CN112504590A - Clamp for vibration fatigue test of aero-engine blade - Google Patents
Clamp for vibration fatigue test of aero-engine blade Download PDFInfo
- Publication number
- CN112504590A CN112504590A CN202011226624.7A CN202011226624A CN112504590A CN 112504590 A CN112504590 A CN 112504590A CN 202011226624 A CN202011226624 A CN 202011226624A CN 112504590 A CN112504590 A CN 112504590A
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- blade
- bolt
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- fatigue test
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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
- G01M7/00—Vibration-testing of structures; Shock-testing of structures
- G01M7/02—Vibration-testing by means of a shake table
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B25—HAND TOOLS; PORTABLE POWER-DRIVEN TOOLS; MANIPULATORS
- B25B—TOOLS OR BENCH DEVICES NOT OTHERWISE PROVIDED FOR, FOR FASTENING, CONNECTING, DISENGAGING OR HOLDING
- B25B11/00—Work holders not covered by any preceding group in the subclass, e.g. magnetic work holders, vacuum work holders
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Investigating Strength Of Materials By Application Of Mechanical Stress (AREA)
- Testing Of Devices, Machine Parts, Or Other Structures Thereof (AREA)
Abstract
The invention relates to a clamp for an aeroengine blade vibration fatigue test, which comprises a clamp (2), a supporting frame (3), a pressure sensor (5), a force application bolt (6), a stepping motor (7), a laser displacement sensor (8) and a vibration controller (9), wherein: anchor clamps (2) are used for centre gripping blade (1), on anchor clamps (2) are fixed in the shaking table excitation plane through braced frame (3), step motor (7) drive application of force bolt (6) rotatory and apply pressure in anchor clamps (2) in order to press from both sides tight blade (1), be provided with pressure sensor (5) between application of force bolt (6) and anchor clamps (2) with the holding down force of synchronous survey application of force bolt (6), laser displacement sensor (8) are arranged in measuring the amplitude of vibration fatigue test in-process blade (1). The clamp can quantitatively give the clamping force of the blade while realizing reliable clamping on the root part of the blade, realizes real-time monitoring on the clamping force of the blade, avoids test invalidity caused by overlarge or undersize of the clamping force of the blade, obviously improves the clamping stability of blade vibration fatigue and improves the reliability of the blade vibration fatigue test.
Description
Technical Field
The invention discloses a multifunctional clamp for an aircraft engine blade vibration fatigue test, and belongs to the technical field of mechanical property test characterization.
Background
Vibration is the most common form of loading during service of an aircraft engine blade. The fatigue failure problem caused by vibration must be considered in the blade strength design process, and the service safety and reliability of the blade can be effectively improved. The vibration fatigue test of the blade of the aero-engine and the performance data acquisition are not only necessary examination links in the blade development and production process, but also important means for blade strength design and service life prediction.
The blade vibration fatigue tests are subject to long test periods (typically lasting days or even weeks) and high loading frequencies (typically requiring hundreds to thousands of hertz), thus placing higher demands on the reliability of the blade grip. Since the vibration fatigue test of the blade is usually performed in a resonance state, the clamping looseness in the test process can obviously reduce the resonance frequency of the test system, and the invalidity of the vibration fatigue test data of the blade is directly caused. In order to avoid looseness in the test process, in the test preparation stage, when the blade is clamped, a tester usually improves the clamping force as much as possible so as to ensure the clamping reliability, and the excessive clamping force can cause the clamping end face of the blade to generate initial plastic deformation or even extrusion damage, and the initial damage is easy to induce macrocracks in the vibration fatigue test process, so that the clamping end of the blade is broken, and the test failure can also be caused. In conclusion, the requirements of the vibration fatigue test on clamping are moderate in size, damage and damage to the clamping end part of the blade can not be caused, and obvious looseness can not occur in the test process. However, the conventional clamping device and clamping method for the blade cannot completely meet the clamping reliability requirement, and failure of the blade vibration fatigue test due to the clamping problem sometimes occurs. Therefore, a need exists to provide a novel clamp for blade vibration fatigue test, so as to realize reliable clamping of the blade in the vibration fatigue test process.
Disclosure of Invention
The invention provides a clamp for an aeroengine blade vibration fatigue test, which is designed aiming at the defects of the clamp for the blade vibration fatigue test, and aims to realize quantitative monitoring, adjustability and controllability of clamping force in the whole flow of the blade vibration fatigue test and ensure the reliability of blade clamping in the vibration fatigue test process.
The technical solution of the invention is as follows:
the clamp for the aeroengine blade vibration fatigue test comprises a clamp 2, a supporting frame 3, a pressure sensor 5, a force application bolt 6, a stepping motor 7, a laser displacement sensor 8 and a vibration controller 9, wherein: the fixture 2 is used for clamping the blade 1, the fixture 2 is fixed on an excitation plane of the vibration table through the supporting frame 3, the stepping motor 7 is installed on a vertical lifting platform, the stepping motor 7 drives the force application bolt 6 to rotate and apply pressure to the fixture 2 so as to clamp the blade 1, the pressure sensor 5 is arranged between the force application bolt 6 and the fixture 2 so as to synchronously measure the downward pressure of the force application bolt 6, and the stepping motor 7 is provided with a torque measurement and amplification device so as to measure the torque output value. The laser displacement sensor 8 is used for measuring the amplitude of the blade 1 in the vibration fatigue test process. The vibration controller 9 is used for controlling the fastening torque output and vibration of the clamp in the test process.
In the implementation, the clamp 2 is in a fish-mouth-shaped structure, mortises 21 are formed on two inner sides of a front opening end for inserting the blade 1 and are matched with the shape of a tenon of the blade of the aircraft engine, a bolt hole 22 is formed at the rear end of the clamp 2, and the horizontal fastening bolt 4 is installed in the bolt hole 22 and is pressed and tightened from the bottom of the blade 1.
Further, a positioning hole 23 is formed in the side surface of the jig 2, the pressure sensor 5 is mounted in the positioning hole 23, and the force application bolt 6 applies force to press the jig 2 and the blade 1 and simultaneously acts on the pressure sensor 5 to measure pressure.
In implementation, the force application bolt 6 is in a hexagon bolt structure, a stud 61 at the bottom of the force application bolt is matched with and installed in the bolt hole 32 of the support frame 3, and a round hole 62 is machined at the top of the force application bolt and coaxially assembled with a torque output shaft of the stepping motor 7. The top of the force application bolt 6 is provided with a rectangular groove 63 matched with a rectangular boss on the torque output shaft of the stepping motor 7, and when torque is applied, the torque output shaft of the stepping motor 7 applies force through the mutual contact of the rectangular boss and the rectangular groove 63 to push the force application bolt 6 to load or unload.
The invention has the following characteristics and beneficial effects:
the multifunctional clamp for the vibration fatigue test of the aero-engine blade can realize reliable clamping of the root of the blade and quantitatively give the clamping force of the blade. The blade clamping force real-time monitoring is realized, the optimized value of the blade clamping force can be determined based on the function, the blade clamping force is always stabilized near the optimized value in the vibration fatigue test process, the test invalidity caused by overlarge or undersize of the blade clamping force is avoided, the clamping stability of the blade vibration fatigue is obviously improved, and the reliability of the blade vibration fatigue test is improved.
Drawings
FIG. 1A schematic view of the construction and assembly of the clamp of the present invention
FIG. 2 blade grip schematic
FIG. 3 schematic view of a support frame
FIG. 4 is a schematic view of a force applying bolt
FIG. 5 determination of the critical pressure value PcrSchematic illustration of
Detailed Description
The technical scheme of the invention is further detailed in the following by combining the drawings and the embodiment:
referring to the attached drawings 1 to 4, the clamp for the vibration fatigue test of the blade of the aircraft engine comprises a clamp 2, a supporting frame 3, a pressure sensor 5, a force application bolt 6, a stepping motor 7, a laser displacement sensor 8 and a vibration control instrument 9, wherein: the fixture 2 is used for clamping the blade 1, the fixture 2 is fixed on an excitation plane of the vibration table through the supporting frame 3, the stepping motor 7 is installed on a vertical lifting platform, the stepping motor 7 drives the force application bolt 6 to rotate and apply pressure to the fixture 2 so as to clamp the blade 1, the pressure sensor 5 is arranged between the force application bolt 6 and the fixture 2 so as to synchronously measure the downward pressure of the force application bolt 6, the stepping motor 7 is provided with a torque measurement and amplification device so as to be capable of measuring the torque output value, and the laser displacement sensor 8 is used for measuring the amplitude of the blade 1 in the vibration fatigue test process. The vibration control instrument 9 is used for the fastening torque output control and the vibration control of the clamp in the test process.
The fixture 2 is of a fish-mouth-shaped structure, mortises 21 are machined on two inner sides of a front opening end used for inserting the blade 1 and are matched with tenons of the blades of the aero-engine, bolt holes 22 are machined in the rear end of the fixture 2, and horizontal fastening bolts 4 are installed in the bolt holes 22 and are tightly jacked from the bottom of the blade 1 through force application. A positioning hole 23 is processed on the side surface of the clamp 2, the pressure sensor 5 is installed in the positioning hole 23, and the force application bolt 6 applies force to press the clamp 2 and the blade 1 and simultaneously acts on the pressure sensor 5 to measure pressure.
The force application bolt 6 is of a hexagon bolt structure, a stud 61 at the bottom of the force application bolt is matched with and installed in the bolt hole 32 of the support frame 3, and a round hole 62 is machined at the top of the force application bolt and coaxially assembled with a torque output shaft of the stepping motor 7. The top of the force application bolt 6 is provided with a rectangular groove 63 matched with a rectangular boss on the torque output shaft of the stepping motor 7, and when torque is applied, the torque output shaft of the stepping motor 7 applies force through the mutual contact of the rectangular boss and the rectangular groove 63 to push the force application bolt 6 to load or unload.
Before the vibration fatigue test is carried out, the vibration controller 9 gives an instruction, the stepping motor 7 starts to work and outputs torque, and the torque output shaft 71 rotates along with the rotation. The rectangular boss 72 on the torque output shaft 71 is in close contact with the rectangular groove 63 on the force application bolt 6, drives the force application bolt 6 to rotate, and presses the pressure sensor 5, the clamp 2 and the blade 1 tightly. The pressure sensor 5 feeds back the pressure actual measurement signal to the vibration controller 9 in real time to form closed-loop control with the stepping motor 7, so that the control of the pre-tightening force of the vibration fatigue test is accurate.
When the vibration fatigue test is carried out, the vibration controller 9 sends out an instruction to control the output torque of the stepping motor 7 to be 0, start the vibration fatigue test program and monitor the output value P of the pressure sensor 5 in the whole process.
The initial resonance frequency is first determined by a frequency sweep and subsequently vibrated at the initial resonance frequency. Controlling the stepping motor, gradually increasing the pressure of the fixture block, synchronously giving a pressure-resonance frequency P-f curve relation through a closed-loop control system, stopping the test when the P-f curve forms an obvious platform area, and determining the minimum stress value in the platform area of the P-f curve as a critical pressure value Pcr. As shown in fig. 5.
Inputting a critical pressure value P into a test control system in the test processcrMonitoring the pressure value P of the output of the pressure sensor 5 during the vibration fatigue test when
In the time, the test control system triggers the starting stepping motor to adjust the force application bolt until the force application bolt is adjusted
The step motor 7 stops operating and makes the output torque 0.
The stepping motor 7 can be installed on a platform which can vertically lift, the platform can correspondingly lift along with the rotation of the motor, and the boss 72 on the stepping motor 7 and the rectangular groove 63 of the force application bolt 6 are kept fixed in relative position.
When the influence of the pretightening force of the horizontal fastening bolt 4 on the clamping state is researched, a set of combination of the stepping motor 7, the force application bolt 6 and the sensor 5 can be added to replace the horizontal fastening bolt, and the critical pressure value P is determined according to the programcr。
Claims (8)
1. The utility model provides a anchor clamps for aeroengine blade vibration fatigue test which characterized in that: this anchor clamps include anchor clamps (2), braced frame (3), pressure sensor (5), application of force bolt (6), step motor (7), laser displacement sensor (8) and vibration control appearance (9), wherein: anchor clamps (2) are used for centre gripping blade (1), on anchor clamps (2) are fixed in the shaking table excitation plane through braced frame (3), step motor (7) drive application of force bolt (6) rotatory and apply pressure in anchor clamps (2) in order to press from both sides tight blade (1), be provided with pressure sensor (5) between application of force bolt (6) and anchor clamps (2) with the holding down force of synchronous survey application of force bolt (6), laser displacement sensor (8) are arranged in measuring the amplitude of vibration fatigue test in-process blade (1).
2. The clamp for the vibration fatigue test of the blades of the aircraft engine according to claim 1, wherein: the vibration control instrument (9) is used for controlling the fastening torque output and vibration of the clamp in the test process.
3. The clamp for the vibration fatigue test of the blades of the aircraft engine according to claim 1, wherein: the fixture (2) is of a fish-mouth-shaped structure, mortises (21) are machined on two inner sides of a front opening end used for inserting the blade (1) and are matched with the tenon shape of the blade of the aircraft engine, bolt holes (22) are machined at the rear end of the fixture (2), and horizontal fastening bolts (4) are installed into the bolt holes (22) and are pushed tightly from the bottom of the blade (1) through force application.
4. The clamp for the vibration fatigue test of the blades of the aeroengine according to the claims 1 and 3, wherein: a positioning hole (23) is machined in the side face of the clamp (2), the pressure sensor (5) is installed in the positioning hole (23), and the force application bolt (6) applies force to press the clamp (2) and the blade (1) and simultaneously acts on the pressure sensor (5) to measure pressure.
5. The clamp for the vibration fatigue test of the blades of the aircraft engine according to claim 1, wherein: the force application bolt (6) is of a hexagon bolt structure, a stud (61) at the bottom of the force application bolt is matched with and installed in a bolt hole (32) of the supporting frame (3), and a round hole (62) is machined at the top of the force application bolt and coaxially assembled with a torque output shaft of the stepping motor (7).
6. The clamp for the vibration fatigue test of the blades of the aircraft engine according to claim 5, wherein: the top of the force application bolt (6) is provided with a rectangular groove (63) matched with a rectangular boss on a torque output shaft of the stepping motor (7), and when torque is applied, the torque output shaft of the stepping motor (7) applies force through mutual contact of the rectangular boss and the rectangular groove (63) to push the force application bolt (6) to load or unload.
7. The clamp for the vibration fatigue test of the blades of the aircraft engine according to claim 1, wherein: the stepping motor (7) is provided with a torque measuring and amplifying device so as to measure the magnitude of a torque output value.
8. The clamp for the vibration fatigue test of the blades of the aircraft engine according to claim 1, wherein: the stepping motor (7) is arranged on a platform which can vertically lift.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011226624.7A CN112504590A (en) | 2020-11-05 | 2020-11-05 | Clamp for vibration fatigue test of aero-engine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011226624.7A CN112504590A (en) | 2020-11-05 | 2020-11-05 | Clamp for vibration fatigue test of aero-engine blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN112504590A true CN112504590A (en) | 2021-03-16 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011226624.7A Pending CN112504590A (en) | 2020-11-05 | 2020-11-05 | Clamp for vibration fatigue test of aero-engine blade |
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| CN (1) | CN112504590A (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113483977A (en) * | 2021-06-28 | 2021-10-08 | 北京强度环境研究所 | Acoustic characteristic testing device for light and thin structure |
| CN114659740A (en) * | 2022-04-11 | 2022-06-24 | 河南理工大学 | High-frequency vibration fatigue test device for thin-wall blade parts |
| CN115096563A (en) * | 2022-06-10 | 2022-09-23 | 中国航发北京航空材料研究院 | Composite material simulation part vibration fatigue test device and test method |
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| CN110514378A (en) * | 2019-08-30 | 2019-11-29 | 中国航发动力股份有限公司 | A kind of engine band convex shoulder fan blade vibration fatigue test device |
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2020
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Patent Citations (8)
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| CN102840968A (en) * | 2012-07-31 | 2012-12-26 | 沈阳黎明航空发动机(集团)有限责任公司 | Detection device and detection method for wide-range vibration amplitude of blade of aviation engine |
| EP2741068A1 (en) * | 2012-12-05 | 2014-06-11 | Industrieanlagen-Betriebsgesellschaft mbH | Test bench for a rotor blade, assembly with such a test bench and a method for operating such a test bench |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113483977A (en) * | 2021-06-28 | 2021-10-08 | 北京强度环境研究所 | Acoustic characteristic testing device for light and thin structure |
| CN114659740A (en) * | 2022-04-11 | 2022-06-24 | 河南理工大学 | High-frequency vibration fatigue test device for thin-wall blade parts |
| CN114659740B (en) * | 2022-04-11 | 2023-05-12 | 河南理工大学 | High-frequency vibration fatigue test device for thin-wall blade parts |
| CN115096563A (en) * | 2022-06-10 | 2022-09-23 | 中国航发北京航空材料研究院 | Composite material simulation part vibration fatigue test device and test method |
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Application publication date: 20210316 |
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