Rapid design method for supersonic bump inlet based on cylindrical fuselage

Abstract

The invention discloses a method for rapid design of a supersonic Bump inlet duct bulge based on a cylindrical fuselage, and relates to the field of air intake systems of aircraft. Specifically, the following steps are included: taking the maximum total pressure after the wave as the optimization target, based on the conical flow theory, establishing the relationship between the Mach number and pressure before and after the shock wave, and solving the conical shock wave angle; establishing the geometric relationship between the cylindrical fuselage and the conical shock wave surface, and solving to obtain the leading edge curve of the bulge on the cylindrical fuselage; starting from the leading edge curve, building the Taylor-Maccoll equation, solving the streamline, and designing the leading edge compression surface of the bulge based on the streamline tracking method; using the trailing edge line of the leading edge compression surface, the intersection line of the trailing edge curved surface and the inlet duct wall, and the transition curve tangent to the two as the guiding curve, designing the trailing edge surface of the bulge based on the NURBS curved surface, and finally forming an external pressure type two-wave system three-dimensional supersonic inlet suitable for a cylindrical fuselage.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft, and in particular to a rapid design method for a supersonic Bump inlet based on a cylindrical fuselage. Background Art

[0002] The air inlet is a key component that connects the aircraft and its power system, and is an important part of the aircraft-engine matching. The air inlet not only provides the required air flow for the engine, completes the transmission of the required air flow from the free flow on the aircraft surface to the engine inlet, and meets the normal working requirements of the compressor and combustion chamber of the power system, but also needs to ensure the uniformity, stability and flow rate of the aircraft's engine inlet airflow in all flight conditions. For supersonic aircraft, the air inlet can slow down and pressurize the airflow.

[0003] The supersonic bump inlet is a type of non-separator external pressure two-wave inlet. The bump inlet's bulge is its most critical part, and its main functions are as follows: 1) compressing the incoming flow, discharging the low-energy boundary layer in front of the inlet lip from both sides of the bulge, reducing airflow distortion; 2) generating oblique shock waves, thereby achieving a deceleration effect on the airflow. The bump inlet has the following advantages in terms of performance, aerodynamics, stealth and structure: First, it adopts the conical flow principle design, and the total pressure recovery is relatively high; second, it reduces the resistance of the aircraft's windward surface and improves the aircraft's stealth performance; third, there is no additional or subsidiary structure, which greatly reduces the structural weight of the aircraft.

[0004] The design of the bump inlet is very challenging. The models that have been successfully applied include my country's J-10B, J-20, Xiaolong fighters, and the United States' F-16, F-35, etc. The design of the bump inlet needs to be integrated with the fuselage and the inlet wall of the aircraft. The present invention proposes a rapid design method for a supersonic bump inlet based on a cylindrical fuselage, which takes into account both design efficiency and performance, solves the problems of complex configuration of the bump inlet wave system and high design difficulty, and provides a rapid design process for the engineering application of supersonic bump inlets. Summary of the invention

[0005] In order to overcome the shortcomings of the prior art, the present invention provides a rapid design method for a supersonic bump inlet based on a cylindrical fuselage. The design method solves the problems of complex wave system configuration, high difficulty and long design cycle in the prior art, and greatly improves the design efficiency and performance of the inlet.

[0006] The specific technical solutions adopted by the present invention are as follows:

[0007] The supersonic bump inlet is a type of external pressure two-wave three-dimensional supersonic inlet. After the front supersonic flow passes through the compression surface of the bump leading edge, a conical shock wave (oblique shock wave) will be generated, and the supersonic flow velocity will be decelerated for the first time. By reasonably designing the position of the conical shock wave, the oblique shock wave is ensured to fall at the lip position of the inlet entrance, and then a normal shock wave is generated, thereby achieving a second deceleration of the airflow, ensuring that most of the airflow entering the inlet duct is subsonic.

[0008] Based on the above principle, the present invention provides a rapid design method for a supersonic bump inlet based on a cylindrical fuselage, which is as follows:

[0009] 1) Based on the conical flow theory, the relationship between the Mach number and pressure before and after the shock wave is established, and the optimal conical shock wave angle is solved with the maximum total pressure recovery after the shock wave as the design goal;

[0010] 2) According to the optimal conical shock wave angle, a geometric relationship between the cylindrical fuselage and the conical shock wave surface is established, and the oblique shock wave is projected onto the lip position by adjusting the position of the conical shock wave surface to obtain the optimal leading edge curve of the bulge on the cylindrical fuselage;

[0011] 3) Starting from the leading edge curve, the Taylor-Maccoll equation is constructed to solve the streamline, and then the compression surface of the leading edge of the bulge is designed based on the streamline tracing method;

[0012] 4) The trailing edge line of the leading edge compression surface, the intersection line of the trailing edge curved surface and the inlet duct wall, and the transition curve tangent to the trailing edge line and the intersection line are used as guide curves to design the bulge trailing edge surface based on the NURBS surface.

[0013] Preferably, the step 1) is as follows:

[0014] Firstly, the functions of the wavefront and post-wave Mach numbers and total pressure of the first oblique shock wave with respect to the shock angle are established. Then, the functions of the wavefront and post-wave Mach numbers and total pressure of the second normal shock wave with respect to the shock angle are established. The shock angle is related to the cone semi-apex angle and the Mach number of the far front incoming flow. When the incoming flow Mach number is known, the shock angle can be changed within a certain range. When the total pressure after the constrained shock wave recovers to the maximum, the unique optimal conical shock angle β can be obtained.

[0015] Preferably, the leading edge curve in step 2) is the intersection line of the conical shock wave surface generated in the conical flow theory and the surface of the cylindrical fuselage in front of the air inlet inlet.

[0016] Preferably, the geometric relationship between the cylindrical fuselage and the conical shock wave surface in step 2) is a function of the optimal conical shock wave angle obtained in step 1); when the geometry of the cylindrical fuselage is known, the equation of the leading edge curve can be solved; in order to form a second normal shock wave, the oblique shock wave needs to be projected at the lip position of the inlet, and with this as a constraint, the optimal leading edge curve can be obtained.

[0017] Preferably, the step 3) is as follows:

[0018] Starting from the optimal discrete points on the leading edge curve obtained in step 2), the Taylor-Maccoll flow field equation for each point is established through the streamline tracing method, and the fourth-order Runge-Kutta method is used to solve the flow field equation, and finally the streamlines starting from the discrete points on the leading edge curve are obtained. When the streamlines and step lengths are dense enough, this is used as a grid to form the compression surface shape of the bulge leading edge.

[0019] Preferably, in the step 4), the shape of the trailing edge surface of the bulge is mainly related to the trailing edge line of the leading edge compression surface and the shape of the inlet duct wall; the outer contour of the NURBS surface is the trailing edge line of the leading edge compression surface and the outer edge line on the inlet duct wall, and the two transition lines inside it are respectively located in the longitudinal symmetry plane and the transverse cross-section, and the two sections of the transition line are respectively tangent to the leading edge compression surface and the inlet duct wall.

[0020] Compared with the prior art, the present invention has the following beneficial effects:

[0021] 1. The method proposed in the present invention can determine the unique shock wave angle and bulge position by constraining the maximum total pressure after the wave and the position of the oblique shock wave and the lip, and can quickly determine the shape and position of the bulge. It eliminates the problems of long design cycle, complex process and strong dependence on designers caused by the traditional multiple iteration wave system configuration design method, and greatly improves the design efficiency of the bump inlet bulge while ensuring the performance of the inlet.

[0022] 2. The present invention provides a rapid design method for a three-dimensional bulge, which can quickly obtain a three-dimensional configuration of the bulge that meets the requirements of a supersonic aircraft. The method proposed in the present invention can be applied not only to cylindrical fuselages, but also to other fuselage shapes that can construct geometric equations, such as elliptical, spindle-shaped, etc., by simply replacing the cylindrical geometric equations in this method. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 It is the relationship diagram of the conical flow theory and the two wave systems of the bump inlet;

[0024] Figure 2 is the geometric relationship of the leading edge curve of the cylindrical fuselage;

[0025] Figure 3It is the coordinate transformation relationship of the streamline tracing method;

[0026] Figure 4 This is the flow chart of numerical solution of the benchmark flow field;

[0027] Figure 5 It is the compression surface of the leading edge of the bulge on the cylindrical fuselage;

[0028] Figure 6 It is a three-dimensional picture of the bulge;

[0029] Figure 7 It is the bump air inlet on the cylindrical fuselage. DETAILED DESCRIPTION

[0030] The present invention is further described and illustrated below in conjunction with the accompanying drawings and specific embodiments. The technical features of each embodiment of the present invention can be combined accordingly without conflicting with each other.

[0031] The rapid design method of a supersonic bump inlet based on a cylindrical fuselage proposed in the present invention specifically comprises the following steps:

[0032] (1) Determination of the cone shock wave angle β (maximum total pressure recovery coefficient after the wave is the optimization goal)

[0033] The theoretical basis of supersonic bump inlet is conical flow theory, such as Figure 1 As shown in the figure, when the supersonic airflow blows toward the cone, a conical shock wave surface coaxial with the cone will be generated at the top of the cone, and the angle between the shock wave and the cone is defined as the conical shock wave angle β. The function of the leading edge compression surface of the supersonic bump inlet is equivalent to that of the cone. After the front supersonic incoming flow passes through the leading edge compression surface of the bump, a conical shock wave (oblique shock wave) will be generated. The supersonic incoming flow velocity is decelerated for the first time, but the speed at this time is still supersonic, which is difficult to meet the intake requirements of the power system. Therefore, the supersonic bump inlet needs to be designed as an external pressure type two-wave system inlet. By reasonably designing the position of the inlet inlet, it is ensured that the oblique shock wave falls at the lip position, and then a positive shock wave is generated, thereby achieving a second deceleration of the airflow, ensuring that most of the airflow entering the inlet pipe is subsonic.

[0034] After an ideal gas passes through a conical shock wave and a normal shock wave, the total pressure decreases. The total pressure recovery coefficient after the airflow passes through two shock waves is a function of the conical shock wave angle β. The first step in the design of a bump inlet is to determine the conical shock wave angle β. Let δ be the half cone angle of the cone that generates the shock wave. For the first oblique shock wave, the relationship between the wavefront Mach number (Ma1) and the waveback Mach number (Ma2) is:

[0035]

[0036] Where K is the Boltzmann constant, the ratio of the stagnation pressure (P1) before the wave to the stagnation pressure (P2) after the wave

[0037]

[0038] For the second normal shock wave, the relationship between the wavefront Mach number Ma2 and the waveback Mach number Ma3 is as follows:

[0039]

[0040] Let the stopping pressure after the positive shock wave be P3, and the pressure relationship before and after is as follows:

[0041]

[0042] From the above relationship, it can be seen that the shock wave angle β and the cone semi-apex angle are related to the Mach number of the far front incoming flow. When the incoming flow Mach number is known, the shock wave angle β is related to the cone semi-apex angle. Obtaining the maximum total pressure recovery is one of the goals of the bump inlet design. Through the wave system configuration, there is a unique shock wave angle β that maximizes the total pressure behind the wave. Taking the incoming flow Mach number of 1.6 as an example, when the total pressure behind the wave is not constrained, the range of the shock wave angle generated is 40° to 70°. Through programming, with the maximum total pressure behind the wave as the optimization goal, the optimal shock wave angle β can be obtained as 52.33°. At this time, the total pressure loss is minimal, and the total pressure after two shock waves is 97.67% of the total pressure of the far front incoming flow.

[0043] (2) Derivation of equation for the leading edge curve of a cylindrical fuselage

[0044] The leading edge curve is the basis for determining the compression surface of the leading edge of the bump inlet, such as Figure 2 As shown in the figure, the leading edge curve is actually the intersection of the shock wave cone generated in the conical flow theory and the cylindrical fuselage surface in front of the inlet entrance.

[0045] A rectangular coordinate system is established with the vertex of the cone as the origin O, the central axis of the cone as the X-axis, the vertical direction as the Z-axis, and Y as determined by the right-hand screw rule. The equation of the cone shock wave can be obtained as follows:

[0046] x 2 tan 2 β=y 2 +z 2 (5)

[0047] The fuselage of the aircraft is a cylinder with a radius of R, the height of the shock cone in the X direction is l, and the coordinate of the center of the cylinder in the Z direction is a, then the equation of the cylinder is:

[0048]

[0049] In the XOZ plane, Z = 0, the cone and the cylinder have an intersection point P, which is the front end point of the leading edge array. The coordinates of the intersection point are [(R+a) / tanβ,0,R+a], so the equation of the leading edge curve can be obtained as:

[0050]

[0051] The leading edge curve can be obtained by solving the equation. In order to form the second normal shock wave, the conical shock wave needs to be projected at the lip position of the inlet duct. With this as a constraint, the optimal leading edge curve can be obtained.

[0052] (3) Design of the compression surface of the bulge front edge based on the streamline tracking method

[0053] The flow field of the conical oblique shock wave surface is axially symmetric, and the flow in each section is a two-dimensional flow. Figure 3 As shown in the figure, after the leading edge curve is determined, for each point A of the leading edge curve, the streamline passing through point A can be solved by a two-dimensional streamline tracing method in the symmetry plane between the vertex O of the conical shock wave and point A. First, coordinate transformation is required. With O as the vertex, it is rotated around the X axis so that the rotated OZ′ passes through point A. At this time, streamline tracing can be performed in the XOZ′ plane. From the XYZ coordinate transformation to XY′Z′, the coordinates of point A after transformation are:

[0054]

[0055] in When z(A)=0, α=90°.

[0056] The definition of a streamline is that at a certain moment, the tangent direction of all points on the streamline is consistent with the velocity direction of the point. The differential equation of the streamline can be expressed as:

[0057]

[0058] Starting from the leading edge curve, the following iterative formula is used:

[0059]

[0060] in The solution is obtained by using the reference flow field after the shock wave cone. The flow field of the conical flow after the shock wave surface follows the Taylor-Maccoll equation. For the ideal gas supersonic cone flow, the Taylor-Maccoll flow model believes that in the flow field after the shock wave on the conical surface, the airflow parameters on the cone surface starting from the same vertex angle are consistent. In the spherical coordinate system, let v r , v θ are the radial and circumferential velocity components respectively, then the first-order ordinary differential equation system of Taylor-Maccoll equation is expressed as follows

[0061]

[0062] The coupled iterative formula of the fourth-order Runge-Kutta method is used to numerically solve the above equation. The flowchart of the solution process of the Taylor-Maccoll equation is as follows: Figure 4 As shown. Get v r ,v θ After that, it is converted into a rectangular coordinate system to obtain the velocity expressions in the X and Z directions:

[0063]

[0064] For different initial points, the corresponding streamlines are tracked and the obtained streamlines are combined into a flow surface, which is the waverider bulge. Since the position of the vertex O of the aforementioned conical shock wave and the center axis of the fuselage cylinder is variable, the bulge shape that best suits the lip can be obtained by adjusting the position of point O. The shape of the leading edge compression surface obtained by the streamline tracking method is shown in Figure 5 shown.

[0065] (4) Design of the rear edge of the bulge

[0066] The bulge trailing edge surface is an important connecting transition component between the leading edge compression surface and the inlet duct. The design of the bulge trailing edge surface is crucial to the stability of the fluid entering the inlet duct. The shape of the trailing edge surface is mainly related to the trailing edge line of the leading edge compression surface and the shape of the inlet duct wall. The present invention proposes a design method based on NURBS surfaces. The construction of the NURBS surface uses the trailing edge line of the leading edge compression surface, the intersection line of the trailing edge surface and the inlet duct wall, and the transition tangent line tangent to the two as guide curves, such as Figure 6 As shown in the axonometric view, the two transition lines are located in the longitudinal symmetry plane and the transverse section respectively, and the two sections of the transition line are tangent to the leading edge compression surface and the inlet duct wall respectively. The final design of the complete shape of the bulge is as follows Figure 6 shown.

[0067] Figure 7 The three-dimensional diagram of the bump air intake on the cylindrical fuselage is shown, in which the compression surface of the leading edge of the bump is smoothly tangent to the trailing edge of the bump, the bump is embedded in the cylindrical fuselage, and the trailing edge and the S-bend air intake duct have a smooth transition, which can effectively ensure the stability of the flow field and reduce the distortion of the flow field. The outer surface of the air intake duct lip of the air intake duct cover is embedded in the cylindrical fuselage, and the inner surface is connected with the S-bend duct, which is mainly used to capture the front flow and play a role in rectification.

[0068] The above-described embodiment is only a preferred solution of the present invention, but it is not intended to limit the present invention. A person skilled in the relevant technical field may make various changes and modifications without departing from the spirit and scope of the present invention. Therefore, any technical solution obtained by equivalent replacement or equivalent transformation falls within the protection scope of the present invention.

Claims

1. A rapid design method for a supersonic bump inlet based on a cylindrical fuselage, characterized in that: The details are as follows: 1) Based on the conical flow theory, the relationship between the Mach number and pressure before and after the shock wave is established, and the optimal conical shock wave angle is solved with the maximum total pressure recovery after the shock wave as the design goal; 2) According to the optimal conical shock wave angle, a geometric relationship between the cylindrical fuselage and the conical shock wave surface is established, and the oblique shock wave is projected onto the lip position by adjusting the position of the conical shock wave surface to obtain the optimal leading edge curve of the bulge on the cylindrical fuselage; 3) Starting from the leading edge curve, the Taylor-Maccoll equation is constructed to solve the streamline, and then the compression surface of the leading edge of the bulge is designed based on the streamline tracing method; 4) The trailing edge line of the leading edge compression surface, the intersection line of the trailing edge curved surface and the inlet duct wall, and the transition curve tangent to the trailing edge line and the intersection line are used as guide curves to design the bulge trailing edge surface based on the NURBS surface; The leading edge curve in step 2) is the intersection line between the conical shock wave surface generated in the conical flow theory and the cylindrical fuselage surface in front of the air inlet entrance; The geometric relationship between the cylindrical fuselage and the conical shock wave surface in step 2) is a function of the optimal conical shock wave angle obtained in step 1); when the geometry of the cylindrical fuselage is known, the equation of the leading edge curve can be solved; in order to form a second normal shock wave, the oblique shock wave needs to be projected at the lip position of the air inlet, and with this as a constraint, the optimal leading edge curve can be obtained.

2. The rapid design method for a supersonic bump inlet based on a cylindrical fuselage according to claim 1 is characterized in that: The step 1) is specifically as follows: Firstly, the functions of the wavefront and post-wave Mach numbers and total pressure of the first oblique shock wave with respect to the shock angle are established. Then, the functions of the wavefront and post-wave Mach numbers and total pressure of the second normal shock wave with respect to the shock angle are established. The shock angle is related to the cone semi-apex angle and the Mach number of the far front incoming flow. When the incoming flow Mach number is known, the shock angle can be changed within a certain range. When the total pressure after the constrained shock wave recovers to the maximum, the unique optimal conical shock angle β can be obtained.

3. The rapid design method of supersonic bump inlet based on cylindrical fuselage according to claim 1 is characterized in that: The step 3) is as follows: Starting from the optimal discrete points on the leading edge curve obtained in step 2), the Taylor-Maccoll flow field equation for each point is established through the streamline tracing method, and the fourth-order Runge-Kutta method is used to solve the flow field equation, and finally the streamlines starting from the discrete points on the leading edge curve are obtained. When the streamlines and step lengths are dense enough, this is used as a grid to form the compression surface shape of the bulge leading edge.

4. The rapid design method for a supersonic bump inlet based on a cylindrical fuselage according to claim 1 is characterized in that: In the step 4), the shape of the trailing edge surface of the bulge is mainly related to the trailing edge line of the leading edge compression surface and the shape of the intake duct wall; the outer contour of the NURBS surface is the trailing edge line of the leading edge compression surface and the outer edge line on the intake duct wall, and the two transition lines inside it are respectively located in the longitudinal symmetry plane and the transverse cross-section, and the two sections of the transition line are respectively tangent to the leading edge compression surface and the intake duct wall.

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