Hypersonic dynamic jet test device and test method based on forced vibration

Abstract

The present invention belongs to the technical field of hypersonic wind tunnel testing, and discloses a hypersonic dynamic jet test device and test method based on forced vibration. The test device includes a forced vibration device located on the central axis of an aircraft model; it also includes a plurality of nozzles fixed to the aircraft model, the nozzle outlets smoothly transitioning with the surface of the aircraft model, and the rear ends of the nozzles are connected to an external air source through jet pipes located in the inner cavity of the aircraft model. The test method includes determining forced vibration simulation parameters; installing a hypersonic jet test device; testing the hypersonic jet test device; and conducting a hypersonic dynamic jet test based on forced vibration. The test device and test method simulate the aerodynamic characteristics of an aircraft under the simultaneous action of maneuvering flight and jet control through a forced vibration method, and the test results have engineering practical value for improving the maneuverability of the aircraft.

Description

Technical Field

[0001] The present invention belongs to the technical field of hypersonic wind tunnel testing, and in particular relates to a hypersonic dynamic jet testing device and a testing method based on forced vibration. Background Art

[0002] High-precision, fast-response guidance and control technology based on direct force / aerodynamic force composite control is the core technology for improving maneuverability and achieving precision guidance in modern advanced aircraft. Aerodynamic force control mainly changes the flight attitude by changing the deflection angle of aerodynamic control surfaces to generate pitch, yaw, and roll torques. The aerodynamic torque generated by aerodynamic control surfaces is related to the dynamic pressure, which is proportional to the atmospheric density and the square of the flight speed. In high-altitude, low-density, or low-speed flight conditions, the efficiency of the control surfaces is significantly reduced and cannot meet control requirements. Direct force control is mainly achieved by generating a transverse jet from a micro-pulse engine, gas generator, or main engine. On the one hand, it directly provides control force and torque. On the other hand, when used within the atmosphere, it utilizes the additional control force and torque generated by the interference of the transverse jet with the free flow.

[0003] Combined direct force / aerodynamic control involves highly maneuverable flight at high angles of attack, large sideslip, and high rudder deflection. This involves not only the mutual interference of the lateral jet and the external flow field, but also the nonlinear coupling between aerodynamics and motion, and the mutual influence between motion and the lateral jet. Rapid maneuvers inevitably cause dramatic changes in the surrounding airflow, triggering lateral jet interference caused by unsteady turbulence under dynamic conditions.

[0004] While much research has previously focused on static models, dynamic aerodynamic characteristics are now receiving increasing attention. Rapid, high-angle-of-attack maneuvers, in particular, inevitably cause dramatic changes in the airflow around the vehicle, which in turn triggers the phenomenon of lateral jet interference, a phenomenon known as unsteady turbulence under dynamic conditions. During forced pitch oscillations in interceptor missile models, significant unsteady hysteresis can occur, with the jet interference amplification factor varying with the pitch angle cycle and deviating from its steady-state value.

[0005] Currently, there is an urgent need to develop a hypersonic dynamic jet test device and test method based on forced vibration. Summary of the Invention

[0006] One technical problem to be solved by the present invention is to provide a hypersonic dynamic jet test device based on forced vibration, and another technical problem to be solved by the present invention is to provide a hypersonic dynamic jet test method based on forced vibration.

[0007] The hypersonic dynamic jet test device and test method based on forced vibration of the present invention simulate the aerodynamic characteristics of an aircraft under the simultaneous effects of maneuvering flight and jet control by means of a forced vibration method.

[0008] The hypersonic dynamic jet test device based on forced vibration of the present invention comprises a forced vibration device located on the central axis of an aircraft model; and further comprises a plurality of nozzles fixed to the aircraft model, wherein the nozzle outlets smoothly transition with the surface of the aircraft model, and the rear ends of the nozzles are connected to an external air source through jet pipes located in the inner cavity of the aircraft model.

[0009] The forced vibration device includes a front cone section, a balance, a yaw frame, and a tail strut, which are sequentially connected from front to back. A cross elastic hinge and a yaw shaft are sequentially fixed in the yaw frame from front to back. A spherical bearing is mounted on the yaw shaft, and the yaw shaft and the spherical bearing are fixed in a cantilever groove. An internal transmission shaft and a rear chamber are provided in the tail strut, with the front end of the internal transmission shaft connected to the yaw shaft, and the rear end of the internal transmission shaft connected to the drive shaft of the drive motor fixed in the rear chamber.

[0010] The inner cavity of the aircraft model is provided with a conical model bushing matching the front cone section of the forced vibration device, a plurality of counterweights are fixed on the inner wall of the aircraft model, and the center of mass of the aircraft model coincides with the center of the cross elastic hinge.

[0011] Furthermore, the balance adopts a four-column beam measuring element for measuring the normal force, lateral force, pitch moment, yaw moment and roll moment of the aircraft model.

[0012] Furthermore, the design and production requirements of the aircraft model are that, while ensuring the strength and rigidity of the aircraft model, the mass and moment of inertia of the aircraft model are gradually reduced, so as to reduce the inertia force and inertia moment of the aircraft model and improve the natural frequency and measurement accuracy of the hypersonic dynamic jet test device.

[0013] The hypersonic dynamic jet test method based on forced vibration of the present invention comprises the following steps:

[0014] S10. Determine forced vibration simulation parameters;

[0015] The simulation parameters are the aerodynamic pitch damping torque of the aircraft model; the aerodynamic pitch damping torque is calculated based on the applied torque amplitude, the maximum pitch angular displacement, the phase angle between the applied torque and the vibration angular displacement, the angular velocity of the maximum pitch motion, and the mechanical damping of the hypersonic jet test device.

[0016] S20. Install the hypersonic jet test device;

[0017] First, the tail strut is fixed to the middle bracket of the hypersonic wind tunnel. Second, the drive motor is installed in the rear chamber of the tail strut. Third, the drive shaft of the drive motor is connected to the internal transmission shaft. Finally, the aircraft model is fixed to the nose cone from front to back.

[0018] S30. Test the hypersonic jet test device;

[0019] The hypersonic jet test device is started. The drive shaft of the drive motor sequentially drives the internal transmission shaft and the yaw shaft. The yaw shaft drives the joint bearing to exhibit cam motion, converting the rotational motion of the drive shaft of the drive motor into the pitch motion of the balance. Under the constraints of the yaw frame, the aircraft model is made to pitch around the center of the cross elastic hinge.

[0020] During the test, if it is found that the pitching motion of the aircraft model around the center of the cross elastic hinge has deviation, the hypersonic jet test device is removed, the model bushing and counterweight are adjusted, and the test is repeated until the aircraft model can achieve pitching motion around the center of the cross elastic hinge;

[0021] S40. Conduct hypersonic dynamic jet flow tests based on forced vibration;

[0022] First, the hypersonic jet test device was activated to achieve pitch motion of the aircraft model around the center of the cross elastic hinge;

[0023] Secondly, open the gas source, and let the jet gas that meets the jet requirements enter the jet pipe through the jet inlet, and then spray it out through the nozzle outlet;

[0024] Finally, the hypersonic wind tunnel is started. After the hypersonic flow field stabilizes, the displacement sensor measures the vibration angular displacement of the aircraft model, and the balance measures the aerodynamic force and torque of the aircraft model.

[0025] S50. Calculate aerodynamic pitch damping moment;

[0026] Calculate the aerodynamic pitch damping moment and evaluate the simulation results.

[0027] Furthermore, the calculation process of the aerodynamic pitch damping moment of S10 is as follows:

[0028] The aircraft model performs simple harmonic vibration with fixed frequency and amplitude in the pitch direction. The differential equation of motion of simple harmonic vibration is:

[0029] ;

[0030] in, is the inertia moment of the aircraft model during pitch motion, 、 and are the angular displacement, angular velocity and angular acceleration of pitch motion respectively, is the mechanical damping torque of the hypersonic jet test device, is the mechanical damping of the hypersonic jet test device, is the aerodynamic pitch damping moment of the hypersonic jet test device, is the aerodynamic pitch damping moment derivative of the hypersonic jet test device, is the elastic recovery torque of the hypersonic jet test device, is the aerodynamic pitch restoring moment of the hypersonic jet test device, is the elastic hinge constant of the hypersonic jet test device, is the aerodynamic pitch recovery damping torque of the hypersonic jet test device, is the imaginary unit, is the angular velocity of the maximum pitch motion, For time, is the external forced torque;

[0031] Equation (1) is a linear, second-order non-homogeneous differential equation with constant coefficients. The solution of equation (1) consists of two parts: one is the general solution of the homogeneous form corresponding to equation (1), and the other is the special solution of the non-homogeneous equation, namely:

[0032] ;

[0033] Where, is the maximum pitch angle displacement, is the phase angle between the applied torque and the vibration angular displacement;

[0034] Substituting equation (2) into equation (1) yields:

[0035] ;

[0036] in, is the amplitude of the applied external torque;

[0037] Since the imaginary parts at both ends of equation (3) are equal, the aerodynamic pitch damping moment derivative of the aircraft model is obtained as:

[0038] ;

[0039] From equation (4), we can see that by measuring the amplitude of the applied torque , maximum pitch angle displacement , the phase angle between the applied torque and the vibration angular displacement , angular velocity of maximum pitch motion and mechanical damping of hypersonic jet test devices , the aerodynamic pitch damping moment derivative can be calculated .

[0040] The present invention's forced-vibration hypersonic dynamic jet test apparatus and test method, based on research on the static, steady lateral jet interference characteristics of aircraft, investigates the impact of unsteady turbulent lateral jet interference characteristics on aircraft maneuvering flight. This study captures the unsteady coupling characteristics of aerodynamics, motion, and lateral jets during direct force / aerodynamic composite control. The test results have practical engineering value for improving the accuracy of direct force / aerodynamic composite control modeling, effectively predicting nonlinear aerodynamic / motion modeling characteristics, and rationally designing flight control laws to achieve effective allocation of control surfaces and lateral jets, thereby enhancing aircraft maneuverability. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] Figure 1 This is a schematic diagram of the installation of a hypersonic dynamic jet test device based on forced vibration according to the present invention;

[0042] Figure 2 Schematic diagram of the structure of the forced vibration device in the hypersonic dynamic jet test device based on forced vibration of the present invention;

[0043] Figure 3 This is a schematic diagram of the aircraft model and the test device adjustment principle in the hypersonic dynamic jet test device based on forced vibration of the present invention.

[0044] In the figure: 1. Forced vibration device; 2. Aircraft model; 3. Nozzle; 4. Jet pipe;

[0045] 101. Balance; 102. Yaw frame; 103. Cross elastic hinge; 104. Yaw axis; 105. Internal transmission shaft; 106. Tail support rod; 107. Front cone section; 108. Rear chamber;

[0046] 201. Model bushing; 202. Model center of mass; 203. Counterweight. DETAILED DESCRIPTION

[0047] The present invention will be described in detail below with reference to the accompanying drawings and embodiments.

[0048] Example: Figures 1 to 3 As shown, the hypersonic dynamic jet test device based on forced vibration of this embodiment includes a forced vibration device 1 located on the central axis of an aircraft model 2; and further includes a plurality of nozzles 3 fixed to the aircraft model 2. The outlets of the nozzles 3 have a smooth transition with the surface of the aircraft model 2. The rear ends of the nozzles 3 are connected to an external air source through jet pipes 4 located in the inner cavity of the aircraft model 2.

[0049] The forced vibration device 1 includes a front cone section 107, a balance 101, a yaw frame 102, and a tail support rod 106, which are sequentially connected from front to back. A cross elastic hinge 103 and a yaw shaft 104 are fixed in sequence from front to back within the yaw frame 102. A spherical bearing is mounted on the yaw shaft 104, and the yaw shaft 104 and the spherical bearing are fixed in a cantilever groove. An internal transmission shaft 105 and a rear chamber 108 are provided within the tail support rod 106. The front end of the internal transmission shaft 105 is connected to the yaw shaft 104, and the rear end of the internal transmission shaft 105 is connected to the drive shaft of the drive motor fixed in the rear chamber 108.

[0050] The inner cavity of the aircraft model 2 is provided with a conical model bushing 201 that matches the front cone section 107 of the forced vibration device 1. Several counterweight blocks 203 are fixed to the inner wall of the aircraft model 2. The model center of mass 202 of the aircraft model 2 coincides with the center of the cross elastic hinge 103.

[0051] Furthermore, the balance 101 adopts a four-column beam measuring element for measuring the normal force, lateral force, pitch moment, yaw moment and roll moment of the aircraft model 2 .

[0052] Furthermore, the design and production requirements of the aircraft model 2 are that, under the premise of ensuring the strength and rigidity of the aircraft model 2, the mass and moment of inertia of the aircraft model 2 are gradually reduced, so as to reduce the inertia force and moment of inertia of the aircraft model 2 and improve the natural frequency and measurement accuracy of the hypersonic dynamic jet test device.

[0053] The hypersonic dynamic jet test method based on forced vibration of this embodiment includes the following steps:

[0054] S10. Determine forced vibration simulation parameters;

[0055] The simulation parameters are the aerodynamic pitch damping torque of the aircraft model 2; the aerodynamic pitch damping torque is calculated by the applied torque amplitude, the maximum pitch angular displacement, the phase angle between the applied torque and the vibration angular displacement, the angular velocity of the maximum pitch motion, and the mechanical damping of the hypersonic jet test device;

[0056] S20. Install the hypersonic jet test device;

[0057] First, the tail strut 106 is fixed to the middle bracket of the hypersonic wind tunnel; second, the drive motor is installed in the rear chamber 108 of the tail strut 106; third, the drive shaft of the drive motor is connected to the internal transmission shaft 105; finally, the aircraft model 2 is fixed to the front cone section 107 from front to back;

[0058] S30. Test the hypersonic jet test device;

[0059] The hypersonic jet test device is started. The drive shaft of the drive motor sequentially drives the internal transmission shaft 105 and the yaw shaft 104. The yaw shaft 104 drives the joint bearing to exhibit cam motion, converting the rotational motion of the drive shaft of the drive motor into the pitch motion of the balance 101. Under the constraints of the yaw frame 102, the aircraft model 2 is made to pitch around the center of the cross elastic hinge 103.

[0060] During the test, if it is found that the pitching motion of the aircraft model 2 around the center of the cross elastic hinge 103 has deviation, the hypersonic jet test device is removed, the model bushing 201 and the counterweight 203 are adjusted, and the test is repeated until the aircraft model 2 can achieve pitching motion around the center of the cross elastic hinge 103;

[0061] S40. Conduct hypersonic dynamic jet flow tests based on forced vibration;

[0062] First, the hypersonic jet test device is started to achieve a pitching motion of the aircraft model 2 around the center of the cross elastic hinge 103;

[0063] Secondly, open the gas source, and let the jet gas that meets the jet requirements enter the jet pipe 4 through the jet inlet, and then spray out through the nozzle 3 outlet;

[0064] Finally, the hypersonic wind tunnel is started. After the hypersonic flow field stabilizes, the displacement sensor measures the vibration angular displacement of the aircraft model 2, and the balance 101 measures the aerodynamic force and torque of the aircraft model 2.

[0065] S50. Calculate aerodynamic pitch damping moment;

[0066] Calculate the aerodynamic pitch damping moment and evaluate the simulation results.

[0067] Furthermore, the calculation process of the aerodynamic pitch damping moment of S10 is as follows:

[0068] Aircraft model 2 performs simple harmonic vibration with fixed frequency and amplitude in the pitch direction. The motion differential equation of simple harmonic vibration is:

[0069] ;

[0070] in, is the inertia moment of aircraft model 2 during pitch motion, 、 and are the angular displacement, angular velocity and angular acceleration of pitch motion respectively, is the mechanical damping torque of the hypersonic jet test device, is the mechanical damping of the hypersonic jet test device, is the aerodynamic pitch damping moment of the hypersonic jet test device, is the aerodynamic pitch damping moment derivative of the hypersonic jet test device, is the elastic recovery torque of the hypersonic jet test device, is the aerodynamic pitch restoring moment of the hypersonic jet test device, is the elastic hinge constant of the hypersonic jet test device, is the aerodynamic pitch recovery damping torque of the hypersonic jet test device, is the imaginary unit, is the angular velocity of the maximum pitch motion, For time, is the external forced torque;

[0071] Equation (1) is a linear, second-order non-homogeneous differential equation with constant coefficients. The solution of equation (1) consists of two parts: one is the general solution of the homogeneous form corresponding to equation (1), and the other is the special solution of the non-homogeneous equation, namely:

[0072] ;

[0073] Where, is the maximum pitch angle displacement, is the phase angle between the applied torque and the vibration angular displacement;

[0074] Substituting equation (2) into equation (1) yields:

[0075] ;

[0076] in, is the amplitude of the applied external torque;

[0077] Since the imaginary parts at both ends of equation (3) are equal, the aerodynamic pitch damping moment derivative of aircraft model 2 is obtained as:

[0078] ;

[0079] From equation (4), we can see that by measuring the amplitude of the applied torque , maximum pitch angle displacement , the phase angle between the applied torque and the vibration angular displacement , angular velocity of maximum pitch motion and mechanical damping of hypersonic jet test devices , the aerodynamic pitch damping moment derivative can be calculated .

[0080] Although the embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the description and embodiments, and can be applied to various fields suitable for the present invention. For those skilled in the art, further improvements and modifications can be easily realized without departing from the principles of the present invention. Therefore, the present invention is not limited to the specific details and illustrations shown and described herein without departing from the general concept defined by the claims and their equivalents.

Claims

1. A hypersonic dynamic jet test device based on forced vibration, characterized in that: The hypersonic dynamic jet test device comprises a forced vibration device (1) located on the central axis of an aircraft model (2); and further comprises a plurality of nozzles (3) fixed on the aircraft model (2), wherein the outlet of the nozzle (3) smoothly transitions with the surface of the aircraft model (2), and the rear end of the nozzle (3) is connected to an external air source through a jet pipe (4) located in the inner cavity of the aircraft model (2); The forced vibration device (1) comprises a front cone section (107), a balance (101), a yaw frame (102) and a tail support rod (106) connected in sequence from front to back; a cross elastic hinge (103) and a yaw shaft (104) are fixed in sequence from front to back in the yaw frame (102); a joint bearing is mounted on the yaw shaft (104); the yaw shaft (104) and the joint bearing are fixed in a cantilever groove; an internal transmission shaft (105) and a rear chamber (108) are arranged in the tail support rod (106); the front end of the internal transmission shaft (105) is connected to the yaw shaft (104), and the rear end of the internal transmission shaft (105) is connected to the drive shaft of a drive motor fixed in the rear chamber (108); The inner cavity of the aircraft model (2) is provided with a conical model bushing (201) that matches the front cone section (107) of the forced vibration device (1), a plurality of counterweight blocks (203) are fixed to the inner wall of the aircraft model (2), and the model center of mass (202) of the aircraft model (2) coincides with the center of the cross elastic hinge (103).

2. The hypersonic dynamic jet test device based on forced vibration according to claim 1, characterized in that: The balance (101) uses a four-column beam measuring element for measuring the normal force, lateral force, pitch moment, yaw moment and roll moment of the aircraft model (2).

3. The hypersonic dynamic jet test device based on forced vibration according to claim 1, characterized in that: The design and production requirements of the aircraft model (2) are that, under the premise of ensuring the strength and rigidity of the aircraft model (2), the mass and moment of inertia of the aircraft model (2) are gradually reduced, so as to reduce the inertial force and moment of inertia of the aircraft model (2) and improve the natural frequency and measurement accuracy of the hypersonic dynamic jet test device.

4. A hypersonic dynamic jet test method based on forced vibration, which is used in the hypersonic dynamic jet test device based on forced vibration according to any one of claims 1 to 3, characterized in that: The test device installation method comprises the following steps: S10. Determine forced vibration simulation parameters; The simulation parameters are the aerodynamic pitch damping moment of the aircraft model (2); the aerodynamic pitch damping moment is calculated by the amplitude of the applied external moment, the maximum pitch angular displacement, the phase angle between the applied moment and the vibration angular displacement, the angular velocity of the maximum pitch motion and the mechanical damping of the hypersonic jet test device; S20. Install the hypersonic jet test device; First, the tail strut (106) is fixed to the middle bracket of the hypersonic wind tunnel; second, the drive motor is installed in the rear chamber (108) of the tail strut (106); third, the drive shaft of the drive motor is connected to the internal transmission shaft (105); finally, the aircraft model (2) is fixed to the front cone section (107) from front to back; S30. Test the hypersonic jet test device; The hypersonic jet test device is started, and the driving shaft of the driving motor sequentially drives the internal transmission shaft (105) and the yaw shaft (104), and the yaw shaft (104) drives the joint bearing to present a cam motion, converting the rotational motion of the driving shaft of the driving motor into the pitch motion of the balance (101), and under the constraint of the yaw frame (102), the aircraft model (2) is made to perform a pitch motion around the center of the cross elastic hinge (103); During the test, if it is found that the pitching motion of the aircraft model (2) around the center of the cross elastic hinge (103) has deviation, the hypersonic jet test device is removed, the model bushing (201) and the counterweight (203) are adjusted, and the test is repeated until the pitching motion of the aircraft model (2) around the center of the cross elastic hinge (103) is achieved; S40. Conduct hypersonic dynamic jet flow tests based on forced vibration; First, the hypersonic jet test device is started to achieve a pitching motion of the aircraft model (2) around the center of the cross elastic hinge (103); Secondly, the gas source is turned on, and the jet gas that meets the jet requirements is introduced into the jet pipe (4) through the jet air inlet, and then ejected through the nozzle (3) outlet; Finally, the hypersonic wind tunnel is started, and after the hypersonic flow field is stabilized, the displacement sensor measures the vibration angular displacement of the aircraft model (2), and the balance (101) measures the aerodynamic force and torque of the aircraft model (2); S50. Calculate aerodynamic pitch damping moment; Calculate the aerodynamic pitch damping moment and evaluate the simulation results.

5. The hypersonic dynamic jet test method based on forced vibration according to claim 4, characterized in that: The calculation process of the aerodynamic pitch damping moment of S10 is as follows: The aircraft model (2) performs simple harmonic vibration with fixed frequency and amplitude in the pitch direction. The differential equation of motion of the simple harmonic vibration is: ； in, is the moment of inertia of the aircraft model (2) during pitch motion, 、 and are the angular displacement, angular velocity and angular acceleration of pitch motion respectively, is the mechanical damping torque of the hypersonic jet test device, is the mechanical damping of the hypersonic jet test device, is the aerodynamic pitch damping moment of the hypersonic jet test device, is the aerodynamic pitch damping moment derivative of the hypersonic jet test device, is the elastic recovery torque of the hypersonic jet test device, is the aerodynamic pitch restoring moment of the hypersonic jet test device, is the elastic hinge constant of the hypersonic jet test device, is the aerodynamic pitch recovery damping torque of the hypersonic jet test device, is the imaginary unit, is the angular velocity of the maximum pitch motion, For time, is the external forced torque; Equation (1) is a linear, second-order non-homogeneous differential equation with constant coefficients. The solution of equation (1) consists of two parts: one is the general solution of the homogeneous form corresponding to equation (1), and the other is the special solution of the non-homogeneous equation, namely: ； Where, is the maximum pitch angle displacement, is the phase angle between the applied torque and the vibration angular displacement; Substituting equation (2) into equation (1) yields: ； in, is the amplitude of the applied external torque; Since the imaginary parts at both ends of equation (3) are equal, the aerodynamic pitch damping moment derivative of the aircraft model (2) is obtained as: ； From equation (4), we can see that by measuring the amplitude of the applied torque , maximum pitch angle displacement , the phase angle between the applied torque and the vibration angular displacement , angular velocity of maximum pitch motion and mechanical damping of hypersonic jet test devices , the aerodynamic pitch damping moment derivative can be calculated .

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