A comprehensive wind tunnel test platform for a tiltrotor aircraft

Abstract

A tilt-rotor aircraft flutter comprehensive wind tunnel test platform belongs to the technical field of wind tunnel test. The platform is a semi-span length modular model: the wing module is fixed at one end to the wind tunnel test table, the other end is connected to the nacelle module, and the inside is installed with a rotor system. The helical flutter test of the platform mainly adopts a non-powered windmill operation mode, the rudder in the nacelle is used as a total pitch control actuator, and does not continuously input rotary driving power to the rotor shaft. The rudder drives the pitch control base and the tilt disc through the rudder rocker arm and the long control rod, and then transmits the variable pitch amount to the rotor blades through the short control rod and the blade clamp; when the flow velocity changes, the rudder input is adjusted to change the total pitch, so that the rotor reaches and maintains the predetermined rotating speed working condition. The nacelle module together with the rotor system obtains two rotating degrees of freedom of pitch and yaw with the elastic deformation of the wing, and the two degrees of freedom are coupled to form the motion form of helical flutter. The platform is used for helical flutter wind tunnel tests of different wing, hub and blade configurations.

Description

Technical Field

[0001] This invention belongs to the field of wind tunnel testing technology and relates to a comprehensive wind tunnel testing platform for flutter of tiltrotor aircraft. Background Technology

[0002] Tiltrotor aircraft achieve dynamic switching between helicopter and fixed-wing modes through the tilting of their rotor system, combining the superior performance of helicopter vertical takeoff and landing (VTOL) with that of propeller-driven aircraft high-speed cruise. This represents a significant development direction in the field of electric vertical takeoff and landing (EVL) aircraft. In fixed-wing flight mode, the rotor / nacelle system is connected to the wing via a flexible structure, resulting in strong aeroelastic coupling between the rotor, wing, and nacelle. When flight speed and rotor speed meet specific conditions, a precession-type aeroelastic divergence phenomenon, i.e., propeller flutter, is induced, where the engine nacelle pitch and yaw are coupled. Propeller flutter directly limits the maximum flight speed of tiltrotor aircraft and seriously threatens flight safety. Accurately predicting its flutter boundary is a crucial aspect of tiltrotor aircraft development.

[0003] Unlike the propeller flutter of conventional propeller aircraft, the flutter problem of tiltrotor aircraft, based on the classic propeller flutter mechanism, further introduces additional degrees of freedom such as flapping, flaring, and pitch changes of the rotor hub and rotor blades. The fluid-structure interaction between the wing, nacelle, and rotor is more complex, and the aerodynamic loads exhibit strong nonlinearity, making it difficult to accurately predict the flutter velocity boundaries through theoretical modeling and numerical simulation. Therefore, flutter wind tunnel tests based on similar scaled-down models have become an important technical approach to obtain real flutter data and verify simulation methods.

[0004] In the scaled-down test of the unpowered wind turbine, the rotor shaft is not continuously driven by an external power unit. The rotor relies on the aerodynamic torque generated by the wind tunnel flow to overcome the mechanical resistance torque and maintain rotation. When the flow velocity changes, the rotor aerodynamic torque and its equilibrium speed change accordingly. If the rotor collective pitch is fixed, wind speed sweeping will simultaneously cause speed drift, making the two variables of flow velocity and rotor speed coupled together, making it difficult to identify the spiral flutter boundary according to the predetermined speed condition. Therefore, it is necessary to adjust the collective pitch during rotor operation to ensure that the rotor reaches the predetermined speed condition at each flow velocity test point.

[0005] However, existing flutter wind tunnel test models are generally designed and manufactured in a one-time scaled-down manner for specific target objects. After the test, they are considered obsolete and scrapped. For flutter tests on different models or with different structural forms, a completely new model needs to be designed and manufactured, resulting in a huge waste of test resources. At the same time, the flutter mechanism of tiltrotor aircraft varies greatly depending on the hub type, rotor flexibility, and tilting mechanism, making it difficult to complete the test verification using a unified standard model. Advanced flutter active control strategies and the application of new materials also require test verification before being put into use in models, and existing single-purpose models cannot provide a universal and reusable verification platform for them. Summary of the Invention

[0006] To address the aforementioned issues, this invention proposes a comprehensive platform suitable for wind tunnel testing of helical flutter in tiltrotor aircraft. This platform employs a modular design, allowing for the replacement of wings, hubs, and blades for different design requirements. It primarily utilizes a non-powered windmill operation mode for helical flutter testing. In this mode, the rotor is driven to rotate by the wind tunnel inflow, with the servo motor only driving the collective pitch control mechanism. By adjusting the collective pitch during the inflow velocity sweep, a predetermined rotor speed is achieved. This eliminates the need for a continuous, through-type power transmission chain that continuously inputs driving torque to the rotor shaft, thereby reducing the impact of non-target transmission components on the model's mass, stiffness, and dynamic boundaries. Furthermore, it supports comparative flutter tests of different configurations.

[0007] The technical solution of the present invention is as follows: a tiltrotor aircraft flutter integrated wind tunnel test platform, comprising a wing module, a nacelle module and a rotor system; the root of the wing module is fixed to the mounting fixture of the wind tunnel test rig, forming a root fixed support boundary; the nacelle module is installed on the wingtip of the wing module via a connecting joint, and the box section at the wingtip is connected to the nacelle joint with bolts.

[0008] The nacelle module has a bearing housing installed inside the nacelle shell, and the rotor shaft is supported in the bearing housing via the main shaft bearing;

[0009] The servo, servo rocker arm and long control stick are arranged inside the nacelle shell. The guide rod and pitch control base are installed at the front end of the bearing housing. The guide rod is arranged along the rotor shaft axis, constraining the pitch control base to only make axial translation and not rotate with the bushing. The bearing housing is provided with an opening that mates with the guide rod.

[0010] One end of the long control stick is connected to the servo arm, and the other end is connected to the pitch control base. The rotation plane of the servo arm is parallel to the rotor axis. The servo drives the servo arm to rotate, and the rotation of the servo arm is converted into axial displacement of the pitch control base along the rotor axis through the long control stick.

[0011] The pitch control base is connected to the bushing via a deep groove ball bearing; the bushing is fitted onto the rotor shaft, which has an axially extending slot; a connector passes through the bushing and engages with the slot, allowing the bushing to rotate with the rotor shaft and move axially relative to the rotor shaft within the range defined by the slot; a swashplate is fixed to the bushing, and the swashplate is connected to the blade clamp via a short control lever.

[0012] The rotor system's hub assembly is mounted at the end of the rotor shaft, and the rotor blades are mounted on the hub assembly via blade clamps. The platform operates in a non-powered windmill mode during the helical flutter wind tunnel test. The rotor shaft is not continuously driven by an external power unit, and the rotor is driven to rotate by the wind tunnel flow. The servo motor is used to adjust the rotor collective pitch without providing rotational drive power to the rotor shaft.

[0013] Furthermore, the servo receives external control signals and drives the servo rocker arm to rotate, causing the pitch control base to move axially along the rotor shaft. The axial displacement of the pitch control base is transmitted to the bushing and swashplate via a deep groove ball bearing, causing the bushing to move axially within the slotted range of the rotor shaft. The swashplate then pushes the blade clamp via a short control lever, causing the blade clamp to drive the rotor blades to rotate around their own pitch axis, thus achieving continuous adjustment of the rotor collective pitch.

[0014] Furthermore, under the condition of no-power wind turbine operation, the collective pitch of the rotor is adjusted by changing the input signal of the servo motor based on the measured rotor speed, so that the rotor reaches the predetermined speed under different incoming flow velocities and remains within the allowable deviation range of the predetermined speed.

[0015] Furthermore, both the rotor hub assembly and rotor blades in the rotor system can be replaced independently; the rotor hub assembly adopts a constant velocity universal joint structure, and the flapping of each rotor blade is determined by the rigid body motion of the rotor hub assembly. The rotor disk, as a rigid body, tilts as a whole around the center of the rotor hub in any direction, and the flapping direction is rigid.

[0016] Furthermore, the rotor hub assembly is a hinged rotor hub or a rigid rotor hub, in order to study the coupling form and corresponding flutter characteristics of the rotor blade flapping, oscillation, pitch change degree of freedom and the wing and nacelle under different rotor hub constraints.

[0017] The rotor blades are mounted on the rotor hub assembly via blade clamps. The blade clamps rotate around their own axis to achieve pitch variation. The blades can be made of different materials, airfoils, spans, or twist angles.

[0018] The beneficial effects of this invention are as follows: 1. The platform mainly adopts a non-powered windmill operation mode. The rotor shaft is not continuously driven by an external power unit. The rotor speed is determined by the aerodynamic torque, mechanical drag torque, and rotor inertia under the action of the incoming flow. When the incoming flow speed changes, the aerodynamic torque changes accordingly, and the equilibrium speed of the fixed collective pitch rotor will drift. This application uses a collective pitch control chain composed of a servo motor, servo motor rocker arm, long control stick, pitch control base, deep groove ball bearing, bushing, swashplate, short control stick, and blade clamp to adjust the collective pitch according to the measured speed during rotor operation, so that the aerodynamic torque and drag torque can be rebalanced near the target speed, thereby enabling the rotor to reach the predetermined speed and stay within the allowable deviation range. The helical flutter boundary is a function of the incoming flow speed and the rotor speed. The above adjustment can reduce the influence of speed drift on flutter boundary identification during wind speed sweep and improve the controllability and repeatability of each test condition. Moreover, compared with the conventional helicopter swashplate, the flutter test model has been simplified and arranged in the space-constrained nacelle.

[0019] 2. The integrated wind tunnel test platform combines a unified platform with controllable collective pitch. The variable pitch degree of freedom is achieved through a variable pitch control mechanism, while blade flapping and twitching degrees of freedom are controlled by changing different hub configurations. The variable pitch output is applied to the rotor blades via a short control lever and blade clamps. In lateral comparative tests of different hub or blade configurations, the incoming flow velocity, target rotational speed, and other key conditions are controlled according to the test objective, and the collective pitch corresponding to the windmill's equilibrium state for each configuration is recorded. When collective pitch is selected as an independent comparison variable, the collective pitch conditions are kept consistent. This allows for the differentiation of the impact of configuration changes and test condition changes on flutter boundaries, improving the comparability of test results for different configurations.

[0020] 3. This invention designs the tiltrotor aircraft flutter wind tunnel test device as a modular integrated test platform. The wing, hub, and rotor blades can all be independently replaced, and each module can be repeatedly disassembled and reused. This avoids the waste of traditional wind tunnel flutter model testing, which involves scrapping the model after each test and the repetitive design and manufacturing, thus saving test costs and shortening the development cycle. The platform can replace different types of hubs and rotor blades with different configurations on the same structure, covering the differences in flutter coupling mechanisms under different hub types and rotor flexibility. It allows for the study of complex flutter problems under different structural combinations using a universal platform, and can pre-verify active flutter control strategies and the application of new materials.

[0021] In summary, this invention provides a unified and reusable integrated test platform for the study of flutter mechanism and model design verification of tiltrotor aircraft. Attached Figure Description

[0022] Figure 1 This is a schematic diagram of the overall assembly of the tiltrotor aircraft flutter integrated wind tunnel test platform.

[0023] Figure 2This is a schematic diagram of the nacelle cross-section.

[0024] Figure 3 This is a schematic diagram of nacelle-rotor pitch control.

[0025] Figure 4 It is the calibration relationship curve between the servo target position command value and the rotor pitch angle.

[0026] Figure 5 These are the first three main mode shapes of the test platform; among them: (a) vertical bending mode; (b) in-plane bending mode; and (c) torsional mode.

[0027] Figure 6 The relationship between the calculated results of helical flutter and the rotor speed is shown; where: (a) incoming flow velocity; (b) helical flutter frequency.

[0028] In the diagram: 1. Wing module, 2. Nacelle module, 3. Rotor system; 21. Nacelle shell, 22. Bearing housing, 23. Servo, 24. Servo rocker arm, 25. Long control stick, 31. Rotor shaft, 32. Main shaft bearing, 33. Guide rod, 34. Pitch control base, 35. Deep groove ball bearing, 36. Bushing, 37. Swashplate, 38. Short control stick, 39. Blade clip. Detailed Implementation

[0029] The specific embodiments of the present invention will be further described below with reference to the accompanying drawings and technical solutions.

[0030] A tiltrotor aircraft flutter integrated wind tunnel test platform adopts a modular design and consists of three independent modules: a wing module, a nacelle module, and a rotor system module. By replacing and switching the modules, the platform can cover flutter wind tunnel tests for various design objects and structural forms.

[0031] The platform is a semi-span model: one end of the wing module is fixed to the installation fixture in the wind tunnel or on the ground, and the other end is connected to the nacelle module; the rotor system is installed inside the nacelle module.

[0032] The wing module is a straight, truss-type wing. The main wing spars are the primary load-bearing components, and they are integrated with the wing ribs using a frame-integrated structure with matching openings to form an overall load-bearing skeleton. Each wing segment has a reinforcing rib in the middle that connects to the main spars at a single point to release excess constraints and avoid static instability and stress concentration. The wing ribs, stringers, leading edge, and trailing edge form a frame, with the skin attached to it, transferring aerodynamic loads to the main spars. The wing module connects to the nacelle module via a connector at the wingtip. Designed as a completely replaceable unit, it allows for direct replacement of the entire wing module to suit different airfoils, spans, planforms, or stiffness requirements.

[0033] The nacelle module is the core of the platform's load-bearing and control system. Its shell consists of two halves bolted together and mounted to the wingtip of the wing module via a connecting joint. The platform's spiral flutter tests primarily employ a windmill-like operation mode, with the rotor driven by airflow from the wind tunnel. The nacelle module does not contain a power motor continuously inputting driving torque to the rotor shaft or a power transmission chain running through the wing. The nacelle module, together with the rotor system, gains pitch and yaw rotational degrees of freedom through the elastic deformation of the wing. These two components couple and precess under the influence of rotor gyroscopic torque and aerodynamic loads, forming the spiral flutter motion. The internal shell houses the pitch control mechanism: the servo motor acts only as a collective pitch actuator, driving the pitch control base and swashplate via the rocker arm and long control stick, and then transmitting the pitch variable to the rotor blades via the short control stick and blade clamp.

[0034] The rotor system consists of a hub assembly and rotor blades, both of which can be independently replaced. The hub assembly primarily employs a constant velocity universal joint structure. The flapping of each blade is determined by the rigid body motion of the hub. The rotor disk, as a rigid body, tilts as a whole around the center of the hub in any direction, exhibiting rigid oscillation. The hub can be entirely replaced with other structural forms such as an articulated hub or a rigid hub to study the coupling forms and corresponding flutter characteristics of blade flapping, oscillation, and pitch variation under different hub constraints, in relation to the wing and nacelle. The rotor blades are mounted to the hub assembly via blade clamp assemblies. The blade clamps rotate around their own axes to achieve pitch variation. The blades can be replaced with different materials, airfoils, spans, or twist angles.

[0035] Example 1: Installation of the test model:

[0036] like Figure 1 As shown, this invention provides a tiltrotor aircraft flutter integrated wind tunnel test platform, including a wing module 1, a nacelle module 2, and a rotor system 3. During installation: First, the root of the wing module 1 is fixed to the mounting fixture of the wind tunnel test rig, forming a root support boundary; second, the nacelle module 2 is installed to the wingtip of the wing module 1 via a connecting joint, with the wingtip section connected to the nacelle joint using bolts; subsequently... Figure 2Complete the internal assembly of nacelle module 2: A bearing housing 22 is installed inside the nacelle outer shell 21. The rotor shaft 31 is supported on the bearing housing 22 via a main shaft bearing 32. A servo motor 23, a servo motor rocker arm 24, and a long control stick 25 are arranged inside the nacelle outer shell 21. A guide rod 33 and a pitch control base 34 are installed at the front end of the bearing housing 22. The bearing housing 22 has openings that mate with the guide rod 33. The pitch control base 34 is connected to a bushing 36 via a deep groove ball bearing 35. The bushing 36 is fitted onto the rotor shaft 31. The rotor shaft 31 has an axially extending slot. The connecting piece passes through the bushing 36 and engages with the slot, so that the bushing 36 rotates with the rotor shaft 31 and can move axially relative to the rotor shaft 31 within the range defined by the slot. The swashplate 37 is fixed to the bushing 36. Finally, the rotor hub assembly of the rotor system 3 is installed at the end of the rotor shaft 31. The rotor blades are installed on the rotor hub assembly via the blade clamp 39. For a detailed installation method of the universal joint rotor hub assembly, please refer to patent CN120553144A.

[0037] Example 2: Variable Pitch Operation Procedure

[0038] like Figure 3 As shown, the rotor blade pitch is controlled by a pitch control mechanism within the nacelle module 2. The servo motor 23 receives external control signals and drives the servo arm 24 to rotate. One end of the long control stick 25 is connected to the servo arm 24, and the other end is connected to the pitch control base 34, converting the rotation of the arm into axial displacement of the pitch control base 34 along the rotor shaft 31. The guide rod 33 is arranged axially along the rotor shaft 31, constraining the pitch control base 34 to only perform axial translation. The pitch control base 34 is connected to the bushing 36 via a deep groove ball bearing 35. The deep groove ball bearing 35 allows the pitch control base 34 to rotate relative to the bushing 36 and transmits axial control force. The bushing 36 is fitted onto the rotor shaft 31 and moves axially within the slotted range of the rotor shaft 31, while rotating with the rotor shaft 31. The swashplate 37 is fixed to the bushing 36. The axial displacement of the pitch control base 34 is transmitted to the bushing 36 and the swashplate 37 via the deep groove ball bearing 35. The swashplate 37 then pushes the blade clamp 39 via the short control lever 38, causing the blade clamp 39 to drive the rotor blades to rotate around their own pitch axis, thereby achieving synchronous and continuous adjustment of the total pitch of each rotor blade.

[0039] During the helical flutter wind tunnel test, the platform primarily operates in a non-powered windmill mode. The rotor shaft 31 is not continuously driven by an external power unit; the rotor blades generate aerodynamic torque under the influence of the wind tunnel airflow, driving the rotor system to rotate. The servo motor 23 only provides displacement input to the collective pitch control mechanism and does not provide rotational drive power to the rotor shaft 31. When the windmill is running stably, the rotor aerodynamic torque and the mechanical resistance torque generated by bearings, connecting components, etc., reach equilibrium at the corresponding speed.

[0040] During the incoming flow velocity sweep, the target rotational speed at the current test point is first set, and the actual rotor speed is measured. When the actual rotational speed deviates from the target speed, the input signal of the servo motor 23 is adjusted. This adjustment, via the servo motor rocker arm 24, long control stick 25, pitch control base 34, deep groove ball bearing 35, bushing 36, swashplate 37, short control stick 38, and blade clamp 39, changes the rotor collective pitch, thereby altering the aerodynamic torque until the rotor speed enters the allowable deviation range of the target speed. Data is then collected while maintaining this rotational speed condition, and the incoming flow velocity is further varied. This adjustment process is used to reduce rotational speed drift caused by wind speed sweep. The target rotational speed, allowable deviation, and collective pitch corresponding to each test point are set and recorded according to the specific test plan.

[0041] Example 3: Measurement of pitch adjustment range:

[0042] In this embodiment, an STS3020 TTL serial bus servo motor is used as servo motor 23. The specific pitch measurement method can be found in patent CN122443713A. The controller sends digital command packets to servo motor 23 via half-duplex asynchronous serial communication. The target position command value P in the digital command packet is used to set the position of the servo motor's output shaft. According to the servo motor's nominal encoding parameters, its target position encoding range is 0 to 4095, containing a total of 4096 position counts, corresponding to 360° of the output shaft rotation, with a nominal resolution of approximately 0.088° / position count. This encoding range is not equivalent to the actual usable mechanical travel of the servo motor after it is installed in the variable pitch control mechanism.

[0043] To verify the adjustment range of the pitch control mechanism and the correspondence between the servo target position command value and the rotor pitch angle, a static geometric calibration was performed on the pitch control mechanism. During calibration, the rotor was kept stationary, and the servo target position command value P was increased from 2470 to 2830. A measurement point was set for every 40 position counts, for a total of 10 measurement points including the two endpoints. The range of 2470 to 2830 is the calibration range of the pitch control mechanism in this embodiment.

[0044] At each target position command value, the heights of the leading and trailing edges of the three rotor blades are measured at the same selected cross-section, with the same reference plane used as the measurement reference. Then, the average heights of the leading and trailing edge measurement points of the three rotor blades are calculated, and the average pitch angle θ of the three rotor blades is calculated based on this. The average pitch angle θ is calculated according to the following formula:

[0045] ;

[0046] In the formula, θ is the average pitch angle of the three rotor blades in degrees; dle and dte (with underlined lines) are the average heights of the leading and trailing edge measurement points of the three rotor blades, respectively; cb is the chordal distance between the leading and trailing edge measurement points at the same measurement section, which is 35 mm in this embodiment.

[0047] Table 1. Servo target position command values ​​and rotor pitch angle measurement results

[0048]

[0049] Measurement results show that as the servo target position command value P increases from 2470 to 2830, the average pitch angle of the three rotor blades monotonically increases from approximately -7.1° to approximately 47.2°, corresponding to a pitch angle adjustment range of approximately 54.3°. Within this calibration range, for every 40 position counts increase in the servo target position command value, the average increment of the average pitch angle across the entire range is approximately 6.03°.

[0050] By performing linear fitting on the above measurement data, the calibration relationship between the servo target position command value and the rotor pitch angle is obtained as follows:

[0051] ;

[0052] In the formula, θ is in degrees, and P is in position counts. The coefficient of determination Rfit 2 The value is 0.9987, and the maximum absolute fitting residual of each measurement point relative to the fitted straight line is approximately 1.12°. These results indicate that within the calibration interval of 2470–2830° in this embodiment, there is a good monotonicity and approximately linear relationship between the servo target position command value and the average pitch angle of the three rotor blades. Based on this static calibration relationship, the rotor collective pitch can be quantitatively set by adjusting the servo target position command value; this result illustrates the static correspondence between digital commands and collective pitch input. The speed maintenance effect during wind turbine operation is determined by the speed measurement and adjustment process during wind tunnel testing. The fitting relationship is only applicable to this embodiment and the aforementioned calibration interval.

[0053] Example 4: Modal characteristics and helical flutter calculation of the test platform:

[0054] Under the boundary condition that the root of wing module 1 is fixed, finite element modal analysis was performed on the test platform. The three main modes most closely related to helical flutter were calculated to be the vertical bending mode, the in-plane bending mode, and the torsional mode, with natural frequencies of 4.2760 Hz, 5.8838 Hz, and 7.8519 Hz, respectively. The mode shapes are shown in the figure below. Figure 5As shown, the modal separation is reasonable and the distinction between each mode is good, which facilitates the observation and identification of the helical flutter mechanism and is suitable for flutter test research. Further calculations using the PK method for conventional flutter were performed. When the structural damping is 0.02, the critical velocity for conventional flutter is 82.49 m / s, corresponding to a flutter frequency of 4.82 Hz, which is higher than the maximum incoming flow velocity of 60 m / s in the test wind tunnel. This indicates that the conventional flutter boundary is outside the working range of the wind tunnel, meaning that the platform will not experience conventional flutter within the entire test range of the wind tunnel, possessing sufficient flutter safety margin and enabling safe and stable conduct of helical flutter wind tunnel tests.

[0055] Based on the above modal calculation results, with the structural damping also taken as 0.02, the rotor speed was changed and the helical flutter boundary of the test platform was calculated. The results are shown in Table 2.

[0056] Table 2 Calculation results of rotor flutter at different rotor speeds

[0057]

[0058] Table 2 shows that, under the adopted calculation model and structural damping of 0.02, when the rotor speed increases from 400 r / min to 1600 r / min, the calculated critical velocity for helical flutter decreases from 49.15 m / s to 35.18 m / s, and the helical flutter frequency varies within the range of 4.38–4.54 Hz. The critical velocity for helical flutter on this platform exhibits a continuous and smooth variation with the rotor speed, and the flutter frequency fluctuates little, indicating that the test platform can stably reproduce the helical flutter boundary at different rotor speeds. Based on the above calculation results, a key test condition can be set near a rotor speed of approximately 1100 r / min and an incoming flow velocity of approximately 38.0 m / s to verify the platform's ability to test the helical flutter boundary. The platform possesses good parameter adjustability and experimental observation capabilities.

[0059] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A tiltrotor aircraft flutter integrated wind tunnel test platform, comprising a wing module (1), a nacelle module (2), and a rotor system (3); the root of the wing module (1) is fixed to the mounting fixture of the wind tunnel test platform, forming a root fixed support boundary; the nacelle module (2) is installed on the wingtip of the wing module (1) via a connecting joint, and the box section at the wingtip is bolted to the nacelle joint; characterized in that: A bearing housing (22) is installed inside the nacelle shell (21) of the nacelle module (2), and the rotor shaft (31) is supported on the bearing housing (22) via the main shaft bearing (32). The servo (23), servo rocker arm (24) and long control stick (25) are arranged inside the nacelle shell (21). The guide rod (33) and pitch control base (34) are installed at the front end of the bearing housing (22). The guide rod (33) is arranged axially along the rotor shaft (31) to constrain the pitch control base (34) to only make axial translation and not rotate with the bushing (36). The bearing housing (22) is provided with an opening that mates with the guide rod (33). One end of the long control stick (25) is connected to the servo rocker arm (24), and the other end is connected to the pitch control base (34). The rotation plane of the servo rocker arm (24) is parallel to the rotor shaft (31). The servo (23) drives the servo rocker arm (24) to rotate, and the rotation of the servo rocker arm (24) is converted into the axial displacement of the pitch control base (34) along the rotor shaft (31) through the long control stick (25). The pitch control base (34) is connected to the bushing (36) via a deep groove ball bearing (35); the bushing (36) is fitted onto the rotor shaft (31), the rotor shaft (31) has an axially extending slot, the connecting piece passes through the bushing (36) and engages with the slot, so that the bushing (36) rotates with the rotor shaft (31) and can move axially relative to the rotor shaft (31) within the range defined by the slot; the swashplate (37) is fixed to the bushing (36), and the swashplate (37) is connected to the blade clamp (39) via a short control lever (38); The rotor system (3) has its hub assembly installed at the end of the rotor shaft (31), and the rotor blades are installed on the hub assembly via blade clamps (39); the platform operates in a non-powered windmill mode during the helical flutter wind tunnel test.

2. The tiltrotor aircraft flutter integrated wind tunnel test platform according to claim 1, characterized in that: The servo motor (23) receives external control signals and drives the servo motor rocker arm (24) to rotate, causing the pitch control base (34) to move axially along the rotor shaft (31). The axial displacement of the pitch control base (34) is transmitted to the bushing (36) and swashplate (37) via the deep groove ball bearing (35), causing the bushing (36) to move axially within the slotted range of the rotor shaft (31). The swashplate (37) then pushes the blade clip (39) via the short control stick (38), causing the blade clip (39) to drive the rotor blades to rotate around their own pitch axis, thereby realizing continuous adjustment of the rotor collective pitch.

3. The tiltrotor aircraft flutter integrated wind tunnel test platform according to claim 1, characterized in that: Under the condition of no-power wind turbine operation, the rotor collective pitch is adjusted by changing the input signal of the servo motor (23) based on the measured rotor speed, so that the rotor reaches the predetermined speed under different incoming flow velocities and is kept within the allowable deviation range of the predetermined speed.

4. The tiltrotor aircraft flutter integrated wind tunnel test platform according to claim 1, characterized in that: The rotor system (3) can be replaced independently with the hub assembly and rotor blades; the hub assembly adopts a constant velocity universal joint structure, and the flapping of each rotor blade is determined by the rigid body motion of the hub assembly. The rotor disk, as a rigid body, is tilted in any direction around the center of the hub, and the flapping direction is rigid.

5. The tiltrotor aircraft flutter integrated wind tunnel test platform according to claim 4, characterized in that: The rotor hub assembly is either a hinged rotor hub or a rigid rotor hub, in order to study the coupling form and corresponding flutter characteristics of the rotor blade flapping, oscillation, pitch change degree of freedom and the wing and nacelle under different rotor hub constraints. The rotor blades are mounted on the rotor hub assembly via blade clamps (39). The blade clamps rotate around their own axis to achieve pitch variation. The blades are made of different materials, airfoils, spans or twist angles.

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