An ultrasonic flow control device based on an array of inductors

Abstract

The application provides an ultrasonic internal flow flow control device based on an inducer array, which comprises an equal-length flow channel and an inducer array, the inducer array comprises two groups of inducers which are symmetrically distributed on both sides of the center line of the equal-length flow channel, each group of inducers comprises a shock generator and a guide vane, the guide vane is fixedly installed in the equal-length flow channel, and the shock generator is rotatably installed in the equal-length flow channel; a boundary layer suction groove is arranged on the side wall of the equal-length flow channel and corresponds to each group of inducers, the boundary layer suction groove is arranged at a position corresponding to the position between the trailing edge point of the shock generator and the leading edge point of the guide vane, and a suction device which is connected with the boundary layer suction groove is arranged outside the equal-length flow channel and is used for adjusting the back pressure of the outlet of the boundary layer suction groove. The application can simultaneously realize the following functions under certain working conditions: improving the total pressure at the outlet, weakening the structural shock of the flow field, and improving the maximum back pressure that can be borne by the ultrasonic flow field in the isolation section.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of flow control, in particular to an ultrasonic internal flow flow control device based on an array of inducers. BACKGROUND

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

[0003] As the main area of oblique shock train generation, the role and influence of the isolation section are irreplaceable for the high-speed aircraft propulsion system. In the supersonic flow field of the isolation section, the oblique shock train is very sensitive to the upstream flow field conditions and the downstream pressure disturbance, thereby causing a series of complex flow phenomena in the isolation section. These complex shock waves and expansion waves form an interference structure with the boundary layer, which continuously reflects in the isolation section, causing strong asymmetry in the overall flow field structure, and ultimately leading to the generation of "surge" phenomenon. In addition, the self-excited oscillation phenomenon and the non-stationary movement characteristics of the oblique shock train in the flow field are very likely to cause the internal flow field structure of the overall flow passage to oscillate. In order to minimize the influence of the oblique shock train on the air intake performance and the stability of the isolation section structure, research on the control of the oblique shock train is a very important topic in the current high-speed aircraft propulsion technology. In supersonic ejector devices and supersonic wind tunnel diffuser structures, there are also complex oblique shock train structures.

[0004] Research shows that using vortex generators (inducers) alone, boundary layer suction alone, or any single method cannot significantly improve the flow field quality of the isolation section, while achieving improved outlet total pressure, weakening of flow field structure oscillation, and effective increase of the maximum outlet back pressure that the supersonic flow field in the isolation section can withstand. Active flow control such as wall bulge and plasma will bring more complex structural design. SUMMARY

[0005] In view of the above problems, the present application provides an ultrasonic internal flow flow control device based on an array of inducers, which can simultaneously achieve improved outlet total pressure, weakening of flow field structure oscillation, and increased maximum outlet back pressure that the supersonic flow field in the isolation section can withstand under certain working conditions.

[0006] The application provides an ultrasonic internal flow flow control device based on an inducer array, which comprises an equal-length flow channel and an inducer array arranged in the equal-length flow channel, the inducer array comprises two groups of inducers symmetrically arranged on both sides of the center line of the equal-length flow channel; each group of the inducers comprises a shock generator and a flow vane, the flow vane is fixedly arranged in the equal-length flow channel, the shock generator is rotatably arranged in the equal-length flow channel, a driving device connected with the shock generator is arranged outside the equal-length flow channel, and the driving device is used for changing the included angle between the leading edge of the shock generator and the flow direction; a boundary layer suction groove is arranged on the side wall of the equal-length flow channel and corresponds to each group of the inducers, the boundary layer suction groove is arranged at the position between the trailing edge point of the shock generator and the leading edge point of the flow vane, and a suction device in communication with the boundary layer suction groove is arranged outside the equal-length flow channel, and the suction device is used for adjusting the back pressure P1 of the outlet of the boundary layer suction groove.

[0007] Specifically, the leading edge point of the shock generator is close to but not in the flow field boundary layer, the local boundary layer thickness is δ0, the distance between the leading edges of the two symmetrically arranged flow vanes is Y1, the inlet height of the equal-length flow channel is H1, the ratio of Y1 to H1 is 0.64-0.7, and the longitudinal distance between the leading edge points of the shock generator and the flow vane in each group of the inducers is Y2, Y2≈0.5*H1+δ0-0.5*Y1.

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

[0009] Specifically, the shock generator has an approximate quadrilateral structure in a two-dimensional plane, the included angle between the leading edge of the shock generator and the flow direction is the airflow turning angle θ0, the airflow turning angle θ0 is determined in the same Mach number flow field, the included angle between the shock wave front induced by the airflow turning angle θ0 and the flow direction is the shock angle β1, and the shock angle β1 is the average of the upper and lower separation shock angles in the flow field without flow control at the same Mach number.

[0010] The included angle between the inner side of the trailing edge of the shock generator and the flow direction is θ1, the included angle between the outer side of the trailing edge of the shock generator and the flow direction is θ2, and the ratio of θ1 to θ2 is 2-2.5.

[0011] The leading edge and the trailing edge of the inner side of the shock generator are smoothly processed by a continuous small external bias angle to form a Prandtl-Meyer flow, so that the continuous scattering expansion wave system is caused in the turning area.

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

[0013] Specifically, the airflow angle θ0 can be adjusted synchronously with the Mach number Ma, and the airflow angle θ0 has an optimal control effect of the inducer array at each Mach number Ma. In the three-dimensional structure diagram of the shock generator, the shock generator can be rotated by a certain angle around the rotation axis, so that the airflow angle θ0 inside the leading edge of the shock generator can be controlled and confirmed.

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

[0015] In two supersonic flow fields with different Mach numbers, the wall boundary layer thickness of the corresponding shock generator leading edge streamwise position coordinates is δ0 and δ1 respectively, and the airflow angle inside the corresponding shock generator leading edge is θ0 and θ1 respectively, and X2 is approximately (δ1-δ0) / arctan(θ1-θ0). Under the same Reynolds number and other factors, X2 under each supersonic Mach number is relatively consistent, so that after the rotation angle, the outer side of the shock generator is close to but does not enter the boundary layer of the corresponding Mach number flow field as much as possible.

[0016] Specifically, the shock generator and the straight section flow passage can also be fixedly installed, and when fixedly installed, the airflow angle θ0 of the leading edge of the shock generator is not adjustable, and at this time, the inducer array has a rated Mach number Ma, i.e. a design Mach number Ma, at which the control effect is optimal. 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 ultrathin quadrilateral structure in a two-dimensional plane, the leading edge outside of the guide vane is parallel to the flow direction, and the trailing edge inside of the guide vane is parallel to the flow direction.

[0018] The ratio of the length L4 of the guide vane to the inlet height H1 of the straight section flow channel is 1, the ratio of the maximum thickness H2 of the guide vane to the inlet height H1 of the straight section flow channel is 1:30, and the ratio of the front edge flow direction length L5 of the guide vane to the rear edge flow direction length L6 of the guide vane is 2.5-3.2, so as to ensure that the front edge inner side of the guide vane is consistent with the flow direction of the flow field near the front edge point of the guide vane, and a strong shock wave is not caused.

[0019] Specifically, the boundary layer suction groove suction inner cavity inlet front edge point flow direction position is downstream of the rear edge point flow direction position of the shock generator.

[0020] The length of the boundary layer suction groove suction inner cavity inlet is set as L7, the rear edge point of the shock generator to the front edge point of the guide vane is X3, and L7 is less than or equal to X3.

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

[0022] The boundary layer suction groove inlet rear edge 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 determined by the suction device. For the same straight section flow channel, under the same incoming flow condition, the outlet back pressure P1 of the boundary layer suction groove is uniquely determined, and the value of P1 is 0.9-1 times the inlet static pressure P2 of the boundary layer suction groove, so as to realize that the flow loss of the boundary layer suction groove is nearly zero under the condition of small back pressure ratio.

[0024] Specifically, the relative two side walls of the straight section flow channel are symmetrical or approximately symmetrical, that is, the upper and lower two side walls of the straight section flow channel are symmetrical or approximately symmetrical in two-dimensional structure, and the two groups of opposite faces of the straight section flow channel are symmetrical or approximately symmetrical in three-dimensional structure, that is, it can be simplified as a two-dimensional symmetrical straight section flow channel structure, so as to realize that the flow field structure is approximately symmetrical when the supersonic flow field is at a back pressure ratio of 0, which is a prerequisite for the inducer array to play an active role; wherein the wall surface of the straight section flow channel is a standard equivalent, and there can be a certain boundary layer correction angle, so as to realize that the boundary layer demarcation line of the straight section flow channel is approximately straight when the supersonic flow field is at a back pressure ratio of 0.

[0025] The supersonic internal flow flow control device based on the inducer array provided by the application can effectively improve the outlet total pressure recovery coefficient of the flow field in a specific back pressure ratio range, and inhibit the flow field unsteadiness caused by shock wave oscillation, by setting the shock generator and the guide vane to form the inducer array; the effective action back pressure ratio range of the inducer array can be widened, and the adaptation range to severe working conditions can be improved, by setting the boundary layer suction groove; the Mach number range of the effective action of the inducer array is improved, by the adjustable design of the airflow deflection angle of the leading edge of the shock generator; the different Mach numbers can be adapted, and the flow loss caused by the flow control design can be reduced, by the adjustable design of the outlet back pressure of the boundary layer suction groove.

[0026] Therefore, the supersonic internal flow flow control device based on the inducer array provided by the application can simultaneously realize the improvement of the outlet total pressure, the weakening of the flow field structure oscillation, and the improvement of the maximum outlet back pressure that can be borne by the supersonic flow field in the isolation section under certain working conditions. BRIEF DESCRIPTION OF DRAWINGS

[0027] In order to more clearly illustrate the specific embodiments of the application or the technical solutions in the prior art, the drawings needed to be used in the specific embodiments or the prior art description will be briefly introduced below. Obviously, the drawings in the following description are some embodiments of the application, and other drawings can also be obtained by those skilled in the art without creative labor on the basis of these drawings.

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

[0029] Figure 2 It is a schematic diagram of the structure of the supersonic internal flow flow control device based on the inducer array of the application in the flow control area.

[0030] Figure 3 It is a schematic diagram of the structure of the shock generator in the supersonic internal flow flow control device based on the inducer array of the application.

[0031] Figure 4 It is a schematic diagram of the structure of the shock generator in the supersonic internal flow flow control device based on the inducer array of the application, which adjusts the airflow deflection angle inside the leading edge.

[0032] Figure 5 It is a schematic diagram of the structure of the guide vane and the boundary layer suction groove in the supersonic internal flow flow control device based on the inducer array of the application.

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

[0034] Reference numerals:

[0035] 1: Shock generator; 2: Guide vane; 3: Boundary layer suction slot; 4: Straight section flow passage; 5: Inducer array. DETAILED DESCRIPTION

[0036] The technical solutions of the present application will be described below in conjunction with the embodiments. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work fall within the scope of protection of the present application.

[0037] In the description of the present application, it should be understood that the terms "center", "longitudinal", "transverse", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise", and the like indicate the orientation or positional relationship based on the orientation or positional relationship shown in the drawings, and are only for the purpose of facilitating the description of the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation of the present application.

[0038] In addition, the terms "first", "second" are only 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 with "first", "second" can explicitly or implicitly include one or more of the features. In the description of the present application, "a plurality of" means two or more, unless otherwise explicitly specified. In addition, the terms "mounting", "connecting", "connecting" should be broadly understood, for example, it can be fixedly connected, or it can be detachably connected, or integrally connected; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium, or it can be the communication between two elements. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0039] As Figures 1 to 6 shown, the supersonic internal flow flow control device based on the inducer array of the embodiment of the present application includes a straight section flow passage 4 and an inducer array 5 arranged in the straight section flow passage 4, and the inducer array 5 includes two groups of inducers symmetrically distributed on both sides of the center line of the straight section flow passage 4.

[0040] Wherein, each group of inducer comprises a shock generator 1 and a guide vane 2, the guide vane 2 is fixedly installed in the straight section flow channel 4, the shock generator 1 is rotatably installed in the straight section flow channel 4, a driving device is arranged outside the straight section flow channel 4 and is drivingly connected with the shock generator 1, and the driving device is used for changing the included angle between the leading edge of the shock generator 1 and the flow direction, the included angle is the airflow deflection angle θ0, and the airflow deflection angle θ0 means the angle of airflow deflection after passing through the oblique shock wave.

[0041] Wherein, a boundary layer suction groove 3 is arranged on the side wall of the straight section flow channel 4 and corresponds to each group of inducer, the boundary layer suction groove 3 is correspondingly arranged at the position between the trailing edge point of the shock generator 1 and the leading edge point of the guide vane 2, and a suction device is arranged outside the straight section flow channel 4 and is connected with the boundary layer suction groove 3, and the suction device is used for adjusting the outlet back pressure P1 of the boundary layer suction groove 3.

[0042] The embodiment of the present application can effectively improve the outlet total pressure recovery coefficient of the flow field in a specific back pressure ratio range, and inhibit the flow field unsteadiness caused by shock wave oscillation by arranging the inducer array 5 composed of the shock generator 1 and the guide vane 2; the effective action back pressure ratio range of the inducer array can be widened, and the adaptation range to severe working conditions can be improved by arranging the boundary layer suction groove 3; the Mach number range of the effective action of the inducer array is improved by the adjustable design of the airflow deflection angle of the leading edge of the shock generator 1; the flow control design can be adapted to different Mach numbers, and the flow loss caused by the flow control design can be reduced by the adjustable design of the outlet back pressure of the boundary layer suction groove 3.

[0043] Therefore, the supersonic internal flow flow control device based on the inducer array can simultaneously realize the improvement of the outlet total pressure, the weakening of the flow field structure oscillation and the improvement of the maximum outlet back pressure that can be borne by the supersonic flow field in the isolation section under certain working conditions.

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

[0045] The ratio of Y1 to H1 is 0.64-0.7, Y1 is the front edge distance between the two guide vanes 2 arranged in a symmetrical manner, and H1 is the inlet height of the straight section flow channel 4.

[0046] Specifically, the flow direction length of the shock generator 1 is L1. X1 is the flow direction distance between the leading edge points of the shock generator 1 and the guide vane 2 in each group of inducer, and the ratio of X1 to L1 is about 1.25.

[0047] Specifically, the shock generator 1 is in the form of a quadrilateral in a two-dimensional plane, and an included angle between a leading edge of the shock generator 1 and a flow direction is set as an airflow deflection angle θ0, wherein the airflow deflection angle θ0 is designed to satisfy a specific shock angle to control stability of an overall flow field structure. In the same Mach number flow field, the airflow deflection angle θ0 is determined, an included angle between a shock wave front induced by the airflow deflection angle θ0 and the flow direction is a shock angle β1, and the shock angle β1 is an average of upper and lower separation shock angles in a flow control-free flow field of the same Mach number.

[0048] An included angle between an inner side of a trailing edge of the shock generator 1 and the flow direction is set as θ1, and an included angle between an outer side of the trailing edge of the shock generator 1 and the flow direction is set as θ2, and a ratio of θ1 to θ2 is 2-2.5.

[0049] In the formula, an outer side wall surface airflow deflection angle of the shock generator 1 is designed as a small angle, and an expansion type flow channel is formed on an upper (lower) wall surface of the main flow channel to suppress a trailing edge shock oscillation from passing through the shock generator 1. An inner side wall surface airflow deflection angle of the shock generator 1 is designed as a certain angle, and an induced oblique shock can effectively comb flow fields on both sides of the main flow channel.

[0050] In the formula, the leading edge and the trailing edge of the shock generator 1 are smoothly processed by a continuous small external bias angle to form a Prandtl-Meyer flow to ensure that a continuous scattered expansion wave system is caused in a turning region.

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

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

[0053] In the formula, the rotation axis of the shock generator 1 is usually uniquely determined, and the rotation axis of the shock generator 1 is arranged at a position upstream of a leading edge point of the shock generator 1. A distance between the rotation axis of the shock generator 1 and the leading edge of the shock generator 1 is set as X2, and X2 is affected by a Reynolds number and other complex factors, and therefore the following approximate algorithm is given.

[0054] Assuming two Mach number supersonic flow fields, the corresponding wall boundary layer thicknesses at the flow direction position coordinates of the leading edge of shock generator 1 are δ0 and δ1 respectively, and the corresponding airflow deflection angles inside the leading edge of 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 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 has 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 direction length L5 of the guide vane 2 to the trailing edge flow direction length L6 of the guide vane 2 is 2.5 to 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 no strong shock wave will be caused.

[0059] That is, the length L4 of the guide vane 2 must be 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 fluid flowing to both sides will be greater. The inner side of the trailing edge is designed to be flat in the direction of flow 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 leading edge of the suction cavity inlet of the boundary layer suction slot 3 is located downstream of the trailing edge of the shock generator 1, suppressing the upstream development of the separation shock wave. When the subsonic flow after the separation shock wave passes through the boundary layer suction slot 3, the high-pressure subsonic flow preferentially flows to the outlet of the boundary layer suction slot 3.

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

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

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

[0064] Specifically, the outlet back pressure P1 of the boundary layer suction slot 3 is adjusted and determined by the suction device. For the same straight section of flow channel 4 and under the same incoming flow conditions, the outlet back pressure P1 of the boundary layer suction slot 3 is uniquely determined. The value of P1 is 0.9 to 1 times the inlet static pressure P2 of the boundary layer suction slot 3, achieving near-zero flow loss due to the boundary layer suction slot 3 under conditions 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 an approximately symmetrical flow field structure when the back pressure ratio of the supersonic flow field passing 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, it first deflects at the inner leading edge of shock generator 1, inducing an oblique shock wave. The airflow is deflected by the induced shock wave, causing the supersonic airflow to concentrate toward the center. Then, through the continuous expansion wave system, the airflow deflection angle decreases. When it reaches the leading edge of guide vane 2, the airflow is split into two. The supersonic main flow flows through the center, passes through the induced shock wave and its multi-stage emission shock wave, and then flows out of the flow control area. Because the induced shock wave is a relatively stable and symmetrical shock wave, the total pressure loss of the central supersonic main flow is small and the oscillation is weak. The fluid on both sides of the main channel is deflected at the trailing edge of shock generator 1, forming a trailing edge shock wave. The trailing edge shock wave and its reflected shock wave effectively organize the flow field structure of the two sides of the channel. Under high back pressure conditions, the high-pressure subsonic flow in the downstream separation zone pushes the separation shock wave to propagate upstream, where it is stably suppressed at the boundary layer suction groove 3, providing back pressure resistance and separation shock wave stability.

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

[0069] Finally, it should be noted that: the above embodiments are only used to illustrate the technical solutions of the present application, and not to limit them; although the present application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that: it can still modify the technical solutions recorded in the foregoing embodiments, or make equivalent replacement to part or all of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the scope of the technical solutions of the embodiments of the present application.

Claims

1. A supersonic inflow control device based on an inductor array, characterized in that: The invention comprises a straight section flow channel (4) and an inductor array (5) arranged in the 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 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 straight section flow channel (4), the shock wave generator (1) is rotatably installed in the straight section flow channel (4), and a guide vane (2) is provided outside the straight section flow channel (4) so ​​as to be symmetrical with 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 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 straight section flow channel (4) for adjusting the outlet back pressure P1 of the boundary layer suction groove (3).

2. The supersonic inflow control device based on an inducer 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 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 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 inducer 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 inducers is set to X1, and the ratio of X1 to L1 is 1.

25.

4. The supersonic inflow control device based on an inducer array according to claim 1, characterized in that: The shock wave generator (1) has an approximately quadrilateral structure in a two-dimensional plane, and 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 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. 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; 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 inside the shock wave generator (1) are processed into a smooth curve by 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.

5. The supersonic inflow control device based on an inducer array according to claim 1, characterized in that: The ratio of the flow direction 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 direction length of the shock wave generator (1) is set to L2, and the trailing edge flow direction 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 strong induced shock wave, and the trailing edge of the shock wave generator (1) will not cause a strong shock wave.

6. The supersonic inflow control device based on an inducer array according to claim 1, characterized in that: The rotation axis of the shock wave generator (1) is arranged at an upstream position of the leading edge point of the shock wave generator (1), and 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; in 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 wave generator (1) are set to δ0 and δ1 respectively, and the corresponding airflow deflection angles on the inner side of the leading edge of the shock wave generator (1) are set to θ0 and θ1 respectively, then X2 is approximately (δ1-δ0) / arctan(θ1-θ0).

7. The supersonic inflow control device based on an inducer array according to claim 1, characterized in that: The guide plate (2) has an ultra-thin quadrilateral structure in a two-dimensional plane, the outer side of the leading edge of the guide plate (2) is parallel to the flow direction, and the inner side of the trailing edge of the guide plate (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 inducer 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 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 X3, then L7 is less than or equal to X3; 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 to 0.8; The rear edge of the inlet of the boundary layer suction groove (3) is rounded, and the ratio of the rounding radius R1 to the inlet length L7 of the suction cavity is 0.25-0.

35.

9. The supersonic inflow control device based on an inducer 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 inducer 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 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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