Model for supersonic natural laminar flow wing calibration

Abstract

A model for supersonic natural laminar flow wing calibration comprises a wing, a cushion block is arranged below the wing, the cushion block is installed in a wind tunnel installation groove, a baffle is arranged between the wing and the cushion block to separate the cushion block from the wing, and it is ensured that incoming flow sensed by the wing is uniform and stable, is the same as wind tunnel incoming flow and has no acceleration or deceleration characteristics; the wing has typical aerodynamic design characteristics of a supersonic NLF wing; according to the model, an observable laminar flow area range can be always kept in a use range, a transition model can be accurately calibrated, and the transition characteristic of an NLF airfoil profile is kept, so that the problem of a blind area of a supersonic laminar flow wing wind tunnel test and transition model calibration is solved.

Description

Technical Field

[0001] The invention belongs to the technical field of high-speed wind tunnel testing, and in particular relates to a model for calibration of supersonic natural laminar flow wings. Background Art

[0002] With rapid economic development, people are increasingly demanding travel time. Currently, civil aircraft in service are high-subsonic, resulting in lengthy flight times on long-haul routes and reduced travel comfort. Supersonic civil aircraft, on the other hand, can significantly shorten long-haul flights, especially on transoceanic flights. In the 20th century, supersonic aircraft like the Soviet Tu-144 and the European Concorde suffered from performance shortcomings such as high fuel consumption and severe sonic boom. Since then, with technological advancements, the supersonic civil aircraft field has been committed to addressing key technologies such as low drag and low sonic boom. Among these key technologies is drag reduction. Laminar flow drag reduction technology significantly reduces drag on supersonic civil aircraft and is a key tool for achieving economical flight. Frictional drag accounts for 40% of total drag on supersonic passenger aircraft. Therefore, implementing laminar flow drag reduction on supersonic passenger aircraft can significantly improve their aerodynamic performance.

[0003] Laminar drag reduction has entered the engineering application stage on subsonic aircraft. The Boeing 787Max has implemented a laminar flow nacelle, the Boeing 787-900 has implemented a hybrid laminar flow vertical tail design, and Airbus has also conducted research on hybrid laminar flow design on the A320 vertical tail. Currently, transition prediction based on subsonic flow has been widely studied, among which the e- N The method has been verified by engineering application. N For the calibration of the transition threshold, researchers have extensive wind tunnel and flight testing in subsonic flows, providing a reference for comparison. Furthermore, design methods for subsonic flows have also made significant progress in recent years. However, for supersonic laminar flows, transition prediction has lagged behind, limiting the development of technologies such as supersonic laminar wing design methods.

[0004] Since the transition mechanism is significantly different from that of subsonic flow, that is, in addition to TS waves (Tomien-Schlichting) and CF (Crossflow) standing waves, supersonic flow also has oblique TS waves and CF traveling waves. Currently, the transition prediction of supersonic flow is also mainly based on the linear stability theory. NThe main method is to calibrate the transition threshold of supersonic flow. There are two main supersonic test models (calibration of transition threshold): the NEXST model jointly designed by Onera and JASA in France, including wind tunnel and flight tests; and the supersonic swept wing designed by NASA in the United States, including wind tunnel and flight tests. For the NEXST model, under typical wind tunnel conditions (Mach number 2, based on chord length, Reynolds number 5×10 6 -11×10 6 ), the calibrated transition threshold is 4.4 to 7.4, with an average of about 5.0. Under flight test conditions (Mach number 2, Reynolds number about 33×10 6 ), the transition threshold is 7.4-15.5. The NEXST supersonic model has not yet been made public and cannot be used for transition prediction model calibration research. In 2020, the AAASciTech Forum published an article on the transition characteristics of supersonic swept wings in ground and flight experiments. The article mentioned that at Mach number 2 and Reynolds number of about 2.77×10 6 The transition threshold is calibrated to 6.0, while the transition threshold referenced in the design is 5.6. Therefore, in other words, the NASA supersonic swept wing under wind tunnel and flight test conditions (Mach number is 2, Reynolds number is about 2.77×10 6 ), the transition threshold is calibrated to 6.0, while the transition threshold used during design is 5.6. While this model can be used to calibrate and study transition prediction models, it only utilizes one operating condition, making it difficult to fully validate and calibrate the transition model. Consequently, there is a lack of wind tunnel testing and transition model calibration research on supersonic laminar wing structures, both domestically and internationally. Summary of the Invention

[0005] In order to overcome the above-mentioned shortcomings of the prior art, the purpose of the present invention is to provide a model for the calibration of supersonic natural laminar flow wings, which can always maintain the observable laminar flow range within the operating range, can calibrate the transition model, and maintain the transition characteristics of the NLF airfoil, thereby solving the blind spot problem in supersonic laminar flow wing wind tunnel testing and transition model calibration.

[0006] In order to achieve the above object, the technical solution of the present invention is:

[0007] A model for supersonic natural laminar flow wing calibration includes a wing 3, a pad 1 is provided below the wing 3, the pad 1 is installed in a wind tunnel installation slot, and a baffle 2 is provided between the wing 3 and the pad 1. The baffle 2 separates the pad 1 and the wing 3 to ensure that the incoming flow felt by the wing 3 is uniform and stable and is the same as the wind tunnel incoming flow, without acceleration or deceleration characteristics.

[0008] The wing 3 includes two structures: a supersonic leading edge wing configuration 3-1 and a subsonic leading edge wing configuration 3-2. Both structures are formed using three control surfaces: the wing root, 40% span-wise position, and wingtip. The maximum thickness is 5% of the chord length and is located at 58% of the chord length. The difference between the three control surfaces is concentrated in the first 30% of the chord length of the upper surface, among which the airfoil thickness corresponding to the wing root increases; the airfoil camber of the two structures is small, among which the supersonic leading edge and airfoil camber of the supersonic leading edge wing configuration 3-1 are both -0.12%, and the maximum camber corresponds to 54.46% of the chord length. The airfoil camber at the subsonic leading edge and wing root of the subsonic leading edge wing configuration 3-2 is -0.26%, and the maximum camber corresponds to 43.94% of the chord length.

[0009] When wing 3 is in supersonic leading edge wing configuration 3-1, its leading edge sweep angle is 30°, the applicable Mach number range is Ma=1.5~1.8, and the corresponding Reynolds number is Re=8×10 6 ~13×10 6 , the angle of attack is AoA = -1°~4°.

[0010] When wing 3 is a subsonic leading edge wing configuration 3-2, its leading edge sweep angle is 60°, the applicable Mach number range is Ma=1.6~2.0, and the corresponding Reynolds number is Re=8×10 6 ~10×10 6 , the angle of attack is AoA = -2°~3°.

[0011] Compared with the prior art, the advantages of the present invention are:

[0012] (1) The design of the pads and baffles of the present invention can keep the incoming flow reaching the wing section uniform and stable without being affected by shock waves.

[0013] (2) The wing section of the present invention is formed by using three control surfaces: the wing root, the 40% span position, and the wing tip. The airfoil used for forming is the result of optimized design, which can ensure that the formed model is within the use range (supersonic leading edge: Mach number range is Ma = 1.5 ~ 1.8, corresponding to Reynolds number Re = 8 × 10 6 ~13×10 6 , the angle of attack range is AoA = -1° ~ 4°; subsonic leading edge: Mach number range is Ma = 1.6 ~ 2.0, corresponding to Reynolds number Re = 8×10 6 ~10×10 6 , angle of attack AoA = -2° ~ 3°) to maintain the observable laminar flow region and the transition characteristics of the NLF airfoil; and it has strong robustness and can reflect the typical transition mechanism of the supersonic NLF wing within the test range.

[0014] In summary, the present invention can always maintain the observable laminar flow region within the scope of use, accurately calibrate the transition model, and maintain the transition characteristics of the NLF airfoil, thereby solving the blind spot problem in the wind tunnel test of supersonic laminar airfoils and the calibration of the transition model. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] Figure 1 The three-view drawing of the final configuration of the supersonic leading edge wing is shown in Figure 1. Figure 1 (a) is the main view, Figure 1 (b) is the side view. Figure 1 (c) is a top view.

[0016] Figure 2 Schematic diagram of detailed parameters of the final configuration of the supersonic leading edge wing of Example 1, wherein Figure 2 (a) is the main view, Figure 2 (b) is a side view, Figure 2 (c) is a top view.

[0017] Figure 3 The three-view diagram of the final configuration of the subsonic leading edge wing is shown in Figure 1. Figure 3 (a) is the main view, Figure 3 (b) is the side view. Figure 3 (c) is a top view.

[0018] Figure 4 This is a schematic diagram of detailed parameters of the final configuration of the subsonic leading edge wing of Example 2, where Figure 4 (a) is the main view, Figure 4 (b) is a side view, Figure 4 (c) is a top view.

[0019] Figure 5 It is a supersonic leading edge wing airfoil design.

[0020] Figure 6 It is a subsonic leading edge wing airfoil design.

[0021] Figure 7 Schematic diagram of the supersonic leading edge cross section.

[0022] Figure 8 The Slice 4 section of the supersonic leading edge is at Ma = 1.5, Re = 8×10 6 Pressure distribution at different angles of attack.

[0023] Figure 9 The Slice 4 section of the supersonic leading edge is at Ma = 1.6, Re = 10×10 6 Pressure distribution at different angles of attack.

[0024] Figure 10The Slice 4 section of the supersonic leading edge is at Ma = 1.8, Re = 13×10 6 Pressure distribution at different angles of attack.

[0025] Figure 11 Schematic diagram of the subsonic leading edge cross section.

[0026] Figure 12 The subsonic leading edge Slice 4 section is at Ma = 1.8, Re = 8×10 6 Pressure distribution at different angles of attack.

[0027] Figure 13 The subsonic leading edge Slice 4 section is at Ma = 2.0, Re = 10×10 6 Pressure distribution at different angles of attack. DETAILED DESCRIPTION

[0028] The present invention will be described in detail below with reference to the accompanying drawings and embodiments.

[0029] Example 1: Supersonic leading edge wing configuration

[0030] use Figure 5 The airfoil shown is formed into supersonic leading edge wing configuration 3-1. Figure 5 The results shown in the figure are the normalized airfoil cross-section data. The cross-sections of the three control surfaces at the wing root, 40% span position, and wing tip position correspond to Figure 5 After forming, follow the Figure 1 The three views shown are used to assemble a calibration model, specifically: a model for supersonic natural laminar flow wing calibration, including a wing 3, a pad 1 is provided under the wing 3, the pad 1 is installed in the wind tunnel installation slot, ensuring that the observed section of the wing 3 is located in the wind tunnel observation window; a baffle 2 is provided between the wing 3 and the pad 1, the baffle 2 separates the pad 1 and the wing 3, ensuring that the incoming flow felt by the wing 3 is uniform and stable and the same as the wind tunnel incoming flow, without acceleration or deceleration characteristics; the wing 3 has a typical aerodynamic design feature of a supersonic NLF wing, and can maintain an observable laminar flow area within the Mach number and angle of attack range of the experimental design) The leading edge sweep angle, Mach number, and angle of attack range can cover the typical low-drag, low-sonic boom supersonic layout design features. This embodiment is a supersonic leading edge wing configuration 3-1, which is formed using three control surfaces: the wing root, 40% span-wise position, and wingtip. The maximum thickness is 5% of the chord length and is located at 58% of the chord length. The difference between the three control surfaces is concentrated in the first 30% of the chord length of the upper surface, among which the airfoil thickness corresponding to the wing root increases; the airfoil camber of the supersonic leading edge wing configuration 3-1 is small, with the supersonic leading edge (SW30) and airfoil camber both being -0.12%. The maximum camber corresponds to a position at 54.46% of the chord length, and its leading edge sweep angle is 30° (see Figure 7 ), the applicable range of Mach number is Ma=1.5~1.8, and the corresponding Reynolds number is Re=8×10 6 ~13×10 6 , the angle of attack is AoA = -1° to 4°. The specific parameters of the baffle and the pad 1 are as follows Figure 2 and Figure 4 As shown in , the supersonic leading edge and the subsonic leading edge have different parameters. For the supersonic leading edge, the specific parameters of the supersonic leading edge wing configuration 3-1 are as follows Figure 2 As shown in the , the wingtip length is 267mm, the corresponding chord length is 306.9mm, the chord length at the wing root is 400mm, and the tip-to-root ratio is 0.767. The wingspan is 600mm, the leading edge sweep angle is 30°, the trailing edge sweep angle is 20.26°, and the angle at the wingtip is 85°.

[0031] The total length of the pad root is 407.9mm, the distance from the root to the turning surface is 100mm, the distance from the turning surface to the interface with the pad is 190mm, and the total length at the interface with the pad is 660mm. The angle between the forward swept portion of the pad and the pad interface is 37°, and the rear side is perpendicular to the pad. The distance between the front end of the pad interface and the front end of the pad is 446.4mm. When viewed from above, the front end of the pad is triangular and the rear end is rectangular. The overall shape is similar to baffle 2. The angle of the triangular part is 24.7°, and the height of the rectangular part is 100mm.

[0032] The overall shape of the baffle is similar to that of the pad. When viewed from above, it consists of a triangle and a rectangle. The front triangle angle of the baffle is 60° and the length is 346.41mm, which is an equilateral triangle with a side of 400mm. The height of the rectangular part is 400mm and the length is 900mm. When viewed from the front, it can be seen that the baffle has a cut corner. The front triangle angle of the equilateral triangle on the cut corner side is 60° and the length is 251.2mm, which is an equilateral triangle with a side length of 290.1mm. The height of the rectangular part is 290.1mm and the length is 830.3mm. The thickness of the baffle is 20mm.

[0033] The calibration model is installed in the high-speed wind tunnel test section, along the incoming flow. The bottom of pad 1 is mounted in the wind tunnel mounting slot. After setting the appropriate incoming flow parameters, such as Mach number and angle of attack (AoA), wind tunnel testing can begin. Depending on the purpose of the wind tunnel experiment, pressure distribution across different cross sections can be measured using pressure taps, or the wing transition can be observed through methods such as thermal paint spraying. Figure 8 、 Figure 9 、 Figure 10The RANS calculation results of the pressure distribution of Slice 4 section at different angles of attack when the Mach number Ma is 1.5, 1.6, and 1.8 can be used as a reference value of the results. Under the corresponding incoming flow conditions and angles of attack, the experimental results or numerical simulation results should not deviate significantly from the given reference results.

[0034] Example 2 Subsonic leading edge wing configuration

[0035] use Figure 6 The airfoil shown is formed into subsonic leading edge wing configuration 3-2. The cross-sections at the wing root, 40% span position, and wing tip position correspond to Figure 6 After the "Root", "Middle", and "Tip" are formed, follow Figure 2 The three views shown are used to assemble a calibration model, specifically: a model for supersonic natural laminar flow wing calibration, including a wing 3, a pad 1 is provided under the wing 3, the pad 1 is installed in the wind tunnel installation slot, ensuring that the observed section of the wing 3 is located in the wind tunnel observation window; a baffle 2 is provided between the wing 3 and the pad 1, the baffle 2 separates the pad 1 and the wing 3, ensuring that the incoming flow felt by the wing 3 is uniform and stable and the same as the wind tunnel incoming flow, without acceleration or deceleration characteristics; the wing 3 has a typical aerodynamic design feature of a supersonic NLF wing, and can maintain an observable laminar flow area within the Mach number and angle of attack range of the experimental design) The leading edge of the wing 3 is behind The sweep angle, Mach number, and angle of attack range can cover the typical low-drag, low-sonic boom supersonic layout design features. This embodiment is a subsonic leading edge wing configuration 3-2, which is formed using three control surfaces: the wing root, 40% span position, and wingtip. The maximum thickness is 5% of the chord length and is located at 58% of the chord length. The difference between the three control surfaces is concentrated in the first 30% of the chord length of the upper surface, among which the airfoil thickness corresponding to the wing root increases; the airfoil camber of the subsonic leading edge wing configuration 3-2 is small, the airfoil camber at the subsonic leading edge (SW60) and the wing root (ROOT) is -0.26%, the maximum camber corresponds to 43.94% of the chord length, and its leading edge sweep angle is 60° ( Figure 11 ), the applicable range of Mach number is Ma=1.6~2.0, and the corresponding Reynolds number is Re=8×10 6 ~10×10 6 , the angle of attack is AoA = -2°~3°.

[0036] For the subsonic leading edge, the specific parameters of the subsonic leading edge wing configuration 3-2 are as follows Figure 4 As shown in the figure, the wingtip length is 174.2mm, the corresponding chord length is 315.6mm, the chord length at the wing root is 375mm, and the tip-to-root ratio is 0.842. The wingspan is 480mm, the leading edge sweep angle is 60°, the trailing edge sweep angle is 57.6°, and the angle at the wingtip is 115°.

[0037] The total length of the pad root is 407.9mm, the distance from the root to the turning surface is 114.4mm, the distance from the turning surface to the interface with the pad is 145.6mm, and the total length at the interface with the pad is 668.7mm. The swept-back angle of the swept-back part of the pad root is 20°, and the swept-back angle at the junction of the triangular section and the rectangular section is 36.8°. The angle between the swept-forward part and the interface with the pad is 35°, and the angle between the rear side of the pad and the baffle is 110°. The distance between the front end of the interface with the pad and the front end of the pad is 340.7mm. When viewed from above, the front end of the pad is triangular and the rear end is rectangular. The overall shape is similar to baffle 2. The angle of the triangular part is 25.5°, and the height of the rectangular part is 100mm.

[0038] During wind tunnel testing, the calibration model is installed in the high-speed wind tunnel test section, along the incoming flow. The base of the pad is mounted in the wind tunnel mounting slot. After setting the appropriate incoming flow parameters, such as Mach number and angle of attack (AoA), the wind tunnel test can begin. Depending on the purpose of the wind tunnel test, pressure distribution across different cross-sections can be measured using pressure taps, or the wing transition can be observed through methods such as thermal paint spraying. Figure 12 、 Figure 13 The RANS calculation results of the pressure distribution of Slice 4 section at different angles of attack under Mach numbers Ma of 1.8 and 2.0 are given, which can be used as reference values of the results. Under the corresponding incoming flow conditions and angles of attack, the experimental results or numerical simulation results should not deviate significantly from the given reference results.

Claims

1. A model for calibration of a supersonic natural laminar airfoil, comprising an airfoil (3), characterized in that: A pad (1) is provided below the wing (3), the pad (1) being installed in a wind tunnel installation slot, a baffle (2) being provided between the wing (3) and the pad (1), the baffle (2) separating the pad (1) and the wing (3), ensuring that the incoming flow felt by the wing (3) is uniform and stable and is the same as the incoming flow in the wind tunnel, without any acceleration or deceleration characteristics.

2. A model for supersonic natural laminar flow wing calibration according to claim 1, characterized in that: The wing (3) includes two structures, a supersonic leading edge wing configuration (3-1) and a subsonic leading edge wing configuration (3-2), both of which are formed using three control surfaces at the wing root, 40% spanwise position, and wingtip. The maximum thickness is 5% of the chord length and is located at 58% of the chord length. The difference between the three control surfaces is concentrated in the first 30% of the chord length of the upper surface, wherein the airfoil thickness corresponding to the wing root increases. The airfoil camber of the two structures is small, wherein the supersonic leading edge wing configuration (3-1) has a supersonic leading edge and airfoil camber of -0.12%, and the maximum camber corresponds to 54.46% of the chord length. The subsonic leading edge wing configuration (3-2) has an airfoil camber at the subsonic leading edge and wing root of -0.26%, and the maximum camber corresponds to 43.94% of the chord length.

3. A model for supersonic natural laminar flow wing calibration according to claim 1, characterized in that: When the wing (3) is a supersonic leading edge wing configuration (3-1), its leading edge sweep angle is 30°, the applicable Mach number range is Ma=1.5~1.8, and the corresponding Reynolds number is Re=8×10 6 ~13×10 6 , the angle of attack is AoA = -1°~4°.

4. A model for supersonic natural laminar flow wing calibration according to claim 1, characterized in that: When the wing (3) is a subsonic leading edge wing configuration 3-2, its leading edge sweep angle is 60°, the applicable Mach number range is Ma=1.6~2.0, and the corresponding Reynolds number is Re=8×10 6 ~10×10 6 , the angle of attack is AoA = -2°~3°.

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