A transient high-speed wind tunnel low supersonic flow field test method and nozzle device

Abstract

A transient high-speed wind tunnel low supersonic flow field test method and a nozzle device thereof belong to the field of wind tunnel test technology; the existing technology cannot realize an ideal low supersonic flow field and meet test requirements; the method uses newly designed M1.3, M1.35 and M1.4 nozzle devices combined with an open-hole wall test section to establish a low supersonic flow field in a transient wind tunnel; the Mach number control method is used to obtain an ideal flow field control quality; through the axial Mach number distribution analysis of the test section at different expansion angles, it is determined that the smaller the expansion angle of the test section at low supersonic speed, the better the uniformity of the flow field, and the expansion angle of 0° is selected as the tunnel body condition; the nozzle device is combined with the open-hole wall test section, and the test section expansion angle and flow field control method are adopted to obtain a qualified flow field that meets the national military standard requirements; the ideal low supersonic flow field is realized and the test requirements are met; the device provides a hardware foundation for realizing an ideal low supersonic flow field and meeting test requirements by adopting the M1.3, M1.35 and M1.4 nozzle aerodynamic surfaces.

Description

Technical Field

[0001] The invention belongs to the technical field of wind tunnel tests, and in particular relates to a temporary high-speed wind tunnel low supersonic flow field test method and a nozzle device thereof. Background Art

[0002] Temporary impulse wind tunnels generally do not have the ability to conduct low supersonic tests. The FL-1 wind tunnel has conducted research and debugging on the realization of low supersonic flow fields using the M1.35 nozzle and the hole wall test section, but the quality of the low supersonic flow field it has established does not meet the test requirements. In general, there has not been a relatively systematic experimental research on temporary impulse low supersonic speed, and the low supersonic flow field established is not very ideal. The flow field M number distribution calibration results do not meet the index requirements of the national military standard σM≤0.011. Therefore, whether the ideal low supersonic flow field can be achieved using the combination of the nozzle and the hole wall test section, and whether the quality of the low supersonic flow field established can meet the test requirements are still to be solved. Summary of the invention

[0003] The present invention overcomes the above-mentioned deficiencies of the prior art and provides a temporary high-speed wind tunnel low supersonic flow field test method and a nozzle device thereof.

[0004] The technical solution of the present invention:

[0005] Technical Solution 1

[0006] A temporary high-speed wind tunnel low supersonic flow field test method comprises the following steps:

[0007] Step a, using the newly designed M1.3, M1.35 and M1.4 nozzle devices and combined with the perforated wall test section to establish a low supersonic flow field in a transient wind tunnel;

[0008] Step b, using the Mach number control method to obtain the ideal flow field control quality;

[0009] Step c, by analyzing the axial Mach number distribution of the test section at different expansion angles, it is determined that the smaller the expansion angle of the test section is at low supersonic speed, the better the uniformity of the flow field is, and the expansion angle of 0° is selected as the cave body condition;

[0010] Step d, using the nozzle device in combination with the open hole wall test section, adopting the test section expansion angle and flow field control method, to obtain a qualified flow field that meets the national military standard requirements, the flow field uniformity of M1.3, M1.35 and M1.4 are all better than the index requirement of the national military standard σM≤0.011, and the standard model test and the aircraft model test verification show that the quality of the established low supersonic flow field fully meets the test requirements, and the wind tunnel has the ability to conduct low supersonic tests.

[0011] Furthermore, the specific method of step a comprises the following steps:

[0012] Step a1, using a nozzle profile design program, by inputting nozzle parameters, the nozzle parameters including nozzle type, nozzle inlet half height, nozzle outlet half height, outlet Mach number and inlet total pressure, to obtain the nozzle aerodynamic profile and the nozzle profile spiral elevator extension;

[0013] Step a2, performing coefficient correction on the boundary layer thickness of the wind tunnel test section;

[0014] Step a3, calibrating the profile to meet the mechanical condition restriction requirements, and obtaining M1.3, M1.35 and M1.4 nozzle aerodynamic profiles;

[0015] Step a4, according to the calibrated M1.3, 1.35, 1.4 nozzle mechanical profiles, the control system drives the electric screw lifter of the nozzle profile flexible plate to operate synchronously, and after the bending shape of the controlled flexible plate matches the nozzle theoretical aerodynamic profile, the sliding end is axially locked;

[0016] Step a5: Use the perforated wall test section and establish a low supersonic flow field using step d.

[0017] Furthermore, the Mach number control method of step b comprises the following steps:

[0018] Step b1, during the test, the flow field M number is calculated using the total pressure of the stable section and the static pressure of the test section;

[0019] Step b2, calculating the difference between the flow field M number and the given M number;

[0020] Step b3: Use a PI controller to adjust the opening of the pressure regulating valve according to the difference.

[0021] Furthermore, the specific method of step c comprises the following steps:

[0022] Step c1, setting the test section expansion angle to 0° to perform flow field calibration, and obtaining the test section axial Mach number distribution data;

[0023] Step c2, setting the test section expansion angle to +0.3° to perform flow field calibration and obtain the test section axial Mach number distribution data;

[0024] Step c3, setting the test section expansion angle to +0.5° to perform flow field calibration and obtain the test section axial Mach number distribution data;

[0025] Step c4, analyzing the axial Mach number distribution data of the test section to determine that different expansion angles have an impact on the uniformity of the flow field in the test section. When the wind tunnel operation mode is the mainstream injection Mach number control mode, the smaller the expansion angle of the test section, the better the uniformity of the flow field; when the expansion angle is 0°, the flow field of the test section is better than the flow field of the test section when the expansion angle is +0.3° and +0.5°.

[0026] Technical Solution 2

[0027] A nozzle device designed in a test method for a temporary impulse high-speed wind tunnel low supersonic flow field based on technical solution 1, comprising a nozzle flexible plate, a support, a frame, a front end plate, a sliding end, a driver, a rear end plate and a main shell; the front end plate, the rear end plate and the main shell constitute a hollow quadrilateral shell, a plurality of frames are arranged at the upper end of the quadrilateral shell, a plurality of supports are arranged at the lower end of the quadrilateral shell, the aerodynamic profile of the nozzle flexible plate is a curved nozzle aerodynamic profile, the left and right ends of the sliding end are both fixed between the front end plate and the rear end plate, the upper surface aerodynamic profile and the lower surface aerodynamic profile of the nozzle flexible plate are both provided with a plurality of drivers, and the plurality of drivers can change the bending shape of the nozzle flexible plate aerodynamic profile, and the bending shapes include curved nozzle aerodynamic profiles M1.3, M1.35 and M1.4.

[0028] Furthermore, the number of the supports and the number of the racks are both greater than four.

[0029] Furthermore, the frame is an I-shaped bracket, including an upper frame plate, a lower bracket plate and a support rod, and the upper frame plate and the lower bracket plate have the same length and width.

[0030] Furthermore, the support includes an upper sliding hollow tube and a lower sliding I-shaped hollow tube; the upper sliding hollow tube and the lower sliding I-shaped hollow tube are both provided with matching holes, and the upper sliding hollow tube and the lower sliding I-shaped hollow tube are tightened and fixed by bolts and nuts.

[0031] Furthermore, the driver includes a spiral lifting rod and a motor; the spiral lifting rod is connected to the motor, and the spiral lifting rod is driven to rise or fall by the rotation of the motor, thereby changing the bending shape of the nozzle flexible plate aerodynamic surface.

[0032] Furthermore, the sliding end is located at the entrance of the nozzle flexible plate, rotates around the horizontal axis installed on the front end plate, and cooperates with the T-slot on the back of the nozzle flexible plate through two T-guide rails to make it slide inside to adapt to the changes in the length of the flexible plate and the slope of the entrance profile when the aerodynamic profile of the nozzle flexible plate changes; an axial locking device is installed inside the sliding end, and the axial force transmitted from the nozzle flexible plate is borne by two screws; the nut slides back and forth on the screw and is positioned by a switch to ensure the profile stability of the nozzle flexible plate.

[0033] Compared with the prior art, the present invention has the following beneficial effects:

[0034] The present invention provides a method for testing low supersonic flow field in a transient high-speed wind tunnel and a nozzle device thereof. The method enables the transient high-speed wind tunnel to be capable of conducting low supersonic tests at M=1.3, 1.35 and M1.4, so as to meet the urgent demand for low supersonic aerodynamic wind tunnel tests in current aircraft development and the demand for refined testing and analysis of transsupersonic aerodynamic characteristics of future advanced aircraft, and provide a good research platform and means for solving the complex aerodynamic characteristics research of aircraft flying in the low supersonic range;

[0035] This device adopts the nozzle aerodynamic surfaces of M1.3, M1.35 and M1.4 to provide the hardware foundation for realizing the ideal low supersonic flow field in the combination of nozzle and open wall test section and ensuring that the quality of the established low supersonic flow field meets the test requirements. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 It is a diagram of the nozzle structure device;

[0037] Figure 2 It is the nozzle aerodynamic profile curve;

[0038] Figure 3 It is the flow field Mach number control principle diagram;

[0039] Figure 4 It is the flow field Mach number homogeneity index curve diagram;

[0040] Figure 5 It is the distribution structure diagram of the model area of ​​M=1.3, 1.35, and 1.4;

[0041] Figure 6 It is the flow field calibration and schlieren diagram;

[0042] Figure 7 These are the test images of the YF-16 standard model and the AGARD-B standard model;

[0043] Figure 8 This is a comparison chart of the aerodynamic derivative curves of the YF-16 standard model.

[0044] In the figure: 1 nozzle flexible plate, 2 support, 3 frame, 4 front end plate, 5 sliding end, 6 driver, 7 rear end plate, 8 main shell. DETAILED DESCRIPTION

[0045] The present invention will be described in detail below with reference to the accompanying drawings. Specific implementation method 1

[0047] A temporary high-speed wind tunnel low supersonic flow field test method comprises the following steps:

[0048] Step a, using the newly designed M1.3, M1.35 and M1.4 nozzle devices and combined with the perforated wall test section to establish a low supersonic flow field in a transient wind tunnel;

[0049] Step b, using the Mach number control method to obtain the ideal flow field control quality;

[0050] Step c, by analyzing the axial Mach number distribution of the test section at different expansion angles, it is determined that the smaller the expansion angle of the test section is at low supersonic speed, the better the uniformity of the flow field is, and the expansion angle of 0° is selected as the cave body condition;

[0051] Step d, using the nozzle device in combination with the open hole wall test section, adopting the test section expansion angle and flow field control method, to obtain a qualified flow field that meets the national military standard requirements, the flow field uniformity of M1.3, M1.35 and M1.4 are all better than the index requirement of the national military standard σM≤0.011, and the standard model test and the aircraft model test verification show that the quality of the established low supersonic flow field fully meets the test requirements, and the wind tunnel has the ability to conduct low supersonic tests.

[0052] Specifically, the specific method of step a comprises the following steps:

[0053] Step a1, using a nozzle profile design program, by inputting nozzle parameters, the nozzle parameters including nozzle type, nozzle inlet half height, nozzle outlet half height, outlet Mach number and inlet total pressure, to obtain the nozzle aerodynamic profile and the nozzle profile spiral elevator extension;

[0054] Step a2, performing coefficient correction on the boundary layer thickness of the wind tunnel test section;

[0055] Step a3: calibrate the profile to meet the mechanical condition restriction requirements, and obtain the M1.3, M1.35 and M1.4 nozzle aerodynamic profiles, such as Figure 2 As shown;

[0056] Step a4: Based on the calibrated M1.3, 1.35, and 1.4 nozzle mechanical profiles, the control system drives the electric screw lifter of the nozzle profile flexible plate to operate synchronously. After the bending shape of the controlled flexible plate matches the nozzle theoretical pneumatic profile, the sliding end is axially locked, such as Figure 1 As shown;

[0057] Step a5: Use the perforated wall test section and establish a low supersonic flow field using step d.

[0058] Specifically, the transient high-speed wind tunnel uses two methods to control the flow field Mach number: total pressure control or Mach number control.

[0059] The total pressure control adopts an incremental PID algorithm with output limiting to complete the feedback control of the total pressure.

[0060] Assuming the control error at time k is e(k), the control output is formula (1):

[0061] u(k)=u(k-1)+K p [e(k)-e(k-1)]+K i e(k) (1)

[0062] This formula is the incremental PID algorithm, where K p is the proportionality coefficient, K i is the integral coefficient. Since the integral output is a cumulative amount, in order to prevent integral saturation and ensure the safe operation of the equipment, u(k) is output limited, taking the linear area of ​​the main pressure regulating valve and limiting it to 3% to 70% of the full displacement.

[0063] In addition, during the blowing process, due to the continuous decrease in the pressure of the spherical tank, in order to compensate for the total pressure change caused by the decrease in the pressure of the spherical tank, static feedforward compensation is added on the basis of feedback control, and the compensation coefficient changes with different given values.

[0064] The control output at time k after considering feedback and feedforward is formula (2):

[0065] u(k)=u(k-1)+K p [e(k)-e(k-1)]+K i e(k)+K co Tc (2)

[0066] Where K co is the compensation coefficient, and Tc is the control period.

[0067] The total pressure is controlled by using the PID closed-loop control of the pressure regulating valve to make the total pressure in the stable section of the wind tunnel meet △p0max / p0≤0.3%, ensuring the stability of the front chamber pressure.

[0068] Specifically, the Mach number control method in step b is as follows: Figure 3 As shown, repeat the following steps to make the number of test sections M meet the control requirements:

[0069] Step b1, during the test, the flow field M number is calculated using the total pressure of the stable section and the static pressure of the test section;

[0070] Step b2, calculating the difference between the flow field M number and the given M number;

[0071] Step b3: Use a PI controller to adjust the opening of the pressure regulating valve according to the difference.

[0072] The low supersonic flow field control can adopt the total pressure control mode and the Mach number control mode. The experimental research of the present invention shows that under the same cave conditions, the flow field M number distribution is uneven under the total pressure control mode. Under the Mach number control, the flow field M number distribution uniformity reaches the qualified index required by the national military, where M=1.4 The flow field M number distribution uniformity index meets the advanced index requirements of the national military standard, such as Figure 4 shown.

[0073] Specifically, the flow field calibration method of step c comprises the following steps:

[0074] Step c1, setting the test section expansion angle to 0° to perform flow field calibration, and obtaining the test section axial Mach number distribution data;

[0075] Step c2, setting the test section expansion angle to +0.3° to perform flow field calibration and obtain the test section axial Mach number distribution data;

[0076] Step c3, setting the test section expansion angle to +0.5° to perform flow field calibration and obtain the test section axial Mach number distribution data;

[0077] Step c4, analyzing the axial Mach number distribution data of the test section to determine that different expansion angles have an impact on the uniformity of the flow field in the test section. When the wind tunnel operation mode is the mainstream injection Mach number control mode, the smaller the expansion angle of the test section, the better the uniformity of the flow field; when the expansion angle is 0°, the flow field of the test section is better than the flow field of the test section when the expansion angle is +0.3° and +0.5°.

[0078] Specifically, the specific steps of step d are as follows:

[0079] Step d1, designing the aerodynamic profiles of the M1.3, M1.35 and M1.4 nozzles according to the method described in step a, and calibrating the mechanical profiles of the nozzles;

[0080] Step d2, based on the research conclusions described in step b, combined with the perforated wall test section, the flow field control method adopts the mainstream ejection Mach number control method;

[0081] Step d3: Based on the conclusion in step c, the flow field calibration study is carried out with the expansion angle of the test section preferably set to 0°.

[0082] When the above method is adopted, the flow field uniformity of M1.3, M1.35 and M1.4 is better than the index requirement of the national military standard σM≤0.011, among which the flow field uniformity of M1.4 reaches the advanced index requirement of the national military standard.

[0083] Based on the flow field calibration research, the YF-16 standard model test and the AGARD-B standard model test were carried out. Figure 5 , 6 , 7, Figure 5Middle: The total length of the test section is 3800mm, the length of the model test area is 1440mm, and the model must be placed in the test area during the experiment. The length of the support interference area is 460mm, that is, the distance between the tail of the model and the support is not less than 460mm. The diameter of the rotating window on the side wall of the model area is 500mm. Figure 6 Middle: The left is a front view of the total pressure rake installed on the wind tunnel support pole, the middle is the test Schlieren diagram when the total pressure rake is placed horizontally, and the right is the test Schlieren diagram when the total pressure rake is placed vertically. Figure 7 Middle: Left is the front view of the YF-16 standard model installed on the wind tunnel strut, right is the front view of the AGARD-B standard model installed on the wind tunnel strut. The key aerodynamic derivatives of the YF-16 standard model and the AGARD-B standard model tests are reasonable, such as Figure 8 As shown in Table 1, it shows that the temporary wind tunnel is capable of undertaking M1.3, M1.35 and M1.4 low supersonic tests.

[0084] This implementation further expands the Mach number range of the FL-60 wind tunnel test, enabling it to have low supersonic test capabilities of M=1.3, 1.35 and 1.4. Through flow field debugging and calibration, a corresponding Mach number nozzle is combined with an open wall test section to establish a low supersonic flow field of M=1.3, 1.35 and M1.4 that meets the requirements of the national military standard, so that the wind tunnel has the ability to carry out low supersonic tests of M=1.3, 1.35 and 1.4, so as to meet the urgent needs of current aircraft development for low supersonic aerodynamic wind tunnel tests and the needs of refined testing and analysis of the transsupersonic aerodynamic characteristics of future advanced aircraft.

[0085] Table 1 Aerodynamic derivatives of AGARD-B standard model

[0086] Specific implementation method 2

[0088] A nozzle device designed based on a test method for a transient high-speed wind tunnel low supersonic flow field as described in a specific implementation method 1, such as Figure 1 As shown, it includes a nozzle flexible plate 1, a support 2, a frame 3, a front end plate 4, a sliding end 5, a driver 6, a rear end plate 7 and a main shell 8; the front end plate 4, the rear end plate 7 and the main shell 8 constitute a hollow quadrilateral shell, the upper end of the quadrilateral shell is provided with a plurality of frames 3, the lower end of the quadrilateral shell is provided with a plurality of supports 2, the aerodynamic profile of the nozzle flexible plate 1 is a curved nozzle aerodynamic profile, the left and right ends of the sliding end 5 are both fixed between the front end plate 4 and the rear end plate 7, the upper surface aerodynamic profile and the lower surface aerodynamic profile of the nozzle flexible plate 1 are both provided with a plurality of drivers 6, and the plurality of drivers 6 can change the bending shape of the aerodynamic profile of the nozzle flexible plate 1, and the bending shapes include curved nozzle aerodynamic profiles M1.3, M1.35 and M1.4.

[0089] Working process: The control system controls several drivers to rise or fall synchronously, thereby changing the bending shape of the nozzle profile 1, thereby realizing different nozzle aerodynamic profiles including M1.3 nozzle aerodynamic profile, M1.3 nozzle aerodynamic profile 5 and M1.4 nozzle aerodynamic profile.

[0090] Specifically, the number of the supports 2 and the number of the racks 3 are both greater than four.

[0091] Specifically, the frame 3 is an I-shaped frame, including an upper frame plate, a lower frame plate and a support rod, and the upper frame plate and the lower frame plate have the same length and width.

[0092] Specifically, the support 2 includes an upper sliding hollow tube and a lower sliding I-shaped hollow tube; the upper sliding hollow tube and the lower sliding I-shaped hollow tube are both provided with matching holes, and the upper sliding hollow tube and the lower sliding I-shaped hollow tube are tightened and fixed by bolts and nuts.

[0093] Specifically, the driver 6 includes a spiral lifting rod and a motor; the spiral lifting rod is connected to the motor, and the spiral lifting rod is driven to rise or fall by the rotation of the motor, thereby changing the bending shape of the aerodynamic profile of the nozzle flexible plate 1.

[0094] Specifically, the sliding end 5 is located at the entrance of the nozzle flexible plate 1, rotates around the horizontal axis installed on the front end plate 4, and cooperates with the T-slot on the back of the nozzle flexible plate 1 through two T-guide rails to make it slide inside to adapt to the changes in the length of the flexible plate and the slope of the entrance profile when the aerodynamic profile of the nozzle flexible plate 1 changes; an axial locking device is installed inside the sliding end 5, and the axial force transmitted from the nozzle flexible plate 1 is borne by two screws; the nut slides back and forth on the screw and is positioned by a switch to ensure the profile stability of the nozzle flexible plate 1.

Claims

1. A temporary high-speed wind tunnel low supersonic flow field test method, characterized in that: A nozzle device comprises a nozzle flexible plate (1), a support (2), a frame (3), a front end plate (4), a sliding end (5), a driver (6), a rear end plate (7) and a main shell (8); the front end plate (4), the rear end plate (7) and the main shell (8) constitute a hollow quadrilateral shell; the upper end of the quadrilateral shell is provided with a plurality of frames (3); the lower end of the quadrilateral shell is provided with a plurality of supports (2); the aerodynamic profile of the nozzle flexible plate (1) is a curved nozzle aerodynamic profile; the left and right ends of the sliding end (5) are both fixed between the front end plate (4) and the rear end plate (7); the upper surface aerodynamic profile and the lower surface aerodynamic profile of the nozzle flexible plate (1) are both provided with a plurality of drivers (6); the plurality of drivers (6) are capable of changing the bending shape of the aerodynamic profile of the nozzle flexible plate (1); the bending shapes include curved nozzle aerodynamic profiles M1.3, M1.35 and M1.4; The test method includes the following steps: Step a, using the newly designed M1.3, M1.35 and M1.4 nozzle devices and combined with the perforated wall test section to establish a low supersonic flow field in a transient wind tunnel; Step b, using the Mach number control method to obtain the ideal flow field control quality; Step c, by analyzing the axial Mach number distribution of the test section at different expansion angles, it is determined that the smaller the expansion angle of the test section is at low supersonic speed, the better the uniformity of the flow field is, and the expansion angle of 0° is selected as the cave body condition; Step d, using the nozzle device in combination with the perforated wall test section, adopting the test section expansion angle and flow field control method, to obtain a qualified flow field that meets the national military standard requirements, the flow field uniformity of M1.3, M1.35 and M1.4 are all better than the index requirement of the national military standard flow field uniformity σM≤0.011, and the standard model test and the aircraft model test verification show that the quality of the established low supersonic flow field fully meets the test requirements, and the wind tunnel has the ability of low supersonic test; the flow field control method adopts the mainstream ejection Mach number control method; The specific method of step a comprises the following steps: Step a1, using a nozzle profile design program, by inputting nozzle parameters, the nozzle parameters including nozzle type, nozzle inlet half height, nozzle outlet half height, outlet Mach number and inlet total pressure, to obtain the nozzle aerodynamic profile and the nozzle profile spiral elevator extension; Step a2, performing coefficient correction on the boundary layer thickness of the wind tunnel test section; Step a3, calibrating the profile to meet the mechanical condition restriction requirements, and obtaining M1.3, M1.35 and M1.4 nozzle aerodynamic profiles; Step a4, according to the calibrated M1.3, 1.35, 1.4 nozzle mechanical profiles, the control system drives the electric screw lifter of the nozzle profile flexible plate to operate synchronously, and after the bending shape of the controlled flexible plate matches the nozzle theoretical aerodynamic profile, the sliding end is axially locked; Step a5, using the perforated wall test section, and establishing a low supersonic flow field using step d; The Mach number control method of step b comprises the following steps: Step b1, during the test, the flow field M number is calculated using the total pressure of the stable section and the static pressure of the test section; Step b2, calculating the difference between the flow field M number and the given M number; Step b3: Use a PI controller to adjust the opening of the pressure regulating valve according to the difference.

2. According to claim 1, a temporary high-speed wind tunnel low supersonic flow field test method is characterized in that: The specific method of step c comprises the following steps: Step c1, setting the test section expansion angle to 0° to perform flow field calibration, and obtaining the test section axial Mach number distribution data; Step c2, setting the test section expansion angle to + 0.3° to perform flow field calibration and obtain the test section axial Mach number distribution data; Step c3, setting the test section expansion angle to + 0.5° to perform flow field calibration and obtain the test section axial Mach number distribution data; Step c4, analyzing the axial Mach number distribution data of the test section to determine that different expansion angles have an impact on the uniformity of the flow field in the test section. When the wind tunnel operation mode is the mainstream injection Mach number control mode, the smaller the expansion angle of the test section, the better the uniformity of the flow field; when the expansion angle is 0°, the flow field of the test section is better than the flow field of the test section when the expansion angle is + 0.3° and + 0.5°.

3. A nozzle device designed based on the temporary high-speed wind tunnel low supersonic flow field test method according to claim 1, characterized in that: The invention comprises a nozzle flexible plate (1), a support (2), a frame (3), a front end plate (4), a sliding end (5), a driver (6), a rear end plate (7) and a main shell (8); the front end plate (4), the rear end plate (7) and the main shell (8) constitute a hollow quadrilateral shell; a plurality of frames (3) are arranged at the upper end of the quadrilateral shell; a plurality of supports (2) are arranged at the lower end of the quadrilateral shell; the aerodynamic profile of the nozzle flexible plate (1) is a curved nozzle aerodynamic profile; the left and right ends of the sliding end (5) are both fixed between the front end plate (4) and the rear end plate (7); the upper surface aerodynamic profile and the lower surface aerodynamic profile of the nozzle flexible plate (1) are both provided with a plurality of drivers (6); the plurality of drivers (6) are capable of changing the bending shape of the aerodynamic profile of the nozzle flexible plate (1); the bending shapes include curved nozzle aerodynamic profiles M1.3, M1.35 and M1.4; The number of the supports (2) and the number of the racks (3) are both greater than 4; The frame (3) is an I-shaped frame, comprising an upper frame plate, a lower frame plate and a support rod, and the upper frame plate and the lower frame plate have the same length and width.

4. The nozzle device according to claim 3, characterized in that: The support (2) comprises an upper sliding hollow tube and a lower sliding I-shaped hollow tube; the upper sliding hollow tube and the lower sliding I-shaped hollow tube are both provided with matching holes, and the upper sliding hollow tube and the lower sliding I-shaped hollow tube are fixed by tightening bolts and nuts.

5. The nozzle device according to claim 4, characterized in that: The driver (6) comprises a spiral lifting rod and a motor; the spiral lifting rod is connected to the motor, and the spiral lifting rod is driven to rise or fall by the rotation of the motor, thereby changing the bending shape of the aerodynamic profile of the nozzle flexible plate (1).

6. The nozzle device according to claim 5, characterized in that: The sliding end (5) is located at the entrance of the nozzle flexible plate (1), rotates around a horizontal axis installed on the front end plate (4), and cooperates with the T-slot on the back of the nozzle flexible plate (1) through two T-guide rails to slide inside to adapt to the changes in the length of the flexible plate and the slope of the entrance profile when the aerodynamic profile of the nozzle flexible plate (1) changes; an axial locking device is installed inside the sliding end (5), and the axial force transmitted from the nozzle flexible plate (1) is borne by two lead screws; the nut slides back and forth on the lead screw and is positioned by a switch to ensure the profile stability of the nozzle flexible plate (1).

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