CN114001906B - Automatic rudder deflection angle changing device for hypersonic wind tunnel hinge moment measurement test and using method thereof - Google Patents
Automatic rudder deflection angle changing device for hypersonic wind tunnel hinge moment measurement test and using method thereof Download PDFInfo
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- CN114001906B CN114001906B CN202111289019.9A CN202111289019A CN114001906B CN 114001906 B CN114001906 B CN 114001906B CN 202111289019 A CN202111289019 A CN 202111289019A CN 114001906 B CN114001906 B CN 114001906B
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- 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
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
- B64F5/60—Testing or inspecting aircraft components or systems
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- 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
- G01M9/06—Measuring arrangements specially adapted for aerodynamic testing
- G01M9/062—Wind tunnel balances; Holding devices combined with measuring arrangements
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
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Abstract
The invention discloses an automatic rudder deflection angle changing device for a hypersonic wind tunnel hinge moment measurement test and a using method thereof, wherein the device comprises the following steps: the aircraft model is provided with a horizontal reference surface and a 0-degree control rudder positioning mark line, a motor assembly is fixedly installed in a motor installation cavity, a hinge moment balance is fixedly installed in a balance installation cavity, and the motor assembly is fixedly connected with the hinge moment balance through a balance adapter; the measuring rudder is fixedly connected with the hinge moment balance; a first heat insulation sleeve assembly is arranged between the motor assembly and the aircraft model, and a second heat insulation sleeve assembly is arranged between the hinge moment balance and the aircraft model and between the hinge moment balance and the measuring rudder. The motor assembly drives the hinge moment balance and the measuring rudder to rotate, and the measured values of the hinge moment balance and the vision measuring system are recorded in real time in the test process. The invention can conveniently and quickly change the rudder deflection angle of the measuring rudder, effectively avoid the operation failure of the motor under the vacuum condition, ensure that the vacuum does not need to be released after each test and improve the test efficiency.
Description
Technical Field
The invention belongs to the technical field of wind tunnel tests, and particularly relates to an automatic rudder deflection angle changing device for a hypersonic wind tunnel hinge moment measurement test and a using method thereof.
Background
The hinge moment test of the control surface of the model is one of important wind tunnel test items in the development stage of an aircraft, and the purpose is to accurately predict the aerodynamic force of each control surface and the hinge moment relative to a rotating shaft and provide a basis for steering engine selection, aerodynamic shape design and structural design.
In the hinge moment test, the change mode of the rudder deflection angle is one of main factors influencing the test efficiency and the data accuracy. The most common rudder deflection angle conversion mode in various wind tunnels at present adopts an angle conversion device with a preset angle, such as an angle sheet, a flange or a preset angle control surface. The angle conversion mode is reliable in connection and low in cost, but the model needs to be assembled and disassembled every time, the test efficiency is low, the test device is easy to damage, and particularly the balance is damaged. Automatic rudder deflection angle changing mechanisms are successively developed in low-speed and supersonic-speed-crossing wind tunnels at home and abroad, and generally, a crank-slider mechanism drives a connecting-rod-slider mechanism to realize deflection of a control surface. The system is complex and large in size, and cannot be applied to a hypersonic wind tunnel test model.
Therefore, the research on the design and the use method of the automatic variable test device suitable for the hypersonic wind tunnel hinge moment test is of great significance to the development of the wind tunnel test in the future. At present, relevant documents do not exist at home and abroad.
Disclosure of Invention
An object of the present invention is to solve at least the above problems and/or disadvantages and to provide at least the advantages described hereinafter.
To achieve these objects and other advantages and in accordance with the purpose of the invention, as embodied and broadly described, there is provided an automatic rudder deflection angle varying device for hypersonic wind tunnel hinge moment measurement test, comprising:
the aircraft model is provided with a horizontal reference surface and a 0-degree control rudder positioning mark line, a balance installation cavity and a motor installation cavity are communicated with each other and are arranged in the aircraft model, a motor assembly is fixedly installed in the motor installation cavity, a hinge torque balance is fixedly installed in the balance installation cavity, and the motor assembly is fixedly connected with the hinge torque balance through a balance adaptor;
the measuring rudder is fixedly connected with the hinge moment balance, and the aircraft model is also fixedly provided with a plurality of interference rudders;
and a first heat insulation sleeve assembly is arranged between the motor assembly and the aircraft model, and a second heat insulation sleeve assembly is arranged between the hinge moment balance and the aircraft model and between the hinge moment balance and the measuring rudder.
Preferably, the structure of the motor assembly includes:
the micro motor is connected with an encoder for feeding back signals of the micro motor;
a brake connected with the micro motor;
the speed reducer is connected with a motor shaft of the micro motor, and the speed reducer is connected with the balance adaptor in a flat key positioning and interference fit mode;
the motor combination body, the hinge moment balance and the measuring rudder are arranged on the same central axis.
Preferably, the inner ring of the balance adapter is connected with the motor combination body in a flat key positioning and interference fit mode, the front end of the balance adapter is fixedly provided with a mounting flange, and the mounting flange is fixedly connected with the hinge torque balance in a pin positioning and screw fastening mode;
the lower end of the measuring rudder is fixedly connected with a rotating shaft, the lower end of the rotating shaft is fixedly connected with a flange, and the flange is fixedly connected with the hinge moment balance in a pin positioning and screw fastening mode.
Preferably, the first insulating jacket assembly has a structure comprising:
the heat insulation sleeve comprises a heat insulation sleeve metal shell and a heat insulation sleeve nonmetal liner arranged in the heat insulation sleeve metal shell, wherein zirconia ceramics are sprayed on the surface of the heat insulation sleeve metal shell;
the cover plate is fixed at the upper end of the heat insulation sleeve through a screw, and comprises a heat insulation sleeve non-metal cover plate and a heat insulation sleeve metal cover plate arranged outside the heat insulation sleeve non-metal cover plate;
the motor combination body is integrally sealed in the first heat insulation sleeve assembly, the cover plate is provided with a through hole and a plurality of small holes, the speed reducing shaft of the speed reducer penetrates through the through hole, and the cover plate is fixedly connected with the first heat insulation sleeve assembly through screws.
Preferably, the cover plate is provided with a limiting block, and gaps are respectively arranged in front of and behind the limiting block and the balance adapter.
Preferably, the structure of the second heat insulating jacket assembly comprises:
the balance fixed end heat insulation sleeve is fixedly arranged at the upper end of the balance adaptor, and the hinge moment balance is arranged in the balance fixed end heat insulation sleeve;
and the balance model end heat insulation sleeve is arranged outside the balance fixed end heat insulation sleeve in a reverse buckling mode.
Preferably, the first heat insulation sleeve assembly faces the tail of the aircraft model and one side of the first heat insulation sleeve assembly, which is far away from the measuring rudder, is provided with a long hole for leading out all wires of the motor assembly.
Preferably, wherein, the tail of aircraft model is seted up and is seted up the through wires hole corresponding with rectangular hole, the through wires hole uses removable sealed glue sealed.
Preferably, the aircraft model is filled with a front-end control surface filling block and a rear-end control surface filling block, a round hole is formed at the joint of the front-end control surface filling block and the rear-end control surface filling block in a surrounding manner, and the rotating shaft of the measuring rudder is rotatably arranged in the round hole in a penetrating manner.
A use method of an automatic rudder deflection angle changing device for a hypersonic wind tunnel hinge moment measurement test comprises the following steps:
step one, mounting a motor assembly, a hinge moment balance, a measuring rudder, a first heat insulation sleeve assembly and a second heat insulation sleeve assembly on an aircraft model;
step two, assembling the aircraft model and other test devices;
step three, mounting the assembled test device on a wind tunnel attack angle mechanism;
step four, electrifying the micro motor, driving the hinge moment balance and the measuring rudder to rotate until the rudder deflection angle is 0 degrees, powering off the micro motor, electrifying the brake, and ensuring that the rudder deflection angle of the measuring rudder is not changed accidentally in the test process;
marking points are pasted on the upper surface and the lower surface of the measuring rudder, and an angle value of the measuring rudder under the no-load condition is measured and recorded by a vision measuring system of the wind tunnel;
step six, finishing the installation of other parts of the test device;
step seven, carrying out a test according to a wind tunnel operation rule, and recording the measured values of the hinge moment balance and the vision measuring system in real time in the test process;
step eight, after the test is finished, powering on the micro motor, driving the hinge moment balance and the measuring rudder to rotate to a required angle, then powering off the micro motor, and powering on the brake;
and step nine, repeating the step seven to the step eight until all the tests are finished.
The invention at least comprises the following beneficial effects: the automatic rudder deflection changing device can conveniently and quickly change the rudder deflection angle of the measuring rudder, effectively avoid the operation failure of the motor under the vacuum condition, ensure that the vacuum does not need to be released after each test and effectively improve the test efficiency.
Additional advantages, objects, and features of the invention will be set forth in part in the description which follows and in part will become apparent to those having ordinary skill in the art upon examination of the following or may be learned from practice of the invention.
Drawings
FIG. 1 is a schematic view of the exterior configuration of an aircraft model provided by the present invention;
FIG. 2 is a schematic cross-sectional view of an aircraft model provided by the present invention;
FIG. 3 is a schematic structural diagram of a motor assembly provided by the present invention;
FIG. 4 is a schematic view of a connection structure of the motor assembly, the balance and the measuring rudder provided by the invention;
FIG. 5 is a schematic cross-sectional view of a first insulating jacket assembly provided in accordance with the present invention;
FIG. 6 is a schematic cross-sectional view of a second insulating jacket assembly provided in accordance with the present invention;
FIG. 7 is a schematic diagram of a hinged moment balance according to the present invention;
fig. 8 is a schematic structural diagram of the balance adapter provided by the invention.
Detailed Description
The present invention is further described in detail below with reference to the attached drawings so that those skilled in the art can implement the invention by referring to the description text.
It will be understood that terms such as "having," "including," and "comprising," as used herein, do not preclude the presence or addition of one or more other elements or groups thereof.
It is to be understood that in the description of the present invention, the terms indicating orientation or positional relationship are based on the orientation or positional relationship shown in the drawings, and are used only for convenience in describing the present invention and for simplification of the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In the description of the present invention, unless otherwise specifically stated or limited, the terms "mounted," "disposed," "sleeved/connected," "connected," and the like are used broadly, and for example, "connected" may be a fixed connection, a detachable connection, or an integral connection, a mechanical connection, an electrical connection, a direct connection, an indirect connection via an intermediate medium, or a communication between two elements, and those skilled in the art will understand the specific meaning of the terms in the present invention specifically.
Further, in the present invention, unless otherwise explicitly specified or limited, a first feature "on" or "under" a second feature may be directly contacted with the first and second features, or indirectly contacted with the first and second features through an intermediate. Also, a first feature "on," "over," and "above" a second feature may be directly or diagonally above the second feature, or may simply indicate that the first feature is at a higher level than the second feature. A first feature being "under," "below," and "beneath" a second feature may be directly under or obliquely under the first feature, or may simply mean that the first feature is at a lesser elevation than the second feature.
As shown in fig. 1-8: the invention relates to an automatic rudder deflection angle changing device for a hypersonic wind tunnel hinge moment measurement test, which comprises:
the aircraft model comprises an aircraft model 1, a balance mounting cavity 101 and a motor mounting cavity 102 which are communicated with each other are arranged in the aircraft model 1, a motor assembly is fixedly mounted in the motor mounting cavity 101, a hinge moment balance 2 is fixedly mounted in the balance mounting cavity 102, and the motor assembly is fixedly connected with the hinge moment balance 2 through a balance adapter 3;
the measuring rudder 4 is fixedly connected with the hinge moment balance 2, and the aircraft model 1 is also fixedly provided with a plurality of interference rudders 5;
a first heat insulation sleeve assembly 6 is arranged between the motor assembly and the aircraft model 1, and a second heat insulation sleeve assembly 7 is arranged between the hinge moment balance 2 and the aircraft model 1 and between the hinge moment balance and the measuring rudder 4.
The structure of the motor assembly includes:
a micro motor 8 connected with an encoder 9 for feeding back a micro motor signal;
the speed reducer 10 is connected with a motor shaft of the micro motor 8, and the speed reducer 10 is connected with the balance adaptor 3 in a flat key positioning and interference fit mode;
a brake (not shown) interfacing with the micro-motor 8;
the motor combination body, the hinge moment balance 2 and the measuring rudder 4 are arranged on the same central axis.
The working principle is as follows: when hypersonic wind tunnel hinge moment measurement is carried out, the micro motor 8 is electrified, the micro motor 8 drives the hinge moment balance 2 and the measuring rudder 4 to rotate until the leveling plane of the hinge moment balance 2 is in a horizontal state; then the micro motor 8 is powered off, and the brake is powered on, so that the rudder deflection angle of the rudder 4 is ensured not to be changed accidentally in the test process; marking points are pasted on the upper surface and the lower surface of the measuring rudder 4, and an angle value of the measuring rudder 4 under the no-load condition is measured and recorded by a vision measuring system of the wind tunnel; the other parts of the test device are installed; carrying out a test according to a wind tunnel operation procedure, and recording the measurement values of the hinge moment balance 2 and the vision measurement system in real time in the test process; after the test is finished, the micro motor 8 is powered on, the micro motor 8 drives the hinge moment balance 2 and the measuring rudder 4 to rotate to a required angle, then the micro motor 8 is powered off, the brake is powered on, and the test is repeated. The first heat insulation sleeve assembly 6 is used for carrying out heat insulation protection on a motor combination body consisting of the encoder 9, the micro motor 8, the speed reducer 10 and the brake, and ensuring that the encoder 9, the micro motor 8, the speed reducer 10 and the brake are in a normal working state; the second heat insulation sleeve assembly 7 is used for providing heat insulation protection for the hinge moment balance 2 and ensuring that the hinge moment balance 2 obtains accurate moment and aerodynamic force measurement values.
In the technical scheme, the inner ring of the balance adapter 3 is connected with the motor assembly in a flat key positioning and interference fit mode, the front end of the balance adapter 3 is fixedly provided with the mounting flange 31, and the mounting flange 31 is fixedly connected with the hinge torque balance 2 in a pin positioning and screw fastening mode;
the lower end of the measuring rudder 4 is fixedly connected with a rotating shaft 41, the lower end of the rotating shaft 41 is fixedly connected with a flange 42, and the flange 42 is fixedly connected with the hinge moment balance 2 in a pin positioning and screw fastening mode.
In the above technical solution, the structure of the first heat insulating jacket assembly 6 includes:
the heat insulation sleeve comprises a heat insulation sleeve metal shell 61 and a heat insulation sleeve nonmetal liner 62 arranged inside the heat insulation sleeve metal shell 61, wherein zirconia ceramics are sprayed on the surface of the heat insulation sleeve metal shell 61 and are used for enhancing the heat insulation performance of the heat insulation sleeve metal shell 61;
the cover plate is fixed at the upper end of the heat insulation sleeve through a screw, and comprises a heat insulation sleeve nonmetal cover plate 63 and a heat insulation sleeve metal cover plate 64 arranged outside the heat insulation sleeve nonmetal cover plate 63;
the motor assembly is integrally sealed in the first heat insulation sleeve assembly 6, a through hole 66 and a plurality of small holes are formed in the cover plate, and a speed reducing shaft of the speed reducer penetrates through the through hole 66. The cover plate integrally seals the motor assembly, and a plurality of small holes are formed in the cover plate to ensure that the cover plate does not interfere with a speed reducer shaft; the cover plate is fixedly connected with the first heat insulation sleeve assembly 6 through screws;
in the above technical scheme, the cover plate is provided with the limiting block 11, gaps of about 0.5mm are respectively arranged in front of and behind the limiting block 11 and the balance adaptor 3, and the limiting block 11 is used for limiting the balance adaptor 3, so that the balance adaptor 3 is ensured not to deviate in the rotating process.
In the above technical solution, the structure of the second heat insulating jacket assembly 7 includes:
the balance fixed end heat insulation sleeve 71 is fixedly arranged at the upper end of the balance adaptor 3, and the hinge moment balance 2 is arranged in the balance fixed end heat insulation sleeve 71;
and the balance model end heat insulation sleeve 72 is arranged outside the balance fixed end heat insulation sleeve 71 in an inverted mode.
In the above technical solution, the first heat insulating sleeve assembly 6 faces the tail of the aircraft model 1 and is provided with a long hole 65 for leading out all wires of the motor assembly on one side away from the measuring rudder 4.
In the technical scheme, the tail of the aircraft model 1 is provided with the threading hole 103 corresponding to the strip hole 65, and the threading hole 103 is sealed by removable sealant.
In the above technical solution, the aircraft model 1 is filled with a front control surface filling block 104 and a rear control surface filling block 105, a circular hole is formed by enclosing a connection part of the front control surface filling block 104 and the rear control surface filling block 105, and the rotating shaft 41 of the measurement rudder 4 is rotatably inserted into the circular hole.
A use method of an automatic rudder deflection angle changing device for a hypersonic wind tunnel hinge moment measurement test comprises the following steps:
step one, mounting a motor assembly, a hinge moment balance, a measuring rudder, a first heat insulation sleeve assembly and a second heat insulation sleeve assembly on an aircraft model;
step two, assembling the aircraft model and other test devices;
step three, mounting the assembled test device on a wind tunnel attack angle mechanism;
step four, electrifying the micro motor, driving the hinge torque balance and the measuring rudder to rotate until the rudder deflection angle is 0 degree, powering off the micro motor, electrifying the brake and ensuring that the rudder deflection angle of the measuring rudder is not changed accidentally in the test process;
marking points are pasted on the upper surface and the lower surface of the measuring rudder, and an angle value of the measuring rudder under the no-load condition is measured and recorded by a vision measuring system of the wind tunnel;
step six, finishing the installation of other parts of the test device;
step seven, carrying out a test according to a wind tunnel operation procedure, and recording the measurement values of the hinge moment balance and the vision measurement system in real time in the test process;
step eight, after the test is finished, powering on the micro motor, driving the hinge moment balance and the measuring rudder to rotate to a required angle, then powering off the micro motor, and powering on the brake;
and step nine, repeating the step seven to the step eight until all the tests are finished.
The number of apparatuses and the scale of the process described herein are intended to simplify the description of the present invention. Applications, modifications and variations of the present invention will be apparent to those skilled in the art.
While embodiments of the invention have been described above, it is not limited to the applications set forth in the description and the embodiments, which are fully applicable in various fields of endeavor to which the invention pertains, and further modifications may readily be made by those skilled in the art, it being understood that the invention is not limited to the details shown and described herein without departing from the general concept defined by the appended claims and their equivalents.
Claims (9)
1. The utility model provides an automatic helm deflection angle changing device for hypersonic wind tunnel hinge moment measurement test which characterized in that includes:
the aircraft model is provided with a horizontal reference surface and a 0-degree control rudder positioning mark line, a balance installation cavity and a motor installation cavity are communicated with each other and are arranged in the aircraft model, a motor assembly is fixedly installed in the motor installation cavity, a hinge torque balance is fixedly installed in the balance installation cavity, and the motor assembly is fixedly connected with the hinge torque balance through a balance adaptor;
the measuring rudder is fixedly connected with the hinge moment balance, and the aircraft model is also fixedly provided with a plurality of interference rudders;
a first heat insulation sleeve assembly is arranged between the motor assembly and the aircraft model, and a second heat insulation sleeve assembly is arranged between the hinge moment balance and the aircraft model and between the hinge moment balance and the measuring rudder;
the structure of the motor assembly includes:
the micro motor is connected with an encoder for feeding back signals of the micro motor;
the speed reducer is connected with a motor shaft of the micro motor, and the speed reducer is connected with the balance adaptor in a flat key positioning and interference fit mode;
a brake connected to the micro motor;
the motor combination body, the hinge moment balance and the measuring rudder are arranged on the same central axis.
2. The automatic rudder deflection angle changing device for the hypersonic wind tunnel hinge moment measurement test according to claim 1, wherein an inner ring of the balance adapter is connected with the motor assembly in a flat key positioning and interference fit manner, a mounting flange is fixedly arranged at the front end of the balance adapter, and the mounting flange is fixedly connected with the hinge moment balance in a pin positioning and screw fastening manner;
the lower end of the measuring rudder is fixedly connected with a rotating shaft, the lower end of the rotating shaft is fixedly connected with a flange, and the flange is fixedly connected with the hinge moment balance in a pin positioning and screw fastening mode.
3. The automatic rudder deflection angle varying device for the hypersonic wind tunnel hinge moment measurement test according to claim 1, wherein the structure of the first heat insulation sleeve component comprises:
the heat insulation sleeve comprises a heat insulation sleeve metal shell and a heat insulation sleeve nonmetal liner arranged in the heat insulation sleeve metal shell, wherein zirconia ceramics are sprayed on the surface of the heat insulation sleeve metal shell;
the cover plate is fixed at the upper end of the heat insulation sleeve through a screw, and comprises a heat insulation sleeve non-metal cover plate and a heat insulation sleeve metal cover plate arranged outside the heat insulation sleeve non-metal cover plate;
the motor combination body is integrally sealed in the first heat insulation sleeve assembly, the cover plate is provided with a through hole and a plurality of small holes, the speed reducing shaft of the speed reducer penetrates through the through hole, and the cover plate is fixedly connected with the first heat insulation sleeve assembly through screws.
4. The automatic rudder deflection angle varying device for the hypersonic wind tunnel hinge moment measurement test according to claim 3, wherein a limiting block is arranged on the cover plate, and gaps are respectively arranged in front of and behind the limiting block and the balance adapter.
5. The automatic rudder deflection angle varying device for the hypersonic wind tunnel hinge moment measurement test according to claim 3, wherein the structure of the second heat insulation sleeve component comprises:
the balance fixed end heat insulation sleeve is fixedly arranged at the upper end of the balance adaptor, and the hinge moment balance is arranged in the balance fixed end heat insulation sleeve;
and the balance model end heat insulation sleeve is arranged outside the balance fixed end heat insulation sleeve in a reverse buckling mode.
6. The automatic rudder deflection angle varying device for the hypersonic wind tunnel hinge moment measurement test according to claim 3, wherein the first heat insulation sleeve component faces the tail part of the aircraft model, and one side of the first heat insulation sleeve component away from the measuring rudder is provided with a strip hole for leading out all wires of the motor assembly.
7. The automatic rudder deflection angle varying device for the hypersonic wind tunnel hinge moment measurement test according to claim 6, wherein threading holes corresponding to the strip holes are formed in the tail of the aircraft model, and the threading holes are sealed by removable sealant.
8. The automatic rudder deflection angle varying device for the hypersonic wind tunnel hinge moment measurement test according to claim 1, wherein a front end control surface filling block and a rear end control surface filling block are filled in the aircraft model, a round hole is formed by enclosing the connection part of the front end control surface filling block and the rear end control surface filling block, and a rotating shaft of the measuring rudder is rotatably arranged in the round hole in a penetrating mode.
9. The use method of the automatic rudder deflection angle changing device for the hypersonic wind tunnel hinge moment measurement test is characterized by comprising the following steps of:
step one, mounting a motor assembly, a hinge moment balance, a measuring rudder, a first heat insulation sleeve assembly and a second heat insulation sleeve assembly on an aircraft model;
step two, assembling the aircraft model and other test devices;
step three, mounting the assembled test device on a wind tunnel attack angle mechanism;
step four, electrifying the micro motor, driving the hinge moment balance and the measuring rudder to rotate until the rudder deflection angle is 0 degrees, powering off the micro motor, electrifying the brake, and ensuring that the rudder deflection angle of the measuring rudder is not changed accidentally in the test process;
marking points are pasted on the upper surface and the lower surface of the measuring rudder, and an angle value of the measuring rudder under the no-load condition is measured and recorded by a vision measuring system of the wind tunnel;
sixthly, finishing the installation of other parts of the test device;
step seven, carrying out a test according to a wind tunnel operation procedure, and recording the measurement values of the hinge moment balance and the vision measurement system in real time in the test process;
step eight, after the test is finished, powering on the micro motor, driving the hinge moment balance and the measuring rudder to rotate to a required angle, then powering off the micro motor, and powering on the brake;
and step nine, repeating the step seven to the step eight until all tests are completed.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202111289019.9A CN114001906B (en) | 2021-11-02 | 2021-11-02 | Automatic rudder deflection angle changing device for hypersonic wind tunnel hinge moment measurement test and using method thereof |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202111289019.9A CN114001906B (en) | 2021-11-02 | 2021-11-02 | Automatic rudder deflection angle changing device for hypersonic wind tunnel hinge moment measurement test and using method thereof |
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| CN115356072B (en) * | 2022-10-19 | 2023-01-13 | 中国航空工业集团公司沈阳空气动力研究所 | Automatic variable-angle aircraft wind tunnel model component load dynamic measurement device and method |
| CN115452308B (en) * | 2022-11-09 | 2023-03-14 | 中国空气动力研究与发展中心高速空气动力研究所 | Deflection angle adjustable structure for measuring control surface manipulation efficiency in wind tunnel |
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