A high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model

Abstract

The application discloses a high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model and belongs to the field of aircraft design. The leading edge of the test model body is single backward swept, the cross sections of the fuselage, the wing and the wing-body transition position all adopt NACA0012 airfoils, the wing and the fuselage are smoothly transitioned in shape and are symmetrically arranged in space, rotating control surfaces I and II are arranged at the outer side of the wing and close to the fuselage, fixed control surface I is arranged between the rotating control surfaces I and II, and rotating control surface III is arranged at the trailing edge of the fuselage. The design structure comprehensively considers the shape of the high-aspect-ratio aircraft, the internal structure design of the aeroelastic characteristics to be simulated, the position space design of the steering engine required by the driving control surface, the position space design of the sensor used for the active aeroelastic test and the like, and provides a reliable and feasible design scheme for the high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model.

Description

Technical Field

[0001] This invention relates to the field of aircraft design. More specifically, this invention relates to a high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model. Background Technology

[0002] To improve key performance indicators such as range and flight time, large aircraft are designed with a high aspect ratio and tailless layout, and extensive use of composite materials to reduce structural mass, thereby increasing payload. This results in increased structural flexibility, leading to severe aerodynamic / structural coupling problems, or aeroelasticity problems, under complex aerodynamic forces. Aeroelasticity directly impacts the flight safety and performance of large aircraft. Active aeroelasticity technology involves adding an active deflection control strategy to the aircraft's control surfaces, thereby altering the aerodynamic force distribution and achieving load reduction and enhanced stability. This represents a feasible approach to solving the aeroelasticity problems of large aircraft.

[0003] For aircraft design to apply active aeroelasticity technology, experimental verification is essential. Currently, wind tunnel testing has become a necessary step in assessing the aerodynamic and structural performance of aircraft during the design phase. Common test model design methods typically involve scaling down the aircraft based on similarity principles to ensure the test model's dimensions meet wind tunnel requirements. In high-speed wind tunnel testing, traditional test models generally use control surfaces at fixed angles to perform tests under relevant conditions. However, test models used for verifying active aeroelasticity technology must incorporate internal drive servos to ensure the control surfaces dynamically deflect according to a pre-defined pattern during testing. On one hand, high-speed wind tunnel tests involve relatively large aerodynamic loads, but the size of the wind tunnel limits the size of the test model, severely restricting the size and power of the internal servos. Furthermore, the drive linkage between the servos and control surfaces cannot extend beyond the model surface and disrupt its shape as in low-speed wind tunnels. On the other hand, the test model used for verification of active aeroelasticity technology must ensure that its structural characteristics are similar to those of the original aircraft. All-metal models cannot meet the requirements for similar structural characteristics, so a combination of metal beams and composite materials must be used. However, the matching design of the main components such as beams, servos, and sensors inside the model is the main challenge.

[0004] Currently, wind tunnel test models with servos are mainly divided into two categories. One category is used to measure the aerodynamic performance and control surface efficiency of aircraft (such as an automatic adjustment device for the control surface angle of a wind tunnel test aircraft model, patent application number CN201510773488.6). This type of test model basically adopts an all-metal design, leaving space for the servos inside the wing. During the test, it is only necessary to lock the control surfaces at a fixed angle. The dynamic performance requirements of the servos are not high, and usually only the design of the servos is considered, without providing the internal structure for simulating the aeroelastic characteristics of the aircraft. Therefore, it is not suitable for the design of test models for the verification of active aeroelastic technology. Another type is the low-speed wind tunnel test model (such as the literature title: Dongqiang ZHAO, Zhichun YANG, Xianang ZENG, et al. Wind tunnel test of gust loadalleviation for a large-scale full aircraft model [J]. Chinese Journal of Aeronautics, 2023, 36(4):201-216.). This type of model is large in size, has low aerodynamic load, and has sufficient space for internal servo motors. Moreover, the control surfaces and servo motors are mostly driven by exposed linkages, so the damage to the shape has little impact on the test measurement. It is not suitable for high-speed wind tunnel models. Summary of the Invention

[0005] One object of the present invention is to solve at least the above-mentioned problems and / or defects, and to provide at least the advantages described below.

[0006] To achieve these objectives and other advantages of the present invention, a high aspect ratio tailless aircraft high-speed wind tunnel active aeroelastic test model is provided. The leading edge of the main body of the test model is single swept back, and the cross sections of the fuselage, wings and wing-body transition position are all made of NACA0012 airfoil. The wings and fuselage are smoothly transitioned in shape and have a vertically symmetrical structure in space. The trailing edge of the wing is provided with three control surfaces. Among them, rotating control surface I and rotating control surface II are respectively provided on the outer side of the wing and near the fuselage. Rotating control surface I is directly driven to rotate by servo motor I arranged in the wing, and rotating control surface II is driven to rotate by servo motor II arranged in the front of the fuselage and equipped with a crank connecting rod. A fixed control surface I is arranged between the rotating control surface I and the rotating control surface II. The angle of the fixed control surface I is changed by changing the angle with the connecting block at the wing. The rotating control surface III, located at the rear edge of the fuselage, is driven to rotate by a servo motor III located inside the fuselage.

[0007] Preferably, two wing leading-edge main spars and two wing trailing-edge main spars are installed inside the wing, parallel to the leading edge of the test model. A rearward main spar parallel to the rearward edge is installed on the side of the wing leading-edge main spar and the wing trailing-edge main spar near the fuselage. Among them, the three main beams are located on the upper and lower symmetry planes of the test model; A side beam is provided at the root of the test model to connect the wing trailing edge main spars and the fuselage trailing edge main spars into an integral structure. The side beam is connected to the wing trailing edge main spars at the wing-fuselage transition position. Multiple ribs are installed on the three main beams at predetermined intervals.

[0008] Preferably, the leading edge main spars of the wing extends through the fuselage, and the trailing edge main spars of the wing are shorter than the leading edge main spars on the side near the wingtip.

[0009] Preferably, acceleration sensor I, acceleration sensor II, and acceleration sensor III are respectively arranged at the wingtip, nose, and tail.

[0010] Preferably, strain gauges I, II, III, VI, and V are respectively arranged at the wing-fuselage transition and the fuselage root of the three main spars.

[0011] Preferably, the rudder shafts corresponding to the rotating rudder surface II and the rotating rudder surface III are respectively equipped with angle encoder I and angle encoder II.

[0012] Preferably, the surface of the test model body is encapsulated with carbon fiber skin, and the interior of the test model body is filled with foam.

[0013] Preferably, as an alternative, the rotating control surface I is replaced with a fixed control surface II.

[0014] The present invention has at least the following beneficial effects: The present invention comprehensively considers the shape of the high aspect ratio tailless aircraft, the structural design that needs to simulate aeroelastic characteristics, the servo position space design required for the drive control surface, and the sensor position space design for active aeroelastic testing, etc., and provides a reliable and feasible design scheme for the high aspect ratio tailless aircraft high-speed wind tunnel active aeroelastic test model, which can effectively solve the problem of matching the structure and servo position design of the high aspect ratio tailless aircraft aeroelastic test model in high-speed wind tunnel.

[0015] Other advantages, objectives and features of the present invention will become apparent in part from the following description, and in part from those skilled in the art through study and practice of the invention. Attached Figure Description

[0016] Figure 1A schematic diagram of the external shape of a high-aspect-ratio tailless aircraft active aeroelastic semi-experimental model in a high-speed wind tunnel. Figure 2 For the present invention Figure 1 A schematic diagram of the cross-section at mark A in the middle; Figure 3 For the present invention Figure 1 A schematic diagram of the cross-section at mark B in the middle; Figure 4 For the present invention Figure 1 A schematic diagram of the cross-section at the C mark; Figure 5 A schematic diagram of the internal structure of a high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic semi-experimental model. Figure 6 For the present invention Figure 5 A schematic diagram of the cross-section at the D mark in the middle; Figure 7 For the present invention Figure 5 A schematic diagram of the cross-section at mark B in the middle; Figure 8 For the present invention Figure 5 A schematic diagram of the cross-section at the C mark; Figure 9 For the present invention Figure 5 A schematic diagram of the cross-section at the E mark in the middle; Figure 10 This is a cross-sectional schematic diagram showing the coordination of the main beams, foam filling, and skin structure within the fuselage of the experimental model of this invention. Figure 11 This is a cross-sectional schematic diagram of the main beam, foam filling, and skin structure fitting together at the wing within the experimental model of this invention. Figure 12 A schematic diagram of the overall shape of a high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model. Figure 13 This is a schematic diagram of the flow field mesh of the semi-modal test model in the simulation experiment of this invention; Figure 14 This is a schematic diagram of the structural mesh of the semi-modal test model in the simulation experiment of this invention; Figure 15 This is a schematic diagram of the bending modes of the semi-modal test model in the simulation experiment of this invention; Figure 16 This is a schematic diagram of the torsional modes of the semi-modal test model in the simulation experiment of this invention; Among them, 1-main body; 2-rotating control surface I, 3-rotating control surface II, 4-rotating control surface III, 5-fixed control surface I, 6-servo motor I, 7-servo motor II, 8-servo motor III, 9-wing leading edge main sparsity, 10-wing trailing edge main sparsity, 11-fleet trailing edge main sparsity, 12-side sparsity, 13-rib plate, 14-accelerometer I, 15-accelerometer II, 16-accelerometer III, 17-strain gauge I, 18-strain gauge II, 19-strain gauge III, 20-strain gauge VI, 21-strain gauge V, 22-angle encoder I, 23-angle encoder II, 24-linear servo linkage, 25-crank linkage, 26-carbon fiber skin, 27-foam, 28-fuselage, 29-wing, 30-fixed control surface II. Detailed Implementation

[0017] The present invention will now be described in further detail with reference to the accompanying drawings, so that those skilled in the art can implement it based on the description.

[0018] It should be understood that terms such as “having,” “comprising,” and “including” as used herein do not exclude the presence or addition of one or more other elements or combinations thereof.

[0019] It should be noted that in the description of this invention, the orientations or positional relationships indicated by terms are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing this invention and simplifying the description. They do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as limiting this invention. In addition, the terms "I" and "II" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0020] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installed", "equipped", "sleeved / connected", "connected", etc., should be interpreted broadly. For example, "connection" can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium; it can be a connection within two components. For those skilled in the art, the specific meaning of the above terms in this invention can be understood according to the specific circumstances.

[0021] like Figures 1-11 As shown, the invention proposes a high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model (hereinafter referred to as the test model).

[0022] The main body 1 of the experimental model designed in this invention mainly includes multiple main beams, ribs, rudder surfaces, rudder motors, acceleration sensors, strain gauges, angle encoders, linear rudder connecting rods 24, crank connecting rods 25, carbon fiber skin 26, and foam for filling, etc.

[0023] The experimental model simulates the left half of a high aspect ratio tailless aircraft (viewed from the nose to the tail). The base of the experimental model is fixed to the side wall support device inside the wind tunnel test section.

[0024] The experimental model features a single swept leading edge. The fuselage 28, wings, and wing-fuselage transition sections all utilize the NACA0012 airfoil, exhibiting a symmetrical structure. The wing sections are identical across the entire fuselage, resulting in a smooth transition between the wing and fuselage 28. The wing trailing edge is designed with three control surfaces. The control surface near the outer edge of the wing is a rotating control surface I2 with a control shaft, which can be directly driven to rotate by a servo motor I6 located within the wing to perform corresponding deflection operations. In practical applications, the rotating control surface I2 is directly driven by the servo motor I6, which is fixed to the wing's leading-edge main sparsity 9 and also connected to the trailing-edge main sparsity on its side. The output shaft of servo motor I6 is connected to the control shaft of the rotating control surface I2 near the wingtip.

[0025] The rotating control surface II3 of the wing, located near the fuselage 28, can be driven to rotate by a servo motor II7 with a crank-connecting rod 25 located at the front of the fuselage 28, thus completing the corresponding deflection operation. In practical applications, the servo motor II7 is designed inside the fuselage 28, fixed between the wing's leading-edge main spars 9 and trailing-edge main spars, and connected to the servo motor's output shaft and the control surface's rotation shaft via the crank-connecting rod 25, driving the control surface to deflect while passing around the wing's trailing-edge main spars. Since the thickness of the fuselage 28 is significantly greater than the thickness of the wing, the space dimensions of the servo motor II7 within the rotating control surface II3 are sufficient to meet the power design requirements. A fixed control surface I5 is located between rotating control surface I2 and rotating control surface II3. The angle of the fixed control surface I5 can only be changed by changing the angle of the connecting block at the wing.

[0026] Additionally, a rotating control surface III4 is designed at the rear edge of the fuselage. It is driven to rotate by a servo motor III7 located within the fuselage 28 to perform the corresponding deflection operation. In practical applications, the rotating control surface III4 is deflected by a linear servo motor III, which is fixed from below to the rear edge main beam. The servo motor III7 is connected to the rotating control surface III4 from above the rear edge main beam 11 via a hinged linear servo linkage 24, allowing it to deflect while avoiding the rear edge main beam 11.

[0027] Inside the test model, there are two main spars parallel to the leading edge and one main spar parallel to the trailing edge, namely the wing leading edge main spar 9, the wing trailing edge main spar 10, and the fuselage trailing edge main spar 11. These three main spars are primarily used to adjust the modal frequency characteristics of the test model. The three main spars are located on the upper and lower symmetrical planes of the model and are fixed to the root side spar 12 of the test model. The wing trailing edge main spar 10 and the fuselage trailing edge main spar 11 connect at the wing-fuselage transition point. The wing leading edge main spar 9 extends through the wing 29 and fuselage 28, while the wing trailing edge main spar 10 is slightly shorter than the wing leading edge main spar 9 near the wingtip, leaving design space for the wingtip servo.

[0028] Several ribs 13 are designed between the three main beams to maintain the shape of the test model and fix the control surface. The spacing between the ribs 13 depends on the length of the control surface and the servo motor.

[0029] Accelerometer I14, Accelerometer II15, and Accelerometer III16 are designed at the wingtip, nose, and tail of the aircraft to sense the buoyancy and pitch motion of the test model during the test.

[0030] Five strain gauges (i.e., strain gauge I17, strain gauge II18, strain gauge III19, strain gauge VI20, and strain gauge V21) are designed at the wing-fuselage transition and fuselage root of the three main spars to sense the elastic deformation of the main spars, facilitating the analysis of the aeroelastic characteristics of the experimental model during the test. Angle encoders I22 and II23 are designed on the control surfaces at the rear edge of the fuselage and the wing control surfaces near the fuselage, respectively, to provide feedback on the actual deflection state of the control surfaces.

[0031] In addition to the space required for the movement of the main beam, ribs, accelerometer, strain gauge, servo motor, and servo motor linkage, the interior of the test model is filled with foam 27, and the surface of the test model is covered with high-strength carbon fiber skin 26 to maintain its shape. The foam 27 and carbon fiber skin 26 are bonded together with adhesive suitable for the test model.

[0032] After the experimental model is designed and manufactured, it is fixed to the side wall support device inside the wind tunnel test section by the fuselage root, and can then be used to carry out research on high-speed wind tunnel active aeroelasticity tests of a high aspect ratio tailless aircraft. It should be noted that traditional pure metal models are too rigid to simulate aeroelastic characteristics. The beam + rib design is the main design method for aeroelastic test models, which facilitates the adjustment of the model's stiffness, modal and other aeroelastic characteristics. However, it is necessary to fully consider the space design between the internal structure and the servo motors, sensors, etc. This invention, through its internal space design, makes it well meet the usage requirements.

[0033] like Figure 12As shown, based on the semi-modal test model designed in this invention, by further reducing the model scale and making a symmetrical design, a full-modal active aeroelastic test model can be obtained. However, due to the further reduction in model size, the space inside the wing will not be sufficient to accommodate the servo motors that can meet the test power and size requirements. Therefore, the rotating control surface I2 near the wingtip needs to be changed to the fixed control surface II30, and only the rotating control surface II3 and rotating control surface III4 near the fuselage 28 and the rear edge of the fuselage are retained.

[0034] Verification example: To verify the effectiveness of this invention, this verification example uses... Figure 5 The semi-modal test model was simulated and verified. The simulation process mainly includes three parts: aerodynamic mesh generation, structural mesh generation, and aeroelastic property analysis. First, the flow field mesh was generated according to the shape of the semi-modal test model, such as... Figure 13 As shown; secondly, the structural mesh is divided according to the structure of the semi-modal test model, as follows. Figure 14 As shown, plate elements were used for meshing in this verification example, but beam elements or other similar elements can be used instead in practice. Finally, by establishing spline interpolation between the flow field mesh and the structural mesh, the aeroelastic properties of the semi-modulus test model can be obtained. Figure 15 This is a schematic diagram of the bending modes of the semi-modal test model. Figure 16 This is a schematic diagram of the torsional modes of the semi-modal test model. It should be noted that... Figures 15-16 The black line in the middle represents the initial position of the wing corresponding to the semi-modal test model, while Figures 15-16 The graphic area constructed from the medium-colored region is used to represent the position of the wing after deformation in the corresponding mode. The farther the graphic area is from the initial position, the greater the deformation. The corresponding colors on the graphic area correspond to the stress distribution on the wing, with blue to red representing stress increasing from small to large. The darker the blue, the smaller the stress, and the darker the red, the greater the stress.

[0035] The aeroelastic properties of the semi-modulus test model are compared with the design target. If the design conditions are met, the structural design and subsequent processing of the semi-modulus test model can be carried out according to the above implementation method to obtain a wind tunnel test model for active aeroelastic testing. Otherwise, the cross-section of the main beam needs to be adjusted to change the aeroelastic properties of the semi-modulus test model to make it closer to the design target.

[0036] The above solution is merely an illustration of a preferred example and is not limited thereto. When implementing this invention, appropriate substitutions and / or modifications can be made according to the user's needs.

[0037] The number of devices and processing scale described herein are for the purpose of simplifying the description of the invention. Applications, modifications, and variations of the invention will be readily apparent to those skilled in the art.

[0038] Although embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. It can be applied to various fields suitable for the present invention. Other modifications can be readily made by those skilled in the art. Therefore, without departing from the general concept defined by the claims and their equivalents, the present invention is not limited to the specific details and examples shown and described herein.

Claims

1. A high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model, characterized in that, The leading edge of the main body of the test model is single swept back. The cross sections of the fuselage, wings and wing-fuselage transition section all adopt the NACA0012 airfoil. The wings and fuselage are smoothly transitioned in shape and have a vertically symmetrical structure in space. The trailing edge of the wing is provided with three control surfaces. Among them, rotating control surface I and rotating control surface II are respectively provided on the outer side of the wing and near the fuselage. Rotating control surface I is directly driven to rotate by servo motor I arranged in the wing, and rotating control surface II is driven to rotate by servo motor II arranged in the front of the fuselage and equipped with a crank connecting rod. A fixed control surface I is arranged between the rotating control surface I and the rotating control surface II. The angle of the fixed control surface I is changed by changing the angle with the connecting block at the wing. The rotating control surface Ⅲ is located at the rear edge of the fuselage and is driven to rotate by the servo motor Ⅲ arranged inside the fuselage; Inside the wing, two main spars, the leading edge spar and the trailing edge spar, are installed parallel to the leading edge of the test model. On the side of the fuselage near the leading edge main spars and trailing edge main spars of the wing, there are trailing edge main spars parallel to the rear edge of the fuselage. Among them, the three main beams are located on the upper and lower symmetry planes of the test model; A side beam is provided at the root of the test model to connect the wing trailing edge main spars and the fuselage trailing edge main spars into an integral structure. The side beam is connected to the wing trailing edge main spars at the wing-fuselage transition position. Among them, multiple ribs are set at predetermined intervals on the three main beams; The wing leading edge main spars extend through the wing and fuselage, while the wing trailing edge main spars are shorter than the wing leading edge main spars on the side near the wingtip.

2. The high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model as described in claim 1, characterized in that, Acceleration sensor I, acceleration sensor II, and acceleration sensor III are respectively installed at the wingtip, nose, and tail.

3. The high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model as described in claim 1, characterized in that, Strain gauges I, II, III, VI, and V are respectively installed at the wing-body transition and fuselage root of the three main spars.

4. The high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model as described in claim 1, characterized in that, Angle encoder I and angle encoder II are respectively installed on the rudder shafts corresponding to the rotating rudder surface II and the rotating rudder surface III.

5. The high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model as described in claim 1, characterized in that, The surface of the test model body is encapsulated with carbon fiber skin, and the interior of the test model body is filled with foam.

6. The high-aspect-ratio tailless aircraft high-speed wind tunnel active aeroelastic test model as described in claim 1, characterized in that, As an alternative, the rotating control surface I is replaced with the fixed control surface II.

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