A design method for the second-stage swept side panels of a supersonic inlet

Abstract

The present invention provides a method for designing a two-stage swept side panel of a supersonic inlet, comprising an inlet lip cover and side panels located on both sides thereof. By designing the side panels to be two-stage swept, the lip cover shock wave surface is located downstream of the first-stage side panel and upstream of the second-stage side panel. A triangular overflow window is formed between the lip cover shock wave surface, the leading edge of the second-stage side panel, and the bottom wall, and the spanwise pressure difference formed between the shock wave and the outside world is utilized to partially discharge the low-energy flow of the wall boundary layer and the flow vortex in the inner flow channel. Compared with conventional inlets with single-stage swept side panels, the present invention can reduce the negative impact of the flow vortex formed by the interference of the lip cover shock wave / side panel boundary layer on the inlet performance, improve the uniformity of the flow field in the inlet throat, and enhance the aerodynamic performance of the inlet.

Description

Technical Field

[0001] The present invention relates to the technical field of supersonic aircraft air inlets, in particular to a design method for a supersonic air inlet side panel. Background Art

[0002] Supersonic inlets decelerate and pressurize high-speed airflow through a shock wave system, converting some of the airflow's kinetic energy into pressure energy, providing high-quality, low-speed, high-pressure gas to the compressor or combustor. Supersonic inlet performance is typically evaluated using parameters such as the total pressure recovery coefficient, flow coefficient, and outlet distortion.

[0003] For two-dimensional supersonic inlets, low-energy flow convergence zones are easily formed at the corners between the side panels and the compression surface due to viscous and three-dimensional effects, inducing flow separation in the corners. Furthermore, the inlet flow field also experiences a swept shock / boundary layer interaction, formed by the interaction between the lip shock wave and the side panel boundary layer. This exacerbates the accumulation of low-energy flow in the corners, increasing the range of influence of low-energy flow in the corners. This reduces the uniformity of the flow field within the inlet, increases flow distortion at the inlet exit, and affects the flow field organization in the downstream combustor. Therefore, reducing the accumulation of low-energy flow in the corners and reducing the streamwise vortex size have become research focuses on supersonic inlet flow field control.

[0004] Current research on shock wave / boundary layer interaction control in supersonic inlets, both domestically and internationally, focuses primarily on the interaction between the inlet lip shock wave and the compression surface boundary layer. Numerous studies have been conducted to suppress boundary layer separation, reduce the size of the separation pocket, and improve inlet aerodynamic performance. These efforts include employing bleed flow, boundary layer extraction, vortex generators, and variable lip structures on the compression surface. However, relatively little research has been conducted on supersonic inlet lip shock wave / side panel boundary layer interaction and flow control in the corners. Summary of the Invention

[0005] To solve the above problems, the present invention proposes a design method for a supersonic inlet two-stage swept side panel. The purpose is to discharge part of the streamwise vortex and boundary layer formed by the interference of the lip mask shock wave / side panel boundary layer while ensuring that the inlet flow capture remains basically unchanged, reduce the scale of the streamwise vortex, improve the uniformity of the flow field in the inlet throat, and enhance the overall aerodynamic performance of the inlet.

[0006] To achieve the above-mentioned object, a method for designing a supersonic inlet two-stage swept side panel according to the present invention can adopt the following technical solutions:

[0007] A method for designing a supersonic inlet secondary swept side panel, the design comprising a lip cover outer surface, a lip cover inner surface, secondary swept side panels located on both sides of the lip cover, a bottom wall connected to the bottom of the secondary swept side panel, an inlet inner channel formed by the lip cover inner surface, the bottom wall and the side panels on both sides, the tail end of the inner channel being the inlet throat; the leading edge lines of the inlet side panels are divided into a first-stage side panel leading edge line and a second-stage side panel leading edge line, both of which are straight lines, the first-stage side panel leading edge line extending obliquely backward from the top of the side panel front end and the end point of the first-stage side panel leading edge line being located between the lip cover inner surface and the bottom wall, the second-stage side panel leading edge line tilts backward at a certain angle from the end point of the first-stage side panel leading edge line and continues to extend until it is connected to the bottom wall.

[0008] Furthermore, the angle between the straight line where the leading edge line of the first-stage side panel is located and the inner profile of the lip cover is greater than the angle between the shock wave front of the lip cover and the inner profile of the lip cover, that is, γ2>γ1.

[0009] Furthermore, the intersection point between the straight line where the leading edge line of the second-stage side panel is located and the bottom wall should be located behind the intersection line between the lip shield shock wave surface and the bottom wall, that is, l>0.

[0010] Beneficial effects: Compared with the existing technology, the present invention discharges low-energy flow through the overflow window. By reasonably designing the size and shape of the overflow window, the flow field uniformity can be improved and the total pressure recovery coefficient of the inlet throat can be increased while slightly reducing the flow coefficient. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] Figure 1 It is a three-dimensional structural diagram of a second-stage swept side panel of a supersonic air inlet in the present invention.

[0012] Figure 2 It is a schematic diagram of the two-dimensional structure of the side panels of the present invention when they are swept back in a single stage.

[0013] Figure 3 This is a schematic diagram of the two-dimensional structure of the side panel in the present invention when it is swept back at the second stage. δ is the lip cover compression angle; γ1 is the angle between the straight line where the leading edge line of the first stage side panel is located and the inner profile of the lip cover; γ2 is the angle between the shock wave front of the lip cover and the inner profile of the lip cover; is the angle between the leading edge line of the first-stage side panel and the vertical line (forward / backward sweep angle); θ is the angle between the leading edge line of the second-stage side panel and the bottom wall (opening angle); l is the distance between the intersection of the leading edge line of the second-stage side panel and the bottom wall and the intersection of the lip cover shock wave front and the bottom wall; H is the inlet capture height.

[0014] Figure 4 It is a Mach number cloud diagram of the symmetry plane of the single-stage swept side panel and the two-stage swept side panel of the air inlet duct of the present invention.

[0015] Figure 5It is a throat Mach number cloud diagram of the single-stage swept side panel and the two-stage swept side panel of the air inlet of the present invention. DETAILED DESCRIPTION

[0016] Please combine Figure 1 and Figure 3 As shown, 1 represents the shock wave surface of the lip cover, 2 represents the side plate, 3 represents the outer surface of the lip cover, 4 represents the overflow window, 5 represents the bottom wall, and 6 represents the inner surface of the lip cover.

[0017] The present invention provides a method for designing a supersonic inlet secondary swept side panel, comprising a lip shield outer surface 3, a lip shield inner surface 6, secondary swept side panels 2 located on both sides of the lip shield, a bottom wall 5 connected to the bottom of the secondary swept side panel 2, and an inlet inner channel formed by the lip shield inner surface 6, the bottom wall 5, and the side panels 2 on both sides, wherein the tail end of the inner channel is the inlet throat. The leading edge lines of the inlet side panels are divided into a first-stage side panel leading edge line 21 and a second-stage side panel leading edge line 22. The first-stage side panel leading edge line 21 and the second-stage side panel leading edge line 22 are both straight lines. The first-stage side panel leading edge line 21 extends backward from the top front end of the side panel 2 (i.e., a swept design), and the end point of the first-stage side panel leading edge line 21 is located between the lip shield inner surface 6 and the bottom wall 5. The second-stage side panel leading edge line 22 extends backward at a certain angle from the end point of the first-stage side panel leading edge line 21 until it connects with the bottom wall.

[0018] The angle between the first-stage side panel leading edge line 21 and the lip mask inner profile 6 is greater than the angle between the lip mask shock wave surface 1 and the lip mask inner profile 6, that is, γ2>γ1. The intersection of the second-stage side panel leading edge line 22 and the bottom wall 5 is located behind the intersection of the lip mask shock wave surface 1 and the bottom wall 5, that is, l>0.

[0019] Application Examples

[0020] The design Mach number is 4, the lip mask compression angle δ = 10°, and the inlet capture height H = 100mm.

[0021] Before designing the two-stage swept side panel, how to design the reference side panel (single-stage side panel) is the first key of the present invention. The mutual interference between the lip shield shock wave and the side panel boundary layer will form a conical vortex. The scale of the vortex is greatly affected by the thickness of the boundary layer. This is why the forward-swept or vertical type is not adopted. The use of a swept side panel configuration can minimize the development space of the boundary layer and reduce the thickness of the boundary layer, thereby reducing the scale of the conical vortex. The scale of the flow direction vortex finally formed is also reduced, which makes the low total pressure area caused in the throat smaller and the total pressure recovery coefficient increased. The present invention has conducted a regularity study on the throat performance parameters of the inlet configuration of the single-stage side panel, and the results are shown in Table 1. It can be seen from the table that since the lip shield shock wave is located downstream of the side panel, the flow coefficients of various configurations are all around 1, among which the throat total pressure recovery coefficient of the 60° swept side panel configuration is significantly higher than that of other configurations. This is also the reference side panel configuration selected by the present invention.

[0022] Table 1 Inlet throat performance parameters of different single-stage side panel configurations

[0023] Single-stage side panel configuration Mach number Flow coefficient Total pressure recovery coefficient Sweep 60° 2.738 1.000 0.660 Sweep 45° 2.686 1.000 0.630 Sweep 30° 2.626 1.000 0.600 Sweep 15° 2.566 1.000 0.570 vertical 2.550 1.000 0.569 15° forward sweep 2.551 1.000 0.568 30° forward sweep 2.570 1.000 0.572 45° forward sweep 2.594 0.999 0.582 60° forward sweep 2.613 0.997 0.592

[0024] The second key of the present invention is the intersection position of the secondary side panel leading edge line and the bottom wall and the angle. These two parameters will affect the size and shape of the overflow window. Therefore, it is necessary to conduct a regularity study on these two parameters to obtain a set of optimal parameters so that the flow loss is small but the total pressure recovery coefficient is improved. First, keep the angle unchanged and change the intersection. The performance parameters are shown in Table 2. The configuration representation method in the table means "l / H-θ". For details, please refer to Figure 3 . It can be seen from the table that when the intersection is exactly on the intersection of the lip mask shock wave surface and the bottom wall, its flow coefficient and total pressure recovery coefficient are very close to the reference configuration, and the overflow window has just been formed. As the intersection moves backward, the overflow window gradually becomes larger, the low-energy flow is discharged, the flow coefficient becomes smaller, and the total pressure recovery coefficient becomes larger. In order to ensure that the flow is not lost too much and the total pressure recovery coefficient is improved, l / H=0.63 is selected for further regularity research. Keeping the intersection unchanged and changing the opening angle, the performance parameters are shown in Table 3. It can be seen from the table that as the opening angle decreases, the flow coefficient becomes larger, and the total pressure recovery coefficient first increases and then decreases. Within the research scope of the present invention, 0.63-6° is the optimal configuration.

[0025] Table 2 Inlet throat performance parameters of different two-stage swept side panel configurations (fixed angle)

[0026] Secondary swept side panel configuration Mach number Flow coefficient Total pressure recovery coefficient 0-10° 2.748 0.998 0.662 0.13-10° 2.753 0.996 0.663 0.33-10° 2.770 0.989 0.668 0.63-10° 2.809 0.972 0.680 0.83-10° 2.848 0.954 0.693 1.03-10° 2.895 0.930 0.703

[0027] Table 3 Inlet throat performance parameters of different two-stage swept side panel configurations (fixed intersection point)

[0028] Secondary swept side panel configuration Mach number Flow coefficient Total pressure recovery coefficient 0.63-12° 2.815 0.966 0.677 0.63-10° 2.809 0.972 0.680 0.63-8° 2.802 0.977 0.681 0.63-6° 2.794 0.982 0.681 0.63-4° 2.782 0.987 0.678

[0029] The results in the table above show that compared with the inlet with a single-stage swept side panel, the inlet with a two-stage swept side panel has a flow coefficient that decreases by 1.8% and a total pressure recovery coefficient that increases by 2.1%.

[0030] Figure 4 and Figure 5 Comparisons of Mach number contours at the inlet symmetry plane and throat for the single-stage and two-stage swept side panels reveal a smaller separation pocket near the lip shock wave impact point and a smaller low-speed zone at the inlet throat corner, improving the uniformity of the inlet throat flow parameters.

[0031] There are many methods and approaches to implement the technical solution of the present invention. The above is merely a preferred embodiment of the present invention. It should be noted that those skilled in the art may make various improvements and modifications without departing from the principles of the present invention, and such improvements and modifications should also be considered within the scope of protection of the present invention. Any components not specified in this embodiment may be implemented using existing technologies.

Claims

1. A method for designing a supersonic inlet two-stage swept side panel, characterized in that: The design includes a lip cover outer surface, a lip cover inner surface, secondary swept side panels on both sides of the lip cover, a bottom wall connected to the bottom of the secondary swept side panels, an inlet channel formed by the lip cover inner surface, the bottom wall and the side panels on both sides, the tail end of the channel is the inlet throat; the leading edge lines of the inlet side panels are divided into the leading edge line of the first-stage side panel and the leading edge line of the second-stage side panel, the leading edge line of the first-stage side panel and the leading edge line of the second-stage side panel are both straight lines, and the leading edge line of the first-stage side panel is from the top of the front end of the side panel to the rear The first-stage side panel extends obliquely and the end point of the leading edge line is located between the inner profile of the lip cover and the bottom wall, and the second-stage side panel leading edge line is tilted backward at a certain angle from the end point of the first-stage side panel leading edge line and continues to extend until it is connected to the bottom wall; the angle between the straight line where the first-stage side panel leading edge line is located and the inner profile of the lip cover is greater than the angle between the lip cover shock wave surface and the lip cover inner profile; the intersection between the straight line where the second-stage side panel leading edge line is located and the bottom wall is located behind the intersection line of the lip cover shock wave surface and the bottom wall.

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