is formed in the central region of the fourth plane 104 through the fourth jet flow channel 206, so as to compensate for the sharp corner structure at the tail of the hypersonic aircraft heat reduction structure, and the hypersonic aircraft heat reduction structure maintains good flow control effect.

[0079] As a preferred embodiment, the jet flow structure further includes a third jet flow channel 205 and a fourth jet flow channel 206 arranged inside the structure 1. The first end of the third jet flow channel 205 is in communication with the gas collection cavity 202, and the second end of the third jet flow channel 205 is located on the fifth plane 105. The first end of the fourth jet flow channel 206 is in communication with the gas collection cavity 202, and the second end of the fourth jet flow channel 206 is located on the sixth plane 106. The second end of the third jet flow channel 205 and the second end of the fourth jet flow channel 206 are inclined towards the top and tail of the structure 1. In this way, after the gas flow enters the gas collection cavity 202 through the air inlet channel 201, a vortex flow is generated in the downstream of the hypersonic aircraft heat reduction structure, and the flow control effect of the hypersonic aircraft heat reduction structure is enhanced.

[0080] It is worth noting that in the jet structure, the first jet channel 203 and the second jet channel 204 are the main jet control parts, and a stronger jet needs to be generated. The third jet channel 205 and the fourth jet channel 206 are secondary jet control parts, mainly playing a reinforcing effect. Therefore, the cross-sectional area of the first jet channel 203 and the second jet channel 204 is basically the same, and the cross-sectional area size is larger than that of the third jet channel 205 and the fourth jet channel 206. Reference Figure 6 It is the schematic diagram of the working principle of the hypersonic aircraft heat reduction structure in the embodiment. When the hypersonic aircraft heat reduction structure having only the first jet channel 203 and the second jet channel 204 is arranged on the surface of the aircraft, the angle of the shock wave can be increased to a certain extent. When the hypersonic aircraft heat reduction structure having only the first jet channel 203, the second jet channel 204, and the third jet channel 205 and the fourth jet channel 206 is arranged on the surface of the aircraft, the angle of the shock wave can be further increased.

[0081] Further preferably, the jet structure further comprises a discharge electrode 207 arranged in the gas collection cavity 202 and electrically connected to an external control device through a pre-embedded lead wire. By arranging the discharge electrode 207 in the gas collection cavity 202, the discharge electrode 207 can be selectively controlled to perform high-voltage discharge in specific application processes, thereby controlling the jet effect of the jet structure. For example, when the incoming flow Mach number of the aircraft is low, the thermal protection requirement of the aircraft is low at this time, and the discharge electrode 207 does not need to perform high-voltage discharge, and only the jet generated by the air inlet channel 201 itself can meet the requirement; when the incoming flow Mach number of the aircraft is high, the thermal protection requirement of the aircraft is high at this time, and it is difficult to meet the requirement only by relying on the jet generated by the air inlet channel 201 itself, that is, the discharge electrode 207 can be controlled to perform high-voltage discharge, and the energy and frequency of the discharge can be modulated according to the requirement, thereby controlling the jet energy and frequency of the jet structure to meet the current thermal protection requirement of the aircraft.

[0082] It is worth noting that in specific applications, the communication position of the air inlet channel 201 and the gas collection cavity 202 is arranged as a stepped structure, and the height of the bottom wall of the gas collection cavity 202 is lower than that of the air inlet channel 201. At the same time, the discharge electrode 207 is fixed on the bottom wall or the stepped wall surface of the gas collection cavity 202, and the height of the discharge electrode 207 is kept lower than that of the air inlet channel 201, so as to prevent the air flow from interfering with the discharge of the discharge electrode 207.

[0083] Embodiment 2

[0084] The embodiment discloses a hypersonic or supersonic aircraft heat reduction method, which changes the shock wave-shock wave interference mode at the upstream shock wave and the downstream shock wave of the aircraft from the fourth type shock wave-shock wave interference or the fourth r type shock wave-shock wave interference to the fifth type shock wave-shock wave interference or the sixth type shock wave-shock wave interference. In the fourth type shock wave-shock wave interference mode, the upstream shock wave and the downstream shock wave interfere with each other in an obvious interaction mode, resulting in a sharp change in the shock wave structure and angle. This interference mode often increases the aircraft wall temperature and aerodynamic drag, and has an adverse effect on the performance of the aircraft. In the fourth r type shock wave-shock wave interference mode, the interference is more complex, involving reflection, refraction and change of propagation direction and other phenomena. In contrast, the fifth type and the sixth type shock wave-shock wave interference modes show less interference effect and are easier to control and optimize. In the fifth type shock wave-shock wave interference mode, the interaction between the upstream shock wave and the downstream shock wave is small, and the shock wave angle changes gently, so it is relatively easier to reduce the wall temperature. The sixth type shock wave-shock wave interference mode further reduces the interference effect, so that the interaction between the shock waves is weakened. Therefore, the heat reduction method in the embodiment can achieve effective heat reduction through interference suppression, and is more feasible and has practical application potential compared with the method of directly eliminating interference, can effectively overcome the complexity of interference, improve the heat reduction effect, and reduce the complexity and cost of technical implementation.

[0085] In the embodiment, the specific implementation mode of changing the shock wave-shock wave interference mode at the upstream shock wave and the downstream shock wave of the aircraft from the fourth type shock wave-shock wave interference or the fourth r type shock wave-shock wave interference to the fifth type shock wave-shock wave interference or the sixth type shock wave-shock wave interference is as follows:

[0086] The aircraft wall temperature monitored based on the sensor system;

[0087] The heat reduction structure of the supersonic or hypersonic aircraft in embodiment 1 is arranged at the upstream shock wave of the aircraft, and the angle of the first shock wave at the upstream shock wave is increased through active and / or passive flow control mode;

[0088] Optionally, active and / or passive flow control devices are arranged at the downstream shock wave, and the angle of the second shock wave at the downstream shock wave is reduced through active and / or passive flow control mode.

[0089] As a preferred real-time mode, the heat reduction method in the embodiment includes a fixed-length working mode, a periodic working mode and a non-steady working mode. Specifically:

[0090] The constant length mode is to arrange the supersonic or hypersonic aircraft heat reduction structure and the active and / or passive flow control device at the upstream and downstream shock waves of the aircraft respectively, and control the discharge electrode of the supersonic or hypersonic aircraft heat reduction structure at the upstream shock wave and the active and / or passive flow control device at the downstream shock wave to be in the open state, i.e. to continuously change the shock-shock interference mode at the upstream and downstream shock waves of the aircraft from the type IV shock-shock interference or the type IVr shock-shock interference to the type V shock-shock interference or the type VI shock-shock interference, so that the heat flux of the aircraft wall surface is always maintained at a low level.

[0091] The periodic mode is to arrange the supersonic or hypersonic aircraft heat reduction structure and the active and / or passive flow control device at the upstream and downstream shock waves of the aircraft respectively, and control the discharge electrode of the supersonic or hypersonic aircraft heat reduction structure at the upstream shock wave and the active and / or passive flow control device at the downstream shock wave to be in the periodic open state with a periodic frequency of 3 kHz to 10 kHz. That is, to periodically change the shock-shock interference mode at the upstream and downstream shock waves of the aircraft from the type IV shock-shock interference or the type IVr shock-shock interference to the type V shock-shock interference or the type VI shock-shock interference, so that the heat flux of the aircraft wall surface is periodically changed, but the time-averaged heat flux is reduced to below the ablation threshold of the thermal protection material.

[0092] The unsteady mode is to arrange the supersonic or hypersonic aircraft heat reduction structure and the active and / or passive flow control device at the upstream and downstream shock waves of the aircraft respectively, and control the discharge electrode of the supersonic or hypersonic aircraft heat reduction structure at the upstream shock wave and the active and / or passive flow control device at the downstream shock wave to be opened or closed according to the real-time detection value of the aircraft wall surface temperature.

[0093] In the embodiment, when the aircraft wall surface temperature monitored by the sensor system is 1300K-1500K, the reference Figure 7 The unsteady mode of the heat reduction method specifically includes the following steps:

[0094] Step 101, after increasing the attack angle of the aircraft by δθ degrees, it is judged whether the aircraft wall surface temperature is less than 1300K:

[0095] If yes, the aircraft is controlled to fly at the current attack angle, and after a preset period of time, the attack angle of the aircraft is adjusted to the initial value;

[0096] Otherwise, step 102 is performed;

[0097] Step 102, it is judged whether the current attack angle of the aircraft reaches the attack angle threshold:

[0098] If yes, step 103 is performed;

[0099] Otherwise, go to step 101;

[0100] Step 103, control the high-voltage discharge of the discharge electrode on the heat-shedding structure of the hypersonic vehicle at the upstream shock wave, and change the shock-shock interference mode of the vehicle at the upstream shock wave and the downstream shock wave from the fourth type shock-shock interference or the fourth r type shock-shock interference to the fifth type shock-shock interference;

[0101] Step 104, judge whether the wall temperature of the vehicle is less than 1300K after a preset period of time:

[0102] If yes, turn off the discharge electrode on the heat-shedding structure of the hypersonic vehicle at the upstream shock wave, and adjust the attack angle of the vehicle to the initial value after a preset period of time;

[0103] Otherwise, go to step 105;

[0104] Step 105, judge whether the wall temperature of the vehicle exceeds 1500K:

[0105] If yes, turn on the active and / or passive flow control device at the downstream shock wave, change the shock-shock interference mode of the vehicle at the upstream shock wave and the downstream shock wave from the fifth type shock-shock interference to the sixth type shock-shock interference, and then turn off the discharge electrode on the heat-shedding structure of the hypersonic vehicle at the upstream shock wave and the active and / or passive flow control device at the downstream shock wave, and adjust the attack angle of the vehicle to the initial value after a preset period of time;

[0106] Otherwise, go to step 104.

[0107] In this embodiment, when the wall temperature of the vehicle monitored by the sensor system exceeds 1500K, the active and / or passive flow control device at the downstream shock wave is turned on, and the shock-shock interference mode of the vehicle at the upstream shock wave and the downstream shock wave is changed from the fifth type shock-shock interference to the sixth type shock-shock interference. Figure 8 The unsteady working mode of the heat-shedding method includes the following steps:

[0108] Step 201, increase the attack angle of the vehicle by δθ degrees, and judge whether the wall temperature of the vehicle is less than 1300K:

[0109] If yes, control the vehicle to fly at the current attack angle, and adjust the attack angle of the vehicle to the initial value after a preset period of time;

[0110] Otherwise, go to step 202;

[0111] Step 202, judge whether the current attack angle of the vehicle reaches the attack angle threshold:

[0112] If yes, go to step 203;

[0113] Otherwise, go to step 201;

[0114] In step 203, high-voltage discharge of the discharge electrode on the hypersonic vehicle heat-shedding structure at the upstream shock wave is controlled, and the active and / or passive flow control device at the downstream shock wave is turned on, so that the shock-shock interference mode of the vehicle at the upstream shock wave and the downstream shock wave is changed from the fourth type shock-shock interference or the fourth r type shock-shock interference to the sixth type shock-shock interference, and after the wall temperature of the vehicle is less than 1300K, the discharge electrode on the hypersonic vehicle heat-shedding structure at the upstream shock wave and the active and / or passive flow control device at the downstream shock wave are turned off, and after a preset period of time, the attack angle of the vehicle is adjusted to the initial value.

[0115] In the specific implementation process, other active and / or passive flow control devices can also be arranged at the upstream shock wave of the vehicle to increase the first shock angle, which has a plurality of different embodiments, respectively as follows:

[0116] In the first embodiment, spark discharge or arc discharge is performed on the surface of the vehicle at the upstream shock wave, so that the gas is pressurized to form a high-temperature and high-speed jet, and then a new shock wave is induced to lift the first shock wave, so that the angle of the first shock wave is increased. In this process, the vehicle does not need to carry a large volume of gas source, only needs to use the on-board battery, not only simple structure, but also small volume, and will not cause additional load to the vehicle.

[0117] In the second embodiment, high-temperature and high-pressure carbon dioxide gas generated by the semi-Brayton cycle active cooling and power generation system of the vehicle is used to form a high-temperature and high-speed jet on the surface of the vehicle at the upstream shock wave, and then a new shock wave is induced to lift the first shock wave, so that the angle of the first shock wave is increased. In this process, the vehicle also does not need to carry a large volume of gas source, and the remaining gas after power generation can be used, which can turn waste into treasure, and the jet pressure and temperature after power generation are high, and have good control effect.

[0118] In the third embodiment, a leading edge attachment is arranged on the surface of the vehicle at the upstream shock wave, the leading edge attachment is a small protruding structure installed on the leading edge of the vehicle, can generate vortex around the leading edge, and affect the formation and propagation of the first shock wave, so as to increase the angle of the first shock wave.

[0119] It is worth noting that the third embodiment of the active and / or passive flow control device at the upstream shock wave is mainly used for the fixed-length working mode, and the first and second embodiments are mainly used for the fixed-length working mode, the periodic working mode and the unsteady working mode. In specific applications, the aircraft does not necessarily remain in one of the fixed-length working mode, the periodic working mode and the unsteady working mode. The three embodiments described above can also be implemented at the upstream shock wave of the aircraft at the same time, that is, the first shock wave angle is continuously increased by the third embodiment, and then the first shock wave angle is periodically or unsteadily increased by the first and / or second embodiments.

[0120] In the process of specific implementation, the active and / or passive flow control device at the downstream shock wave of the aircraft also has multiple different embodiments for reducing the second shock wave angle, which are as follows:

[0121] In the first embodiment, the local high temperature is generated by abnormal glow discharge and laser-induced plasma on the surface of the aircraft at the downstream shock wave, and then the local virtual solid wall is generated at the slope of the second shock wave under the action of the local high temperature, so that the turning angle of the slope of the second shock wave is reduced, thereby reducing the angle of the second shock wave. This active flow control method is flexible, can be turned on and off, and can adjust the control strength at any time. It only needs to provide electric energy and does not need to add an additional magnetic field, which will not cause additional load to the aircraft.

[0122] In the second embodiment, the mixture of free electrons and positive ions is generated by normal glow discharge on the surface of the aircraft at the downstream shock wave, and the Lorentz force is generated by the coupling of the magnetic field and the mixture, which exerts the same volume force as the flow direction on the fluid at the bottom of the boundary layer, locally accelerates the fluid, and then reduces the angle of the second shock wave. This active flow control method is also flexible, can be turned on and off, and can also adjust the control strength at any time. Compared with the local high temperature control, it has stronger control ability.

[0123] In the third embodiment, the adjustable surface structure is arranged as a shock wave control surface on the surface of the aircraft at the downstream shock wave, and the angle of the second shock wave is reduced by adjusting the shape and position of the shock wave control surface.

[0124] It is worth noting that the third embodiment of the active and / or passive flow control device at the downstream shock wave is mainly used in the fixed-length working mode, and the first and second embodiments are mainly used in the fixed-length working mode, the periodic working mode and the unsteady working mode. In specific applications, the aircraft does not necessarily remain in one of the fixed-length working mode, the periodic working mode and the unsteady working mode. The three embodiments described above can also be implemented at the upstream shock wave of the aircraft, that is, the second shock wave angle is continuously reduced by the third embodiment, and then the second shock wave angle is periodically or unsteadily reduced by the first and / or second embodiments.

[0125] As shown in Figure 9 , Figure 10 is a specific application example of the heat reduction method in the embodiment, by which the head shock wave / side wing shock wave interference control technology based on the flow field pressurized plasma high-energy synthetic jet is verified by experiments, and the active heat reduction of the shock wave interference region is realized. The test incoming flow Mach number is 7-9, the incoming flow total temperature is 3000-5000K, the incoming flow total pressure is 6-8MPa, and the incoming flow conditions completely reproduce the real flight flow field conditions of the near space hypersonic vehicle. The results show that the plasma synthetic jet exciter has strong control ability on the shock wave / boundary layer under the high-enthalpy hypersonic real flow field conditions, can weaken or even eliminate the shock wave and the mutual interference of the shock wave, and can reduce the wall average heat flow caused by the shock wave interference by more than 40%.

[0126] The above only describes the preferred embodiments of the present application, and does not limit the patent scope of the present application, and any equivalent structural transformation made under the inventive concept of the present application, or direct / indirect application in other related technical fields is included in the patent protection scope of the present application.

Claims

1. A hypersonic or a supersonic aircraft heat reduction structure, characterized by, The structure includes a first plane of a bottom, a second plane and a third plane of a top, a fourth plane of a tail, and a fifth plane and a sixth plane of a side; The first plane, the second plane, the third plane and the fourth plane are sequentially connected end to end and form a closed ring structure, and the fifth plane and the sixth plane cover both sides of the ring structure; The first plane and the third plane are perpendicular to the fourth plane; The structure further includes a fluidic structure, which includes an air inlet channel, a gas collection cavity, a first fluidic channel and a second fluidic channel arranged inside the structure; the gas collection cavity is located inside the structure and close to the third plane and the fourth plane, the first end of the air inlet channel is located on the second plane, and the second end of the air inlet channel communicates with the gas collection cavity; The first end of the first fluidic channel communicates with the gas collection cavity, and the second end of the first fluidic channel is located on the third plane; the first end of the second fluidic channel communicates with the gas collection cavity, and the second end of the second fluidic channel is located on the fourth plane.

2. The hypersonic or supersonic aircraft heat reduction structure according to claim 1, characterized in that, The fluidic structure further includes a third fluidic channel and a fourth fluidic channel arranged inside the structure; The first end of the third fluidic channel communicates with the gas collection cavity, and the second end of the third fluidic channel is located on the fifth plane; The first end of the fourth fluidic channel communicates with the gas collection cavity, and the second end of the fourth fluidic channel is located on the sixth plane.

3. The hypersonic or supersonic aircraft heat reduction structure according to claim 2, characterized in that, The second end of the third fluidic channel and the second end of the fourth fluidic channel are inclined towards the top and tail of the structure.

4. The hypersonic or supersonic aerocraft heat reduction structure according to any one of claims 1 to 3, characterized in that, The fluidic structure further includes a discharge electrode; The discharge electrode is arranged in the gas collection cavity.

5. A method of heat reduction for a hypersonic or a hypervelocity vehicle, characterized in that, Comprising: The hypersonic aircraft heat reduction structure of any one of claims 1 to 4 is arranged at the upstream shock wave of the aircraft, the angle of the first shock wave at the upstream shock wave is increased by active and / or passive flow control mode, so that the shock wave-shock wave interference mode at the upstream shock wave and the downstream shock wave of the aircraft is changed from the fourth type shock wave-shock wave interference or the fourth r type shock wave-shock wave interference to the fifth type shock wave-shock wave interference or the sixth type shock wave-shock wave interference.

6. The hypersonic vehicle heat reduction method according to claim 5, characterized in that, Further comprising: Active and / or passive flow control devices are arranged at the downstream shock wave of the aircraft to reduce the angle of the second shock wave at the downstream shock wave.

7. The method of claim 6, wherein, When the aircraft wall temperature monitored by the sensor system is 1300K-1500K, the heat reduction method comprises the following steps: Step 101, increasing the angle of attack of the aircraft δθ After the temperature of the aircraft wall surface is increased by 100 degrees, it is determined whether the temperature of the aircraft wall surface is less than 1300 K: If yes, the aircraft is controlled to fly at the current attack angle, and the attack angle of the aircraft is adjusted to the initial value after a preset period of time; Otherwise, step 102 is performed; Step 102, judge whether the current attack angle of the aircraft reaches the attack angle threshold: If yes, step 103 is performed; Otherwise, return to step 101; Step 103, control the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock wave to discharge at high voltage, so that the shock wave-shock wave interference mode at the upstream shock wave and the downstream shock wave of the aircraft is changed from the fourth type shock wave-shock wave interference or the fourth r type shock wave-shock wave interference to the fifth type shock wave-shock wave interference; Step 104: judging whether the temperature of the aircraft wall is less than 1300K after a preset period of time: If yes, turn off the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock wave, and adjust the attack angle of the aircraft to the initial value after a preset period of time; Otherwise, proceed to step 105; Step 105: judging whether the temperature of the aircraft wall exceeds 1500K: If yes, turn on the active and / or passive flow control device at the downstream shock wave, and change the shock-shock interference mode at the upstream and downstream shock waves of the aircraft from the Ⅴ-type shock-shock interference to the Ⅵ-type shock-shock interference until the temperature of the aircraft wall is less than 1300K, then turn off the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock wave and the active and / or passive flow control device at the downstream shock wave, and adjust the attack angle of the aircraft to the initial value after a preset period of time; Otherwise, return to step 104.

8. The method of claim 6, wherein, When the temperature of the aircraft wall monitored by the sensor system exceeds 1500K, the heat reduction method comprises the following steps: Step 201, increasing the angle of attack of the aircraft δθ After the temperature of the wall surface of the aircraft is increased by 100 degrees, it is determined whether the temperature of the wall surface of the aircraft is less than 1300 K: If yes, control the aircraft to keep the current attack angle, and adjust the attack angle of the aircraft to the initial value after a preset period of time; Otherwise, proceed to step 202; Step 202: judging whether the current attack angle of the aircraft reaches the attack angle threshold: If yes, proceed to step 203; Otherwise, return to step 201; Step 203: controlling the high-voltage discharge of the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock wave, and turning on the active and / or passive flow control device at the downstream shock wave, changing the shock-shock interference mode at the upstream and downstream shock waves of the aircraft from the Ⅳ-type shock-shock interference or the Ⅳr-type shock-shock interference to the Ⅵ-type shock-shock interference until the temperature of the aircraft wall is less than 1300K, then turning off the discharge electrode on the hypersonic aircraft heat reduction structure at the upstream shock wave and the active and / or passive flow control device at the downstream shock wave, and adjusting the attack angle of the aircraft to the initial value after a preset period of time.

9. The hypersonic vehicle heat reduction method according to any one of claims 6 to 8, characterized in that, The process of reducing the second shock angle by the active and / or passive flow control device at the downstream shock wave is as follows: By performing abnormal glow discharge and laser-induced plasma on the surface of the aircraft at the downstream shock wave to generate local high temperature, and then generating a local virtual solid wall at the slope of the second shock under the action of the local high temperature, the turning angle of the slope of the second shock is reduced, thereby reducing the angle of the second shock; By performing normal glow discharge on the surface of the aircraft at the downstream shock wave to generate a mixture of free electrons and positive ions, and generating Lorentz force by the coupling of the magnetic field and the mixture, the boundary layer fluid is subjected to the same volume force as the flow direction, and the fluid is locally accelerated, thereby reducing the angle of the second shock; Or By arranging an adjustable surface structure as a shock control surface on the surface of the aircraft at the downstream shock wave, and reducing the angle of the second shock by adjusting the shape and position of the shock control surface.

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