Design method for two-dimensional flexible nozzle with variable exit area across supersonic wind tunnel

Abstract

The present application belongs to the field of transonic wind tunnel design, and discloses a design method of a binary flexible nozzle for a variable exit section of a transonic wind tunnel. The design method comprises determining a model area length; determining a position of a transonic nozzle aerodynamic theory exit section; determining a minimum distance between the transonic nozzle aerodynamic theory exit section and a supersonic nozzle aerodynamic theory exit section; determining a position of the supersonic nozzle aerodynamic theory exit section; determining a height of the supersonic nozzle aerodynamic theory exit section; and determining an expansion angle. The design method of the binary flexible nozzle for the variable exit section of the transonic wind tunnel solves the problem of shock wave interference with the flow field caused by the connection step or gap between the nozzle section and the supersonic test section, improves the uniformity index of the flow field in the full speed range of the transonic wind tunnel, improves the quality of the supersonic flow field and keeps it stable for a long time, improves the quality of the test data, and has engineering practical value.

Description

Technical Field

[0001] This invention belongs to the field of transonic wind tunnel design, specifically relating to a two-dimensional flexible nozzle design method for a transonic wind tunnel with a variable exit cross-section. Background Technology

[0002] Conventional integrated transonic and supersonic wind tunnels typically employ flexible nozzle sections combined with different types of test sections to establish subsonic, transonic, and supersonic flow fields. Specifically, the combination of a flexible nozzle section and a permeable wall test section establishes the subsonic and transonic flow field; the combination of a flexible nozzle section and a solid wall test section establishes the supersonic flow field. Conventional integrated transonic and supersonic wind tunnels offer the following advantages: First, for subsonic, transonic, and supersonic speeds, only different types of test sections need to be replaced to simulate flow fields across different velocity ranges. Second, apart from the test sections, all other tunnel equipment can be shared, resulting in low construction costs. Third, only the test sections need to be replaced to achieve interchangeability between transonic and supersonic conditions, leading to high replacement efficiency. Fourth, a single model can complete all wind tunnel tests across the subsonic, transonic, and supersonic speed ranges, resulting in high model utilization, test preparation efficiency, and wind tunnel testing efficiency. Therefore, they have been widely adopted.

[0003] However, conventional transonic integrated wind tunnels have significant shortcomings. First, the uniformity of the supersonic flow field is not high, typically only meeting the acceptable requirements of GJB1179A-2012, falling far short of advanced standards. Second, the location of the model zone during supersonic flight is unsuitable; the entrance to the supersonic test section is usually used as the starting point of the model zone, with the actual model zone used in the wind tunnel located between the first and second rhomboid regions. Third, it significantly impacts the accuracy of pitching moment measurements. Because the model zone is located between the first and second rhomboid regions, the shock wave from the nozzle exit almost inevitably hits the model, easily causing deviations in aerodynamic forces, especially pitching moment, and reducing the quality of test data. Fourth, the manufacturing and installation / commissioning difficulties of the nozzle and test sections are significantly increased. Fifth, the requirements for the step and gap at the connection between the test and nozzle sections are extremely high, requiring only a positive difference, not exceeding 0.05 mm, while also demanding extremely high sealing performance at the connection.

[0004] The root cause of the above problems lies in the mechanism by which conventional integrated transonic and supersonic wind tunnels establish transonic and supersonic flow fields. Conventional integrated transonic and supersonic wind tunnels require a segmented design for the test section and nozzle section, inevitably resulting in significant shock wave interference from the connection step and gaps between the test section and nozzle section. This shock wave interference negatively impacts flow field uniformity and reduces supersonic flow field performance. To reduce the intensity of shock waves at the connection points, stringent requirements have been imposed on the connection step, gaps, and sealing between the nozzle and test section. However, achieving satisfactory supersonic flow field quality remains challenging. Based on the experience of using conventional transonic integrated wind tunnels, the reasons why good flow field quality cannot be obtained despite significant costs are as follows: First, under supersonic conditions, the profile changes of flexible nozzles vary greatly at different Mach numbers. Although the nozzle outlet uses a fixed flange connection, the internal stress of the flexible plate makes it difficult to ensure that the geometric dimensions of the nozzle outlet are completely consistent under all Mach numbers due to the large profile changes. This makes it difficult to ensure that the step difference between the exit position and the inlet of the test section for all supersonic nozzle profiles meets the requirement of a surplus of no more than 0.05 mm. Second, the long-term use and frequent replacement of the test section makes it difficult to guarantee the long-term stability of the structure, thus making it difficult to ensure that the axial clearance and step difference between the test section and the nozzle section are completely consistent. Third, air leakage at the connection sealing strip interferes with the flow field.

[0005] Currently, there is an urgent need to develop a design method for two-dimensional flexible nozzles with variable exit cross-sections in transsupersonic wind tunnels. Summary of the Invention

[0006] The technical problem to be solved by the present invention is to provide a design method for a two-dimensional flexible nozzle with a variable exit cross section for transsupersonic wind tunnels, so as to overcome the defects of the prior art.

[0007] The present invention provides a two-dimensional flexible nozzle design method for a transsupersonic wind tunnel with a variable exit cross-section, comprising the following steps:

[0008] S10. Determine the length of the model area;

[0009] Given the height H and area A of the transonic and supersonic test sections, determine the length of the model region. L , ;

[0010] S20. Determine the location of the aerodynamically theoretical exit section of the transonic nozzle;

[0011] The aerodynamic profile of the transonic nozzle is a sonic profile. The theoretical aerodynamic exit section position of the transonic nozzle coincides with the physical exit section position of the transonic nozzle. Based on the numerical simulation results, the length of the increased acceleration zone of the transonic test section is determined.

[0012] S30. Determine the minimum distance between the theoretical aerodynamic exit section of the supersonic nozzle and the theoretical aerodynamic exit section of the transonic nozzle;

[0013] Using the wind tunnel axis as the axis of symmetry, and taking the vertex C of the supersonic nozzle's aerodynamically theoretical exit section as the starting point, draw a line with an angle of θ to the axis. hypotenuse, α Let the shock angle be Mach number 1.5; then draw a line segment AB parallel to the wind tunnel axis and at a height of 1 / 3H from the wind tunnel axis. The intersection point of line segment AB and the hypotenuse is A. Draw a perpendicular line from the midpoint F of line segment AB to the wind tunnel axis. The wind tunnel axis and the perpendicular line intersect at O. Point E is the point symmetrical to point C about FO. The length of line segment CE is the minimum distance D between the theoretical aerodynamic exit section of the supersonic nozzle and the theoretical aerodynamic exit section of the transonic nozzle.

[0014] S40. Determine the position of the aerodynamically theoretical exit section of the supersonic nozzle;

[0015] Using the theoretical aerodynamic exit section position of the transonic nozzle as a reference, the theoretical aerodynamic exit section of the transonic nozzle is shifted by a minimum distance D in the upstream direction of the incoming flow, and this is used as the theoretical aerodynamic exit section position of all supersonic Mach number nozzles; the region between the physical exit section and the theoretical exit section of the supersonic nozzle is taken as part of the supersonic test section, realizing the integrated design of the supersonic nozzle and the supersonic test section, resulting in the integrated supersonic test section; within the integrated supersonic test section, the shock wave avoidance test model is used.

[0016] S50. Determine the theoretical aerodynamic exit section height of the supersonic nozzle;

[0017] Based on the characteristics of supersonic boundary layer thickness growth, the boundary layer correction amount from the exit of each supersonic Mach number nozzle to the theoretical aerodynamic exit section of the supersonic nozzle is determined as follows: Based on the linear expansion method, the theoretical aerodynamic exit section height of the supersonic nozzle corresponding to each supersonic Mach number was calculated. , ;

[0018] S60. Determine the opening angle;

[0019] Setting the physical exit section height H of the supersonic nozzle and the aerodynamic theoretical exit section height of the supersonic nozzle at various Mach numbers. The values ​​increase linearly; calculate the divergence angle corresponding to each supersonic Mach number nozzle. , The linearity of each supersonic Mach number nozzle wall panel is ensured by adding support nodes.

[0020] The present invention provides a two-dimensional flexible nozzle design method for a transsonic wind tunnel with variable exit cross-section, employing different nozzle exit cross-sections across two different velocity ranges: transsonic and supersonic. Specifically, the transsonic nozzle has a sonic profile, and its theoretical aerodynamic exit cross-section is the same as the physical exit cross-section of the flexible nozzle. The supersonic nozzle has a supersonic profile at each Mach number, with its theoretical aerodynamic exit cross-section located at a fixed position upstream of the physical exit cross-section. The region between the theoretical and physical exit cross-sections constitutes the supersonic test section.

[0021] The present invention provides a two-dimensional flexible nozzle design method for a transonic wind tunnel with a variable exit cross-section. This method alters the aerodynamic theoretical nozzle exit cross-section position, fundamentally resolving the issue that establishing transonic and supersonic flow fields requires a segmented design for the test section and nozzle section. In transonic conditions, the aerodynamic theoretical nozzle exit cross-section is located at the physical exit cross-section; in supersonic conditions, it is located upstream of the physical exit cross-section. This effectively resolves the contradiction between establishing transonic and supersonic flow fields and the nozzle exit cross-section. Furthermore, the region between the aerodynamic theoretical and physical exit cross-sections is incorporated as part of the wind tunnel's supersonic test section, enabling an integrated design of the nozzle and the supersonic test section.

[0022] The two-dimensional flexible nozzle design method for variable exit cross-section of transsupersonic wind tunnels of the present invention fundamentally solves the problem of shock wave interference on the flow field caused by the connection step or gap between the nozzle section and the supersonic test section. It improves the uniformity index of the flow field across the entire velocity range of the transsupersonic wind tunnel, improves the quality of the supersonic flow field and maintains its stability over a long period of time, improves the quality of test data, and has practical engineering value. Attached Figure Description

[0023] Figure 1 This is a flowchart of the design method for a two-dimensional flexible nozzle with a variable exit cross-section for a transsupersonic wind tunnel according to the present invention.

[0024] Figure 2 This is a schematic diagram showing the connection between the transonic nozzle section and the transonic test section of a conventional transonic integrated wind tunnel.

[0025] Figure 3 This is a schematic diagram showing the connection between the supersonic nozzle section and the supersonic test section in a conventional transsupersonic integrated wind tunnel.

[0026] Figure 4 A schematic diagram of a transonic wind tunnel structure obtained by the binary flexible nozzle design method for a variable exit section of a transonic wind tunnel according to the present invention.

[0027] Figure 5 A schematic diagram of a supersonic wind tunnel structure obtained by the binary flexible nozzle design method for a variable exit section of a supersonic wind tunnel according to the present invention.

[0028] Figure 6 for Figure 5 A magnified view of a portion of the image.

[0029] In the diagram, 1. Transonic nozzle; 2. Transonic nozzle exit; 3. Transonic test section; 4. Supersonic nozzle; 5. Support section; 6. Support mechanism; 7. Connecting step or gap; 8. Model area; 9. Supersonic test section; 10. Interference wave system; 11. Supersonic nozzle exit; 12. Transonic nozzle physical exit section; 13. Transonic nozzle aerodynamic theoretical exit section; 14. Supersonic nozzle aerodynamic theoretical exit section; 15. Integrated supersonic test section; 16. Supersonic nozzle physical exit section; 17. Shock wave. Detailed Implementation

[0030] The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0031] Comparative example: such as Figure 2 As shown, a conventional transonic integrated wind tunnel uses a combination of a flexible nozzle and a transonic permeable wall test section. The transonic nozzle 1 is a two-dimensional flexible nozzle, and the transonic test section 3 is a transonic permeable wall test section to establish a subsonic and transonic flow field. The conventional transonic integrated wind tunnel includes a transonic nozzle 1, a transonic test section 3, and a support section 5 connected sequentially from front to back. There is a connection step or gap 7 at the transonic nozzle outlet 2 of the transonic nozzle 1. A model area 8 is set in the transonic test section 3, and a support mechanism 6 is set in the support section 5. The support mechanism 6 supports the test model from back to front, and the test model is located in the model area 8.

[0032] like Figure 3 As shown, a conventional transonic integrated wind tunnel uses a combination of a flexible nozzle and a solid-wall test section. The supersonic nozzle 4 and the transonic nozzle 1 use the same two-dimensional flexible nozzle. The supersonic test section 9 is a solid-wall test section to establish a supersonic flow field. The conventional transonic integrated wind tunnel includes the supersonic nozzle 4, the supersonic test section 9, and the support section 5 connected sequentially from front to back. Similarly, there is a connection step or gap 7 at the supersonic nozzle outlet 11 of the supersonic nozzle 4. A model area 8 is set in the supersonic test section 9, and a support mechanism 6 is set in the support section 5. The support mechanism 6 supports the test model from back to front. The test model is located in the model area 8, and there is an interference wave system 10 in the model area 8.

[0033] Example: Figure 1 As shown, the design method for a two-dimensional flexible nozzle with a variable exit cross-section across a supersonic wind tunnel in this embodiment includes the following steps:

[0034] S10. Determine the length of model area 8;

[0035] Given the height H and area A of transonic test section 3 and supersonic test section 9, determine the length of model region 8. L , ;

[0036] According to GJB1179A-2012, the length L of the model area 8 of transonic test section 3 and supersonic test section 9 is generally not less than [a certain value]. Furthermore, given that transonic test section 3 and supersonic test section 9 typically employ square cross-sections, the length of model area 8 is taken as... ;

[0037] S20. Determine the position of the aerodynamically theoretical exit section 13 of the transonic nozzle;

[0038] like Figure 4 As shown, the aerodynamic profile of the transonic nozzle 1 is a sonic profile. The position of the theoretical aerodynamic exit section 13 of the transonic nozzle coincides with the position of the physical exit section 12 of the transonic nozzle. Based on the numerical simulation results, the length of the increased acceleration zone of the transonic test section is determined.

[0039] For transonic flow fields, it is generally required that the transonic nozzle 1 adopts a sound velocity profile so that the airflow reaches the speed of sound at the physical exit section 12 of the transonic nozzle under the action of pressure ratio and nozzle. Then, by utilizing the acceleration characteristics of the transonic permeable wall test section, the length of the acceleration zone of the transonic test section is appropriately increased to generate a hyposonic flow field within the transonic test section 3. Based on the above mechanism, it is determined that the transonic nozzle 1 adopts a sound velocity profile, and the aerodynamic theoretical exit section 13 of the transonic nozzle coincides with the physical exit section 12 of the transonic nozzle. Meanwhile, in order to prevent the airflow from reaching supersonic speed in the transonic nozzle 1 and from overexpansion in the acceleration zone of the transonic permeable wall test section, which would increase the length of the acceleration zone of the transonic test section, the increased length of the transonic test section acceleration zone must be limited. Moreover, the increased length of the transonic test section acceleration zone will reduce the length of the model zone 8 and reduce the flow field uniformity index of the model zone 8. A balance between the increased length of the transonic test section acceleration zone and the flow field uniformity index of the model zone 8 must be obtained based on the numerical simulation results.

[0040] S30. Determine the minimum distance between the aerodynamic theoretical exit section 14 of the supersonic nozzle and the aerodynamic theoretical exit section 13 of the transonic nozzle;

[0041] The aerodynamic theoretical exit section 14 of the supersonic nozzle and the aerodynamic theoretical exit section 13 of the transonic nozzle are two different locations; for example Figure 5 As shown, with the wind tunnel axis as the axis of symmetry and the vertex C of the supersonic nozzle aerodynamic theoretical exit section 14 as the starting point, draw a line with an angle of θ with the axis. hypotenuse, αThe shock angle is Mach 1.5; then draw a line segment AB parallel to the wind tunnel axis and at a height of 1 / 3H from the wind tunnel axis. The intersection point of line segment AB and the hypotenuse is A. Draw a perpendicular line from the midpoint F of line segment AB to the wind tunnel axis. The wind tunnel axis and the perpendicular line intersect at O. Point E is the point symmetrical to point C about FO. The length of line segment CE is the minimum distance D between the theoretical aerodynamic exit section 14 of the supersonic nozzle and the theoretical aerodynamic exit section 13 of the transonic nozzle.

[0042] S40. Determine the position of the aerodynamic theoretical exit section 14 of the supersonic nozzle;

[0043] like Figure 5 As shown, in order to eliminate the adverse disturbances of shock waves or expansion waves generated by the connection step or gap 7 between the supersonic nozzle 4 and the supersonic test section 9 to the supersonic flow field, the supersonic nozzle 4 and the supersonic test section 9 are designed and manufactured as a whole to form the integrated supersonic test section 15. Specifically, taking the position of the transonic nozzle aerodynamic theoretical exit section 13 as a reference, the transonic nozzle aerodynamic theoretical exit section 13 is translated by a minimum distance D in the upstream direction of the incoming flow, and this is used as the position of the supersonic nozzle aerodynamic theoretical exit section 14 for all supersonic Mach number nozzles; the area between the physical exit section 16 and the aerodynamic theoretical exit section 14 of the supersonic nozzle is taken as part of the supersonic test section, realizing the integrated design of the supersonic nozzle 4 and the supersonic test section 9, resulting in the integrated supersonic test section 15; within the integrated supersonic test section 15, the shock wave 17 avoids the test model;

[0044] S50. Determine the theoretical aerodynamic exit section 14 height of the supersonic nozzle;

[0045] like Figure 6 As shown, based on the characteristics of supersonic boundary layer thickness growth, the boundary layer correction amount from the exit of each supersonic Mach number nozzle to the aerodynamic theoretical exit section 14 of the supersonic nozzle is determined as follows: Based on the linear expansion method, the theoretical aerodynamic exit section height 14 of the supersonic nozzle corresponding to each supersonic Mach number was calculated. , ;

[0046] In supersonic applications, to reduce the axial Mach number gradient, the supersonic nozzle's physical exit section 16 to its theoretical aerodynamic exit section 14 must have an adjustable opening angle to compensate for the increase in boundary layer thickness. Since the height of the physical exit section 16 is fixed at H, the height of the theoretical aerodynamic exit section 14 must be determined based on the boundary layer thickness growth characteristics of nozzles at different supersonic Mach numbers. Assuming that the supersonic boundary layer displacement thickness increment between the physical exit section 16 and the theoretical aerodynamic exit section 14 for each supersonic Mach number nozzle is... Let i = 1 to n, where n is the number of supersonic Mach number nozzles. Then, the height of the aerodynamic theoretical exit section 14 corresponding to each supersonic Mach number nozzle is... Displacement thickness needs to be deducted. Otherwise, it will affect the spatial uniformity of the supersonic flow field;

[0047] S60. Determine the opening angle;

[0048] The physical exit section height H of the supersonic nozzle is set to the aerodynamic theoretical exit section height H of the supersonic nozzle at various Mach numbers. The values ​​increase linearly; calculate the divergence angle corresponding to each supersonic Mach number nozzle. , The linearity of each supersonic Mach number nozzle wall panel is ensured by adding support nodes.

[0049] The 1-meter-class transonic wind tunnel of the China Aerodynamics Research and Development Center was designed and manufactured using the two-dimensional flexible nozzle design method for variable exit cross-section of transonic wind tunnels of the present invention. According to the flow field calibration results, the 1-meter-class transonic wind tunnel significantly improved the uniformity index of the supersonic flow field while ensuring good subsonic and transonic flow field quality. The flow field quality has reached the advanced index of GJB1179A-2012.

[0050] Although the embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. For those skilled in the art, all features disclosed in the present invention, or all steps in all methods or processes disclosed, except for mutually exclusive features and / or steps, can be combined in any way without departing from the principles of the present invention. The present invention is not limited to the specific details and illustrations shown and described herein.

Claims

1. A method for designing a two-dimensional flexible nozzle with a variable exit cross-section for transsupersonic wind tunnels, characterized in that, Includes the following steps: S10. Determine the length of the model region (8); Given the height H and area A of the transonic test section (3) and the supersonic test section (9), determine the length of the model region (8). L , ; S20. Determine the location of the transonic nozzle aerodynamic theory exit section (13); The aerodynamic profile of the transonic nozzle (1) is a sonic profile. The position of the theoretical aerodynamic exit section (13) of the transonic nozzle coincides with the position of the physical exit section (12) of the transonic nozzle. Based on the numerical simulation results, the length of the increased acceleration zone of the transonic test section is determined. S30. Determine the minimum distance between the aerodynamic theoretical exit section (14) of the supersonic nozzle and the aerodynamic theoretical exit section (13) of the transonic nozzle; With the wind tunnel axis as the axis of symmetry, and the vertex C of the supersonic nozzle aerodynamic theory exit section (14) as the starting point, draw a line with an angle of θ with the axis. hypotenuse, α The shock angle is Mach number 1.5; draw a line segment AB parallel to the wind tunnel axis and at a height of 1 / 3H from the wind tunnel axis. The intersection of line segment AB and the hypotenuse is A. Draw a perpendicular line from the midpoint F of line segment AB to the wind tunnel axis. The wind tunnel axis and the perpendicular line intersect at O. Point E is the symmetrical point of point C about FO. The length of line segment CE is the minimum distance D between the aerodynamic theoretical exit section (14) of the supersonic nozzle and the aerodynamic theoretical exit section (13) of the transonic nozzle. S40. Determine the position of the aerodynamic theoretical exit section (14) of the supersonic nozzle; Based on the position of the transonic nozzle aerodynamic theory exit section (13), the transonic nozzle aerodynamic theory exit section (13) is shifted by a minimum distance D in the upstream direction of the incoming flow, and this is used as the position of the supersonic nozzle aerodynamic theory exit section (14) for all supersonic Mach number nozzles; the area between the supersonic nozzle physical exit section (16) and the supersonic nozzle aerodynamic theory exit section (14) is used as part of the supersonic test section, realizing the integrated design of the supersonic nozzle (4) and the supersonic test section (9), resulting in the supersonic integrated test section (15); within the supersonic integrated test section (15), the shock wave (17) avoids the test model; S50. Determine the height of the aerodynamic theoretical exit section (14) of the supersonic nozzle; Based on the characteristics of the boundary layer thickness growth in supersonic applications, the boundary layer correction for the section (14) from the exit of each supersonic Mach number nozzle to the theoretical aerodynamic exit section of the supersonic nozzle is determined as follows: Based on the linear expansion method, the height of the aerodynamic theoretical exit section (14) of the supersonic nozzle corresponding to each supersonic Mach number nozzle was calculated. , ; S60. Determine the opening angle; Set the height H of the physical exit section (16) of the supersonic nozzle and the height of the aerodynamic theoretical exit section (14) of the supersonic nozzle for each Mach number. The values ​​increase linearly; calculate the divergence angle corresponding to each supersonic Mach number nozzle. , The linearity of each supersonic Mach number nozzle wall panel is ensured by adding support nodes.

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