Supersonic internal flow control device based on induction unit array

Abstract

The invention provides an induction unit array-based supersonic internal flow control device, which comprises an equal straight section flow channel and an induction unit array, and is characterized in that the induction unit array comprises two groups of induction units symmetrically distributed on two sides of a center line of the equal straight section flow channel, and each group of induction units comprises a shock wave generator and a flow deflector; the flow deflector is fixedly mounted in the equal straight section flow channel, and the shock wave generator is rotatably mounted in the equal straight section flow channel; the side wall of the equal straight section flow channel is provided with a boundary layer suction groove corresponding to each induction unit, the boundary layer suction grooves are correspondingly formed in the positions between the rear edge point of the shock wave generator and the front edge point of the flow deflector, and a suction device communicated with the boundary layer suction grooves is arranged outside the equal straight section flow channel. And the outlet back pressure of the boundary layer suction groove is adjusted. Under a certain working condition, the outlet total pressure can be improved, flow field structure oscillation can be weakened, and the maximum outlet back pressure capable of being borne by a supersonic flow field in the isolation section can be improved at the same time.

Description

Technical Field

[0001] The present invention relates to the technical field of flow control, and in particular to a supersonic internal flow control device based on an inducer array. Background Art

[0002] Flow control is often used in supersonic flows. Flow control methods are divided into active control and passive control. Vortex generators are a common passive control method, boundary layer suction, vortex jets, wall bulges and plasma are common active flow control methods. Various combinations of the above basic flow control methods are also derived.

[0003] As the main area where oblique shock trains are generated, the role and influence of the isolation section are irreplaceable for the propulsion system of hypersonic aircraft. In the supersonic flow field of the isolation section, the oblique shock train is very sensitive to the flow field conditions upstream and the pressure disturbance downstream, which causes a series of complex flow phenomena in the isolation section. These complex shock waves and expansion waves form a mutual interference structure with the boundary layer, and are constantly reflected in the isolation section, which will make the overall flow field structure strongly asymmetric, and will eventually lead to the "surge" phenomenon. In addition, the self-excited oscillation phenomenon of the shock train in the flow field and the unsteady motion characteristics of the forward and backward movement are very likely to cause the internal flow field structure of the overall flow channel to oscillate. In order to minimize the influence of the oblique shock train on the air intake performance of the inlet and the stability of the isolation section structure, it is a very important topic for the current hypersonic aircraft propulsion technology to carry out research on shock train control. This complex oblique shock train structure is also widely present in structures such as supersonic ejectors and supersonic wind tunnel expansion sections.

[0004] Studies have shown that the use of either vortex generators (i.e. inducers) or boundary layer suction alone cannot significantly improve the flow field quality of the isolation section, while simultaneously increasing the total outlet pressure, weakening the flow field structure oscillation, and effectively increasing the maximum outlet back pressure that the supersonic flow field in the isolation section can withstand. Active flow control such as wall bulges and plasma will lead to a more complex structural design. Summary of the invention

[0005] In view of the above problems, the present invention provides a supersonic internal flow control device based on an inducer array, which can simultaneously increase the total outlet pressure, weaken the flow field structure oscillation, and increase the maximum outlet back pressure that the supersonic flow field in the isolation section can withstand under certain working conditions.

[0006] The present invention provides a supersonic internal flow control device based on an inducer array, comprising an equal straight section flow channel and an inducer array arranged in the equal straight section flow channel, wherein the inducer array comprises two groups of inducers symmetrically distributed on both sides of the center line of the equal straight section flow channel; each group of the inducers comprises a shock wave generator and a guide vane, wherein the guide vane is fixedly installed in the equal straight section flow channel, and the shock wave generator is rotatably installed in the equal straight section flow channel; a driving device connected to the shock wave generator for driving is arranged outside the equal straight section flow channel, and is used to change the angle between the leading edge of the shock wave generator and the flow direction; a boundary layer suction groove is respectively arranged on the side wall of the equal straight section flow channel corresponding to each group of the inducers, and the boundary layer suction groove is correspondingly arranged at a position between the trailing edge point of the shock wave generator and the leading edge point of the guide vane; a suction device connected to the boundary layer suction groove is arranged outside the equal straight section flow channel, and is used to adjust the outlet back pressure P1 of the boundary layer suction groove.

[0007] Specifically, the leading edge point of the shock generator is close to but does not enter the boundary layer of the flow field, and the local boundary layer thickness is set to δ0; the leading edge spacing of the two symmetrically distributed guide vanes is set to Y1, and the inlet height of the equal straight section flow channel is H1, then the ratio of Y1 to H1 is 0.64~0.7; in each group of the inducers, the longitudinal spacing between the leading edge points of the shock generator and the guide vanes is set to Y2, then Y2≈0.5*H1+δ0-0.5*Y1.

[0008] Specifically, the flow length of the shock wave generator is set to L1; the flow distance between the leading edge points of the shock wave generator and the guide vane in each group of the inducers is set to X1, and the ratio of X1 to L1 is 1.25.

[0009] Specifically, the shock wave generator is approximately quadrilateral in structure in a two-dimensional plane, and the angle between the leading edge of the shock wave generator and the flow direction is set as the airflow deflection angle θ0. In the flow field with the same Mach number, the airflow deflection angle θ0 is determined, and the angle between the shock wave front induced by the airflow deflection angle θ0 and the incoming flow direction is the shock wave angle β1. The shock wave angle β1 is the average value of the upper and lower separation shock wave angles in the flow-free control flow field with the same Mach number.

[0010] The angle between the inner side of the shock wave generator trailing edge and the incoming flow direction is set to θ1, and the angle between the outer side of the shock wave generator trailing edge and the incoming flow direction is set to θ2, and the ratio of θ1 to θ2 is 2 to 2.5;

[0011] The leading edge and the trailing edge on the inner side of the shock generator are processed by curve smoothing through continuous small external deflection angles to form a Prandtl-Meyer flow, so as to ensure that a continuous scattering expansion wave system is caused in the turning area.

[0012] Specifically, the ratio of the flow length L1 of the shock wave generator to the inlet height H1 of the equal straight section flow channel is 4:15; the leading edge flow length of the shock wave generator is set to L2, and the trailing edge flow length of the shock wave generator is set to L3, then the ratio of L2 to L3 is 2.0-2.2, so as to ensure that the leading edge of the shock wave generator can induce a stronger induced shock wave, and the trailing edge of the shock wave generator will not cause a stronger shock wave.

[0013] Specifically, the airflow deflection angle θ0 can be adjusted synchronously with the Mach number Ma, and the induction array has an airflow deflection angle θ0 with the best control effect under each Mach number Ma. In the three-dimensional structure diagram, the shock generator can be rotated around the rotation axis by a certain angle to achieve that the airflow deflection angle θ0 on the inner side of the leading edge of the shock generator can be controlled and confirmed.

[0014] The rotation axis of the shock wave generator is uniquely determined under normal working conditions, and the rotation axis of the shock wave generator is set at the upstream position of the leading edge point of the shock wave generator. The distance between the rotation axis of the shock wave generator and the leading edge of the shock wave generator is set to X2. X2 is affected by complex factors such as the Reynolds number, so an approximate algorithm is given:

[0015] In the two Mach number supersonic flow fields, the corresponding wall boundary layer thicknesses of the flow direction position coordinates of the leading edge of the shock generator are δ0 and δ1 respectively, and the corresponding airflow deflection angles on the inner side of the leading edge of the shock generator are θ0 and θ1 respectively. Then X2 is approximately (δ1-δ0) / arctan(θ1-θ0). Under the same Reynolds number and other factors, X2 at each supersonic Mach number is relatively consistent, so that after the rotation angle, the outer edge of the shock generator is close to but tries not to enter the boundary layer of the corresponding Mach number flow field.

[0016] Specifically, the shock wave generator and the equal straight section flow channel can also be fixedly installed. When fixedly installed, the leading edge airflow angle θ0 of the shock wave generator cannot be adjusted. At this time, the inducer array has a rated Mach number Ma for optimal control, that is, the design Mach number Ma. The present invention is based on the working condition of Mach number Ma being 2.5. As the flight Mach number deviates from the rated Mach number Ma, the control effect will decrease.

[0017] Specifically, the guide vane has an ultra-thin quadrilateral structure in a two-dimensional plane, the outer side of the leading edge of the guide vane is parallel to the flow direction, and the inner side of the trailing edge of the guide vane is parallel to the flow direction;

[0018] The ratio of the length L4 of the guide plate to the entrance height H1 of the equal straight section flow channel is 1, the ratio of the maximum thickness H2 of the guide plate to the entrance height H1 of the equal straight section flow channel is 1:30, and the ratio of the leading edge flow length L5 of the guide plate to the trailing edge flow length L6 of the guide plate is 2.5-3.2, so as to ensure that the flow direction of the flow field inside the leading edge of the guide plate and near the leading edge point of the guide plate is relatively consistent, and will not cause strong shock waves.

[0019] Specifically, the flow direction position of the leading edge point of the suction cavity inlet of the boundary layer suction groove is downstream of the flow direction position of the trailing edge point of the shock wave generator;

[0020] The inlet length of the suction cavity of the boundary layer suction groove is set to L7, and the flow distance from the trailing edge point of the shock wave generator to the leading edge point of the guide vane is X3, then L7 is less than or equal to X3;

[0021] The suction cavity depth of the boundary layer suction groove is set to D1, and the ratio of L7 to D1 is 0.5 to 0.8;

[0022] The rear edge of the boundary layer suction groove inlet is rounded, and the ratio of the rounding radius R1 to the suction inner cavity inlet length L7 is 0.25-0.35.

[0023] Specifically, the outlet back pressure P1 of the boundary layer suction groove is adjusted and determined by the suction device. For the same straight section flow channel, under the same incoming flow conditions, the outlet back pressure P1 of the boundary layer suction groove is uniquely determined, and the value of P1 is 0.9 to 1 times the inlet static pressure P2 of the boundary layer suction groove, so as to achieve the flow loss brought by the boundary layer suction groove to be almost zero under the condition of small back pressure ratio.

[0024] Specifically, the opposite side walls of the equal straight section flow channel are set to be symmetrical or approximately symmetrical, that is, the upper and lower side walls of the equal straight section flow channel are symmetrical or approximately symmetrical in the two-dimensional structure, and the two groups of opposite sides around the equal straight section flow channel are symmetrical or approximately symmetrical in the three-dimensional structure, that is, it can be simplified into a two-dimensional symmetrical equal straight section flow channel structure to achieve approximately symmetrical flow field structure when the back pressure ratio of the supersonic flow field is 0, which is a prerequisite for the inducer array to play a positive role; wherein, the wall surface of the equal straight section flow channel is of standard equivalue, and may also have a certain boundary layer correction angle, so as to achieve approximately equistraight boundary line of the boundary layer of the equal straight section flow channel when the back pressure ratio of the supersonic flow field is 0.

[0025] The supersonic internal flow control device based on the inducer array provided by the present invention can effectively improve the outlet total pressure recovery coefficient of the flow field within a specific back-pressure ratio range and suppress the non-steadiness of the flow field caused by shock wave oscillation by arranging a shock wave generator and a guide vane to form an inducer array; the back-pressure ratio range of the effective effect of the inducer array can be widened by arranging a boundary layer suction groove, and the adaptability to severe working conditions can be improved; the Mach number range of the effective effect of the inducer array can be improved by the adjustable design of the leading edge airflow angle of the shock wave generator; the adjustable design of the outlet back pressure of the boundary layer suction groove can adapt to different Mach numbers and reduce the flow loss caused by the flow control design.

[0026] Therefore, the supersonic internal flow control device based on the inducer array provided by the present invention can simultaneously achieve, under certain working conditions, increasing the total outlet pressure, weakening the flow field structure oscillation, and increasing the maximum outlet back pressure that the supersonic flow field in the isolation section can withstand. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] In order to more clearly illustrate the specific implementation methods of the present invention or the technical solutions in the prior art, the drawings required for use in the specific implementation methods or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some implementation methods of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0028] Figure 1 It is a schematic diagram of the overall structure of the supersonic internal flow control device based on the inducer array of the present invention;

[0029] Figure 2 It is a schematic structural diagram of the supersonic internal flow control device based on the inducer array in the flow control area of ​​the present invention;

[0030] Figure 3 It is a schematic structural diagram of a shock wave generator in a supersonic internal flow control device based on an inducer array of the present invention;

[0031] Figure 4 It is a schematic structural diagram of the shock wave generator in the supersonic inflow control device based on the inducer array of the present invention for adjusting the airflow angle inside the leading edge;

[0032] Figure 5 It is a schematic structural diagram of the guide vanes and boundary layer suction grooves in the supersonic internal flow control device based on the inducer array of the present invention;

[0033] Figure 6 It is a schematic diagram of the main shock wave flow field structure in the flow control area of ​​the supersonic internal flow control device based on the inducer array of the present invention.

[0034] Description of reference numerals:

[0035] 1: shock wave generator; 2: guide vane; 3: boundary layer suction groove; 4: straight section flow channel; 5: inducer array. DETAILED DESCRIPTION

[0036] The technical solution of the present invention will be clearly and completely described below in conjunction with the embodiments. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0037] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", "clockwise", "counterclockwise" and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as limiting the present invention.

[0038] In addition, the terms "first" and "second" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Therefore, the features defined as "first" and "second" may explicitly or implicitly include one or more of the features. In the description of the present invention, "multiple" means two or more, unless otherwise clearly and specifically defined. In addition, the terms "installed", "connected" and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, or it can be the internal connection of two components. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0039] like Figures 1 to 6 As shown, a supersonic inflow flow control device based on an inducer array according to an embodiment of the present invention comprises an equal straight section flow channel 4 and an inducer array 5 arranged in the equal straight section flow channel 4. The inducer array 5 comprises two groups of inducers symmetrically distributed on both sides of the center line of the equal straight section flow channel 4.

[0040] Among them, each group of inducers includes a shock wave generator 1 and a guide vane 2. The guide vane 2 is fixedly installed in the straight section flow channel 4. The shock wave generator 1 can be rotatably installed in the straight section flow channel 4. A driving device connected to the shock wave generator 1 is provided outside the straight section flow channel 4, which is used to change the angle between the leading edge of the shock wave generator 1 and the flow direction. The angle is the airflow deflection angle θ0, which means the angle at which the airflow is deflected after passing through the oblique shock wave.

[0041] Among them, a boundary layer suction groove 3 is provided on the side wall of the equal straight section flow channel 4 corresponding to each group of inducers, and the boundary layer suction groove 3 is correspondingly arranged at a position between the trailing edge point of the shock wave generator 1 and the leading edge point of the guide vane 2. A suction device connected to the boundary layer suction groove 3 is provided outside the equal straight section flow channel 4, which is used to adjust the outlet back pressure P1 of the boundary layer suction groove 3.

[0042] The embodiment of the present invention can effectively improve the outlet total pressure recovery coefficient of the flow field within a specific back pressure ratio range and suppress the non-steadiness of the flow field caused by shock wave oscillation by setting an inducer array 5 composed of a shock wave generator 1 and a guide vane 2; the back pressure ratio range of the effective effect of the inducer array can be widened by setting a boundary layer suction groove 3, and the adaptability to severe working conditions can be improved; the Mach number range of the effective effect of the inducer array can be improved by the adjustable design of the leading edge airflow angle of the shock wave generator 1; and the outlet back pressure of the boundary layer suction groove 3 can be adjusted to adapt to different Mach numbers and reduce the flow loss caused by the flow control design.

[0043] Therefore, the supersonic internal flow control device based on the inducer array in the embodiment of the present invention can simultaneously achieve, under certain working conditions, increasing the total outlet pressure, weakening the flow field structure oscillation, and increasing the maximum outlet back pressure that the supersonic flow field in the isolation section can withstand.

[0044] Specifically, the leading edge point of the shock generator 1 is close to but does not enter the boundary layer of the flow field, and the thickness of the local boundary layer is δ0.

[0045] The leading edge spacing of the two symmetrically distributed guide vanes 2 is set to Y1, and the inlet height of the straight section flow channel 4 is set to H1, then the ratio of Y1 to H1 is 0.64 to 0.7. In each group of inducers, the longitudinal spacing between the leading edge points of the shock wave generator 1 and the guide vane 2 is set to Y2, then Y2 is approximately 0.5*H1+δ0-0.5*Y1.

[0046] Specifically, the flow direction length of the shock wave generator 1 is set to L1. In each group of inductors, the flow direction spacing between the leading edge points of the shock wave generator 1 and the guide vane 2 is set to X1, and the ratio of X1 to L1 is about 1.25.

[0047] Specifically, the shock generator 1 is approximately quadrilateral in a two-dimensional plane, and the angle between the leading edge of the shock generator 1 and the flow direction is set as the airflow deflection angle θ0, wherein the design of the airflow deflection angle θ0 must satisfy the induction of a specific shock wave angle to control the stability of the overall flow field structure. In the same Mach number flow field, the airflow deflection angle θ0 is determined, and the angle between the shock wave front induced by the airflow deflection angle θ0 and the incoming flow direction is the shock wave angle β1, and the shock wave angle β1 is the average value of the upper and lower separation shock wave angles in the no-flow control flow field of the same Mach number.

[0048] The angle between the inner side of the trailing edge of the shock wave generator 1 and the incoming flow direction is set to θ1, and the angle between the outer side of the trailing edge of the shock wave generator 1 and the incoming flow direction is set to θ2, then the ratio of θ1 to θ2 is 2 to 2.5.

[0049] The airflow angle of the outer wall of the trailing edge of the shock generator 1 is designed to be small, forming an expansion flow channel with the upper (lower) wall of the main channel, which inhibits the shock wave of the trailing edge from oscillating over the shock generator 1. The airflow angle of the inner wall of the trailing edge of the shock generator 1 is designed to be a certain angle, and the induced oblique shock wave can effectively comb the flow field on both sides of the main channel.

[0050] The leading edge and the trailing edge on the inner side of the shock generator 1 are processed into a curve smoothing process through continuous small external deflection angles to form a Prandtl-Meyer flow, so as to ensure that a continuous scattering expansion wave system is caused in the turning area.

[0051] Specifically, the ratio of the flow length L1 of the shock generator 1 to the inlet height H1 of the straight section flow channel 4 is about 4:15. The leading edge flow length of the shock generator 1 is set to L2, and the trailing edge flow length of the shock generator 1 is set to L3, then the ratio of L2 to L3 is 2.0-2.2, so as to ensure that the leading edge of the shock generator 1 can induce a strong induced shock wave, and the trailing edge of the shock generator 1 will not cause a strong shock wave.

[0052] Specifically, the airflow deflection angle θ0 can be adjusted synchronously with the Mach number Ma, and the induction array 5 has an airflow deflection angle θ0 with the best control effect at each Mach number Ma. In the three-dimensional structure diagram, the shock generator 1 can be rotated around the rotation axis by a certain angle to achieve that the airflow deflection angle θ0 on the inner side of the leading edge of the shock generator 1 can be controlled and confirmed.

[0053] Among them, the rotation axis of the shock wave generator 1 is uniquely determined under normal working conditions, and the rotation axis of the shock wave generator 1 is set at the upstream position of the leading edge point of the shock wave generator 1. The distance between the rotation axis of the shock wave generator 1 and the leading edge of the shock wave generator 1 is set to X2. X2 is affected by complex factors such as the Reynolds number, so the following approximate algorithm is given:

[0054] In the two Mach number supersonic flow fields, the corresponding wall boundary layer thicknesses of the flow direction position coordinates of the leading edge of the shock generator 1 are δ0 and δ1 respectively, and the corresponding airflow deflection angles on the inner side of the leading edge of the shock generator 1 are θ0 and θ1 respectively. Then X2 is approximately (δ1-δ0) / arctan(θ1-θ0). Under the same Reynolds number and other factors, X2 at each supersonic Mach number is relatively consistent, so that after the rotation angle, the outer edge of the shock generator 1 is close to but tries not to enter the boundary layer of the corresponding Mach number flow field.

[0055] In this embodiment, when the design is based on Ma2.5, when the incoming flow is Ma2.9, the shock wave generator 1 is deflected inward by a certain small angle to achieve a larger airflow angle and a thicker boundary layer at Ma2.9.

[0056] In addition, the shock generator 1 and the straight section flow channel 4 can also be fixedly installed. When fixedly installed, the leading edge airflow angle θ0 of the shock generator 1 cannot be adjusted. At this time, the inducer array 5 has a rated Mach number Ma for optimal control, that is, the design Mach number Ma. The present invention is based on the working condition of Mach number Ma being 2.5. As the flight Mach number deviates from the rated Mach number Ma, the control effect will decrease.

[0057] Specifically, the guide vane 2 is an ultra-thin quadrilateral structure in a two-dimensional plane, the outer side of the leading edge of the guide vane 2 is parallel to the flow direction, and the inner side of the trailing edge of the guide vane 2 is parallel to the flow direction.

[0058] Among them, the ratio of the length L4 of the guide vane 2 to the inlet height H1 of the equal straight section flow channel 4 is 1, the ratio of the maximum thickness H2 of the guide vane 2 to the inlet height H1 of the equal straight section flow channel 4 is approximately 1:30, and the ratio of the leading edge flow length L5 of the guide vane 2 to the trailing edge flow length L6 of the guide vane 2 is 2.5-3.2, so as to ensure that the flow direction of the flow field inside the leading edge of the guide vane 2 and near the leading edge point of the guide vane 2 is relatively consistent, and will not cause strong shock waves.

[0059] That is, the length L4 of the guide vane 2 is designed to meet certain requirements to balance the pressure between the central supersonic mainstream and the flow on both sides. The leading edge of the guide vane 2 is folded inward at a small angle, which neither causes shock waves nor guides more fluid to flow through the center. Because there are many wave systems on both sides, the total pressure loss of the fluid flowing to both sides will be greater. The inner side of the trailing edge is designed to be flat in the flow direction in order to suppress the oblique shock wave caused by the deflection of the airflow in the area behind the trailing edge point from spreading to the inner upstream side of the guide vane 2.

[0060] Specifically, the flow direction position of the leading edge point of the suction cavity inlet of the boundary layer suction groove 3 is downstream of the flow direction position of the trailing edge point of the shock wave generator 1, which inhibits the separation shock wave from developing upstream. When the subsonic flow after the separation shock wave passes through the boundary layer suction groove 3, the high-pressure subsonic flow will preferentially flow to the outlet of the boundary layer suction groove 3.

[0061] Assuming that the entrance length of the suction cavity of the boundary layer suction groove 3 is L7, and the flow distance from the trailing edge point of the shock generator 1 to the leading edge point of the guide vane 2 is X3, then L7 is less than or equal to X3.

[0062] The suction cavity depth of the boundary layer suction groove 3 is set to D1, and the ratio of L7 to D1 is 0.5-0.8.

[0063] The rear edge of the entrance of the boundary layer suction groove 3 is rounded, and the ratio of the rounding radius R1 to the entrance length L7 of the suction inner cavity is 0.25-0.35.

[0064] Specifically, the outlet back pressure P1 of the boundary layer suction groove 3 is adjusted and determined by the suction device. For the same straight section flow channel 4, under the same incoming flow conditions, the outlet back pressure P1 of the boundary layer suction groove 3 is uniquely determined, and the value of P1 is 0.9 to 1 times the inlet static pressure P2 of the boundary layer suction groove 3, so as to achieve a flow loss of nearly zero in the boundary layer suction groove 3 under the condition of a small back pressure ratio.

[0065] Specifically, the opposite side walls of the straight section flow channel 4 are set to be symmetrical or approximately symmetrical, that is, the upper and lower side walls of the straight section flow channel 4 are symmetrical or approximately symmetrical in the two-dimensional structure, and the two groups of opposite sides around the straight section flow channel 4 are symmetrical or approximately symmetrical in the three-dimensional structure, that is, it can be simplified to a two-dimensional symmetrical straight section flow channel 4 structure to achieve approximately symmetrical flow field structure when the back pressure ratio of the supersonic flow field flowing through the flow is 0, which is a prerequisite for the inducer array 5 to play a positive role.

[0066] The wall surface of the straight section flow channel 4 is of standard equal value and may also have a certain boundary layer correction angle to achieve that when the back pressure ratio of the supersonic flow field is 0, the boundary layer boundary line of the straight section flow channel 4 is approximately equal straight.

[0067] like Figure 6 As shown, when the airflow flows from the upstream inlet to the flow control area, the airflow is first deflected at the leading edge of the inner side of the shock generator 1, and an oblique shock wave is induced at the same time. The airflow is deflected by the induced shock wave, so that the supersonic airflow is concentrated in the center, and then the airflow deflection angle decreases after the continuous expansion wave system. When it reaches the leading edge of the guide vane 2, the airflow is divided into two, and the supersonic mainstream flows through the center, and flows out of the flow control area after passing through the induced shock wave and its multi-stage emission shock wave. Since the induced shock wave is a relatively stable and symmetrical shock wave, the total pressure loss of the central supersonic mainstream is small and the oscillation is weak. The fluid on both sides of the main channel is deflected at the trailing edge of the shock generator 1, forming a trailing edge shock wave. The trailing edge shock wave and its reflected shock wave effectively comb the flow field structure of the flow channels on both sides. Under high back pressure conditions, the high-pressure subsonic flow in the downstream separation zone pushes the separation shock wave to spread upstream, and is stably suppressed at the boundary layer suction groove 3, providing anti-back pressure capability and stability of the separation shock wave.

[0068] In summary, the supersonic internal flow control device based on the inducer array of the embodiment of the present invention, through the combined flow control design of the inducer array 5 and the boundary layer suction groove 3, mainly realizes the effective influence on the supersonic flow field by creating, controlling and stabilizing the shock wave, thereby improving the total pressure recovery coefficient at the flow field outlet, improving the anti-back pressure ability of the overall flow channel, suppressing the development of shock waves and separation zones upstream, weakening the oscillation of shock waves, and realizing effective adaptation to multiple Mach numbers by adjusting the leading edge airflow angle of the shock wave generator 1 and the outlet pressure of the boundary layer suction groove 3.

[0069] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or replace some or all of the technical features therein with equivalents. However, these modifications or replacements 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 supersonic inflow flow control device based on an inducer array, characterized in that: The invention comprises an even straight section flow channel (4) and an inductor array (5) arranged in the even straight section flow channel (4), wherein the inductor array (5) comprises two groups of inductors symmetrically distributed on both sides of the center line of the even straight section flow channel (4); each group of inductors comprises a shock wave generator (1) and a guide vane (2), wherein the guide vane (2) is fixedly installed in the even straight section flow channel (4), the shock wave generator (1) is rotatably installed in the even straight section flow channel (4), and a rotatable inductor (2) is provided outside the even straight section flow channel (4) and is rotatably connected to the shock wave generator (1). A driving device connected to the drive is used to change the angle between the leading edge of the shock wave generator (1) and the flow direction; a boundary layer suction groove (3) is provided on the side wall of the equal straight section flow channel (4) corresponding to each group of the inducers, and the boundary layer suction groove (3) is correspondingly arranged at a position between the trailing edge point of the shock wave generator (1) and the leading edge point of the guide vane (2); a suction device connected to the boundary layer suction groove (3) is provided outside the equal straight section flow channel (4) for adjusting the outlet back pressure P1 of the boundary layer suction groove (3).

2. The supersonic inflow flow control device based on an inductor array according to claim 1, characterized in that: The leading edge point of the shock wave generator (1) is close to but does not enter the boundary layer of the flow field, and the local boundary layer thickness is set to δ0; the leading edge spacing of the two symmetrically distributed guide vanes (2) is set to Y1, and the inlet height of the equal straight section flow channel (4) is set to H1, then the ratio of Y1 to H1 is 0.64-0.7; in each group of the inducers, the longitudinal spacing between the leading edge points of the shock wave generator (1) and the guide vanes (2) is set to Y2, then Y2≈0.5*H1+δ0-0.5*Y1.

3. The supersonic inflow control device based on an inductor array according to claim 1, characterized in that: The flow direction length of the shock wave generator (1) is set to L1; the flow direction spacing between the leading edge points of the shock wave generator (1) and the guide vane (2) in each group of the inductors is set to X1, and the ratio of X1 to L1 is 1.

25.

4. The supersonic inflow control device based on an inductor array according to claim 1, characterized in that: The shock wave generator (1) is approximately quadrilateral in structure in a two-dimensional plane. The angle between the leading edge of the shock wave generator (1) and the flow direction is set as the airflow deflection angle θ0. In a flow field with the same Mach number, the airflow deflection angle θ0 is determined. The angle between the shock wave front induced by the airflow deflection angle θ0 and the incoming flow direction is the shock wave angle β1. The shock wave angle β1 is the average value of the upper and lower separation shock wave angles in a flow-free control flow field with the same Mach number. The angle between the inner side of the trailing edge of the shock wave generator (1) and the incoming flow direction is set to θ1, and the angle between the outer side of the trailing edge of the shock wave generator (1) and the incoming flow direction is set to θ2, and the ratio of θ1 to θ2 is 2 to 2.5; The leading edge and the trailing edge on the inner side of the shock wave generator (1) are processed into a curve smoothing process through continuous small external deflection angles to form a Prandtl-Meyer flow, so as to ensure that a continuous scattering expansion wave system is induced in the turning region.

5. The supersonic inflow flow control device based on an inductor array according to claim 1, characterized in that: The ratio of the flow length L1 of the shock wave generator (1) to the inlet height H1 of the straight section flow channel (4) is 4:15; the leading edge flow length of the shock wave generator (1) is set to L2, and the trailing edge flow length of the shock wave generator (1) is set to L3, then the ratio of L2 to L3 is 2.0-2.2, so as to ensure that the leading edge of the shock wave generator (1) can induce a relatively strong induced shock wave, and the trailing edge of the shock wave generator (1) will not cause a relatively strong shock wave.

6. The supersonic inflow control device based on an inductor array according to claim 1, characterized in that: The rotation axis of the shock generator (1) is arranged at an upstream position of the leading edge point of the shock generator (1), and the distance between the rotation axis of the shock generator (1) and the leading edge of the shock generator (1) is set to X2; the wall boundary layer thicknesses of the corresponding shock generator (1) leading edge flow direction position coordinates in two Mach number supersonic flow fields are set to be δ0 and δ1 respectively, and the corresponding airflow deflection angles on the inner side of the leading edge of the shock generator (1) are set to be θ0 and θ1 respectively, then X2 is approximately (δ1-δ0) / arctan(θ1-θ0).

7. The supersonic inflow control device based on an inductor array according to claim 1, characterized in that: The guide vane (2) presents an ultra-thin quadrilateral structure in a two-dimensional plane, the outer side of the leading edge of the guide vane (2) is parallel to the flow direction, and the inner side of the trailing edge of the guide vane (2) is parallel to the flow direction; The ratio of the length L4 of the guide plate (2) to the inlet height H1 of the straight section flow channel (4) is 1, the ratio of the maximum thickness H2 of the guide plate (2) to the inlet height H1 of the straight section flow channel (4) is 1:30, and the ratio of the leading edge flow direction length L5 of the guide plate (2) to the trailing edge flow direction length L6 of the guide plate (2) is 2.5 to 3.

2.

8. The supersonic inflow control device based on an inductor array according to claim 1, characterized in that: The flow direction position of the leading edge point of the suction cavity inlet of the boundary layer suction groove (3) is downstream of the flow direction position of the trailing edge point of the shock wave generator (1); The length of the suction cavity entrance of the boundary layer suction groove (3) is set to be L7, and the flow distance from the trailing edge point of the shock wave generator (1) to the leading edge point of the guide vane (2) is set to be X3, then L7 is less than or equal to X3; The suction inner cavity depth of the boundary layer suction groove (3) is set to D1, and the ratio of L7 to D1 is 0.5 to 0.8; The rear edge of the entrance of the boundary layer suction groove (3) is rounded, and the ratio of the rounding radius R1 to the entrance length L7 of the suction inner cavity is 0.25-0.

35.

9. The supersonic inflow control device based on an inductor array according to claim 1, characterized in that: For the same straight section flow channel (4), under the same incoming flow conditions, the outlet back pressure P1 of the boundary layer suction groove (3) is uniquely determined, and the value of P1 is 0.9 to 1 times the inlet static pressure P2 of the boundary layer suction groove (3), so as to achieve a flow loss of nearly zero caused by the boundary layer suction groove (3) under the condition of a small back pressure ratio.

10. The supersonic inflow control device based on an inductor array according to claim 1, characterized in that: The walls of the equal straight section flow channel (4) on opposite sides are set to be symmetrical or approximately symmetrical, so as to achieve an approximately symmetrical flow field structure when the back pressure ratio of the supersonic flow field flowing through the flow is 0; the wall surface of the equal straight section flow channel (4) is of standard equal value, so as to achieve an approximately equal straight boundary line of the boundary layer of the equal straight section flow channel (4) when the back pressure ratio of the supersonic flow field is 0.

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