A system and method for transition delay and drag reduction and heat reduction for high speed vehicles
By constructing a self-sustaining flow control loop on a high-speed aircraft using the pressure difference between the transition zone and the turbulent zone, and combining it with a mechanical pressure regulating valve and porous media materials, the coordinated control of transition delay and turbulence drag reduction is achieved. This solves the problems of system complexity and high energy consumption in existing technologies, and improves the thermal protection and endurance performance of the aircraft.
Patent Information
- Application Number
- CN202511799017.2
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-12-02
- Publication Date
- 2026-11-03
AI Technical Summary
Existing transition delay and turbulence control technologies for high-speed aircraft suffer from problems such as system complexity, high energy consumption, poor adaptability, and lack of collaborative design, making it difficult to effectively reduce drag and heat under different flight conditions.
A high-speed aircraft transition delay and drag reduction/heat reduction coordinated control system was designed. It utilizes the pressure difference between the transition zone and the turbulent zone on the aircraft surface and constructs a self-sustaining flow control loop through a mechanical pressure regulating valve to achieve the coordinated effect of air intake control in the transition zone and air blowing control in the turbulent zone. Porous media materials and gradually expanding channels are used for gas delivery to ensure that the system adapts to different flight conditions.
It significantly reduces system complexity and energy consumption without increasing external energy input, improves the payload capacity and endurance of high-speed aircraft, achieves coordinated control of transition delay and turbulence drag reduction, and enhances the reliability and adaptability of thermal protection.
Smart Images
Figure CN123025932A_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of aircraft thermal protection and drag reduction, specifically relating to a coordinated control system and method for transition delay and drag reduction and heat dissipation in high-speed aircraft. Background Technology
[0002] High-speed aircraft with Mach numbers of 5 or higher occupy a crucial position in modern aerospace. With continuous technological advancements, high-speed aircraft are constantly striving for higher speeds, longer flight times, and more complex missions. However, this also leads to extreme thermal loads and strong aerodynamic drag on the surface of high-speed aircraft, posing significant challenges to material selection, structural design, and thermal protection systems. When a high-speed aircraft flies at high Mach numbers within the atmosphere, its surface boundary layer undergoes a transition from laminar to turbulent flow. This transition process has a decisive impact on the aerodynamic thermal environment and drag characteristics of high-speed aircraft. Studies have shown that the wall heat flux and friction coefficient of the turbulent boundary layer can reach 3-5 times that of the laminar state, significantly increasing the burden on thermal protection systems and fuel consumption.
[0003] Currently, drag reduction and heat reduction control technologies for high-speed aircraft are mainly divided into two categories: one is transition delay technology, which expands the laminar flow region with low heat flux by delaying the transition position; the other is turbulence control technology, which reduces the wall heat flux and friction of the transition region through boundary layer blowing and sucking. However, existing technologies generally have the following limitations: Regarding transition delay technologies, while passive control methods (such as wall roughness elements and acoustic metasurfaces) can postpone transition to some extent, their effectiveness is limited and their adaptability is poor. Active air intake control, although more effective, requires an external vacuum pump or ejector to provide intake power, increasing system complexity and weight, consuming additional energy, and making it difficult to sustain during long-duration flights. In terms of turbulence control technologies, boundary layer blowing can effectively reduce wall heat flux in the turbulent region, but it also faces the problem of air supply. Traditional solutions require carrying high-pressure gas cylinders or bleed air from the engine; the former increases the weight of the high-speed aircraft, while the latter affects propulsion system performance. Furthermore, existing technologies often treat transition control and turbulence control as independent problems, lacking a systematic and coordinated design approach. Summary of the Invention
[0004] The technical problem to be solved by this invention is to provide a coordinated control system and method for transition delay and drag reduction / heat reduction in high-speed aircraft. It fully utilizes the pressure differences existing between different regions on the surface of a high-speed aircraft to simultaneously achieve coordinated control of transition delay and drag reduction / heat reduction.
[0005] This invention provides a coordinated control system for transition delay and drag reduction / heat reduction in high-speed aircraft, including a thermal protection structure in the transition zone, a thermal protection structure in the turbulence zone, and a mechanical pressure regulating valve; The transition zone thermal protection structure includes an air intake wall located at the transition zone position on the outer surface of the aircraft and an air intake channel located inside the aircraft and communicating with the air intake wall. The turbulence zone thermal protection structure includes an air outlet wall located at the turbulence zone position on the outer surface of the aircraft and an air outlet channel located inside the aircraft and communicating with the air outlet wall. The intake channel and the outlet channel are interconnected by the mechanical pressure regulating valve, which has a minimum bypass opening and the opening degree of the mechanical pressure regulating valve is proportional to the pressure difference between the intake channel and the outlet channel.
[0006] Furthermore, the mechanical pressure regulating valve is set with an opening pressure threshold. When the pressure difference between the air inlet channel and the air outlet channel reaches the opening pressure threshold, the valve of the mechanical pressure regulating valve opens.
[0007] Furthermore, the mechanical pressure regulating valve is a spring-loaded pressure regulating valve or a diaphragm pressure regulating valve.
[0008] Furthermore, the air intake channel is a gradually expanding channel along the gas flow direction.
[0009] Furthermore, the air intake wall is made of a porous medium material, and the micropores of the porous medium material form the air inlet.
[0010] Furthermore, the porous medium material has a pore size range of 0.1 mm to 0.5 mm and a porosity of 2% to 8%.
[0011] Furthermore, the intake wall is positioned at a flow direction distance corresponding to 0.2 to 0.5 times the transition Reynolds number upstream of the predicted boundary layer transition location.
[0012] Furthermore, the outlet wall is arranged in a region where the boundary layer completely transitions to turbulence.
[0013] Furthermore, the air outlet wall surface is provided with air outlet holes, which adopt a slit-type or discrete hole-type structure.
[0014] This invention also provides a method for coordinated control of transition delay and drag reduction / heat reduction in high-speed aircraft. Using the aforementioned coordinated control system for transition delay and drag reduction / heat reduction in high-speed aircraft, the method includes the following steps: The surface of a high-speed aircraft has pressure gradients in different regions, with lower pressure in the transition zone and relatively higher pressure in the turbulent zone; The intake wall draws some of the gas in the boundary layer of the transition zone into the intake channel. This intake can effectively reduce the boundary layer disturbance growth rate, stabilize the laminar boundary layer structure, and delay the transition position from moving downstream, thereby expanding the range of low heat flux and low friction laminar flow region on the surface of the high-speed aircraft and achieving large-area drag reduction and heat reduction. The intake and exhaust passages are connected by a mechanical pressure regulating valve with a minimum bypass opening. The gas in the intake passage flows to the exhaust passage by pressure difference. The gas entering the outlet channel is injected into the boundary layer of the turbulent zone through the outlet wall, forming a low-temperature, low-velocity protective layer in the near-wall region of the turbulent boundary layer. This effectively blocks the direct contact between the external high-temperature, high-speed airflow and the wall, reduces the heat flux density and frictional resistance coefficient of the wall, and achieves active drag reduction and heat reduction control in the turbulent zone. When the pressure difference between the intake and exhaust channels is lower than the set threshold, it indicates that the flight Mach number is lower than the set value or the altitude is higher than the set altitude. The pressure difference between the transition zone and the turbulent zone is insufficient. The valve maintains the minimum bypass opening, reducing the flow area of the intake and exhaust channels. The dynamic pressure generated by the increased flow velocity compensates for the insufficient static pressure, maintaining the basic operation of the system and preventing gas stagnation in the intake and exhaust channels, thus putting the system in a pre-start state. When the pressure difference between the intake and exhaust channels is higher than the opening pressure threshold, it indicates that the flight Mach number is higher than the set value or the altitude is lower than the set altitude. The pressure difference is sufficient. The mechanical pressure regulating valve gradually opens, increasing the flow area of the intake and exhaust channels, ensuring the transition delay and drag reduction and heat reduction effects, while preventing excessive pressure in the intake and exhaust channels from causing excessive intake rate or structural damage. The opening mechanism of the mechanical pressure regulating valve enables the aircraft to adaptively adjust within a wide speed range or altitude range.
[0015] The beneficial effects of this invention include: 1. Traditional boundary layer active control technologies often rely on external power equipment to provide the driving force for air intake or exhaust. This invention utilizes the pressure difference that objectively exists between the transition zone and the turbulent zone on the surface of a high-speed aircraft to construct a self-driven flow control loop. This design eliminates auxiliary equipment such as vacuum pumps, ejectors, or high-pressure gas cylinders, significantly reducing system complexity, weight, and energy consumption, and significantly improving the payload capacity and endurance of high-speed aircraft.
[0016] 2. Dual passive self-sustaining operation: This system not only utilizes aerodynamic pressure difference to passively drive airflow but also employs a cleverly designed mechanical pressure regulating valve to passively regulate flow. This is a dual passive system that completely eliminates external power sources, sensors, and controllers. This self-sustaining cycle mechanism possesses inherent adaptive characteristics; the pressure difference automatically adjusts according to changes in flight conditions, enabling the control effect to adapt to different Mach numbers and altitudes. This significantly reduces system weight, energy consumption, and structural complexity, and substantially improves system reliability and engineering feasibility.
[0017] 3. This invention achieves coordinated control of transition delay and turbulence drag reduction and heat dissipation. Unlike existing technologies that treat transition control and turbulence control as independent issues, this invention achieves an organic combination and synergistic effect between the two. Boundary layer intake control is implemented in the transition region to stabilize the laminar flow structure and delay the transition point, expanding the low-heat-flux laminar flow region. Simultaneously, the intake gas is guided to the turbulent region for blowing control, forming a cooling protective layer in the turbulent region and further reducing the wall thermal load. This series-based coordinated control strategy, which delays transition through front-end intake and cools with turbulent blowing at the rear, achieves full-process, large-area thermal protection coverage from the transition point to the fully turbulent region. Its drag reduction and heat dissipation effect is far superior to the simple superposition of single control methods.
[0018] 4. This invention fully utilizes the energy and pressure characteristics of high-speed incoming flow, enabling the gas drawn in from the transition zone to not only perform the boundary layer stabilization control function, but also to act as a cooling medium. After being pressurized in the self-sustaining pressurization channel, it plays an active cooling role in the turbulent region, realizing the multiple functions of a single gas medium. Without increasing the additional energy input, the control effect is multiplied through ingenious system integration, reflecting an important development direction of hypersonic high-speed vehicle flow control technology.
[0019] 5. The mechanical pressure regulating valve features a unique minimum bypass opening design, ensuring the system's immediate responsiveness and continuous stability under all flight conditions. This structural minimum opening allows the self-sustaining pressurization channel to maintain a minimal (which can be determined by settings) gas flow under all circumstances, effectively preventing gas stagnation and condensation within the channel, while keeping the system in a pre-start standby state. This design enables the system to respond without delay to the aircraft's instantaneous acceleration from subsonic to hypersonic speeds or sudden changes in flight attitude, completely avoiding thermal protection gaps caused by system restarts or pressurization processes, and greatly improving the thermal protection reliability of high-speed aircraft in complex ballistic and maneuvering flight. In addition, this minimum bypass flow also provides continuous system self-cleaning, effectively blowing away trace impurities or condensates that may accumulate in the self-sustaining pressurization channel or mechanical pressure regulating valve, ensuring the unobstructed flow and functional integrity of the channels under long-term service conditions.
[0020] 6. The opening of the mechanical pressure regulating valve is proportional to the pressure difference between the intake and exhaust channels, achieving precise and adaptive matching between the system flow supply and the aerodynamic and thermal environment requirements of the aircraft. The proportional control characteristic of the mechanical pressure regulating valve allows the mass flow rate of intake and exhaust to automatically increase with the increase of the flight Mach number (i.e., external aerodynamic load). At high Mach numbers, it provides stronger transition delay control and more sufficient cooling in the turbulent zone; at low Mach numbers, it automatically reduces the flow rate to avoid premature transition caused by excessive intake, thus achieving optimal performance over a wide speed range. It achieves on-demand regulation and, without introducing any active control or external energy, achieves dynamically optimal transition delay and drag reduction / heat reduction effects that are matched to the flight status in real time, raising the level of automation and adaptability of the passive control system to a new height. Increasing the opening of the mechanical pressure regulating valve also prevents excessive pressure in the intake and exhaust channels from causing excessive intake rate or structural damage. More importantly, this opening characteristic, which is proportional to the pressure difference, essentially converts complex flight state parameters such as flight Mach number and altitude directly and reliably into precise control commands for system flow through the intermediate variable of aerodynamic pressure difference. It is not only equivalent in effect to a controller that requires no external energy source, but also, from a physical perspective, ensures that the system's workload remains synchronized with the severity of the aerodynamic and thermal loads faced by the aircraft, fundamentally avoiding problems of under-control or over-control, and achieving a balance between system performance and flight safety. This physically-based adaptive mechanism is the theoretical foundation for the system's ability to maintain efficient and robust operation across a wide speed range.
[0021] The aforementioned high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system is applicable to various high-speed aircraft, including cruise missiles, high-speed flight test platforms, reusable launch vehicles, and space shuttle flights. This system provides effective technical support for improving the aerodynamic performance, enhancing thermal protection, and improving the endurance of these high-speed aircraft under high-speed flight conditions, and has broad engineering application prospects. Attached Figure Description
[0022] Figure 1 This is a schematic diagram of the structure of the present invention.
[0023] In the diagram, 1-Transition zone thermal protection structure; 2-Intake wall; 3-Outtake wall; 4-Turbulent zone thermal protection structure; 5-Self-sustaining pressurization channel; 51-Intake channel; 52-Outtake channel; 6-Mechanical pressure regulating valve; 7-Schematic diagram of gas flow direction. Detailed Implementation
[0024] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.
[0025] In this invention, unless otherwise explicitly specified and limited, the terms "connection," "fixed," etc., should be interpreted broadly. For example, "fixed" can mean a fixed connection, a detachable connection, or an integral part; it can mean a mechanical connection, an electrical connection, a physical connection, or a wireless communication connection; it can mean a direct connection or an indirect connection through an intermediate medium; it can mean the internal communication of two elements or the interaction between two elements, unless otherwise explicitly limited. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.
[0026] Furthermore, the technical solutions of the various embodiments of the present invention can be combined with each other, but only if they are feasible for those skilled in the art. If the combination of technical solutions is contradictory or cannot be implemented, it should be considered that such combination of technical solutions does not exist and is not within the scope of protection claimed by the present invention.
[0027] like Figure 1 As shown, the present invention provides a high-speed aircraft transition delay and drag reduction and heat reduction coordinated control system, including a transition zone thermal protection structure 1, a turbulence zone thermal protection structure 4 and a mechanical pressure regulating valve 6; The transition zone thermal protection structure 1 includes an air intake wall 2 located at the transition zone position on the outer surface of the aircraft and an air intake channel 51 located inside the aircraft and communicating with the air intake wall 2. The turbulence zone thermal protection structure 4 includes an air outlet wall 3 located at the turbulence zone position on the outer surface of the aircraft and an air outlet channel 52 located inside the aircraft and communicating with the air outlet wall 3. The intake channel 51 and the outlet channel 52 are interconnected by the mechanical pressure regulating valve 6, which has a minimum bypass opening and whose opening is proportional to the pressure difference between the intake channel 51 and the outlet channel 52. The intake channel 51, the mechanical pressure regulating valve 6, and the outlet channel 52 form a self-sustaining pressurization channel 5. Specifically, for pressurization: high-speed airflow (high kinetic energy) rushes into the intake channel 51 through the suction wall 2, colliding with the mechanical pressure regulating valve 6 and the gas already "slowing down and queuing" in front of it, forcing the airflow to decelerate drastically. The airflow does not flow in a free-flowing vacuum. It is drawn into a closed system already filled with gas, with a bottleneck downstream (at the mechanical pressure regulating valve 6). Subsequent airflow continuously surges in (rams), while the airflow ahead is obstructed by the mechanical pressure regulating valve 6 (throttling), inevitably leading to gas "accumulation" and "blockage" in the intake channel 51. The macroscopic manifestation of this "congestion" is a significant decrease in gas velocity and a significant increase in static pressure, which in turn enables the airflow to achieve self-sustaining pressurization.
[0028] In one embodiment, the mechanical pressure regulating valve 6 is set with an opening pressure threshold. When the pressure difference between the air intake channel 51 and the air outlet channel 52 reaches the opening pressure threshold, the valve of the mechanical pressure regulating valve 6 opens.
[0029] The beneficial effects of this invention include: 1. Traditional boundary layer active control technologies often rely on external power equipment to provide the driving force for air intake or exhaust. This invention utilizes the pressure difference that objectively exists between the transition zone and the turbulent zone on the surface of a high-speed aircraft to construct a self-driven flow control loop. This design eliminates auxiliary equipment such as vacuum pumps, ejectors, or high-pressure gas cylinders, significantly reducing system complexity, weight, and energy consumption, and significantly improving the payload capacity and endurance of high-speed aircraft.
[0030] 2. Dual passive self-sustaining operation: This system not only utilizes aerodynamic pressure difference to passively drive airflow, but also employs a cleverly designed mechanical pressure regulating valve 6 to passively regulate flow. This is a dual passive system that completely eliminates external power sources, sensors, and controllers. This self-sustaining cycle mechanism possesses inherent adaptive characteristics; the pressure difference automatically adjusts according to changes in flight conditions, enabling the control effect to adapt to different Mach numbers and altitudes. This significantly reduces the system's weight, energy consumption, and structural complexity, and substantially improves system reliability and engineering feasibility.
[0031] 3. This invention achieves coordinated control of transition delay and turbulence drag reduction and heat dissipation. Unlike existing technologies that treat transition control and turbulence control as independent issues, this invention achieves an organic combination and synergistic effect between the two. Boundary layer intake control is implemented in the transition region to stabilize the laminar flow structure and delay the transition point, expanding the low-heat-flux laminar flow region. Simultaneously, the intake gas is guided to the turbulent region for blowing control, forming a cooling protective layer in the turbulent region and further reducing the wall thermal load. This series-based coordinated control strategy, which delays transition through front-end intake and cools with turbulent blowing at the rear, achieves full-process, large-area thermal protection coverage from the transition point to the fully turbulent region. Its drag reduction and heat dissipation effect is far superior to the simple superposition of single control methods.
[0032] 4. This invention fully utilizes the energy and pressure characteristics of high-speed incoming flow, enabling the gas drawn in from the transition zone to not only perform the boundary layer stabilization control function, but also to act as a cooling medium. After being pressurized in the self-sustaining pressurization channel 5, it plays an active cooling role in the turbulent region, realizing the multiple functions of a single gas medium. Without increasing the additional energy input, the control effect is multiplied through ingenious system integration, reflecting an important development direction of hypersonic high-speed vehicle flow control technology.
[0033] 5. The mechanical pressure regulating valve 6 features a unique minimum bypass opening design, ensuring the system's immediate responsiveness and continuous stability under all flight conditions. This structural minimum opening allows the self-sustaining pressurization channel 5 to maintain a minimal (which can be determined by settings) gas flow under all circumstances, effectively preventing gas stagnation and condensation within the channel, while keeping the system in a pre-start standby state. This design enables the system to respond without delay to the instantaneous acceleration or sudden changes in flight attitude of the aircraft from subsonic to hypersonic speeds, completely avoiding thermal protection gaps caused by system restarts or pressurization processes, and greatly improving the thermal protection reliability of high-speed aircraft in complex ballistic and maneuvering flight. In addition, this minimum bypass flow also provides continuous system self-cleaning, effectively blowing away any trace impurities or condensates that may accumulate in the self-sustaining pressurization channel 5 or the mechanical pressure regulating valve 6, ensuring the unobstructed flow and functional integrity of the channels under long-term service conditions.
[0034] 6. The opening degree of the mechanical pressure regulating valve 6 is proportional to the pressure difference between the intake channel 51 and the exhaust channel 52, achieving precise and adaptive matching between the system flow supply and the aerodynamic and thermal environment requirements of the aircraft. The proportional control characteristic of the mechanical pressure regulating valve 6 allows the mass flow rate of intake and exhaust to automatically increase with the increase of the flight Mach number (i.e., external aerodynamic load), providing stronger transition delay control and more sufficient turbulent cooling under high Mach number conditions; under low Mach number conditions, it automatically reduces the flow rate to avoid premature transition caused by excessive intake, thereby achieving performance optimization over a wide speed range. It achieves on-demand regulation and, without introducing any active control or external energy, achieves dynamically optimal transition delay and drag reduction / heat reduction effects that are matched to the flight state in real time, raising the level of automation and adaptability of the passive control system to a new height. Increasing the opening degree of the mechanical pressure regulating valve 6 also prevents excessive pressure in the intake channel 51 and the exhaust channel 52 from causing excessive intake rate or structural damage. More importantly, this opening characteristic, which is proportional to the pressure difference, essentially converts complex flight state parameters such as flight Mach number and altitude directly and reliably into precise control commands for system flow through the intermediate variable of aerodynamic pressure difference. It is not only equivalent in effect to a controller that requires no external energy source, but also, from a physical perspective, ensures that the system's workload remains synchronized with the severity of the aerodynamic and thermal loads faced by the aircraft, fundamentally avoiding problems of under-control or over-control, and achieving a balance between system performance and flight safety. This physically-based adaptive mechanism is the theoretical foundation for the system's ability to maintain efficient and robust operation across a wide speed range.
[0035] The aforementioned high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system is applicable to various high-speed aircraft, including cruise missiles, high-speed flight test platforms, reusable launch vehicles, and space shuttle flights. This system provides effective technical support for improving the aerodynamic performance, enhancing thermal protection, and improving the endurance of these high-speed aircraft under high-speed flight conditions, and has broad engineering application prospects.
[0036] In one embodiment, the mechanical pressure regulating valve 6 is either a spring-loaded pressure regulating valve or a diaphragm pressure regulating valve. Spring-loaded pressure regulating valves have a simple and robust structure, excellent high-temperature resistance, and long-term stable operation. Diaphragm valves, on the other hand, have higher pressure sensing sensitivity and can respond quickly to small pressure differential changes. The specific choice can be made according to requirements.
[0037] In one embodiment, the intake channel 51 is a gradually expanding channel along the gas flow direction. This configuration improves the pressurization effect of the self-sustaining pressurization channel 5. Specifically, when a high-kinetic-energy airflow enters the intake channel 51 from the intake wall 2, the gradually expanding shape of the intake channel 51 guides the airflow to decelerate. The kinetic energy lost by the airflow is converted into internal energy and pressure energy, resulting in a significant increase in its static pressure.
[0038] In one embodiment, the air intake wall 2 is made of a porous medium material, and the micropores of the porous medium material form air inlets.
[0039] In this embodiment, the densely distributed micropores within the porous medium create an extremely uniform and stable surface-type intake across the entire intake region, completely avoiding the local flow inhomogeneity problems that may arise from traditional discrete-pore intake methods. This more effectively suppresses boundary layer disturbances and maximizes the transition delay effect. Secondly, the porous medium, made of high-temperature alloys or ceramic matrix composites, is itself part of the thermal protection structure, achieving a high degree of integration between function and structure. This integrated design not only simplifies the system structure and reduces assembly interfaces, but its internal microporous structure also significantly increases the heat exchange area. When external high-temperature airflow passes through, the porous medium can efficiently diffuse and dissipate heat through the solid framework, effectively providing an additional cooling layer for the transition zone and further reducing the heat flow entering the aircraft. Furthermore, the tortuous flow channels within the porous medium naturally homogenize and filter the intake gas, effectively smoothing intake pressure fluctuations caused by external flow pulsations. This provides more stable inlet conditions for the downstream self-sustaining pressurization channel 5 and the mechanical pressure regulating valve 6, improving the overall system stability. This porous media-based intake wall design integrates flow control, thermal protection, and structural functionality, representing a key technological approach to improving the overall performance of the system.
[0040] In one embodiment, the porous medium material has a pore size range of 0.1 mm to 0.5 mm and a porosity of 2% to 8%.
[0041] In this embodiment, the pore size range effectively suppresses disturbance waves of different scales within the boundary layer. Micropores of 0.1 mm stabilize high-frequency disturbances, while larger pore sizes of 0.3 mm to 0.5 mm effectively control low-frequency disturbances, forming a transition suppression effect across the entire frequency band. Secondly, the 2%-8% porosity design ensures sufficient air intake while maintaining the material's structural integrity and impact resistance, avoiding strength reduction due to excessive porosity. In particular, this parameter combination exhibits unique advantages in high-temperature environments: the appropriately dense pore distribution (porosity ≤ 8%) significantly increases the heat exchange area, enabling the porous media layer to function as a micro-heat exchanger, achieving efficient heat dissipation through heat conduction via the solid framework and convective heat transfer within the pores.
[0042] In one embodiment, the intake wall 2 is located at a flow direction distance corresponding to 0.2 to 0.5 times the transition Reynolds number upstream of the predicted boundary layer transition location.
[0043] In this embodiment, the suction wall 2 is positioned 0.2 to 0.5 times the transition Reynolds number upstream of the predicted boundary layer transition location. This distance range ensures that the suction effect intervenes at its most significant location, effectively suppressing disturbances before they amplify and cause transition. Positioning it at 0.2 times the distance allows the system to achieve effective flow control with minimal necessary suction, resulting in higher economic efficiency; while positioning it at 0.5 times the distance provides more development space for boundary layer adjustment, enhancing the system's robustness against different disturbance types. This positioning allows the system to establish a stable laminar velocity profile during the critical phase before transition, maximizing the effect of suction delaying transition while avoiding ineffective suction or structural space conflicts caused by positioning it too far forward.
[0044] In one embodiment, the outlet wall 3 is positioned in the region where the boundary layer completely transitions to turbulence. In this embodiment, firstly, the blowing action is precisely applied to the turbulence initiation region where surface heat flux and frictional resistance increase sharply, enabling the timely formation of a near-wall cryogenic protective layer at the location of the most severe thermal load, maximizing the thermal barrier effect. Secondly, the cooling gas injected at this location effectively intervenes in the development of the turbulent boundary layer's bottom layer, achieving both drag reduction and heat reduction. Specifically, this arrangement perfectly complements the upstream intake delay transition control; the further upstream the intake successfully delays the transition, the larger the downstream protected turbulent region becomes, with both working together to form a comprehensive "front-end suppression, rear-end control" thermal protection system. Furthermore, placing the blowing outlet in the transition region avoids potential interference from the blowing jet to the upstream laminar flow stability, ensuring the transition delay effect remains unaffected. This optimized location selection allows limited cooling gas to exert maximum effectiveness in the most needed area, significantly improving the overall protection efficiency of the entire system.
[0045] In one embodiment, the exhaust wall 3 is provided with exhaust holes, which are either slit-type or discrete-hole type structures. In this embodiment, the slit-type outlet is a slit-type outlet, which can form a continuous and uniform cooling gas film in the spanwise direction. Its gas film protection range is wide and seamless, making it particularly suitable for large, flat areas with relatively uniform heat flow distribution. The discrete-hole type outlet has higher structural strength and adaptability to complex curved surfaces. The multiple independent cooling jets generated by it merge with each other as they develop downstream, eventually forming a covering layer. This arrangement is highly flexible and facilitates avoidance arrangement around structural stiffeners or other internal components. The specific arrangement structure can be selected according to actual needs.
[0046] In one embodiment, the cross-sectional area of the air intake channel 51 is 8%-15% of the total area of the air intake wall 2.
[0047] In this embodiment, this ratio range ensures that the intake airflow has a moderate velocity when entering the channel: if the cross-sectional area is too small (below 8%), the inlet velocity will be too high, which will not only significantly increase flow losses and reduce pressure recovery efficiency, but may also induce new flow instability factors near the porous wall due to excessively high local intake velocity, which will have an adverse effect on transition delay; if the cross-sectional area is too large (above 15%), although it can reduce velocity and losses, it will unnecessarily increase the structural volume and weight of the channel, which violates the lightweight design requirements of high-speed aircraft. The cross-sectional area ratio of 8%-15% is just right to enable the intake airflow to complete a smooth turning and acceleration from the wall normal flow to the axial flow in the channel when entering the channel, creating the optimal initial flow conditions for subsequent pressure recovery in the self-sustaining pressurization channel 5 and effective delivery to the turbulent region.
[0048] The present invention also provides a specific embodiment, see reference. Figure 1 , Figure 1This is the overall structural layout of this embodiment. The system is mainly applied to high-speed aircraft, utilizing the pressure difference between the transition zone and turbulent zone on the surface of the high-speed aircraft to achieve self-sustaining flow control without an external power source. A high-speed incoming flow (Ma≥5) flows from the left side towards the surface of the high-speed aircraft. Here, 1 is the thermal protection structure for the transition zone, and 2 is the intake wall. The intake wall 2 is located 0.2 to 0.5 times the transition Reynolds number upstream of the predicted boundary layer transition location, corresponding to a flow direction distance. This intake wall 2 is made of a porous medium material processed by laser sintering or additive manufacturing processes, with a pore size range of 0.1 mm to 0.5 mm and a porosity of 2% to 8%. The pores are distributed in a uniform or gradually varying pattern. When high-speed airflow passes through this region, due to the relatively low pressure in the transition zone (typically 1.5 to 3 times the incoming static pressure), some gas within the boundary layer is naturally drawn into the porous medium material. The ratio of the inhaled mass flow rate to the incoming mass flow rate (inhalation rate) is controlled within the range of 0.1% to 1.0%. This micro-inhalation can significantly reduce the growth rate of disturbance waves within the boundary layer, stabilize the laminar flow structure, and delay the transition position downstream by 10% to 30% of the characteristic length of the high-speed aircraft. 3 represents the exhaust wall, located in the region where the boundary layer completely transitions to turbulence. The outlet of exhaust wall 3 adopts a slit-type or discrete-hole structure. The slit width of the slit-type outlet is 0.5 mm to 2 mm, and the aperture of the discrete-hole outlet is 1 mm to 3 mm, with a hole spacing of 5 to 15 times the aperture, arranged in a staggered or linear pattern. 4 represents the thermal protection structure for the turbulent region. 5 is a self-sustaining pressurization channel, including an interconnected intake channel 51 and an exhaust channel 52. The design of this channel needs to comprehensively consider factors such as pressure recovery, flow loss, and structural strength. The inlet cross-sectional area of the self-sustaining pressurization channel 5 is determined based on the expected intake flow rate and is 8%-15% of the total area of the intake wall 2. The channel adopts a gradually expanding or constant cross-section design along its length to achieve gradual pressure recovery and a moderate reduction in flow velocity. The length of the self-sustaining pressurization channel 5 depends on the distance between the transition zone and the turbulent zone, and is 0.3-0.8 times the characteristic length of the high-speed aircraft. The inner wall surface of the self-sustaining pressurization channel 5 is smoothed to reduce frictional losses along its length. 6 is a mechanical pressure regulating valve. This mechanical pressure regulating valve 6 is located in the middle and rear section of the self-sustaining pressurization channel 5 and adopts a spring-loaded or diaphragm-type pressure regulating valve, which can automatically adjust without external power or control signals. The working principle of the pressure valve is as follows: When the pressure difference between the intake channel 51 and the outlet channel 52 is lower than the opening pressure threshold (indicating that the flight Mach number is lower than the set value or the altitude is higher than the set altitude, and the pressure difference between the transition zone and the turbulent zone is insufficient), the mechanical pressure regulating valve 6 maintains a minimum bypass opening, reducing the flow area in the channel. The increased flow velocity generates dynamic pressure to compensate for insufficient static pressure, maintaining the basic operation of the system. When the pressure in the self-sustaining booster channel 5 is higher than the set threshold (indicating that the flight Mach number is higher than the set value or the altitude is lower than the set altitude, and the pressure difference is sufficient), the valve gradually opens wider, increasing the flow area.This design achieves enhanced transition delay control and more thorough turbulent cooling at high Mach numbers, while automatically reducing flow rate at low Mach numbers to prevent premature transition due to excessive intake. This optimizes performance across a wide speed range and prevents excessive pressure in the self-sustaining booster channel 5 at high Mach numbers, which could lead to excessive intake rate or structural damage. Taking the mechanical pressure regulating valve 6 as an example (a spring-loaded type), the spring preload of the valve is calibrated to correspond to a preset opening pressure threshold. In supersonic design conditions (Ma=5-10), the large pressure difference ΔP overcomes the spring preload, pushing the piston and opening the valve. Furthermore, the larger ΔP is, the larger the valve opening and the larger the flow cross-sectional area of the self-sustaining booster channel 5. This allows the mass flow rate of intake and exhaust to increase with the flight Mach number (i.e., pressure difference), achieving fully automatic, externally energy-free adaptive proportional control. At low Mach numbers or under non-design conditions: When the aircraft is in subsonic or transonic speeds or Ma < 5, the pressure difference ΔP is insufficient, or even pressure reversal may occur in some conditions. In this case, the mechanical pressure regulating valve 6 remains nearly closed under spring force. The piston or seat of this mechanical pressure regulating valve 6 is designed with a structurally minimal bypass opening. This means that even when the valve is pressed to its "most closed" state by spring force, the self-sustaining pressurization channel 5 is not completely blocked, always allowing a very small amount of fluid to pass through. This minimal bypass airflow prevents gas stagnation within the self-sustaining pressurization channel 5, keeping the system in a pre-start state at all times. When the aircraft instantaneously accelerates to a high Mach number, the system can be activated instantaneously without delay, which is crucial for hypersonic aircraft with drastic changes in flight attitude. 7 shows a schematic diagram of the gas flow direction, representing the airflow from the transition zone to the turbulent zone.
[0049] The air intake wall 2 and the wall of the air intake channel 51 of the transition zone thermal protection structure 1 form a multi-layer composite structure. The surface layer, i.e., the air intake wall 2, is made of porous medium material, while the wall of the air intake channel 51 is made of heat-resistant material with a thickness of 5mm-15mm. It has low thermal conductivity and high temperature stability.
[0050] The exhaust wall 3 and exhaust channel 52 of the thermal protection structure 4 in the turbulent region are made of heat-resistant material, which has low thermal conductivity and high temperature stability.
[0051] This invention also provides a method for coordinated control of transition delay and drag reduction / heat reduction in high-speed aircraft. Using the aforementioned coordinated control system for transition delay and drag reduction / heat reduction in high-speed aircraft, the method includes the following steps: The surface of a high-speed aircraft has pressure gradients in different regions, with lower pressure in the transition zone and relatively higher pressure in the turbulent zone; The intake wall 2 draws some of the gas in the boundary layer of the transition zone into the intake channel 51. This intake can effectively reduce the boundary layer disturbance growth rate, stabilize the laminar boundary layer structure, and delay the transition position from moving downstream, thereby expanding the range of low heat flux and low friction laminar flow region on the surface of the high-speed aircraft and achieving large-area drag reduction and heat reduction. The intake passage 51 and the exhaust passage 52 are connected by a mechanical pressure regulating valve 6 with a minimum bypass opening. The gas in the intake passage 51 flows to the exhaust passage 52 by pressure difference. The gas entering the outlet channel 52 is injected into the boundary layer of the turbulent region through the outlet wall 3, forming a low-temperature, low-velocity protective layer in the near-wall region of the turbulent boundary layer. This effectively blocks the direct contact between the external high-temperature, high-speed airflow and the wall, reduces the heat flux density and frictional resistance coefficient of the wall, and achieves active drag reduction and heat reduction control in the turbulent region. When the pressure difference between the intake channel 51 and the exhaust channel 52 is lower than the set threshold, it indicates that the flight Mach number is lower than the set value or the altitude is higher than the set altitude. The pressure difference between the transition zone and the turbulent zone is insufficient. The valve maintains the minimum bypass opening, reducing the flow area of the intake channel 51 and the exhaust channel 52. The dynamic pressure generated by the increased flow velocity compensates for the insufficient static pressure, maintaining the basic operation of the system and preventing gas stagnation in the intake channel 51 and the exhaust channel 52, thus putting the system in a pre-start state. When the pressure difference between the intake channel 51 and the exhaust channel 52 is higher than the opening pressure threshold, it indicates that the flight Mach number is higher than the set value or the altitude is lower than the set altitude. The pressure difference is sufficient, and the mechanical pressure regulating valve 6 gradually opens, increasing the flow area of the intake channel 51 and the exhaust channel 52. This ensures the transition delay and drag reduction and heat reduction effects, while preventing excessive pressure in the intake channel 51 and the exhaust channel 52 from causing excessive intake rate or structural damage. The opening mechanism of the mechanical pressure regulating valve 6 enables the aircraft to adaptively adjust within a wide speed range of flight speed or flight altitude.
[0052] The above description is merely an embodiment and does not constitute any limitation on the present invention. Any person skilled in the art can make many possible variations, modifications, or alterations to the technical solutions of the present invention without departing from the scope of the present invention. Therefore, any simple modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the scope of the present invention, should fall within the protection scope of the present invention.
Claims
1. A coordinated control system for transition delay and drag reduction / heat reduction in high-speed aircraft, characterized in that, It includes a transition zone thermal protection structure (1), a turbulent zone thermal protection structure (4), and a mechanical pressure regulating valve (6); The transition zone thermal protection structure (1) includes an air intake wall (2) located at the transition zone position on the outer surface of the aircraft and an air intake channel (51) connected to the air intake wall (2) inside the aircraft. The turbulence zone thermal protection structure (4) includes an air outlet wall (3) located at the turbulence zone position on the outer surface of the aircraft and an air outlet channel (52) connected to the air outlet wall (3) inside the aircraft. The intake passage (51) and the outlet passage (52) are interconnected by the mechanical pressure regulating valve (6), which has a minimum bypass opening and the opening of the mechanical pressure regulating valve (6) is proportional to the pressure difference between the intake passage (51) and the outlet passage (52).
2. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in claim 1, characterized in that, The mechanical pressure regulating valve (6) is set with an opening pressure threshold. When the pressure difference between the air inlet channel (51) and the air outlet channel (52) reaches the opening pressure threshold, the valve of the mechanical pressure regulating valve (6) opens.
3. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in claim 2, characterized in that, The mechanical pressure regulating valve (6) is a spring-loaded pressure regulating valve or a diaphragm pressure regulating valve.
4. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in claim 1, characterized in that, The air intake channel (51) is a gradually expanding channel along the gas flow direction.
5. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in any one of claims 1-4, characterized in that, The air intake wall (2) is made of porous medium material, and the micropores of the porous medium material form the air inlet.
6. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in claim 5, characterized in that, The porous medium material has a pore size range of 0.1 mm to 0.5 mm and a porosity of 2% to 8%.
7. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in any one of claims 1-4 and 6, characterized in that, The air intake wall (2) is located at a flow direction distance corresponding to 0.2 to 0.5 times the transition Reynolds number upstream of the predicted boundary layer transition location.
8. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in any one of claims 1-4, characterized in that, The outlet wall (3) is arranged in the region where the boundary layer is completely transitioned to turbulence.
9. The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in any one of claims 1-4, characterized in that, The air outlet wall (3) is provided with an air outlet hole, which adopts a slit-type or discrete hole-type structure.
10. A method for coordinated control of transition delay and drag reduction / heat reduction in high-speed aircraft, characterized in that, The high-speed aircraft transition delay and drag reduction / heat reduction coordinated control system as described in any one of claims 1-9 includes the following steps: The surface of a high-speed aircraft has pressure gradients in different regions, with lower pressure in the transition zone and relatively higher pressure in the turbulent zone; The intake wall (2) draws some of the gas in the boundary layer of the transition zone into the intake channel (51). This intake can effectively reduce the boundary layer disturbance growth rate, stabilize the laminar boundary layer structure, and delay the transition position from moving downstream, thereby expanding the range of low heat flux and low friction laminar flow region on the surface of the high-speed aircraft and achieving large-area drag reduction and heat reduction. The intake passage (51) and the outlet passage (52) are connected by a mechanical pressure regulating valve (6) with a minimum bypass opening. The gas in the intake passage (51) flows to the outlet passage (52) by pressure difference. The gas entering the outlet channel (52) is injected into the boundary layer of the turbulent zone through the outlet wall (3), forming a low-temperature and low-speed protective layer in the near-wall region of the turbulent boundary layer. This effectively blocks the direct contact between the external high-temperature and high-speed airflow and the wall, reduces the heat flux density and friction resistance coefficient of the wall, and achieves active drag reduction and heat reduction control in the turbulent zone. When the pressure difference between the intake channel (51) and the exhaust channel (52) is lower than the set threshold, it indicates that the flight Mach number is lower than the set value or the altitude is higher than the set altitude. The pressure difference between the transition zone and the turbulent zone is insufficient. The valve maintains the minimum bypass opening, reduces the flow area of the intake channel (51) and the exhaust channel (52), and uses the dynamic pressure generated by the increased flow rate to compensate for the insufficient static pressure, maintains the basic operation of the system, prevents the gas in the intake channel (51) and the exhaust channel (52) from stagnating, and puts the system in a pre-start state. When the pressure difference between the intake channel (51) and the exhaust channel (52) is higher than the opening pressure threshold, it indicates that the flight Mach number is higher than the set value or the altitude is lower than the set altitude. The pressure difference is sufficient, and the valve of the mechanical pressure regulating valve (6) gradually opens, increases the flow area of the intake channel (51) and the exhaust channel (52), ensures the effect of transition delay and drag reduction and heat reduction, and at the same time prevents the pressure in the intake channel (51) and the exhaust channel (52) from being too high, resulting in excessive intake rate or structural damage. The opening mechanism of the mechanical pressure regulating valve (6) enables the aircraft to adaptively adjust within a wide speed range of flight speed or flight altitude.