Hypersonic velocity blunt body boundary layer delayed transition control device and method

Abstract

The invention provides a control device and method for delayed transition of a boundary layer of a hypersonic blunt body. The device comprises a blunt part shell, a continuous gap structure, a low-pressure cavity, a suction pump, a compressor and a control unit, continuous gap structures are uniformly distributed on the head part of the blunt head part shell; the continuous gap structure is designed into a narrow gap structure which is continuous in the spanwise direction and has no spanwise periodic distribution characteristic; the low-pressure cavity is used for collecting sucked high-pressure gas; the suction pump and the compressor are arranged in the blunt head part shell; the suction pump is connected with the low-pressure cavity through a pipeline; the compressor is positioned at the downstream of the suction pump and is used for further compressing the sucked gas; the control unit is electrically connected with the suction pump and the compressor. The method is reasonable in conception, and weakens the sensibility of non-modal disturbance in a near stationary point area and subsequent transient growth from the source through a targeted wall surface control strategy, so that boundary layer transition is obviously delayed in a relatively large blunt degree range, and the surface friction resistance and aerodynamic thermal load of an aircraft are reduced.

Description

Technical Field

[0001] This invention relates to the field of hypersonic vehicle technology, specifically to a hypersonic blunt-nosed body boundary layer delay transition control device and method. Background Technology

[0002] Hypersonic vehicles typically face severe aerodynamic heating and frictional drag during flight, and the transition of the boundary layer from laminar to turbulent flow leads to a sharp increase in wall heat flux and drag. Therefore, how to delay the hypersonic boundary layer transition is one of the key technical issues in vehicle aerodynamic design. Generally, when external disturbances are small, boundary layer transition is mainly dominated by modal instability (such as the high-frequency Mack second mode wave), which is called natural transition; however, under strong external disturbances, the boundary layer may bypass these modal paths and instead transition rapidly through the transient growth of non-modal disturbances (composed of non-orthogonal modes), which is called bypass transition. Furthermore, under the same external disturbances, the amplitude of non-modal disturbances increases with the increase of leading-edge bluntness. Nonmodal disturbances manifest as elongated high-speed and low-speed stripes within the boundary layer. Although their initial amplitude is small, they can be significantly amplified through transient mechanisms, inducing subsequent secondary instabilities and ultimately leading to turbulence. In the boundary layer of a hypersonic blunt-nosed body, the presence of strong shock waves and an entropy layer complicates the transition process: the bow-shaped shock wave at the blunt-nosed leading edge causes free-flow disturbances to generate acoustic, entropy, and vortex components behind the shock wave. These disturbances penetrate the entropy layer and enter the boundary layer, exciting various modal and nonmodal instabilities. Moderate leading-edge blunting (increasing the blunt-nosed radius) can delay transition by thickening the boundary layer and reducing secondary mode instabilities; however, when the bluntness increases to a certain threshold, transition occurs earlier, exhibiting a "transition reversal" phenomenon. Research indicates that this is because the entropy layer generated by the blunt-nosed body weakens high-frequency modes while enhancing the disturbance sensitivity in the near-stagnation region, triggering the rapid growth of nonmodal disturbances. Therefore, how to suppress non-modal perturbations in the boundary layer of hypersonic blunt bodies has become an urgent technical problem to be solved.

[0003] To delay hypersonic boundary layer transition, existing research mainly focuses on suppressing the evolution of modal perturbations during the linear phase of natural transition. For example, for instability in the second mode, methods such as wall cooling and passive porous coatings for sound absorption have been used to reduce high-frequency perturbation energy. However, since the flow-oriented scale of non-modal perturbations is much longer than that of linear instability modes, commonly used control methods for natural transition are not applicable to bypass transition. Previous studies have attempted to apply spanwise vibrations to the wall to disrupt the strip structure, but this is complex to implement and suffers from parameter sensitivity issues, making it difficult to apply to practical engineering problems. Direct wall suction or blowing is a more direct method to influence boundary layer perturbations: classic steady-state suction techniques have long been used in laminar flow control at low speeds to remove low-energy fluid from the boundary layer, thereby delaying transition; appropriate wall blowing or suction has also been shown to influence instability growth and delay transition under certain conditions in hypersonic situations. However, existing technologies do not clearly define how to design efficient control schemes for the non-modal stripes of hypersonic blunt bodies. For example, wall blowing may have negative effects at different locations, enhancing sensitivity or inducing new unstable structures, while simple downstream suction often results in limited efficiency due to improper parameter settings. In summary, existing technologies lack a mechanism-based, targeted wall control method to suppress the formation and amplification of the aforementioned non-modal stripe disturbances, resulting in a lack of effective measures to delay the boundary layer transition of hypersonic blunt bodies. Summary of the Invention

[0004] To address the shortcomings of existing technologies in controlling boundary layer transition in hypersonic blunt-nosed bodies, particularly the inability to effectively address the challenge of nonmodal stripe growth induced by the interaction of entropy layers and shock waves in the "transition reversal" phenomenon, this invention proposes a boundary layer delay transition control device and method for hypersonic blunt-nosed bodies. Through targeted wall control strategies, it weakens the sensitivity to nonmodal disturbances in the near-stagnation region and their subsequent transient growth at the source, thereby significantly delaying boundary layer transition over a large bluntness range and reducing the surface friction drag and aerodynamic thermal load of the aircraft.

[0005] To solve the above-mentioned technical problems, the present invention provides a hypersonic blunt-nosed body boundary layer delay transition control device, which includes a blunt-nosed body shell, a continuous slit structure, a low-pressure chamber, a suction pump, a compressor and a control unit;

[0006] The blunt-headed shell has an accommodating space inside, and its head is a stationary area with the continuous slit structure evenly distributed in a matching manner.

[0007] The continuous slot structure is designed as a narrow slot structure that is continuous in the spanwise direction and has no spanwise periodic distribution characteristics, or a high-temperature resistant insert with such slots is used.

[0008] The low-pressure chamber is located inside the blunt-headed shell, adjacent to the inner side of the continuous slit structure, and is used to collect the inhaled high-pressure gas.

[0009] The suction pump is matched with the compressor and installed inside the blunt-head housing; the suction pump is connected to the low-pressure chamber through a pipeline to provide the pressure difference required for suction; the compressor is located downstream of the suction pump and is used to further compress the suction gas to overcome the back pressure and discharge it from the machine body;

[0010] The control unit is located in the electronics bay inside the aircraft fuselage and is electrically connected to the suction pump and compressor respectively. It is used to output control signals to adjust the speed or power of the suction pump and compressor according to the received flight parameters.

[0011] As a preferred embodiment of the present invention: the control unit includes a main control processor, a signal acquisition module, and a power drive module; the main control processor is electrically connected to the signal acquisition module and the power drive module respectively and transmits electrical signals simultaneously; the signal acquisition module is electrically connected to the flight control computer or sensor group of the aircraft for acquiring real-time flight parameters; the power drive module is electrically connected to the suction pump and the compressor respectively.

[0012] As a preferred embodiment of the present invention: the airflow generated by the suction pump is gathered into the low-pressure chamber, and discharged outside the aircraft after cooling or throttling; by adjusting the working state of the suction pump and the compressor, the suction flow rate can be changed to adapt to the optimal control requirements under different flight conditions; at the beginning of flight or when no control is required, the suction pump and the compressor are turned off or their power is reduced to reduce the occupation of system resources.

[0013] A hypersonic blunt-nosed body boundary layer delay transition control method is provided, based on the aforementioned hypersonic blunt-nosed body boundary layer delay transition control device. It implements local steady-state wall suction in the stagnation point region at the leading edge of the hypersonic blunt-nosed body, thereby suppressing the initial sensitivity of free flow disturbances to the boundary layer at the source of non-modal strip disturbances by removing low-velocity fluid near the wall of the boundary layer.

[0014] As a preferred embodiment of the present invention, the control method specifically includes the following steps:

[0015] 1) Real-time monitoring of flight operating conditions

[0016] During flight, the control unit receives the aircraft's operating parameters in real time.

[0017] 2) Determine the start-up control conditions

[0018] The control unit determines whether the current flight status is within the preset hypersonic transition sensitive zone;

[0019] 3) Real-time stationary steady-state suction

[0020] The control unit sends commands to the suction pump and compressor. The suction pump operates to create a low-pressure zone in the low-pressure chamber inside the blunt-head housing relative to the external flow field. Under the action of the pressure difference, the external airflow is drawn into the low-pressure chamber through the continuous slit structure at the stagnation point.

[0021] 4) Gas emission and thermal management

[0022] The inhaled high-temperature gas enters the compressor through the pipeline, and after being pressurized and cooled by heat exchange, it is transported to the leeward side of the aircraft or other low-pressure areas through the exhaust channel and discharged outside the aircraft body to avoid interfering with the main aerodynamic surfaces.

[0023] 5) Dynamic adjustment of suction parameters

[0024] The control unit adjusts the speed or power of the suction pump and compressor according to flight parameters to control the suction flow rate. When the flight speed or wall temperature increases, the control unit increases the power of the suction pump and compressor to increase the suction volume. When the wall temperature decreases or the hypersonic range is exited, the control unit adjusts the power of the suction pump and compressor to reduce the power or shut down the suction pump and compressor.

[0025] As a preferred embodiment of the present invention: when the aircraft is in hypersonic flight, the wall of the leading edge stagnation area continuously draws in the airflow in the surrounding boundary layer; the small suction force generated at the suction pump removes a portion of the low-speed fluid near the wall of the boundary layer, forming a stable suction flow along the wall and the flow rate generated by the suction is very small.

[0026] By adopting the above technical solution, the present invention has the following beneficial effects:

[0027] The hypersonic blunt-nosed body boundary layer delay transition control device and method of the present invention are reasonably conceived and have the following advantages:

[0028] (1) Significantly suppresses initial gain of disturbance and exhibits high robustness: This invention directly intervenes in the sensitive process of free-flow disturbance transforming into non-modal stripes within the boundary layer by implementing local suction in the stagnation region at the leading edge of the blunt-nosed body. Numerical verification (calculation model and method as follows) has been performed. Figure 2 As shown in the figure, without control, the disturbance amplitude increases monotonically along the flow direction, while stagnation point intake causes the disturbance growth curve to shift significantly downward, with the peak amplitude being significantly lower than the reference value. Furthermore, for different spanwise wavenumbers... ,frequency and angle of incidence Nonmodal perturbations and pumping at stagnation points both exhibit significant stabilizing effects.

[0029] (2) The scheme has high control efficiency and low energy consumption: Under typical wind tunnel conditions of incoming Mach number 5.96, incoming temperature 87K, and wall temperature 290K, numerical verification results show that when the suction velocity at the stagnation point accounts for 5‰ of the free incoming flow (mass flow rate is about 2% of the free incoming flow), the dominant nonmodal disturbance (spanwise wavenumber) can be controlled. The linear evolution amplitude of the ) is reduced by nearly 50%. More importantly, the corresponding amplification factor of the strip secondary instability disturbance amplitude can be reduced from about 13 to below 1, thereby delaying the transition point to the far downstream region and significantly reducing the surface friction drag and thermal load of the aircraft.

[0030] (3) Highly targeted and safe engineering: This invention avoids the negative side effects that may result from non-targeted intervention methods in the prior art. This invention is a precise intervention based on flow mechanism, avoiding the side effects that may be caused by other non-targeted methods, such as downstream injection inducing new unstable structures. At the same time, the suction section is concentrated in a small area of ​​the leading edge stagnation point, eliminating the need to arrange complex actuators or porous media over a large area on the spacecraft wall, reducing the design difficulty of the thermal protection system, and enhancing the structural integrity and reliability of the system in extreme hypersonic environments.

[0031] In summary, this invention provides a passive-active hybrid control device integrated into the nose of an aircraft. It does not significantly interfere with the incoming flow under normal conditions, but plays a key role in suppressing boundary layer disturbances.

[0032] This invention provides a stationary suction control method for nonmodal disturbance source sensitivity, which differs from existing control methods that mainly target the linear and nonlinear evolution of Mack's second mode (high-frequency acoustic mode), filling the gap in hypersonic blunt body transition control technology.

[0033] This invention analyzes the influence of wall suction parameters on disturbances and finds that the physical location of the control section is the dominant factor determining the suppression effect, i.e., the leading edge stagnation region is the optimal control location. At the same time, it finds that the specific shape of the suction inlet (when the flux is constant) has a robust effect on the effect, thus establishing a technical route that is mainly based on position control and has strong shape adaptability.

[0034] This invention, through the integrated design of a hypersonic boundary layer active control system, is further extended into an integrated device. Specifically, it incorporates a spanwise continuous suction slit and an internal suction pump at the blunt nose of the spacecraft, forming a boundary layer transition control device. The device delays transition by suctioning boundary layer gas through the slit according to set parameters. It utilizes steady-state mass flux through a specific slit structure to control hypersonic boundary layer stability. Attached Figure Description

[0035] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0036] Figure 1 This is a schematic diagram of the physical model and flow structure involved in the present invention;

[0037] Figure 2 This is a schematic diagram of the numerical simulation verification calculation model and method involved in the present invention.

[0038] Figure 3 The diagram shows a comparison of the linear evolution of nonmodal disturbance velocity amplitude along the flow direction under different control schemes involved in this invention.

[0039] Figure 4 The diagram shows the variation curves of the amplitude amplification factor N of the secondary instability mode along the flow direction under different control schemes involved in this invention.

[0040] Figure 5 This is a block diagram illustrating the structure and connection relationships of the control unit involved in this invention.

[0041] The labels in the diagram:

[0042] V0 - No control reference condition; V1 - Incoming flow velocity of 5‰ (blowing speed), suction width of 2 times blunt tip radius; V2 - Incoming flow velocity of 5‰ (suction speed), suction width of 2 times blunt tip radius; V3 - Incoming flow velocity of 2% (suction speed), suction width of 0.5 times blunt tip radius. Detailed Implementation

[0043] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0044] The present invention will be further explained below with reference to specific embodiments.

[0045] like Figure 1 As shown in the figure, this embodiment provides a hypersonic blunt-nosed body boundary layer delay transition control device, which includes a blunt-nosed body housing 1, a continuous slit structure 2, a low-pressure chamber 3, a suction pump 4, a compressor 5, and a control unit 6.

[0046] The blunt-headed shell 1 has an internal accommodating space and a streamlined arc-shaped structure at the front; wherein, the head of the blunt-headed shell, i.e. the stagnation area, is uniformly provided with continuous slit structures 2.

[0047] The continuous slot structure 2 can be designed as a narrow slot structure that is continuous in the spanwise direction and has no spanwise periodic distribution characteristics, or it can be a high-temperature resistant insert with such slots to meet the thermal protection requirements in hypersonic environments.

[0048] The low-pressure chamber 3 is located inside the blunt-head housing 1, adjacent to the inner side of the continuous slit structure 2, and is used to collect the high-pressure gas that is drawn in.

[0049] The suction pump 4 is matched with the compressor 5 and installed inside the blunt head housing 1. The suction pump 4 is connected to the low pressure chamber 3 through a pipeline to provide the pressure difference required for suction.

[0050] The compressor 5 is located downstream of the suction pump 4 and is used to further compress the intake gas to overcome the back pressure and expel it from the machine.

[0051] The control unit 6 can adjust the operating state of the suction pump 4 to change the suction flow rate to adapt to the optimal control requirements under different flight conditions (for example, the suction flow rate can be appropriately increased when the intensity of environmental disturbance increases). At the beginning of flight or when no control is required, the suction pump 4 can also be turned off or its power reduced to reduce the occupation of system resources.

[0052] The control unit 6 is located in the electronics bay inside the aircraft fuselage and includes a signal acquisition module, a main control processor, and a power drive module. The signal acquisition module is electrically connected to the aircraft's flight control computer or sensor array (such as an air data sensor) to acquire real-time flight parameters (Mach number, wall temperature, etc.). The power drive module is electrically connected to both the suction pump 4 and the compressor 5. Based on the received flight parameters, the control unit 6 outputs control signals to adjust the speed or power of the suction pump 4 and the compressor 5.

[0053] The hypersonic blunt-nosed body boundary layer delay transition control device of the present invention achieves the function of delay transition by drawing boundary layer gas through a continuous slit structure 2 according to set parameters. It utilizes steady-state mass flux through the continuous slit structure 2 to control the stability of the hypersonic boundary layer. The control principle and parameter design are based on the following:

[0054] I. Stagnant Pump Control for Non-modal Disturbance Source Sensitivity

[0055] Local steady-state wall suction is implemented near the stagnation point at the leading edge of the hypersonic blunt-nosed body to suppress the energy of nonmodal strip disturbances at their source, i.e., to suppress the initial receptive process of free-flow disturbances entering the boundary layer.

[0056] Figure 1 The figure illustrates the bow-shaped shock wave generated at the leading edge of the hypersonic blunt wedge and the strip structure excited by environmental disturbances, as well as the boundary layer development. The locations of the local suction and blowing control sections on the wall are marked (red areas on the wall), and the sub-figure shows a magnified view of the head region and its internal components. The blunt body is simplified to a blunt wedge model with leading-edge bluntness; the red area indicates the location of the local suction application, and the sub-figure shows a magnified view of the head region. The direction of the hypersonic incoming flow is schematically marked in the figure; the bow-shaped shock wave forms in front of the blunt head and refracts the free-flow disturbance into the boundary layer. Subsequent numerical simulations validate this model under typical wind tunnel incoming flow conditions (free-flow Mach number 5.96, free-flow temperature 87 K, wall temperature 290 K).

[0057] When the aircraft is in hypersonic flight, the walls of the leading-edge stagnation region continuously draw airflow from the surrounding boundary layer. Specifically, the small suction generated at pump 4 removes a portion of the low-velocity fluid near the boundary layer wall, forming a stable suction flow along the wall. The suction flow is typically very small (e.g., only about 2% of the free-flow flow). Because this suction acts directly on the stagnation region, it significantly weakens the initial sensitivity of free-flow disturbances entering the boundary layer through the stagnation point. Thus, the intensity of high-speed / low-speed strip disturbances that would normally form downstream is greatly reduced, and the initial amplitude of non-modal disturbances in the boundary layer is effectively suppressed.

[0058] like Figure 2 As shown, the harmonic linearized Navier-Stokes (SF-HLNS) method of shock assembly is used to calculate the sensitivity and linear evolution of nonmodal disturbances in the nose blunt-nose flow field region. Then, in the downstream flat plate region, the nonlinear parabolic stability equation (NPSE) and quadratic instability analysis are used to describe the nonlinear evolution, achieving efficient prediction of the sensitivity and nonlinear evolution of hypersonic blunt body boundaries. Simulation results show that without control, the disturbance amplitude monotonically increases along the flow direction, while stagnation point intake significantly shifts the disturbance growth curve downwards, with the peak amplitude significantly lower than the baseline value. Furthermore, for different spanwise wavenumbers... ,frequency and angle of incidence Nonmodal perturbations and pumping at stationary points both exhibit significant stabilizing effects. The core of this strategy lies in significantly reducing the excitation and transient amplification of nonmodal stripes at the source with minimal pumping costs, thereby effectively suppressing subsequent secondary instability evolution.

[0059] II. Analysis of the Influence of Wall Suction Parameters on Disturbance

[0060] The physical position of the continuous slit structure 2 is the dominant factor determining the suppression effect, that is, the leading edge stagnation point region is the optimal control position; at the same time, it was found that the specific shape of the suction inlet (when the flux is constant) has a robust effect on the effect, thus establishing a technical route with position control as the main focus and strong shape adaptability.

[0061] From a mechanistic perspective, this invention weakens the sensitive phase of nonmodal disturbances by introducing wall normal velocity at the stagnation point. This not only directly reduces the transient amplification of nonmodal disturbances but also reduces the distortion of the boundary layer mean flow. Further NPSE calculations and secondary instability analysis (experiments to verify the effectiveness of this invention) show that, under uncontrolled conditions, secondary instability caused by nonmodal stripes (dominated by bending modes) grows rapidly, with its amplitude amplification factor N reaching approximately 13 downstream, sufficient to induce turbulence. However, after applying stagnation point suction with a mass flow rate of approximately 2% of the free flow, the growth rate and instability parameter range of the secondary instability disturbance decrease significantly, with its N factor remaining below 1 throughout the entire computational domain, and almost no further significant instability amplification occurs. This means that the boundary layer can maintain near-laminar flow even far downstream, achieving the purpose of delaying transition. It is worth noting that this invention achieves the above effects at minimal cost: because the suction section is located in the leading-edge stagnation point region, only a small suction mass flow rate is needed to achieve significant control effects. Furthermore, this control scheme has proven effective under different parameters and different types of incoming flow disturbances, indicating that stagnation point suction achieves control at the source by changing the basic flow profile and reducing disturbance sensitivity, and has good engineering robustness and wide adaptability to operating conditions.

[0062] Based on the above-mentioned control principles, parameter design criteria, and control device, the hypersonic blunt-nosed body boundary layer delay transition control method of the present invention specifically includes the following steps:

[0063] (1) Real-time monitoring of flight operating conditions

[0064] During flight, the control unit 6 receives real-time operating parameters of the aircraft, including flight Mach number and incoming flow temperature.

[0065] (2) Determine the start-up control conditions

[0066] Control unit 6 determines whether the current flight state is within a preset hypersonic transition sensitive zone. For example, it activates control when the flight Mach number is detected to be greater than 5 and the wall temperature is close to the severe ablation threshold.

[0067] (3) Real-time stationary steady-state suction

[0068] Control unit 6 sends commands to suction pump 4 and compressor 5. Suction pump 4 operates to create a low-pressure zone in low-pressure chamber 3 inside blunt-head housing 1 relative to the external flow field. Under the action of pressure difference, external airflow is drawn into low-pressure chamber 3 through continuous slit structure 2 at stagnation point.

[0069] (4) Gas emission and thermal management

[0070] The inhaled high-temperature gas enters the compressor 5 through the pipeline. After being pressurized and cooled by necessary heat exchange (heat exchange cooling is achieved through the heat exchange system on the pipeline. The cooling medium can be the aircraft's own fuel cold source or the onboard active cooling circulation system. The purpose is to cool the inhaled high-temperature gas to a range that the exhaust channel can withstand), it is transported to the leeward side of the aircraft or other low-pressure areas and discharged outside the aircraft body to avoid interfering with the main aerodynamic surfaces.

[0071] (5) Dynamic adjustment of suction parameters

[0072] To achieve optimal control efficiency and reduce energy consumption, the control unit 6 can adjust the rotation speed or power of the suction pump (4) and compressor (5) according to flight parameters to control the suction flow rate. When the flight speed or wall temperature increases, the control unit (6) increases the power of the suction pump (4) and compressor (5) to increase the suction volume. When the wall temperature decreases or the hypersonic range is exited, the control unit (6) adjusts the power of the suction pump (4) and compressor (5), reduces the power or shuts down the suction pump (4) and compressor (5), thus reducing the power or shutting down the system.

[0073] Figure 3 Comparison curves of the linear evolution of nonmodal disturbance velocity amplitude along the flow direction under different control schemes ( (Situation); the black solid line represents the no-control baseline condition, the red dashed line represents the blowing control comparison condition, and the green dotted line and the blue solid line with solid dots represent two suction control schemes with the same flow rate.

[0074] Figure 4 The graph shows the variation curves of the amplitude amplification factor N of the secondary instability mode along the flow direction under different control schemes; the black solid line in the figure represents the no-control reference condition, and the red dashed line and the blue solid line with solid dots represent two suction controls with the same flow rate.

[0075] Figure 5 The diagram shows the structure and connection relationship of the control unit in this embodiment of the invention (the diagram shows the signal connection relationship between the control unit and the external flight control computer or sensor group, as well as the logical composition of its internal signal acquisition module, main control processor and power drive module, and drives the downstream suction pump and compressor to work through control signals / electric power).

[0076] This invention provides a passive-active hybrid control device integrated into the nose of an aircraft. It does not significantly interfere with the incoming flow under normal conditions, but plays a key role in suppressing boundary layer disturbances.

[0077] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A hypersonic blunt-nosed body boundary layer delay transition control device, characterized in that: The control device includes a blunt head housing (1), a continuous slit structure (2), a low-pressure chamber (3), a suction pump (4), a compressor (5), and a control unit (6). The blunt-headed shell (1) has an accommodating space inside, and its head is a stationary area and is uniformly provided with the continuous slit structure (2). The continuous slot structure (2) is designed as a narrow slot structure that is continuous in the spanwise direction and has no spanwise periodic distribution characteristics, or a high-temperature resistant insert with such slots is used. The low-pressure chamber (3) is located inside the blunt-head shell (1), adjacent to the inner side of the continuous slit structure (2), and is used to collect the high-pressure gas that is drawn in. The suction pump (4) and the compressor (5) are matched and installed inside the blunt head housing (1); the suction pump (4) is connected to the low pressure chamber (3) through a pipeline to provide the pressure difference required for suction; the compressor (5) is located downstream of the suction pump (4) to further compress the gas inhaled to overcome the back pressure and discharge it from the machine body; The control unit (6) is located in the electronic equipment bay inside the aircraft fuselage and is electrically connected to the suction pump (4) and compressor (5) respectively. It is used to output control signals to adjust the speed or power of the suction pump (4) and compressor (5) according to the received flight parameters.

2. The hypersonic blunt-nosed body boundary layer delay transition control device as described in claim 1, characterized in that: The control unit (6) includes a main control processor, a signal acquisition module, and a power drive module; The main control processor is electrically connected to the signal acquisition module and the power drive module respectively and transmits electrical signals simultaneously. The signal acquisition module is electrically connected to the flight control computer or sensor group of the aircraft to acquire real-time flight parameters; The power drive module is electrically connected to the suction pump (4) and the compressor (5) respectively.

3. The hypersonic blunt-nosed body boundary layer delay transition control device as described in claim 1, characterized in that: The airflow generated by the suction pump (4) is gathered into the low-pressure chamber (3) and discharged outside the aircraft after cooling or throttling. By adjusting the working state of the suction pump (4) and the compressor (5), the suction flow rate can be changed to adapt to the optimal control requirements under different flight conditions. At the beginning of flight or when no control is required, the suction pump (4) and the compressor (5) are turned off or their power is reduced to reduce the occupation of system resources.

4. A method for controlling the boundary layer delay transition of a hypersonic blunt-nosed body, based on the hypersonic blunt-nosed body boundary layer delay transition control device according to any one of claims 1 to 3; characterized in that: Local steady-state wall suction is implemented in the stagnation region at the leading edge of the hypersonic blunt-nosed body. By removing the low-velocity fluid near the wall of the boundary layer, the initial sensitivity of free flow disturbances to entering the boundary layer is suppressed at the source of nonmodal strip disturbances.

5. The hypersonic blunt-nosed body boundary layer delay transition control method as described in claim 4, characterized in that, The control method specifically includes the following steps: 1) Real-time monitoring of flight operating conditions During the flight of the aircraft, the control unit (6) receives the operating parameters of the aircraft in real time; 2) Determine the start-up control conditions The control unit (6) determines whether the current flight status is within the preset hypersonic transition sensitive zone; 3) Real-time stationary steady-state suction The control unit (6) sends instructions to the suction pump (4) and the compressor (5). The suction pump (4) works to form a low-pressure zone in the low-pressure chamber (3) inside the blunt head housing (1) relative to the external flow field. Under the action of pressure difference, the external airflow is drawn into the low-pressure chamber (3) through the continuous slit structure (2) at the stagnation point. 4) Gas emission and thermal management The high-temperature gas that is sucked in enters the compressor (5) through the pipeline. After being pressurized and cooled by heat exchange, it is transported to the leeward side of the aircraft or other low-pressure areas through the exhaust channel and discharged outside the aircraft body to avoid interference with the main aerodynamic surfaces. 5) Dynamic adjustment of suction parameters The control unit (6) adjusts the rotation speed or power of the suction pump (4) and compressor (5) according to the flight parameters to control the suction flow rate. When the flight speed or wall temperature increases, the control unit (6) increases the power of the suction pump (4) and compressor (5) to increase the suction volume. When the wall temperature decreases or the hypersonic range is removed, the control unit (6) adjusts the power of the suction pump (4) and compressor (5) to reduce the power or shut down the suction pump (4) and compressor (5).

6. The hypersonic blunt-nosed body boundary layer delay transition control method as described in claim 5, characterized in that: When the aircraft is in hypersonic flight, the wall of the leading edge stagnation area continuously draws airflow from the surrounding boundary layer; the small suction force generated at the suction pump (4) removes a portion of the low-speed fluid near the wall of the boundary layer, forming a stable suction flow along the wall and the flow rate generated by the suction is very small.

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