CN110411825B - Test device for aircraft skin fatigue test - Google Patents
Test device for aircraft skin fatigue test Download PDFInfo
- Publication number
- CN110411825B CN110411825B CN201910842429.8A CN201910842429A CN110411825B CN 110411825 B CN110411825 B CN 110411825B CN 201910842429 A CN201910842429 A CN 201910842429A CN 110411825 B CN110411825 B CN 110411825B
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- China
- Prior art keywords
- clamping
- aircraft skin
- block
- test
- frame
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- 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.)
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- 238000012360 testing method Methods 0.000 title claims abstract description 33
- 238000009661 fatigue test Methods 0.000 title claims abstract description 13
- 230000005484 gravity Effects 0.000 claims abstract description 14
- 230000007246 mechanism Effects 0.000 claims abstract description 11
- 238000007906 compression Methods 0.000 claims description 19
- 230000006835 compression Effects 0.000 claims description 16
- 230000000712 assembly Effects 0.000 claims description 10
- 238000000429 assembly Methods 0.000 claims description 10
- 238000003825 pressing Methods 0.000 claims description 9
- 230000003014 reinforcing effect Effects 0.000 claims description 4
- 230000000149 penetrating effect Effects 0.000 claims description 3
- 230000009471 action Effects 0.000 abstract description 4
- 238000000034 method Methods 0.000 description 9
- 230000008569 process Effects 0.000 description 9
- 230000000694 effects Effects 0.000 description 5
- 238000005056 compaction Methods 0.000 description 4
- 230000002411 adverse Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000005489 elastic deformation Effects 0.000 description 1
- 210000001503 joint Anatomy 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000037380 skin damage Effects 0.000 description 1
- 238000005728 strengthening Methods 0.000 description 1
- 238000003466 welding Methods 0.000 description 1
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N3/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N3/02—Details
- G01N3/04—Chucks
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N3/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N3/32—Investigating strength properties of solid materials by application of mechanical stress by applying repeated or pulsating forces
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/0001—Type of application of the stress
- G01N2203/0005—Repeated or cyclic
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/0058—Kind of property studied
- G01N2203/0069—Fatigue, creep, strain-stress relations or elastic constants
- G01N2203/0073—Fatigue
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N2203/00—Investigating strength properties of solid materials by application of mechanical stress
- G01N2203/02—Details not specific for a particular testing method
- G01N2203/04—Chucks, fixtures, jaws, holders or anvils
- G01N2203/0405—Features allowing alignment between specimen and chucks
Landscapes
- Physics & Mathematics (AREA)
- Health & Medical Sciences (AREA)
- Life Sciences & Earth Sciences (AREA)
- Chemical & Material Sciences (AREA)
- Analytical Chemistry (AREA)
- Biochemistry (AREA)
- General Health & Medical Sciences (AREA)
- General Physics & Mathematics (AREA)
- Immunology (AREA)
- Pathology (AREA)
- Investigating Strength Of Materials By Application Of Mechanical Stress (AREA)
Abstract
The invention discloses a test device for an aircraft skin fatigue test, which comprises a base, a frame-shaped upright post arranged on the base, a gravity eliminating mechanism respectively connected to the upper end and the lower end of the frame-shaped upright post and matched with the aircraft skin, an upper clamping head connected above the frame-shaped upright post, a first clamping body connected to the upper clamping head and a second clamping body arranged on the base and corresponding to the first clamping body; the device has the advantages of light overall structure dead weight, large clamping force and high positioning precision, and effectively eliminates the eccentricity and vibration of the arc skin test sample piece under the action of alternating load.
Description
Technical Field
The invention relates to the technical field of aircraft skin fatigue tests, in particular to a test device for an aircraft skin fatigue test.
Background
The aircraft skin fatigue test mainly carries out the tensile and compressive test of material, and wherein needs to carry out firm centre gripping to the aircraft skin in the test process, current clamping jig structure is complicated, uses inconveniently, and overall structure is heavy, and because the peculiar arc structure of aircraft skin, centre of clamping stress is in decentration, easily leads to the aircraft skin test piece to produce the swing in the centre gripping process, and clamping force is little, and positioning accuracy is low, very big influence test effect.
Disclosure of Invention
The invention aims to provide a test device for an aircraft skin fatigue test, which aims to solve the problem that the existing clamp is poor in clamping stability of the aircraft skin.
The technical scheme for solving the technical problems is as follows: the device comprises a base, a frame-shaped upright post arranged on the base, a gravity eliminating mechanism respectively connected to the upper end and the lower end of the frame-shaped upright post and matched with an aircraft skin, a cross beam connected above the frame-shaped upright post, an upper clamping head connected on the cross beam, a first clamping body connected on the upper clamping head, a lower clamping head arranged on the base and a second clamping body connected on the lower clamping head and corresponding to the first clamping body;
The first clamp body and the second clamp body comprise a supporting plate, clamping assemblies and pressing assemblies, wherein the clamping assemblies and the pressing assemblies are arranged at two ends of the supporting plate respectively, and a gap for placing and clamping an aircraft skin is reserved between the clamping assemblies and the pressing assemblies.
Further, the clamping assembly comprises side plates arranged on the supporting plate, reinforcing ribs and clamping blocks which are respectively connected to two end faces of the side plates, and the clamping blocks are located on the end faces, close to the pressing assembly, of the side plates.
Further, the compressing assembly comprises a fixing plate, a sleeve arranged on the fixing plate, a screwing locking piece penetrating through the sleeve and a compressing block connected with the end part of the screwing locking piece, wherein the compressing block is opposite to the clamping block and is arranged at intervals.
Further, the compaction block comprises a fixing section and a compaction section connected with the fixing section, wherein the fixing section is far away from the end part of the compaction section and is connected with the screwing locking piece, and the fixing section is connected with the compaction section to form a T-shaped structure.
Further, a threaded hole matched with the screwing locking piece is formed in the pressing block, and a disc spring is arranged in the threaded hole.
Further, the gravity eliminating mechanism comprises a clamping hoop, a clamping seat, a supporting rod and a flexible supporting piece, wherein the clamping hoops are respectively arranged at the upper end and the lower end of the frame-shaped upright post, the clamping seat is arranged on the clamping hoop, the supporting rod is arranged on the clamping seat, the flexible supporting piece is arranged at the end part of the supporting rod away from the clamping seat, and the flexible supporting piece is in butt joint with the aircraft skin.
Further, the clamping block is arranged in the middle of the side plate, and a boss for the aircraft skin to lean against is formed between the clamping block and the upper part of the side plate.
Further, friction knurls are arranged on the end face, close to the clamping block, of the clamping block and the end face, close to the clamping block, of the clamping block.
Further, the sleeve is connected to the fixing plate through an inner hexagon head screw.
The invention has the following beneficial effects: the test device for the aircraft skin fatigue test has the advantages of light dead weight of the whole structure, large clamping force, high positioning precision, reliable and stable clamping and good clamping effect; the eccentricity and vibration of the arc skin test sample piece under the action of alternating load are effectively eliminated, the limited space of the fatigue testing machine is reasonably utilized through the gravity eliminating mechanism, the adverse effect on the test caused by gravity eccentricity in the test process of the aircraft skin test piece is effectively eliminated through the cooperation of the upper flexible supporting piece and the lower flexible supporting piece, meanwhile, the vibration of the test sample caused by high-frequency vibration in the test process is also limited to a certain extent, and the test effect is improved.
Drawings
FIG. 1 is a schematic diagram of the structure of the present invention;
FIG. 2 is a schematic view of the structure of a first clamp body and a second clamp body according to the present invention;
FIG. 3 is a schematic view of a clamping assembly according to the present invention;
The reference numerals shown in fig. 1 to 3 are respectively expressed as: 1-base, 2-frame-shaped upright post, 3-gravity eliminating mechanism, 4-crossbeam, 20-upper chuck, 21-lower chuck, 5-first chuck body, 6-second chuck body, 50-backup pad, 51-clamping component, 52-compression component, 510-curb plate, 511-strengthening rib, 512-grip block, 513-boss, 520-fixed plate, 521-sleeve, 522-screwing locking piece, 523-compression block, 524-fixed section, 525-compression section, 30-clampband, 31-clamping seat, 32-branch, 33-flexible support piece.
Detailed Description
The principles and features of the present invention are described below with reference to the drawings, the examples are illustrated for the purpose of illustrating the invention and are not to be construed as limiting the scope of the invention.
As shown in fig. 1 to 2, the device comprises a base 1, a frame-shaped upright 2 arranged on the base 1, a gravity eliminating mechanism 3 respectively connected to the upper end and the lower end of the frame-shaped upright 2 and matched with an aircraft skin, a cross beam 4 connected above the frame-shaped upright 2, an upper clamping head 20 connected on the cross beam 4, a first clamping head body 5 connected on the upper clamping head 20, a lower clamping head 21 arranged on the base 1 and a second clamping head body 6 connected on the lower clamping head 21 and corresponding to the first clamping head body 5;
The upper end and the lower extreme of aircraft skin are fixed respectively on the first anchor clamps body 5 and the second anchor clamps body 6, and frame shape stand is the important supporting mechanism of whole device, provides reliable support, guarantees that test process is stable and reliable to go on. The gravity eliminating mechanism 3 is mainly used for eliminating gravity eccentricity and vibration in the test process of the aircraft skin test piece, and guaranteeing stable clamping of the aircraft skin.
The first clamp body 5 and the second clamp body 6 each comprise a supporting plate 50, and a clamping assembly 51 and a pressing assembly 52 which are respectively arranged at two ends of the supporting plate 50, and a gap for placing and clamping the aircraft skin is arranged between the clamping assembly 51 and the pressing assembly 52.
As shown in fig. 3, the clamping assembly 51 includes a side plate 510 disposed on the support plate 50, and a reinforcing rib 511 and a clamping block 512 respectively connected to two end surfaces of the side plate 510, where the clamping block 512 is located on an end surface of the side plate 510 near the compressing assembly 52. The side plates 510 and the clamping blocks 512 are connected in a welding mode, reinforcing ribs 511 are uniformly distributed on the back surfaces of the side plates 510 (namely, the end surfaces of the side plates 510 far away from the clamping blocks 512), and the strength and rigidity of the clamping assembly 51 are effectively improved. And the effective stress area of the clamping assembly 51 on the clamping end face of the aircraft skin is effectively increased through the clamping blocks 512, so that the aircraft skin damage caused by stress concentration is avoided. Specifically, the clamping block 512 is disposed in the middle of the side plate 510, and a boss 513 for the aircraft skin to lean against is formed between the clamping block and the upper side of the side plate 510. The boss 513 plays a role in supporting and fastening the aircraft skin, so that the relative stability of the aircraft skin in the test process is improved.
The compressing assembly 52 includes a fixing plate 520, a sleeve 521 disposed on the fixing plate 520, a tightening locking member 522 penetrating the sleeve 521, and a compressing block 523 connected to an end of the tightening locking member, wherein the compressing block 523 is disposed opposite to and spaced apart from the clamping block 512. The sleeve 521 is attached to the fixing plate 520 by socket head screws. The tightening locking piece 522 adopts a square-head long cylindrical spherical end set screw, and the adjustment of the interval distance between the compression block 523 and the clamping block 512 is realized by locking the compression block 523 by tightening the locking piece 522, so that the aircraft skin with different size and thickness at different positions is clamped. The compression block 523 is internally provided with a threaded hole matched with the screwing locking piece 522, and the threaded hole is internally provided with a belleville spring, so that the automatic compensation of the clamping force of the compression block in the process of compressing and clamping the aircraft skin is ensured through the elastic deformation of the belleville spring, and the clamping force is improved. The compression block 523 comprises a fixed section 524 and a compression section 525 connected with the fixed section 524, wherein the end part of the fixed section 524 far away from the compression section 525 is connected with the screwing locking piece 522, and the fixed section 524 is connected with the compression section 525 to form a T-shaped structure.
In order to improve friction between the compression block 523 and the contact end of the aircraft skin test piece, friction knurls are arranged on the end face of the compression block 523 close to the clamping block 512 and the end face of the clamping block 512 close to the compression block 523. The friction coefficient of the contact surface is increased through friction of the knurled surface, the clamping force is improved, and the clamping reliability of the aircraft skin is further improved.
In order to eliminate gravity eccentricity and vibration during the test of the aircraft skin test piece, the gravity eliminating mechanism 3 comprises a clampband 30, a clamping seat 31, a supporting rod 32 and a flexible supporting piece 33, wherein the clampband 30 is arranged above and below the frame-shaped upright post 2, the clamping seat 31 is arranged on the clamping seat 30, the supporting rod 32 is arranged on the clamping seat 31, the flexible supporting piece 33 is arranged at the end part of the supporting rod 32 away from the clamping seat 31, and the flexible supporting piece 33 is rigidly connected with the aircraft skin. The flexible supporting piece 33 is a main gravity eliminating flexible supporting piece, the preferable flexible supporting piece adopts a spring plunger, and during the test, the clamping hoop 30 and the clamping seat 31 slide along the frame-shaped upright post 2 to determine the action position, so that the upper end and the lower end of the aircraft skin of the flexible supporting piece 33 are respectively abutted, and under the action of the flexible supporting piece 33, the gravity eccentricity and vibration in the test process of the aircraft skin test piece are eliminated, the clamping reliability is enhanced, and the test effect is improved.
The foregoing description of the preferred embodiments of the invention is not intended to limit the invention to the precise form disclosed, and any such modifications, equivalents, and alternatives falling within the spirit and scope of the invention are intended to be included within the scope of the invention.
Claims (4)
1. The test device for the aircraft skin fatigue test is characterized by comprising a base (1), a frame-shaped upright post (2) arranged on the base (1), a gravity eliminating mechanism (3) respectively connected to the upper end and the lower end of the frame-shaped upright post (2) and matched with the aircraft skin, a cross beam (4) connected above the frame-shaped upright post (2), an upper clamping head (20) connected on the cross beam (4), a first clamping head body (5) connected on the upper clamping head (20), a lower clamping head (21) arranged on the base (1) and a second clamping head (6) connected on the lower clamping head (21) and corresponding to the first clamping body (5);
the first clamp body (5) and the second clamp body (6) comprise a supporting plate (50), clamping assemblies (51) and pressing assemblies (52) which are respectively arranged at two ends of the supporting plate (50), and a gap for placing and clamping an aircraft skin is arranged between the clamping assemblies (51) and the pressing assemblies (52);
The clamping assembly (51) comprises a side plate (510) arranged on the supporting plate (50), and reinforcing ribs (511) and clamping blocks (512) which are respectively connected to two end faces of the side plate (510), and the clamping blocks (512) are positioned on the end face, close to the compression assembly (52), of the side plate (510);
The compressing assembly (52) comprises a fixed plate (520), a sleeve (521) arranged on the fixed plate (520), a screwing locking piece (522) penetrating through the sleeve (521) and a compressing block (523) connected to the end part of the screwing locking piece, wherein the compressing block (523) is opposite to the clamping block (512) and is arranged at intervals;
The gravity eliminating mechanism (3) comprises a clamping hoop (30) arranged at the upper end and the lower end of the frame-shaped upright post (2), a clamping seat (31) arranged on the clamping hoop (30), a supporting rod (32) arranged on the clamping seat (31) and a flexible supporting piece (33) arranged at the end part of the supporting rod (32) far away from the clamping seat (31), wherein the flexible supporting piece (33) is abutted against the aircraft skin;
Friction knurls are arranged on the end face, close to the clamping block (512), of the clamping block (523) and on the end face, close to the clamping block (523), of the clamping block (512);
the sleeve (521) is connected to the fixing plate (520) by a socket head screw.
2. The test device for aircraft skin fatigue testing according to claim 1, wherein the compression block (523) comprises a fixing section (524) and a compression section (525) connected to the fixing section (524), wherein an end of the fixing section (524) remote from the compression section (525) is connected to the screw lock (522).
3. The test device for aircraft skin fatigue test according to claim 2, wherein the compression block (523) is provided with a threaded hole which is matched with the screwing locking piece (522), and a belleville spring is arranged in the threaded hole.
4. A test device for aircraft skin fatigue testing according to claim 3, characterised in that the clamping block (512) is arranged in the middle of the side plate (510) and a boss (513) against which the aircraft skin abuts is formed between the clamping block and the side plate (510).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910842429.8A CN110411825B (en) | 2019-09-06 | 2019-09-06 | Test device for aircraft skin fatigue test |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910842429.8A CN110411825B (en) | 2019-09-06 | 2019-09-06 | Test device for aircraft skin fatigue test |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN110411825A CN110411825A (en) | 2019-11-05 |
| CN110411825B true CN110411825B (en) | 2024-04-19 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201910842429.8A Active CN110411825B (en) | 2019-09-06 | 2019-09-06 | Test device for aircraft skin fatigue test |
Country Status (1)
| Country | Link |
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| CN (1) | CN110411825B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113138072B (en) * | 2021-04-20 | 2024-02-09 | 中国飞机强度研究所 | Reinforced thin-wall curved plate static test supporting device with frame structure |
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| CN110411825A (en) | 2019-11-05 |
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