A hypersonic flow control structure

Abstract

The present application discloses a hypersonic flow control structure, including: a plurality of small rib unit array structures arranged in parallel, the characteristic height of the small rib unit array structure is less than 0.1δ, δ is the boundary layer thickness, the center of each small rib unit array structure is provided with a divergent line parallel to the test flow direction along the length direction, the surface of each small rib unit array structure is symmetrically provided with a plurality of micro groove structures parallel to each other and at a set angle to the test flow direction on both sides of the divergent line length direction, and a convergence groove parallel to the test flow direction is provided between adjacent small rib unit array structures. The characteristic height of this embodiment is much smaller than the boundary layer thickness, and the additional resistance brought by the hypersonic flow condition can be basically ignored; this embodiment has a simple structure, high reliability, and does not require external energy input; this embodiment can effectively delay or eliminate the separation of the boundary layer flow on the surface of the hypersonic aircraft.

Description

Technical Field

[0001] The present application relates to the technical field of hypersonic aircraft, and in particular, to a hypersonic flow control structure. Background Art

[0002] The separation of boundary layer flow on the surface of hypersonic aircraft often leads to increased drag, and the unsteady motion caused by the separation may also cause high extreme thermal flux loads and pulsating pressure loads, which will have serious consequences for the structure and flight safety of the aircraft. In order to delay or eliminate the separation of boundary layer flow on the surface of hypersonic aircraft, certain methods need to be adopted to control the flow.

[0003] Traditional flow control methods include active and passive control. Active control generally includes wall heating / cooling, boundary layer suction, acoustic excitation, jet, plasma, etc. Passive control generally includes ventilation wall, flow bulge, vortex generator, synthetic jet, etc. Among the many flow control methods, vortex generators are the most popular due to their excellent robustness and control effectiveness. However, on the one hand, the characteristic height of the vortex generator is usually at the same order of magnitude as the thickness of the boundary layer, and its blocking effect on the mainstream will bring high additional drag. On the other hand, when the Mach number rises above supersonic speed, the control efficiency of the vortex generator on flow separation will be greatly reduced. This is because the counter-rotating vortex pairs induced by the vortex generator will decay quickly in the hypersonic boundary layer. Therefore, the development of hypersonic aircraft requires a new control method that can simultaneously meet the following two requirements: (1) effectively delay or eliminate flow separation under hypersonic conditions; (2) the structure used for hypersonic flow control produces as little additional drag as possible. Summary of the invention

[0004] The present application provides a hypersonic flow control structure to solve the technical problem that the existing vortex generator has high additional resistance and the control efficiency is greatly reduced when the Mach number rises above supersonic speed.

[0005] The technical solutions adopted in this application are as follows:

[0006] A hypersonic flow control structure, comprising:

[0007] A plurality of small rib unit array structures are arranged in parallel, wherein the characteristic height of the small rib unit array structure is less than 0.1δ, where δ is the boundary layer thickness, a divergent line parallel to the test flow direction is arranged at the center of each small rib unit array structure along the length direction, a plurality of micro groove structures parallel to each other and forming a set angle with the test flow direction are symmetrically arranged on both sides of the divergent line in the length direction on the surface of each small rib unit array structure, and a converging groove parallel to the test flow direction is arranged between adjacent small rib unit array structures.

[0008] Furthermore, the width of each of the small rib unit array structures is w=1δ-3δ, wherein δ is the boundary layer thickness.

[0009] Furthermore, the length of each of the small rib unit array structures is l=3δ~9δ, wherein δ is the boundary layer thickness.

[0010] Furthermore, the total thickness of the small rib unit array structure is less than 0.2δ.

[0011] Furthermore, the set angle α between the micro-groove structure and the test flow direction is 30° to 60°.

[0012] Furthermore, the cross section of the micro-groove structure is trapezoidal, and the groove depth h=0.02δ~0.1δ, and the groove width s=1.5h.

[0013] Furthermore, the cross section of the micro-groove structure is triangular, and the groove depth h=0.02-0.1δ, and the groove width s=h.

[0014] Furthermore, the cross section of the micro-groove structure is rectangular, and the groove depth h=0.02δ~0.1δ, and the groove width s=h.

[0015] Furthermore, the spacing g between the micro-groove structures on both sides of the divergent line is 1.5h, where h is the groove depth of the micro-groove structure.

[0016] Furthermore, the width of the converging groove is 2h-3h, and h is the groove depth of the micro-groove structure.

[0017] Compared with the prior art, this application has the following beneficial effects:

[0018] The present application provides a hypersonic flow control structure, including: a plurality of small rib unit array structures arranged in parallel, the characteristic height of the small rib unit array structure is less than 0.1δ, δ is the boundary layer thickness, the center of each small rib unit array structure is provided with a divergent line parallel to the test flow direction along the length direction, the surface of each small rib unit array structure is symmetrically provided with a plurality of micro groove structures parallel to each other and at a set angle to the test flow direction on both sides of the divergent line length direction, and a convergence groove parallel to the test flow direction is provided between adjacent small rib unit array structures. The characteristic height of the hypersonic flow control structure provided in this embodiment is much smaller than the boundary layer thickness, and the additional resistance brought under the hypersonic flow condition can be basically ignored; the hypersonic flow control structure of this embodiment has a simple structure, high reliability, and does not require external energy input, and has low energy saving and consumption reduction costs; the hypersonic flow control structure of this embodiment can effectively delay or eliminate the separation of the boundary layer flow on the surface of the hypersonic aircraft.

[0019] In addition to the above-described purposes, features and advantages, the present application also has other purposes, features and advantages. The present application will be further described in detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] The drawings constituting a part of the present application are used to provide a further understanding of the present application. The illustrative embodiments and descriptions of the present application are used to explain the present application and do not constitute an improper limitation on the present application. In the drawings:

[0021] Figure 1 It is a schematic diagram of the overall structure of the hypersonic flow control structure of the preferred embodiment of the present application.

[0022] Figure 2 yes Figure 1 Schematic diagram of the AA section view.

[0023] As shown in the figure:

[0024] 1. Divergent lines; 2. Micro-groove structure; 3. Converging grooves. DETAILED DESCRIPTION

[0025] It should be noted that, in the absence of conflict, the embodiments and features in the embodiments of the present application can be combined with each other. The present application will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0026] Reference Figure 1 to Figure 2 , a preferred embodiment of the present application provides a hypersonic flow control structure, comprising:

[0027] A plurality of small rib unit array structures are arranged in parallel, wherein the characteristic height of the small rib unit array structures is less than 0.1δ, where δ is the boundary layer thickness, and a divergent line 1 parallel to the test flow direction is arranged at the center of each of the small rib unit array structures along the length direction, and a plurality of micro groove structures 2 which are parallel to each other and form a set angle with the test flow direction are symmetrically arranged on both sides of the divergent line in the length direction, and a converging groove 3 parallel to the test flow direction is arranged between adjacent small rib unit array structures.

[0028] The present embodiment provides a hypersonic flow control structure, which is a directional surface microstructure imitating bird feathers, including: a plurality of small rib unit array structures arranged in parallel, the characteristic height of the small rib unit array structure is less than 0.1δ, δ is the boundary layer thickness, the center of each small rib unit array structure is provided with a divergent line parallel to the test flow direction along the length direction, the surface of each small rib unit array structure is symmetrically provided with a plurality of micro groove structures parallel to each other and at a set angle to the test flow direction on both sides of the divergent line length direction, and a convergence groove parallel to the test flow direction is provided between adjacent small rib unit array structures. The characteristic height of the hypersonic flow control structure provided in the present embodiment is much smaller than the boundary layer thickness, and the additional resistance brought under the hypersonic flow condition can be basically ignored; the hypersonic flow control structure of the present embodiment has a simple structure, high reliability, and does not require external energy input, and has low energy saving and consumption reduction costs; the hypersonic flow control structure of the present embodiment can effectively delay or eliminate the separation of the boundary layer flow on the surface of the hypersonic aircraft.

[0029] Specifically, the main reason why the hypersonic flow control structure of this embodiment can induce a large-scale vortex structure in the boundary layer is due to the flow-guiding effect of the special converging-diverging groove structure on the bottom flow in the boundary layer:

[0030] At the divergent line 1, the flow inside the micro-groove structure 2 will flow along the groove direction, forming a lateral flow velocity directed from the divergent line 1 to both sides. In order to supplement the mass lost by the lateral flow, the mass of the boundary layer located in the upper layer will be sucked into the bottom layer of the boundary layer, forming a flow velocity from top to bottom in the vertical direction; at the converging groove 3, due to the flow-guiding effect of the micro-groove structure 2, a lateral flow directed from both sides to the converging groove 3 is formed. The converged flow moves toward the boundary layer located in the upper layer due to the overflow effect, forming a flow velocity from bottom to top in the vertical direction. The vertical flows in opposite directions above the divergent line 1 and above the converging groove 3 jointly form a flow-direction rotating vortex structure. Existing research has found that the size of the flow-direction rotating vortex structure mainly depends on the lateral wavelength Λ of the hypersonic flow control structure of this embodiment, and has no correlation with its characteristic height h, and the flow-direction rotating vortex structure will evolve with the flow direction.

[0031] The size of the flow direction rotating vortex structure induced by the hypersonic flow control structure of this embodiment in the boundary layer is at the same order of magnitude as the thickness of the hypersonic boundary layer, so the flow direction rotating vortex structure generated by the hypersonic flow control structure under hypersonic conditions can enhance the energy mixing effect between the boundary layer and the mainstream, and improve the ability of the boundary layer to resist flow separation caused by the adverse pressure gradient. In addition, the characteristic height h of the hypersonic flow control structure of this embodiment is often less than 0.1δ, so the additional resistance generated can be basically ignored. Therefore, the hypersonic flow control structure can reduce or eliminate the boundary layer flow separation existing in the hypersonic flow without significantly increasing the flow resistance.

[0032] The hypersonic flow control structure in this embodiment has been verified through multiple wind tunnel tests. This new flow control structure is stable and the solution is feasible. It can delay or eliminate the boundary layer flow separation on the surface of the hypersonic aircraft, and plays a key role in the wall boundary layer flow control test research. The results have achieved the expected goals.

[0033] In a preferred embodiment of the present application, the width w of each of the small rib unit array structures is w=1δ~3δ, wherein δ is the boundary layer thickness. Experiments have found that the spanwise dimension of the flow-directional rotating vortex structure induced by the device in the boundary layer is basically consistent with the width w of the small rib unit array structure. Therefore, in order to allow the flow-directional rotating vortex structure to generate energy exchange in the entire boundary layer, the width w of the small rib unit array structure should not be less than the boundary layer thickness. At the same time, in order to allow the flow-directional rotating vortex structure to develop in the flow field as much as possible without breaking up, w is usually not greater than 3 times the boundary thickness.

[0034] In a preferred embodiment of the present application, the length l of each of the small rib unit array structures is 3δ to 9δ, where δ is the boundary layer thickness. The control structure needs a flow distance to induce the formation of a flow-direction rotating vortex structure. Experiments have found that a flow distance of more than 3 times the boundary layer thickness is sufficient to induce the formation of a flow-direction rotating vortex structure, and the flow control effect increases with the increase of the length l, while a length l of more than 9 times the boundary layer thickness no longer changes the control effect.

[0035] In a preferred embodiment of the present application, the total thickness of the small rib unit array structure is less than 0.2δ. The reason and benefit of this setting is that the flow control structure is controlled in the low-speed bottom layer where the boundary layer develops, the boundary layer flow control effect is good, and it will not affect the mainstream flow field.

[0036] In a preferred embodiment of the present application, the set angle α between the micro-groove structure and the test flow direction is 30°~60°. The angle between the micro-groove structure and the test flow direction has a diversion effect on the boundary layer, can generate lateral secondary flow, and then induce a flow-directional rotating vortex structure in the entire boundary layer. The formation of the diversion effect requires that the angle α is greater than 30°. At the same time, in order to minimize the increase in the additional resistance generated by the control structure, the angle needs to be controlled below 60°.

[0037] In a preferred embodiment of the present application, the cross-section of the micro-groove structure is a trapezoid, and the groove depth h = 0.02δ ~ 0.1δ, the groove width s = 1.5h. In this embodiment, the cross-section of the micro-groove structure is a trapezoid. The advantage of setting a trapezoid is that there is a velocity gradient when the boundary layer develops from the bottom layer to the normal direction. The flow channel corresponding to the trapezoid is set from small to large, which just meets the velocity change and has the best flow control effect. At the same time, in order to minimize the wave resistance generated by the structure in the hypersonic flow field, its depth should not exceed 10% of the boundary layer thickness, but the experimental effect will decrease with the decrease of the groove depth h, so h should not be less than 2% of the boundary layer thickness.

[0038] In a preferred embodiment of the present application, the cross-section of the micro-groove structure is a triangle, and the groove depth h=0.02-0.1δ, the groove width s=h. The advantage of setting a triangle is that the structure is simple and the processing cost is low.

[0039] In a preferred embodiment of the present application, the cross-section of the micro-groove structure is rectangular, and the groove depth h=0.02δ~0.1δ, the groove width s=h, and the advantage of setting a rectangular groove is that the construction is simple and easy to implement.

[0040] In a preferred embodiment of the present application, the spacing g of the micro-groove structures located on both sides of the divergent line is 1.5h, which is consistent with the groove width s. The advantage is that it can better improve the velocity profile characteristics of the hypersonic boundary layer and achieve the purpose of surface boundary layer flow control.

[0041] In a preferred embodiment of the present application, the width of the convergence groove is 2h to 3h, thereby providing sufficient volume to collect the lateral flow flowing in from the micro-groove structures 2 on both sides, so that the converged flow moves toward the boundary layer located in the upper layer due to the overflow effect, forming a flow velocity from bottom to top in the vertical direction.

[0042] The above embodiments have been verified through wind tunnel tests, and the solutions are feasible and the results have achieved the expected goals.

[0043] In summary, the hypersonic flow control structure provided in the present application can generate counter-rotating flow vortex pairs in the wall boundary layer, thereby improving the velocity profile characteristics of the hypersonic boundary layer, enhancing the ability of the boundary layer to resist separation, and achieving the purpose of delaying or eliminating boundary layer flow separation on the surface of a hypersonic aircraft. It is conducive to conducting wall boundary layer flow control experimental research and promoting the practical application of related technical engineering.

[0044] The above description is only the preferred embodiment of the present application and is not intended to limit the present application. For those skilled in the art, the present application may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present application shall be included in the protection scope of the present application.

Claims

1. A hypersonic flow control structure, It is characterized in that include: A plurality of small rib unit array structures are arranged in parallel, wherein the characteristic height of the small rib unit array structure is less than 0.1δ, where δ is the boundary layer thickness, a divergence line parallel to the test flow direction is arranged at the center of each small rib unit array structure along the length direction, a plurality of micro groove structures parallel to each other and forming a set angle with the test flow direction are symmetrically arranged on both sides of the divergence line in the length direction on the surface of each small rib unit array structure, and a convergence groove parallel to the test flow direction is arranged between adjacent small rib unit array structures; The width w of each of the small rib unit array structures is 1δ~3δ, where δ is the boundary layer thickness; The length of each of the small rib unit array structures l =3δ~9δ; The total thickness of the small rib unit array structure is less than 0.2δ; The spacing g between the micro-groove structures on both sides of the divergent line is 1.5h, where h is the groove depth of the micro-groove structure; The width of the converging groove is 2h-3h, where h is the groove depth of the micro-groove structure.

2. The hypersonic flow control structure according to claim 1, It is characterized in that The set angle α between the micro-groove structure and the test flow direction is 30°~60°.

3. The hypersonic flow control structure according to claim 1, It is characterized in that The cross section of the micro-groove structure is a trapezoid, and the groove depth h=0.02δ~0.1δ, and the groove width s=1.5h.

4. The hypersonic flow control structure according to claim 1, It is characterized in that The cross section of the micro-groove structure is triangular, and the groove depth h=0.02~0.1δ, and the groove width s=h.

5. The hypersonic flow control structure according to claim 1, It is characterized in that The cross section of the micro-groove structure is rectangular, and the groove depth h=0.02δ~0.1δ, and the groove width s=h.

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