Design method of hypersonic integrated inlet with given forebody shape

Abstract

The present invention belongs to the field of hypersonic inlet design, and specifically relates to a method for designing a hypersonic integrated inlet with a given forebody shape. The method comprises the following steps: according to a given forebody shape of an aircraft, the forebody leading edge shock wave is numerically fitted as the shock wave profile, and after selecting a suitable axisymmetric reference flow field, the spanwise design of the inlet is performed according to the close principle, and the flow direction design is performed using the streamline tracking technology. The obtained inlet profile is geometrically spliced ​​with the forebody of the aircraft, and an inlet configuration that meets the requirements of the hypersonic forebody can be obtained. The method makes the leading edge line of the forebody of the aircraft fit with the leading edge line of the inlet. Under the design working condition, the shock wave generated by the airflow passing through the forebody of the aircraft can completely seal the inlet and achieve full flow capture. At the same time, the streamline tracking technology can be used to freely design the inlet capture profile, thereby meeting the integration requirements of forebodies with different shapes.

Description

Technical Field

[0001] The invention belongs to the field of hypersonic inlet design, and in particular relates to a hypersonic integrated inlet design method with a given forebody shape. Background Art

[0002] In the past decade, hypersonic vehicles have become the main direction of technological competition among major scientific and technological powers. Among them, air-breathing hypersonic vehicles use scramjet engines or combined engines as their main power, which can make full use of oxygen in the atmosphere as a fuel oxidant, so that the aircraft has continuous and controllable thrust, higher flight flexibility and economy, and can achieve hypersonic flight in the atmosphere. As a key component of air-breathing hypersonic vehicles, the air inlet has been studied in depth by researchers at home and abroad, taking into account key performance factors such as its flow coefficient and compression efficiency.

[0003] So far, relatively mature inlets with different compression forms have been developed, including axisymmetric inlets, dual inlets, and three-dimensional compression inlets. At the same time, various integrated design schemes for the forebody / inlet of aircraft that combine internal and external flows have also been proposed. The HTV-3X hypersonic technology verification aircraft in the Falcon project of the United States adopts a wing-body fusion double inward-turned inlet layout. The inlets on both sides of the front of the aircraft adopt a projection shape similar to a water droplet shape, and the leading edge profile of the inlet and the leading edge profile of the aircraft are directly integrated as the forebody of the aircraft, achieving the wave-riding effect in the inlets on both sides of the leading edge of the fuselage. He Xuzhao et al. "Integrated design and performance analysis of curved outer cone waverider inlet [J]. Propulsion Technology, 2018, 39(10): 2313-2319." A curved surface cone waverider forebody inlet integrated configuration was designed using integrated streamline tracking technology and osculating axisymmetry theory. By numerically simulating the configuration, a schematic diagram of its flow field characteristics was obtained, and the results showed that it was consistent with the design expectations.

[0004] However, the above design method still has various problems, including: major changes to the existing forebody structure and wave interference between the forebody and the air inlet. Summary of the invention

[0005] The object of the present invention is to provide a method for designing a hypersonic integrated air inlet with a given forebody shape.

[0006] The technical solution to achieve the purpose of the present invention is: a method for designing a hypersonic integrated inlet with a given forebody shape, comprising the following steps:

[0007] Step (1) determining a reference flow field: selecting an initial axisymmetric reference flow field according to the inlet duct length, the total contraction ratio, the internal contraction ratio, the leading edge shock wave angle, the shape of the aircraft front body, and the flight condition requirements;

[0008] Step (2) Preparing a shock wave profile: obtaining a leading edge shock wave surface through numerical simulation, intercepting the leading edge shock wave surface of a given aircraft front body along the axial direction, obtaining a desired shock wave profile through approximate processing of the obtained intersection line, and determining its profile equation, and then modifying the initial axisymmetric reference flow field according to the shock wave profile according to the osculating theory procedure;

[0009] Step (3) generates an inlet profile: draw a capture profile within the reference flow field range, and export the capture profile discrete point data; apply streamline tracking in the modified reference flow field to obtain a series of profile data points of the required inlet profile, and obtain the inlet profile by point fitting.

[0010] Furthermore, the method further includes post-processing step (4): coupling the air inlet and the forebody of the aircraft together through geometric splicing, so that the leading edge line of the air inlet is substantially aligned with the leading edge line of the forebody, and making a smooth transition at the joint through geometric shaping, thereby obtaining a complete integrated configuration.

[0011] Furthermore, the compression law of the reference flow field in step (1) is consistent with the compression profile of the precursor.

[0012] A hypersonic integrated inlet with a given forebody shape is designed using the above method.

[0013] Compared with the prior art, the present invention has the following significant advantages:

[0014] (1) Compared with various hypersonic integrated configuration design schemes that have been published, this method can design a highly integrated hypersonic inlet that meets performance requirements based on the specific aircraft forebody shape, with little modification to the existing forebody structure. In actual engineering, the aircraft forebody can be given priority in design.

[0015] (2) From the perspective of aerodynamic characteristics, the air inlet and the front body of the aircraft share the same shock wave surface, which reduces the wave interference between the two and can ensure the high initial design performance of both. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Figure 1 It is a diagrammatic illustration of the modification principle of the reference flow field using the close principle applied by the present invention.

[0017] Figure 2 It is a diagram of the streamline tracing method used in the present invention.

[0018] Figure 3 1 is a diagram of a hypersonic integrated inlet design for a given precursor of the present invention, wherein (a) is a spanwise view, and (b) is a flowwise view.

[0019] Figure 4 It is a three-dimensional model diagram of a hypersonic forebody / inlet integrated configuration generated by an embodiment of the present invention. DETAILED DESCRIPTION

[0020] The present invention is further described in detail below in conjunction with the accompanying drawings.

[0021] A method for designing a hypersonic integrated inlet with a given forebody shape, characterized in that it comprises the following steps:

[0022] Step (1) Determine the reference flow field: Select a suitable axisymmetric reference flow field according to the inlet length, total contraction ratio, internal contraction ratio, leading edge shock wave angle, aircraft front body shape and flight condition requirements.

[0023] Step (2) Preparing the shock wave profile: The leading edge shock wave surface of the given aircraft front body is intercepted along the axial direction (the leading edge shock wave surface can be obtained by numerical simulation), and the intersection line is obtained by approximation processing in the modeling software to obtain the required shock wave profile, and its profile equation (polynomial equation) is determined. Subsequently, the initial axisymmetric reference flow field is modified according to the shock wave profile according to the osculating theory program.

[0024] Step (3) Generate the inlet profile: Draw the capture profile within the above-mentioned reference flow field range and export its discrete point data. Apply streamline tracing technology in the above-mentioned modified reference flow field to obtain a series of profile data points of the required inlet profile, and obtain the inlet profile by point fitting in the modeling software.

[0025] Step (4) post-processing: In the modeling software, the air inlet and the forebody are coupled together through geometric splicing, the leading edge line of the air inlet is basically in line with the leading edge line of the forebody, and the joint is made smooth through geometric shaping to obtain a complete integrated configuration.

[0026] The method for modifying the reference flow field based on the shock wave profile is shown in the attached figure. Figure 1 As shown in the figure, according to the straight line where the radius of curvature of each point on the shock wave profile lies, several osculating surfaces are generated to capture the differential of the profile. The osculating surface is perpendicular to the paper surface, such as the plane where ABCO and DEFO' are located in the figure. The osculating method assumes that the osculating surface is the axisymmetric surface of the axisymmetric reference flow field. The reference axisymmetric flow field is scaled by the ratio of the radius of curvature to the radius of the original axisymmetric reference flow field, so that O(O') coincides with the symmetry axis. Then, the corresponding relationship between the corresponding points in the osculating surface and the axisymmetric reference flow field can be established, and then the streamlines on each section can be obtained through streamline tracking technology, as shown in the attached figure. Figure 2 The obtained streamlines are combined and fitted in the modeling software to obtain the complete inlet profile. The three-dimensional diagram of this design method is shown in the attached figure. Figure 3 shown.

[0027] Example

[0028] Taking the forebody of the X-51A hypersonic aircraft as an example, an integrated air inlet design is performed according to the design method of the present invention.

[0029] (1) Based on the characteristics of the X-51A forebody compression profile, a conical flow was selected as the reference axisymmetric flow field. The incoming flow condition was set to 6 Ma, and the first shock wave angle was 12.7°.

[0030] (2) According to the shock wave shape of the leading edge of the aircraft, the shock wave profile is fitted as a general quadratic curve, and the curve equation is: y = -0.00031x 2 +1157.5245. The axisymmetric reference flow field in step (1) can be transformed into a reference flow field consistent with the quadratic curve shape by means of a close theoretical procedure.

[0031] (3) Then draw a rectangular capture profile. The data points on the rectangle should be between the shock wave profile and its centerline of curvature, and export it into a finite number of discrete points at equal angles (half of them can be selected for drawing according to the symmetry of the flow field characteristics). Generate the inlet profile in the reference flow field described in step (2) through the streamline tracing program.

[0032] (4) In the modeling software UG, the inlet profile is combined with the aircraft through geometric splicing. It can be found that the leading edge line of the inlet fits the leading edge line of the wing, and the compression surface of the inlet and the compression surface of the forebody are basically coincident. From a structural point of view, it meets the design requirements. The final aircraft forebody / inlet integrated configuration is as follows Figure 4 shown.

Claims

1. A method for designing a hypersonic integrated inlet with a given forebody shape, characterized in that: The steps include: Step (1) determining a reference flow field: selecting an initial axisymmetric reference flow field according to the inlet duct length, the total contraction ratio, the internal contraction ratio, the leading edge shock wave angle, the shape of the aircraft front body, and the flight condition requirements; Step (2) Preparing a shock wave profile: obtaining a leading edge shock wave surface through numerical simulation, intercepting the leading edge shock wave surface of a given aircraft front body along the axial direction, obtaining a desired shock wave profile through approximate processing of the obtained intersection line, and determining its profile equation, and then modifying the initial axisymmetric reference flow field according to the shock wave profile according to the osculating theory procedure; Step (3) generating an inlet profile: drawing a capture profile within the reference flow field range, and exporting the capture profile discrete point data; applying streamline tracking in the modified reference flow field to obtain a series of profile data points of the required inlet profile, and obtaining the inlet profile by point fitting; Step (4) post-processing: the air inlet and the forebody of the aircraft are coupled together by geometric splicing, the leading edge line of the air inlet is basically in line with the leading edge line of the forebody, and the joint is made smooth through geometric shaping to obtain a complete integrated configuration.

2. The method according to claim 1, characterized in that The compression law of the reference flow field in step (1) is consistent with the compression profile of the precursor.

3. A hypersonic integrated inlet with a given front body shape, characterized in that: The method is designed by any one of claims 1 to 2.

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