CN108303227B - Static aeroelastic wind tunnel test semi-model system and test method - Google Patents
Static aeroelastic wind tunnel test semi-model system and test method Download PDFInfo
- Publication number
- CN108303227B CN108303227B CN201810152069.4A CN201810152069A CN108303227B CN 108303227 B CN108303227 B CN 108303227B CN 201810152069 A CN201810152069 A CN 201810152069A CN 108303227 B CN108303227 B CN 108303227B
- Authority
- CN
- China
- Prior art keywords
- wind tunnel
- model
- semi
- model system
- main body
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Fee Related
Links
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
- G01M9/02—Wind tunnels
Landscapes
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- General Physics & Mathematics (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
The invention provides a static pneumatic elastic wind tunnel test half model system and a test method, wherein the static pneumatic elastic wind tunnel test half model system comprises: the measuring device is arranged outside the wind tunnel wallboard; the half model comprises a main body and a pneumatic airfoil, wherein a thickening part is further arranged on one side of the symmetrical surface, facing the wind tunnel wall plate, of the main body, the thickening part is adjacent to the wind tunnel wall plate, and the thickening part is used for enabling the distance between the symmetrical surface and the wind tunnel wall plate to be larger than or equal to the thickness of a boundary layer of the wind tunnel; the measuring device is used for measuring the static pneumatic elastic effect on the pneumatic airfoil. The static pneumatic elastic wind tunnel test half model system and the test method provided by the invention can effectively isolate the boundary layer of the tunnel wall and eliminate the pneumatic interference of gap cross flow to the half model, and have high practicability and economy.
Description
Technical Field
The invention relates to the technical field of wind tunnel tests, in particular to a static pneumatic elastic wind tunnel test semi-model system and a test method.
Background
Static aeroelasticity is a phenomenon in which an aircraft is elastically deformed under aerodynamic action to cause a change in its aerodynamic characteristics. The static aeroelasticity has a great influence on aerodynamic characteristics of the aircraft, and can possibly change the geometric shape of the aircraft, influence the control surface efficiency and the stability of the aircraft, and cause the phenomena of airfoil damage and the like caused by the divergence of the lifting surface. Therefore, it is necessary to investigate the aerostatic elasticity of an aircraft by experiments. The static aeroelasticity of the aircraft can be studied through a high-speed wind tunnel test, and the flow field of the aircraft is bilaterally symmetrical, so that a half-model test method is developed, and the longitudinal aerodynamics of the whole aircraft is studied by taking the aircraft as a general object. The half-model test method has obvious disadvantages: because the half model is installed near the wind tunnel wallboard on the longitudinal symmetry plane, the low-speed and low-energy air flow in the boundary layer of the tunnel wall can influence the test result of the half model, and a certain gap is reserved between the longitudinal symmetry plane of the model and the tunnel wall during the force measurement test, and air flow is required to flow through the gap, so that the cross flow can influence the test result.
In the prior art, interference of a cavity wall boundary layer on a semi-model test method is eliminated by a cushion block method, a boundary layer suction method or a boundary layer blowing method. Fig. 1 is a schematic structural diagram of a half-model system of a spacer block method provided in the prior art, and referring to fig. 1, the spacer block method is to pad a spacer block 13 with equal thickness between a symmetry plane of a half-model 11 and a wind tunnel wall plate 12, the spacer block 13 is fixed on a rotating window 14, the half-model 11 is connected with a force measuring balance 15 through a connecting piece 16, the angle of attack of the half-model 11 can be changed by rotating the rotating window 14, and the influence of a low kinetic energy and low flow velocity region of a tunnel wall boundary layer on a test result is reduced by the arrangement of the spacer block 13. Fig. 2 is a schematic structural diagram of a half model system of a boundary layer suction method provided in the prior art, referring to fig. 2, the boundary layer suction method is to suck all or part of a boundary layer of a cavity wall at a position where the half model 21 is located and a vicinity thereof by arranging a suction device 22 at an upstream of the half model 21, wherein a symmetry plane of the half model 21 is arranged close to a wind tunnel wall plate 23, the half model 21 is connected with a force balance 25 through a connecting piece 24, an attack angle of the half model 21 can be changed by rotating a rotating window 26, and the suction device 22 sucks the boundary layer of the cavity wall to eliminate or reduce an influence of the boundary layer on a test result. Fig. 3 is a schematic structural diagram of a half-model system of a boundary layer blowing method provided in the prior art, and referring to fig. 3, the boundary layer blowing method is implemented by blowing high-pressure air into a boundary layer along a wall surface of a wind tunnel wall plate 33 through a blowing device 32 at an upstream of a half-model 31, so that the air flow speed is increased, and the boundary layer is thinned, so that the influence of low-energy flow of the boundary layer on the half-model test result is reduced, wherein a symmetrical surface of the half-model 31 is arranged close to the wind tunnel wall plate 33, the half-model 31 is connected with a force balance 35 through a connecting piece 34, and the attack angle of the half-model 31 can be changed by rotating a rotating window 36.
In the spacer block method, since a gap with a certain width is formed between the symmetry plane of the half-model 11 and the spacer block 13, the channeling in the gap can interfere the flow around the half-model 11, resulting in distortion of the test result. When the boundary layer suction method is adopted, the suction rate of the suction device 22 is large enough to ensure that the boundary layer of the cavity wall is sufficiently sucked, and for a longer half model, the suction device 22 is also required to be respectively arranged at the front section, the middle section and the rear section of the half model to uniformly suck the boundary layer, and meanwhile, the flow of the main flow is not influenced. When the boundary layer blowing method is adopted, a proper blowing device 23 is selected, and proper blowing pressure and blowing amount are determined, so that high-pressure air flow can be uniformly blown in, the flow velocity in the boundary layer can be close to or equal to but not exceed the main flow velocity, the technical difficulty is high, and the geometric dimension of the half model is limited.
Disclosure of Invention
The invention provides a static pneumatic elastic wind tunnel test half model system and a test method, which can effectively isolate a boundary layer of a tunnel wall and eliminate pneumatic interference of gap cross flow to the half model, and have high practicability and economy.
In one aspect, the present invention provides a static pneumatic elastic wind tunnel test semi-model system, comprising: the measuring device is arranged outside the wind tunnel wallboard;
the half model comprises a main body and a pneumatic airfoil, wherein a thickening part is further arranged on one side of the symmetrical surface, facing the wind tunnel wall plate, of the main body, the thickening part is adjacent to the wind tunnel wall plate, and the thickening part is used for enabling the distance between the symmetrical surface and the wind tunnel wall plate to be larger than or equal to the thickness of a boundary layer of the wind tunnel;
the measuring device is used for measuring the static pneumatic elastic effect on the pneumatic airfoil.
According to the static pneumatic elastic wind tunnel test half model system provided by the embodiment of the invention, the adverse effect of a wind tunnel side wall boundary layer and gap channeling on the static pneumatic elastic half model is avoided by arranging the thickening part, and meanwhile, the measuring device only measures aerodynamic force received by the pneumatic airfoil, so that the static pneumatic elasticity of the pneumatic airfoil of the half model can be accurately measured; compared with the traditional boundary layer suction method and boundary layer blowing method, the method has the advantages of no need of complex air suction or blowing devices, high economy, simple structure, no technical problems, easy operation and no limitation on the geometric dimension of the half model; the setting of the half model and the measurement mode of the pneumatic airfoil surface of the half model are easy to realize, and the pneumatic airfoil surface measuring device is applicable to different kinds of half models and has strong practicability.
In the half-model system described above, the thickness of the thickened portion in the direction perpendicular to the wind tunnel wall plate is equal to the boundary layer thickness.
In the half-model system described above, the contour of the thickened portion and the shape of the symmetry plane are identical.
The semi-model system as described above, said measuring means comprising a rotating window, a force measuring level and a connecting piece; the rotating window is fixedly connected with the main body, the force measuring balance is connected with the pneumatic airfoil surface through the connecting piece, and the force measuring balance is connected with the rotating window.
In the semi-model system, the main body is provided with a first avoidance hole, and the connecting piece penetrates through the first avoidance hole to be connected with the pneumatic airfoil.
According to the semi-model system, the rotary window is provided with the second avoidance hole, and the connecting piece penetrates through the second avoidance hole and the first avoidance hole to be connected with the pneumatic airfoil.
In the half-model system, the connecting pieces, the rotating window and the main body are arranged at intervals.
In the semi-model system described above, a gap is provided between the main body and the aerodynamic airfoil.
In the half-model system, the rotating window drives the half-model, the force measuring balance and the connecting piece to rotate.
According to the static pneumatic elastic wind tunnel test half model system provided by the embodiment of the invention, the adverse effect of a wind tunnel side wall boundary layer and gap channeling on the static pneumatic elastic half model is avoided by arranging the thickening part, and meanwhile, the measuring device only measures aerodynamic force received by the pneumatic airfoil, so that the static pneumatic elasticity of the pneumatic airfoil of the half model can be accurately measured; compared with the traditional boundary layer suction method and boundary layer blowing method, the method has the advantages of no need of complex air suction or blowing devices, high economy, simple structure, no technical problems, easy operation and no limitation on the geometric dimension of the half model; the arrangement of the half model and the measurement mode of the pneumatic airfoil surface of the half model are easy to realize, and the pneumatic airfoil surface measuring device is applicable to different types of half models and has strong practicability; the attack angle of the half model can be changed by controlling the rotation of the rotation window, and aerodynamic force born by the aerodynamic wing surface can be directly measured by the force measuring balance, so that the structure is simple and the realization is easy.
The invention also provides a static pneumatic elastic wind tunnel test half model test method, which is applied to the static pneumatic elastic wind tunnel test half model system, and comprises the following steps:
changing the angle of attack of the semi-model;
and under a preset attack angle, measuring the static pneumatic elastic effect suffered by the pneumatic airfoil in the static pneumatic elastic wind tunnel test half model system.
According to the test method of the static pneumatic elastic wind tunnel test half model, provided by the embodiment of the invention, the adverse effect of the boundary layer of the wind tunnel side wall and the gap channeling on the static pneumatic elastic half model test is avoided by arranging the thickening part on the half model, and meanwhile, the measuring device only measures aerodynamic force born by the pneumatic airfoil, so that the static pneumatic elasticity of the pneumatic airfoil of the half model can be accurately measured; compared with the traditional boundary layer suction method and boundary layer blowing method, the method has the advantages of no need of complex air suction or blowing devices, high economy, simple structure, no technical problems, easy operation and no limitation on the geometric dimension of the half model; the arrangement of the half model and the measurement mode of the pneumatic airfoil surface of the half model are easy to realize, and the pneumatic airfoil surface measuring device is applicable to different types of half models and has strong practicability; the attack angle of the half model can be changed by controlling the rotation of the rotation window, and aerodynamic force born by the aerodynamic wing surface can be directly measured by the force measuring balance, so that the method is simple and easy to realize.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions of the prior art, the drawings that are needed in the embodiments or the description of the prior art will be briefly described below, it will be obvious that the drawings in the following description are some embodiments of the present invention, and that other drawings can be obtained according to these drawings without inventive effort to a person skilled in the art.
FIG. 1 is a schematic diagram of a half-model system of a pad method according to the prior art;
FIG. 2 is a schematic diagram of a prior art boundary layer pumping half model system;
FIG. 3 is a schematic diagram of a half model system of boundary layer blowing provided in the prior art;
fig. 4 is a schematic structural diagram of a static pneumatic elastic wind tunnel test half model system according to an embodiment of the present invention.
Reference numerals:
11-half model 12-wind tunnel wallboard
13-cushion 14-rotating window
15-force balance 16-connector
21-half-mould 22-aspirator
23-wind tunnel wall 24-connector
25 force balance 26 rotation window
31-half model 32-blowing device
33-wind tunnel wall 34-connector
35 force balance 36 rotary window
41-half model 411-body
4111 first relief holes 412-aerodynamic airfoil
413-thickening 42-measuring device
421-Rotary Window 4211-second dodge hole
422-force balance 423-connector
43-wind tunnel wall plate
Detailed Description
For the purpose of making the objects, technical solutions and advantages of the present invention more apparent, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings, and it is apparent that the described embodiments are some embodiments of the present invention, not all embodiments. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
Example 1
Fig. 4 is a schematic structural diagram of a static pneumatic elastic wind tunnel test half model system provided by the embodiment of the present invention, and referring to fig. 4, the embodiment of the present invention provides a static pneumatic elastic wind tunnel test half model system, including: a half-model 41 arranged in the wind tunnel and a measuring device 42 arranged outside the wind tunnel wall plate 43;
the half-model 41 comprises a main body 411 and a pneumatic airfoil 412, wherein the side of the main body 411, facing the wind tunnel wall plate 43, of the symmetry plane is further provided with a thickening part 413, the thickening part 413 is adjacent to the wind tunnel wall plate 43, and the thickening part 413 is used for enabling the distance between the symmetry plane and the wind tunnel wall plate 43 to be larger than or equal to the boundary layer thickness of the wind tunnel;
the measurement device 42 is used to measure the aeroelasticity on the aero-airfoil 412.
In the embodiment of the invention, the actual shape of the half model 41 is outside the influence area of the boundary layer and the gap channeling of the wind tunnel wall plate 43 due to the arrangement of the thickening part 413, which is equivalent to adding a cushion block between the half model 41 and the wind tunnel wall plate 43, but the cushion block and the half model 41 are integrated, so that a new gap can be avoided between the cushion block and the actual half model 41 by the design, and the influence of the gap channeling generated by the traditional cushion block method is eliminated.
Meanwhile, the measuring device 42 is only used for measuring the aeroelasticity of the aerodynamic surface 412, which means that the main body 411 and the aerodynamic surface 412 of the half model 41 are detached, the main body 411 is only used for providing real flow around the aerodynamic surface 412, and the measuring device 42 does not measure the aerodynamic force received by the main body 411, but only the aerodynamic force received by the aerodynamic surface 412. Since the thickening 413 increases the thickness of the body 411 of the half-model 41, the shape of the half-model 41 is distorted, and if aerodynamic forces received by the body 411 and the airfoil 412 of the half-model 41 are measured together in a conventional half-model force measuring manner, the measurement result is distorted. Thus, the provision of the measurement device 42 to measure only the aeroelasticity of the aerofoil 412 eliminates the negative effects of the thickened portion 413, resulting in more accurate aerodynamic data of the aerofoil 412.
Specifically, the half-model 41 disposed in the wind tunnel is impacted by the high-speed airflow, the measuring device 42 disposed outside the wind tunnel wall plate 43 does not affect the flow around the half-model 41, and the measuring device 42 also has the function of changing the attack angle of the half-model 41, so as to obtain the aeroelasticity values of the aerodynamic airfoils 412 under different angles of airflow.
Furthermore, the thickening 413 is used to make the distance between the symmetry plane and the tunnel wall 43 greater than or equal to the boundary layer thickness of the tunnel, which depends on the boundary layer characteristics of the tunnel, on the characteristics of the structure, the way of operation and the total pressure of the air flow, etc., and varies with the test method and conditions. The thickness of the boundary layer is generally within 50 mm, and thus the thickness of the thickened portion 413 is also about 50 mm, and the thickness of the thickened portion 413 is small compared with the length dimension of the half model 41, and the influence of the distortion of the shape of the half model 41 is small, but the thickened portion 413 has a remarkable effect on separating the boundary layer to weaken the influence of the boundary layer on the test result of the half model 41.
In this example, the specific kind and shape of the half model 41 are not limited. The half-model 41 may be an airplane type half-model, in which case the main body is the fuselage of an airplane and the aerodynamic airfoil is the wing or tail of the airplane; the half model 41 may be a missile type half model, in which case the main body is a missile body of the missile, and the aerodynamic airfoil is a missile wing of the missile. The structural types of the half model 41 serving as a main body and the aerodynamic wing surface are met, and meanwhile, the purpose of the aerostatic wind tunnel test is to obtain the condition of the aerostatic elastic data of the aerodynamic wing surface.
According to the static pneumatic elastic wind tunnel test half model system provided by the embodiment of the invention, the adverse effect of a wind tunnel side wall boundary layer and gap channeling on the static pneumatic elastic half model is avoided by arranging the thickening part, and meanwhile, the measuring device only measures aerodynamic force received by the pneumatic airfoil, so that the static pneumatic elasticity of the pneumatic airfoil of the half model can be accurately measured; compared with the traditional boundary layer suction method and boundary layer blowing method, the method has the advantages of no need of complex air suction or blowing devices, high economy, simple structure, no technical problems, easy operation and no limitation on the geometric dimension of the half model; the setting of the half model and the measurement mode of the pneumatic airfoil surface of the half model are easy to realize, and the pneumatic airfoil surface measuring device is applicable to different kinds of half models and has strong practicability.
On the basis of the above embodiment, the thickness of the thickened portion 413 in the direction perpendicular to the wind tunnel wall plate 43 is equal to the boundary layer thickness. The boundary layer thickness of the wind tunnel depends on the boundary layer characteristics of the wind tunnel, and is related to the characteristics of the wind tunnel, such as the structure, the operation mode, the total air flow pressure and the like, and the change of the test method and the condition also has an effect on the boundary layer thickness. In practice, the theoretical boundary layer thickness may be calculated according to the wind tunnel characteristics, so that the thickness of the thickened portion 413 in the direction perpendicular to the wind tunnel wall plate 43 is equal to the boundary layer thickness, and the thickened portion 413 may separate the boundary layer to weaken the test result of the boundary layer half model 41. If the thickness of the thickened portion 413 is smaller than the boundary layer thickness, the boundary layer still has a large influence on the test result of the half model 41; if the thickness of the thickened portion 413 is greater than the thickness of the boundary layer, the thickened portion 413 causes serious model distortion of the half model 41, thereby affecting the test result, and too thick thickened portion 413 causes waste of model material.
In order to further avoid negative effects of model distortion caused by the thickenings 413 on the test results, the contours of the thickenings 413 and the shape of the symmetry plane are identical. If the contour of the thickening 413 is different from the shape of the symmetry plane, it is necessary to have a very different flow-around behavior of the model half 41 from that of the actual model, which results in severely inaccurate test results. When the profile of the thickening 413 is exactly the same as the shape of the plane of symmetry, the thickening 413 is only distorted in thickness, with less negative effects, and the measuring device only measures the aerodynamic forces of the aerodynamic airfoil, which are eliminated to some extent.
Specifically, the measuring device 42 includes a rotary window 421, a load cell 422, and a connector 423; the rotary window 421 is fixedly connected with the main body 411, the force measuring balance 422 is connected with the pneumatic airfoil 412 through a connecting piece 423, and the force measuring balance 422 is connected with the rotary window 421.
The rotary window 421 is used for driving the half-model 41 to rotate so as to change the attack angles of the half-model 41 and the airflow, and obtain the static pneumatic elasticity data under different attack angles. In this embodiment, the rotary window 421 is fixedly connected to the main body 411, and the aerodynamic wing 412 is connected to the force-measuring balance 422 through the connecting piece 423, that is, the main body 411 is not connected to the aerodynamic wing 412, and the force-measuring balance 422 only measures the aerodynamic force of the aerodynamic wing 412. Because the force measuring balance 422 is connected with the rotating window 421, the rotating window 421 drives the force measuring balance 422 to rotate and further drives the pneumatic airfoil 412 to rotate, so that the main body 411 and the pneumatic airfoil 412 are ensured to synchronously rotate, and the main body 411 provides real detouring for the pneumatic airfoil 412.
In order to realize the separation of the main body 411 and the aerodynamic surface 412, a first relief hole 4111 needs to be formed in the main body 411, and the connecting piece 423 is connected to the aerodynamic surface 412 through the first relief hole 4111.
The first relief hole 4111 should be sufficiently sized to receive the connector 423, the connector 423 being connected to the aerodynamic airfoil 412 through the first relief hole 4111, the connector 423 being out of contact with the body 411. The first relief hole 4111 cannot be too large, otherwise the overall weight imbalance of the phantom 41 may result in distortion of the measurement of the static aeroelasticity.
Further, a second avoidance hole 4211 is provided on the rotation window 421, and the connecting member 423 passes through the second avoidance hole 4211 and the first avoidance hole 4111 to be connected with the aerodynamic airfoil 412. The force balance 422 is connected with the aerodynamic wing 412 through the connection piece 423, and since the force balance 422 is disposed in the rotation window 421, the rotation window 421 is provided with the second avoidance hole 4211 so that the connection piece 423 passes through. The first relief hole 4111 and the second relief hole 4211 should have the same cross-sectional shape perpendicular to the wind tunnel wall plate 43, and the second relief hole 4211 should be sufficiently large to accommodate the connecting member 423 without the connecting member 423 contacting the rotary window 421.
In order to ensure the correctness of the measured load, the connecting piece 423 is arranged at intervals with the rotary window 421 and the main body 411, namely, certain gaps are reserved between the connecting piece 423 and the rotary window 421 and between the connecting piece 423 and the main body 411, so that the connecting piece 423 is prevented from contacting with the rotary window 421 and the main body 411 under the action of pneumatic load, and the aerodynamic data errors measured by the measuring device 42 due to collision contact between the connecting piece 423 and the rotary window 421 and the main body 411 in the test process are avoided.
In addition, there is a gap between the body 411 and the aerodynamic airfoil 412. Because the aerodynamic force of the main body 411 and the aerodynamic airfoil 412 can be measured only by the aerodynamic force of the aerodynamic airfoil 412 without measuring the aerodynamic force of the main body 411 in the half-mold static aerodynamic elasticity test, a certain gap is reserved between the main body 411 and the aerodynamic airfoil 412 so as to ensure that no contact exists between the main body 411 and the aerodynamic airfoil 412 under the action of aerodynamic load, thereby avoiding the aerodynamic force data error of the aerodynamic airfoil 412 measured by the measuring device 42.
In actual operation, the gap between the connecting piece 423 and the rotary window 421 and the main body 411, and the gap between the main body 411 and the aerodynamic wing 412 are about 2-3 mm.
During the test, the rotary window 421 drives the half-model 41, the load cell 422 and the connecting piece 423 to rotate. The rotary window 421 can be driven to rotate by a driving device, under the action of the driving force of the driving device, the rotary window 421 drives the main body 411 to rotate, meanwhile, the rotary window 421 drives the force measuring balance 422 to rotate, and further the pneumatic airfoil 412 is driven to rotate, so that the main body 411 and the pneumatic airfoil 412 are guaranteed to synchronously rotate.
According to the static pneumatic elastic wind tunnel test half model system provided by the embodiment of the invention, the adverse effect of a wind tunnel side wall boundary layer and gap channeling on the static pneumatic elastic half model is avoided by arranging the thickening part, and meanwhile, the measuring device only measures aerodynamic force received by the pneumatic airfoil, so that the static pneumatic elasticity of the pneumatic airfoil of the half model can be accurately measured; compared with the traditional boundary layer suction method and boundary layer blowing method, the method has the advantages of no need of complex air suction or blowing devices, high economy, simple structure, no technical problems, easy operation and no limitation on the geometric dimension of the half model; the arrangement of the half model and the measurement mode of the pneumatic airfoil surface of the half model are easy to realize, and the pneumatic airfoil surface measuring device is applicable to different types of half models and has strong practicability; the attack angle of the half model can be changed by controlling the rotation of the rotation window, and aerodynamic force born by the aerodynamic wing surface can be directly measured by the force measuring balance, so that the structure is simple and the realization is easy.
Example two
The embodiment of the invention provides a static pneumatic elastic wind tunnel test half model test method, which is applied to the static pneumatic elastic wind tunnel test half model system, and comprises the following steps:
changing the angle of attack of the half model 41;
aerodynamic forces experienced by aerodynamic airfoil 412 in a static aeroelastic wind tunnel test semi-model system are measured at a preset angle of attack.
Specifically, the angle of attack of the half-model 41 is achieved by rotation of the rotary window 421; under the action of the driving force of the driving device, the rotary window 421 drives the main body 411 to rotate, meanwhile, the rotary window 421 drives the force measuring balance 422 to rotate, and further the pneumatic airfoil 412 is driven to rotate, so that the main body 411 and the pneumatic airfoil 412 are guaranteed to synchronously rotate, and the main body 411 provides real streaming for the pneumatic airfoil 412.
In the case of a determined angle of attack, the aerodynamic forces experienced by the aerodynamic airfoil 412 are readily available through the force balance 422. Because the force-measuring balance 422 is only directly connected with the aerodynamic wing 412, the aerodynamic wing 412 is separated from the main body 411, so that the data measured by the force-measuring balance 422 does not comprise aerodynamic force of the main body 411, and the measurement result is accurate and reliable.
According to the test method of the static pneumatic elastic wind tunnel test half model, provided by the embodiment of the invention, the adverse effect of the boundary layer of the wind tunnel side wall and the gap channeling on the static pneumatic elastic half model test is avoided by arranging the thickening part on the half model, and meanwhile, the measuring device only measures aerodynamic force born by the pneumatic airfoil, so that the static pneumatic elasticity of the pneumatic airfoil of the half model can be accurately measured; compared with the traditional boundary layer suction method and boundary layer blowing method, the method has the advantages of no need of complex air suction or blowing devices, high economy, simple structure, no technical problems, easy operation and no limitation on the geometric dimension of the half model; the arrangement of the half model and the measurement mode of the pneumatic airfoil surface of the half model are easy to realize, and the pneumatic airfoil surface measuring device is applicable to different types of half models and has strong practicability; the attack angle of the half model can be changed by controlling the rotation of the rotation window, and aerodynamic force born by the aerodynamic wing surface can be directly measured by the force measuring balance, so that the method is simple and easy to realize.
In the description of the present invention, it should be understood that the terms "center", "length", "width", "thickness", "top", "bottom", "upper", "lower", "left", "right", "front", "rear", "vertical", "horizontal", "inner", "outer", "axial", "circumferential", etc. are used to indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, merely to facilitate description of the present invention and to simplify the description, and do not indicate or imply that the referred location or element must have a specific orientation, in a specific configuration and operation, and therefore should not be construed as limiting the present invention.
Furthermore, the terms "first," "second," and the like, are used for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include one or more such feature. In the description of the present invention, the meaning of "plurality" means at least two, for example, two, three, etc., unless specifically defined otherwise.
In the present invention, unless explicitly specified and limited otherwise, the terms "mounted," "connected," "secured," and the like are to be construed broadly and may be, for example, fixedly attached, detachably attached, or integrally formed; can be mechanically connected, electrically connected or can be communicated with each other; can be directly connected or indirectly connected through an intermediate medium, and can lead the interior of two elements to be communicated or lead the two elements to be in interaction relationship. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
In the present invention, unless expressly stated or limited otherwise, a first feature "above" or "below" a second feature may include both the first and second features being in direct contact, as well as the first and second features not being in direct contact but being in contact with each other through additional features therebetween. Moreover, a first feature being "above," "over" and "on" a second feature includes the first feature being directly above and obliquely above the second feature, or simply indicating that the first feature is higher in level than the second feature. The first feature being "under", "below" and "beneath" the second feature includes the first feature being directly under and obliquely below the second feature, or simply means that the first feature is less level than the second feature.
Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present invention, and not for limiting the same; although the invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some or all of the technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit of the invention.
Claims (10)
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810152069.4A CN108303227B (en) | 2018-02-14 | 2018-02-14 | Static aeroelastic wind tunnel test semi-model system and test method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810152069.4A CN108303227B (en) | 2018-02-14 | 2018-02-14 | Static aeroelastic wind tunnel test semi-model system and test method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN108303227A CN108303227A (en) | 2018-07-20 |
| CN108303227B true CN108303227B (en) | 2024-04-05 |
Family
ID=62865209
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201810152069.4A Expired - Fee Related CN108303227B (en) | 2018-02-14 | 2018-02-14 | Static aeroelastic wind tunnel test semi-model system and test method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN108303227B (en) |
Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110907125A (en) * | 2018-09-17 | 2020-03-24 | 中国空气动力研究与发展中心低速空气动力研究所 | Method for testing power influence of separated half-mold injection type nacelle |
| CN112098035A (en) * | 2020-11-19 | 2020-12-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Aircraft test system |
| CN112378620B (en) * | 2020-12-09 | 2023-04-18 | 中国航天空气动力技术研究院 | Flexible aircraft wind tunnel static aeroelasticity test model and manufacturing method |
| CN112304564A (en) * | 2020-12-16 | 2021-02-02 | 大连理工大学 | A static aeroelastic wind tunnel test wing model |
| CN116222450B (en) * | 2022-12-29 | 2025-12-23 | 成都凯迪精工科技有限责任公司 | Wind tunnel test model airfoil surface installation accuracy control method |
| CN118130035B (en) * | 2024-04-29 | 2024-09-13 | 中国航空工业集团公司哈尔滨空气动力研究所 | A static aeroelastic test device and method for a large-scale elastic wing section model in a low-speed wind tunnel |
| CN118392441B (en) * | 2024-07-01 | 2024-10-01 | 中国航空工业集团公司沈阳空气动力研究所 | Design method suitable for high-speed wind tunnel test mixed laminar flow vertical tail model |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2010243400A (en) * | 2009-04-08 | 2010-10-28 | Japan Aerospace Exploration Agency | Substation support interference correction method in subsonic half model wind tunnel test |
| CN102012307A (en) * | 2010-11-18 | 2011-04-13 | 中国人民解放军国防科学技术大学 | Supersonic speed boundary layer wind tunnel |
| CN102305699A (en) * | 2011-05-19 | 2012-01-04 | 北京航空航天大学 | Wind tunnel experiment system for free flight model |
| CN102818693A (en) * | 2012-08-17 | 2012-12-12 | 中国航天空气动力技术研究院 | Half-module pivoted window mechanism applied to sub transonic speed wind tunnel |
| CN105021370A (en) * | 2015-07-30 | 2015-11-04 | 中国航空工业集团公司哈尔滨空气动力研究所 | Low speed high Reynolds number wind tunnel semi model force balance and force-measuring method |
| CN105564666A (en) * | 2014-10-11 | 2016-05-11 | 中国航空工业集团公司西安飞机设计研究所 | Aircraft component force measuring wind tunnel test model gap structure design method |
| CN105716827A (en) * | 2014-12-03 | 2016-06-29 | 中航通飞研究院有限公司 | Amphibious aircraft blown flap wind tunnel test model |
| CN208476493U (en) * | 2018-02-14 | 2019-02-05 | 中国空气动力研究与发展中心高速空气动力研究所 | Aeroelastic effect wind tunnel test half model system |
-
2018
- 2018-02-14 CN CN201810152069.4A patent/CN108303227B/en not_active Expired - Fee Related
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2010243400A (en) * | 2009-04-08 | 2010-10-28 | Japan Aerospace Exploration Agency | Substation support interference correction method in subsonic half model wind tunnel test |
| CN102012307A (en) * | 2010-11-18 | 2011-04-13 | 中国人民解放军国防科学技术大学 | Supersonic speed boundary layer wind tunnel |
| CN102305699A (en) * | 2011-05-19 | 2012-01-04 | 北京航空航天大学 | Wind tunnel experiment system for free flight model |
| CN102818693A (en) * | 2012-08-17 | 2012-12-12 | 中国航天空气动力技术研究院 | Half-module pivoted window mechanism applied to sub transonic speed wind tunnel |
| CN105564666A (en) * | 2014-10-11 | 2016-05-11 | 中国航空工业集团公司西安飞机设计研究所 | Aircraft component force measuring wind tunnel test model gap structure design method |
| CN105716827A (en) * | 2014-12-03 | 2016-06-29 | 中航通飞研究院有限公司 | Amphibious aircraft blown flap wind tunnel test model |
| CN105021370A (en) * | 2015-07-30 | 2015-11-04 | 中国航空工业集团公司哈尔滨空气动力研究所 | Low speed high Reynolds number wind tunnel semi model force balance and force-measuring method |
| CN208476493U (en) * | 2018-02-14 | 2019-02-05 | 中国空气动力研究与发展中心高速空气动力研究所 | Aeroelastic effect wind tunnel test half model system |
Non-Patent Citations (5)
| Title |
|---|
| Wilhelm Becker等.Tests for vehicle aerodynamics in the cryogenic wind tunnel Cologne DNW-KKK.《20th International Congress on Instrumentation in Aerospace Simulation Facilities, 2003. ICIASF '03.》.2003,第54-58页. * |
| 左培初,焦予秦,贺家驹.高速柔壁自适应壁风洞中半模型试验技术研究.流体力学实验与测量.2000,第14卷(第3期),第25-31段. * |
| 路波等.跨声速风洞全模颤振试验技术.航空学报.2014,第36卷(第4期),第1086-1092页. * |
| 郭洪涛等.大展弦比机翼跨声速静气动弹性风洞试验.《空气动力学学报》.2017,第第35卷卷(第第6期期),第842-845段. * |
| 郭洪涛等.某战斗机高速全模颤振风洞试验研究.《航空学报》.2012,第33卷(第10期),第1065-1071页. * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN108303227A (en) | 2018-07-20 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN108303227B (en) | Static aeroelastic wind tunnel test semi-model system and test method | |
| CN102680201B (en) | Buffeting wind tunnel test method based on video measurement | |
| CN107757879B (en) | Wingtip device for the wing of an aircraft, aircraft and use | |
| CN105314096B (en) | Individual gas sources supply without rudder face aircraft | |
| CN109918764B (en) | Method for measuring rolling risk degree of aircraft after encountering wake vortex | |
| US10730607B2 (en) | Circulation control system for aerial vehicles | |
| CN105258915B (en) | Become yaw angle blade belly stay device in a kind of high-speed wind tunnel | |
| CN112146839B (en) | Upper surface air blowing power simulation ground test device | |
| CN109131833A (en) | A kind of high aspect ratio wing of high lift-rising | |
| CN112098035A (en) | Aircraft test system | |
| CN106599353A (en) | Dynamic numerical simulation method for external flow field of multi-element airfoil of airplane | |
| CN111017185A (en) | Laminar flow technology verification machine | |
| CN205209733U (en) | A torque measuring balance | |
| Wood et al. | Reduction of tiltrotor download | |
| CN208476493U (en) | Aeroelastic effect wind tunnel test half model system | |
| CN105424971A (en) | Static pressure probe used for gyroplane low speed measurement | |
| Paschal et al. | Evaluation of tunnel sidewall boundary-layer-control systems for high-lift airfoil testing | |
| CN209008845U (en) | A kind of high aspect ratio wing of high lift-rising | |
| CN205246692U (en) | A static pressure probe for little speed measurement of gyroplane | |
| CN205175660U (en) | Become yaw angle blade belly stay device in high -speed wind tunnel | |
| CN208683096U (en) | A kind of scale bridging type covering | |
| Mitchell Jr | Effects of Varying the Size and Location of Trailing-edge Flap-type Controls on the Aerodynamic Characteristics of a Unswept Wing at a Mach Number of 1.9 | |
| CN109625316B (en) | Method for measuring inner side hinge moment of control surface of wing with super-large aspect ratio | |
| CN112660416A (en) | Laminar flow control technology verification machine | |
| CN218865450U (en) | Shape-preserving counterweight device for airplane airfoil flutter model |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20240405 |