CN114199691A - Airplane fuselage wallboard strength test device - Google Patents
Airplane fuselage wallboard strength test device Download PDFInfo
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- CN114199691A CN114199691A CN202210065357.2A CN202210065357A CN114199691A CN 114199691 A CN114199691 A CN 114199691A CN 202210065357 A CN202210065357 A CN 202210065357A CN 114199691 A CN114199691 A CN 114199691A
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- 238000012360 testing method Methods 0.000 title claims abstract description 117
- 238000006073 displacement reaction Methods 0.000 claims description 38
- 229910000831 Steel Inorganic materials 0.000 claims description 35
- 239000010959 steel Substances 0.000 claims description 35
- 238000012544 monitoring process Methods 0.000 claims description 19
- 230000005484 gravity Effects 0.000 claims description 16
- 230000002159 abnormal effect Effects 0.000 claims description 6
- 238000009530 blood pressure measurement Methods 0.000 claims 1
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- 230000009286 beneficial effect Effects 0.000 abstract description 2
- 238000007789 sealing Methods 0.000 description 12
- 238000010586 diagram Methods 0.000 description 4
- 238000007689 inspection Methods 0.000 description 4
- 238000004088 simulation Methods 0.000 description 4
- 239000002131 composite material Substances 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 238000011161 development Methods 0.000 description 2
- 238000002474 experimental method Methods 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 238000011160 research Methods 0.000 description 2
- 125000004122 cyclic group Chemical group 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000007599 discharging Methods 0.000 description 1
- 238000009432 framing Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
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- 230000003068 static effect Effects 0.000 description 1
- 239000003351 stiffener Substances 0.000 description 1
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- 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
- G01N3/08—Investigating strength properties of solid materials by application of mechanical stress by applying steady tensile or compressive forces
- G01N3/10—Investigating strength properties of solid materials by application of mechanical stress by applying steady tensile or compressive forces generated by pneumatic or hydraulic pressure
- G01N3/12—Pressure testing
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- 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
- G01N2203/0014—Type of force applied
- G01N2203/0016—Tensile or compressive
- G01N2203/0019—Compressive
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- 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
- G01N2203/003—Generation of the force
- G01N2203/0042—Pneumatic or hydraulic means
- G01N2203/0044—Pneumatic means
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- 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
- G01N2203/003—Generation of the force
- G01N2203/0042—Pneumatic or hydraulic means
- G01N2203/0048—Hydraulic means
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Abstract
The invention relates to a strength test device for a wall plate of an airplane body. This aircraft fuselage wallboard strength test device is used for carrying out the strength test to the fuselage wallboard under pressurization load and hoop load combined action, includes: the loading assembly comprises a pressurizing loading assembly, a skin hoop loading assembly and a frame ring direction loading assembly, wherein the pressurizing loading assembly is used for applying pressurizing load to the wall plate of the fuselage, the skin hoop loading assembly is used for applying hoop load to the skin of the wall plate of the fuselage, the frame ring direction loading assembly is used for applying hoop load to the frame of the wall plate of the fuselage, the pressurizing loading assembly, the skin hoop loading assembly and the frame ring direction loading assembly are mutually independent, and the frame ring direction loading assembly is mutually independent for loading each frame of the wall plate of the fuselage. According to the technical scheme, the invention can achieve the following beneficial technical effects: the method can simulate the pressurizing load and apply different circumferential loads to skins and frames at different frame positions, so that the load borne by the fuselage wall panel can be simulated more truly.
Description
Technical Field
The invention relates to a strength test device for a wall plate of an airplane body.
Background
The civil aircraft fuselage is similar to the barrel section, and the fuselage wallboard is a part of the fuselage barrel section, in operation, experiences from ground to aerial and then to ground, and the fuselage wallboard bears along the annular load of fuselage and inside pressure charge load, and wherein the pressure charge load is even in whole fuselage barrel section, because the annular load size of the different positions of fuselage is different, and the relative size of the annular load that frame and covering bore respectively also can change. The quality of the static, fatigue and damage tolerance of the fuselage panel under the combined action of hoop load and internal pressurization needs to be studied.
In the prior art, various fuselage panel test devices are published at home and abroad, and typically, the fuselage panel test device of America Boeing company (US20060101921A1 and US7246527B2) and D-Box device, the fuselage panel test device of Germany IMA and a fuselage panel composite loading test device (CN104807694A) and a fuselage panel pressurizing test device (CN105388002A) of China institute of aircraft strength are disclosed. The american boeing company discloses a test apparatus (E-texture apparatus) for testing curved panels which applies a hoop load through independent actuators, but cannot adjust the relative magnitude of the hoop load applied to the frame and skin separately; the fuselage panel test device of D-Box device IMA, Germany, of Boeing, USA, cannot actively apply a circumferential load; a fuselage wall plate composite loading test device of the Chinese airplane strength research institute can apply circumferential loads, but cannot respectively adjust the circumferential loads of frames and cannot respectively adjust the relative magnitudes of the circumferential loads applied to the frames and a skin; a fuselage wallboard pressurizing test device of the Chinese airplane strength research institute can only realize the simulation of pressurizing load and can not actively apply annular load.
Disclosure of Invention
One object of the present invention is to provide an aircraft fuselage panel strength test apparatus, which can overcome the defects existing in the prior art, and can apply different circumferential loads to skins and frames at different frame positions while simulating a pressurized load, so as to more truly simulate the load borne by the fuselage panel.
The above object of the present invention is achieved by an aircraft fuselage panel strength test apparatus for performing a strength test on a fuselage panel under a combined action of a pressurizing load and a circumferential load, the aircraft fuselage panel strength test apparatus comprising: the loading assembly comprises a pressurizing loading assembly, a skin hoop loading assembly and a frame ring, wherein the pressurizing loading assembly is used for applying pressurizing load to the fuselage wall plate, the skin hoop loading assembly is used for applying hoop load to the skin of the fuselage wall plate, the frame hoop loading assembly is used for applying hoop load to the frame of the fuselage wall plate, the pressurizing loading assembly, the skin hoop loading assembly and the frame ring are mutually independent to the loading assembly, and the frame hoop loading assembly is mutually independent to the loading of each frame of the fuselage wall plate.
According to the technical scheme, the strength test device for the aircraft fuselage wall plate has the following beneficial technical effects: the method can simulate the pressurizing load and apply different circumferential loads to skins and frames at different frame positions, so that the load borne by the fuselage wall panel can be simulated more truly.
Preferably, the pressurizing and loading assembly comprises a pressure box supporting frame and a pressure box, wherein the pressure box is mounted on the pressure box supporting frame, and the pressure box comprises a pressure box side wall, a wall plate supporting piece, an inflation pipe, an exhaust pipe, a pressure measuring pipe, a pressure relief pipe, a lead pipe, a pressure box bottom plate and a cover.
Preferably, the pressurizing and loading assembly further comprises a pressurizing protection device, wherein the pressurizing protection device comprises a pressure sensor and a pressure relief valve, the pressure sensor is installed on the pressure measuring pipe, and the pressure relief valve is installed on the exhaust pipe.
Preferably, the skin circumferential loading assembly comprises a skin loading lug, a skin loading pull rod, a first group of load cells, a first group of hydraulic cylinders, a first group of bearing beams and a first group of base adjusting plates, wherein one end of the skin loading lug is connected with a skin loading point, the other end of the skin loading lug is connected with a multi-stage lever, the multi-stage lever is connected with the skin loading pull rod and one ends of the first group of load cells and the first group of hydraulic cylinders, the other end of the first group of hydraulic cylinders is hinged with the lug of the first group of base adjusting plates, and a bottom plate of the first group of base adjusting plates is connected with the first group of bearing beams.
Preferably, the frame circumferential loading assembly comprises a frame loading lug, a frame loading resistance pull rod, a frame loading power pull rod, a second group of load cells, a second group of hydraulic actuating cylinders, a second group of bearing beams, a second group of base adjusting plates, a V-shaped support lever, a middle bearing beam and a third group of base adjusting plates, wherein the frame loading lug is connected with a frame loading point, the other end of the frame loading resistance pull rod is connected with a frame supporting point of the V-shaped support lever, a base supporting point of the V-shaped support lever is hinged with a lug of the third group of base adjusting plates, a bottom plate of the third group of base adjusting plates is connected with the middle bearing beam, a hydraulic actuating cylinder loading point of the V-shaped support lever is connected with one end of the frame loading power pull rod, the other end of the frame loading power pull rod is connected with one end of the second group of load cells and one end of the second group of hydraulic actuating cylinders, the other end of the second group of hydraulic actuating cylinders is hinged with a lug of the second group of base adjusting plates, and a bottom plate of the second group of base adjusting plates is connected with the second group of bearing beams.
Preferably, the hydraulic cylinder loading point is disposed at a middle position of the V-shaped support lever, and a ratio of a distance from the hydraulic cylinder loading point to the base support point to a distance from the frame loading point to the base support point is 1: 2.
preferably, the strength testing device for the aircraft fuselage panel further comprises a test piece abnormal displacement monitoring assembly, wherein the test piece abnormal displacement monitoring assembly comprises a displacement monitoring frame, a laser displacement sensor and a displacement alarm, the displacement monitoring frame is mounted on a skin loading assembly counterweight frame, and the laser displacement sensor and the displacement alarm are mounted on the displacement monitoring frame.
Preferably, the aircraft fuselage panel strength test apparatus further comprises a counterweight assembly including a skin loading assembly counterweight subassembly, a frame loading assembly counterweight subassembly, a V-shaped support lever counterweight subassembly, and a test piece counterweight subassembly.
Preferably, the skin loading assembly counterweight assembly comprises a skin loading assembly counterweight frame, a first fixed pulley, a first steel wire rope and a first counterweight block, wherein the skin loading assembly counterweight frame is built above a first group of hydraulic actuating cylinders, the first fixed pulley is arranged on the skin loading assembly counterweight frame, one end of the first steel wire rope is connected with the gravity center position of the first group of hydraulic actuating cylinders, and the other end of the first steel wire rope bypasses the first fixed pulley to be connected with the first counterweight block, so that the skin loading assembly counterweight is realized.
Preferably, the frame loading assembly counterweight assembly comprises a frame loading assembly counterweight frame, a second fixed pulley, a second steel wire rope and a second counterweight block, wherein the frame loading assembly counterweight frame is built above the second group of hydraulic actuating cylinders, the second fixed pulley is arranged on the frame loading assembly counterweight frame, one end of the second steel wire rope is connected with the gravity center position of the second group of hydraulic actuating cylinders, and the other end of the second steel wire rope bypasses the second fixed pulley to be connected with the second counterweight block, so that the frame loading assembly counterweight is realized.
Preferably, the V-shaped supporting lever counterweight assembly includes a third fixed pulley, a third steel wire rope and a third counterweight, wherein two sets of the third fixed pulleys are respectively installed above the V-shaped supporting lever and on the left side and the right side of the lower surface of the opening cover, one end of the third steel wire rope is connected to the gravity center position of the V-shaped supporting lever, and the other end of the third steel wire rope bypasses the third fixed pulley to be connected to the third counterweight, so as to realize counterweight of the V-shaped supporting lever.
Preferably, the test piece counterweight assembly comprises a test piece counterweight frame, a fourth fixed pulley, a fourth steel wire rope and a fourth counterweight block, wherein the test piece counterweight frame is built above a test piece loading point, the fourth fixed pulley is arranged on the test piece counterweight frame, one end of the fourth steel wire rope is connected with a counterweight hanging ring positioned at the outermost side loading point of the test piece, and the other end of the fourth steel wire rope bypasses the fourth fixed pulley to be connected with the fourth counterweight block, so that the test piece counterweight is realized.
Drawings
Fig. 1 is a general schematic diagram of an aircraft fuselage panel strength testing apparatus according to an embodiment of the invention.
Fig. 2 is a schematic diagram of a test piece of the strength testing device for the aircraft fuselage panel according to an embodiment of the invention.
Fig. 3 is a side view of an aircraft fuselage panel strength testing apparatus according to an embodiment of the invention.
Fig. 4 is a schematic view of a pressurization loading assembly of the aircraft fuselage panel strength testing apparatus according to an embodiment of the present invention.
Fig. 5 is another schematic view of a pressurization loading assembly of the aircraft fuselage panel strength testing apparatus of an embodiment of the present invention.
FIG. 6 is a schematic view of a skin hoop loading assembly of an aircraft fuselage panel strength testing apparatus in accordance with an embodiment of the present invention.
Fig. 7 is a schematic view of a frame ring loading assembly of the aircraft fuselage panel strength testing apparatus according to an embodiment of the invention.
Fig. 8 is a schematic view of a counterweight assembly of an aircraft fuselage panel strength testing apparatus in accordance with an embodiment of the invention.
List of reference numerals
100: a pressurizing and loading assembly;
101: a pressure cell sidewall;
102: a pressure cell end plate;
103: reinforcing ribs;
104: a pressure cell floor;
105: a flap;
106: a sealing groove;
107: sealing the boss;
108: a through hole;
109: a rubber pad;
110: a rubber strip;
111: a piezometric tube;
112: a pressure relief pipe;
113: a lead tube;
114: a rubber hose;
115: a pressure cell support frame;
116: an inflation tube;
117: an exhaust pipe;
118: a manual inspection opening;
119: a frame loading aperture;
120: a pressure cell;
200: a skin hoop loading assembly;
201: skin loading lugs;
202: a skin loading tie rod;
203: a multi-stage lever;
204: a first set of load cells;
205: a first set of hydraulic rams;
206: a first set of base adjustment plates;
207: a first group of bearing beams;
300: a frame ring is used for loading the assembly;
301: a resistance pull rod is loaded on the frame;
302: a V-shaped support lever;
303: a third set of base adjustment plates;
304: a middle bearing beam;
305: a frame loading power link;
306: a second set of load cells;
307: a second set of hydraulic rams;
308: a second set of base adjustment plates;
309: a second group of bearing beams;
400: the abnormal displacement monitoring component of the test piece;
401: a displacement monitoring frame;
402: a laser displacement sensor;
403: a displacement alarm;
500: a counterweight assembly;
501: a skin loading assembly counterweight frame;
502: a first fixed pulley;
503: a first wire rope;
504: a first weight block;
505: a frame loading assembly counterweight frame;
506: a second fixed pulley;
507: a second wire rope;
508: a second counterweight block;
509: a third fixed pulley;
510: a third wire rope;
511: a third counterweight block;
512: a limiting frame;
513: a limiting beam;
514: a test piece counterweight frame;
515: a fourth wire rope;
516: a counterweight hoisting ring;
517: a fourth fixed pulley;
518: a fourth counterweight block;
p: a test piece;
b: framing;
s: covering a skin;
BP: a frame load point;
SP: skin load points.
Detailed Description
While specific embodiments of the invention will be described below, it should be noted that in the course of the detailed description of these embodiments, in order to provide a concise and concise description, all features of an actual implementation may not be described in detail. It should be appreciated that in the development of any such actual implementation, as in any engineering or design project, numerous implementation-specific decisions are made to achieve the developers' specific goals, such as compliance with system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be further appreciated that such a development effort might be complex and tedious, but would nevertheless be a routine undertaking of design, fabrication, and manufacture for those of ordinary skill having the benefit of this disclosure, and thus should not be construed as a complete understanding of this disclosure.
Unless otherwise defined, technical or scientific terms used in the claims and the specification should have the ordinary meaning as understood by those of ordinary skill in the art to which the invention belongs. The use of "first," "second," and similar terms in the description and claims of the present application do not denote any order, quantity, or importance, but rather the terms are used to distinguish one element from another. The terms "a" or "an," and the like, do not denote a limitation of quantity, but rather denote the presence of at least one. The word "comprise" or "comprises", and the like, means that the element or item listed before "comprises" or "comprising" covers the element or item listed after "comprising" or "comprises" and its equivalent, and does not exclude other elements or items. The terms "connected" or "coupled" and the like are not restricted to physical or mechanical connections, nor are they restricted to direct or indirect connections.
Fig. 1 is a general schematic diagram of an aircraft fuselage panel strength testing apparatus according to an embodiment of the invention. Fig. 2 is a schematic diagram of a test piece of the strength testing device for the aircraft fuselage panel according to an embodiment of the invention. Fig. 3 is a side view of an aircraft fuselage panel strength testing apparatus according to an embodiment of the invention. Fig. 4 is a schematic view of a pressurization loading assembly of the aircraft fuselage panel strength testing apparatus according to an embodiment of the present invention. Fig. 5 is another schematic view of a pressurization loading assembly of the aircraft fuselage panel strength testing apparatus of an embodiment of the present invention. FIG. 6 is a schematic view of a skin hoop loading assembly of an aircraft fuselage panel strength testing apparatus in accordance with an embodiment of the present invention. Fig. 7 is a schematic view of a frame ring loading assembly of the aircraft fuselage panel strength testing apparatus according to an embodiment of the invention. Fig. 8 is a schematic view of a counterweight assembly of an aircraft fuselage panel strength testing apparatus in accordance with an embodiment of the invention.
In some embodiments, as shown in fig. 1 to 8, an aircraft fuselage panel strength testing apparatus for performing a strength test on a fuselage panel under combined action of a pressurization load and a hoop load, the aircraft fuselage panel strength testing apparatus comprising: the loading assembly comprises a pressurization loading assembly 100, a skin annular loading assembly 200 and a frame ring direction loading assembly 300, wherein the pressurization loading assembly 100 is used for applying pressurization load to a fuselage wall panel, the skin annular loading assembly 200 is used for applying annular load to a skin S of the fuselage wall panel, the frame ring direction loading assembly 300 is used for applying annular load to a frame B of the fuselage wall panel, the pressurization loading assembly 100, the skin annular loading assembly 200 and the frame ring direction loading assembly 300 are mutually independent, and the frame ring direction loading assembly 300 is mutually independent for loading each frame B of the fuselage wall panel.
The invention realizes that the relative sizes of the annular loads applied to the frame and the skin are respectively and independently adjusted by adopting an independent loading system, the annular loads with different sizes are applied to different frames by adopting independent actuating cylinders and V-shaped supporting levers, and the simulation of the pressurizing load in the operation of the civil aircraft is realized by sealing a fuselage wall plate test piece and a pressure box to form a sealed cavity and inflating and deflating the cavity.
In some embodiments, as shown in fig. 4-5, the charge loading assembly 100 includes a pressure cell support frame 115, a pressure cell 120, and a charge protection device. The pressure cell support frame 115 includes channel steel beams, bearing beam connections, and base adjustment plates. The pressure cell side walls 101 and pressure cell end plates 102 are provided with triangular stiffeners 103 (wall plate supports) to enhance lateral stiffness, and the pressure cell floor 104 is provided with a manual access port 118 and a door 105. An inflation tube 116 and an exhaust tube 117 are mounted on the pressure cell side wall 101, and a pressure measuring tube 111, a pressure relief tube 112, and a lead tube 113 are mounted on the pressure cell end plate 102. Pressure cell side wall 101, pressure cell end plate 102, and pressure cell bottom plate 104 are welded together to form pressure cell 120. The outer ring of the manual inspection opening 118 is provided with a sealing groove 106, the outer ring of the cover 105 is provided with a sealing boss 107 and a through hole 108, a rubber pad 109 is arranged between the sealing groove 106 and the sealing boss 107, and a bolt hole is formed in the overlapping position of the manual inspection opening 118 and the cover 105 and used for placing a bolt for installing the cover 105. The contact position of the test piece P and the pressure box 120 is sealed by the rubber strip 110, and the rubber strip 110 is respectively bonded with the test piece P and the pressure box 120. A pressure measuring pipe 111 on the pressure box end plate 102 is used for installing a pressure sensor, a lead pipe 113 is used for leading out a strain gauge lead, and a pressure relief pipe 112 is used for installing an emergency pressure relief valve. A frame loading hole 119 is formed in the position, corresponding to the end of a frame P of a test piece, of the side wall 101 of the pressure box, the frame loading hole 119 is larger than the section of a frame loading resistance pull rod 301, the frame loading resistance pull rod 301 penetrates through the frame loading hole 119 to be connected with a frame loading point BP (bolt hole) at the end of the frame, the frame loading resistance pull rod 301 is sleeved in a sealing hose 114 and penetrates through the frame loading hole 119 to be connected with the frame, one end of a rubber hose 114 is connected with the frame loading resistance pull rod 301, and the other end of the rubber hose 114 is connected with the side wall 101 of the pressure box. The pressure cell 120 is mounted on the pressure cell support frame 115, the vent pipe 117 is provided with a pressure relief valve, the pressure measuring pipe 111 is provided with a pressure sensor, and the pressure relief pipe 112 is provided with an emergency pressure relief valve.
In some embodiments, as shown in fig. 6, skin hoop loading assembly 200 includes a skin loading tab 201, a skin loading tension rod 202, a first set of load cells 204, a first set of hydraulic rams 205, a first set of outrigger beams 207, a first set of base adjustment plates 206. One end of a skin loading lug 201 is connected with a skin loading point SP, the other end of the skin loading lug is connected with a multi-level lever 203, the multi-level lever 203 and a skin loading pull rod 202 are connected with a first group of load cells 204 and one end of a first group of hydraulic actuating cylinders 205, the other end of the first group of hydraulic actuating cylinders 205 is hinged with a lug of a first group of base adjusting plates 206, and a bottom plate of the first group of base adjusting plates 206 is connected with a first group of bearing beams 207. The number of stages of the multi-stage lever 203 is determined according to the number of skin loading points SP.
In some embodiments, as shown in fig. 7, the frame cyclic loading assembly 300 comprises a frame loading tab, a frame loading drag link 301, a frame loading power link 305, a second set of load cells 306, a second set of hydraulic rams 307, a second set of outrigger beams 309, a second set of base adjustment plates 308, a V-shaped support lever 302, a middle outrigger beam 304, and a third set of base adjustment plates 303. A lug (frame loading lug) of the frame loading resistance tension rod 301 is connected to a frame loading point BP (bolt hole), and the other end of the frame loading resistance tension rod 301 is connected to a frame support point (bolt hole) of the V-shaped support lever 302. The base supporting points (bolt holes) of the V-shaped supporting levers 302 are hinged with the lugs of the third group of base adjusting plates 303, and the bottom plates of the third group of base adjusting plates 303 are connected with the middle bearing beam 304. A hydraulic cylinder loading point (bolt hole) is provided at the middle position of the V-shaped support lever 302, and the ratio of the distance from the hydraulic cylinder loading point to the base support point to the distance from the frame loading point to the base support point is 1: 2. the loading point of the hydraulic actuator cylinder of the V-shaped support lever 302 is connected with one end of the frame loading power pull rod 305, the other end of the frame loading power pull rod 305 is connected with the second dynamometer 306 and one end of the second group of hydraulic actuator cylinders 307, the other end of the second group of hydraulic actuator cylinders 307 is hinged with the lug of the second group of base adjusting plate 308, and the bottom plate of the second group of base adjusting plate 308 is connected with the second group of bearing beams 309. The V-shaped support lever 302 is a truss structure and is designed using an equal strength method.
In some embodiments, as shown in fig. 8, the aircraft fuselage panel strength testing apparatus further comprises a test piece abnormal displacement monitoring assembly 400, the test piece abnormal displacement monitoring assembly 400 comprising a displacement monitoring frame 401, a laser displacement sensor 402, and a displacement alarm 403, wherein the displacement monitoring frame 401 is mounted on a skin loading assembly counterweight frame 501, and the laser displacement sensor 402 and the displacement alarm 403 are mounted on the displacement monitoring frame 401.
In some embodiments, as shown in fig. 7 and 8, the aircraft fuselage panel strength testing apparatus further comprises a weight assembly 500, the weight assembly 500 comprising a skin loading assembly weight subassembly, a frame loading assembly weight subassembly, a V-shaped support lever weight subassembly, and a trial weight subassembly.
In some embodiments, as shown in fig. 7 and 8, the skin loading assembly counterweight assembly includes a skin loading assembly counterweight frame 501, a first fixed pulley 502, a first steel cable 503, and a first counterweight block 504, wherein the skin loading assembly counterweight frame 501 is built above the first set of hydraulic rams 205, the first fixed pulley 502 is disposed on the skin loading assembly counterweight frame 501, one end of the first steel cable 503 is connected to the position of the center of gravity of the first set of hydraulic rams 205, and the other end of the first steel cable 503 is connected to the first counterweight block 504 around the first fixed pulley 502, so as to achieve skin loading assembly counterweight (i.e., to counteract the influence of gravity of the skin loading assembly during testing).
In some embodiments, as shown in fig. 7 and 8, the frame loading assembly weight subassembly includes a frame loading assembly weight frame 505, a second fixed pulley 506, a second wire rope 507, and a second weight block 508, wherein the frame loading assembly weight frame 505 is built above the second set of hydraulic rams 307, the second fixed pulley 506 is disposed on the frame loading assembly weight frame 505, one end of the second wire rope 507 is connected to the position of the center of gravity of the second set of hydraulic rams 307, and the other end of the second wire rope 507 is connected to the second weight block 508 around the second fixed pulley 506, so as to realize frame loading assembly weight balancing (i.e., to counteract the influence of gravity of the frame loading assembly during the test.
In some embodiments, as shown in fig. 7 and 8, the V-shaped support lever counterweight assembly includes a third fixed pulley 509, a third steel wire rope 510, and a third counterweight 511, wherein two sets of the third fixed pulleys 509 are respectively installed above the V-shaped support lever 302 and on the left side and the right side of the lower surface of the cover 105, one end of the third steel wire rope 510 is connected to the position of the center of gravity of the V-shaped support lever 302, and the other end of the third steel wire rope 510 is connected to the third counterweight 511 by bypassing the third fixed pulley 509, so as to realize the V-shaped support lever counterweight (i.e., to counteract the gravity effect of the V-shaped support lever during the test). A limiting frame 512 is installed on the pressure cell supporting frame 115, two limiting beams 513 are installed on the limiting frame 512, and a third balancing weight 511 is located between the two limiting beams 513.
In some embodiments, as shown in fig. 7 and 8, the test piece weight subassembly includes a test piece weight frame 514, a fourth fixed pulley 517, a fourth steel wire rope 515, and a fourth counterweight block 518, wherein the test piece weight frame 514 is built above a test piece loading point, the fourth fixed pulley 517 is arranged on the test piece weight frame 514, one end of the fourth steel wire rope 515 is connected to a weight lifting ring 516 located at an outermost loading point of the test piece P, and the other end of the fourth steel wire rope 515 is connected to the fourth counterweight block 518 around the fourth fixed pulley 517, so as to realize the test piece weight balancing (i.e., to counteract the gravity effect of the test piece during the test).
As shown in fig. 3 to 8, before the test, the pressure box supporting frame 115, the first group of carrier beams 207 and the second group of carrier beams 309 are installed on the laboratory ground rail. The pressure cell 120 is mounted on the pressure cell support frame 115. The test piece P is placed on the pressure cell 120, sealed by the rubber strip 110 at the contact position of the test piece P and the pressure cell 120, and the rubber strip 110 is bonded to the test piece P and the pressure cell 120, respectively. The gauge lead of the test piece P on the pressure cell 120 side was led out from the lead tube 113, and the lead tube 113 was sealed with a soft rubber plug. A relief valve is installed in the exhaust pipe 117, and an emergency relief valve is installed in the relief pipe 112. The test piece P clamping end is prefabricated with a loading point (bolt hole), the clamping end is placed between double lugs of the skin loading double-lug clamp 201, the bolt hole is aligned, and fastening is carried out through bolts. The other end of the skin loading double-lug clamp is connected with a skin loading pull rod 202, and the skin loading pull rod 202 is connected with a multi-stage lever 203.
As shown in fig. 6 and 8, a skin loading assembly counterweight frame 501 is built above the first set of hydraulic rams 205 to be installed, and skin loading assembly counterweight frame 501 is fixed to a laboratory ground rail. A first fixed pulley 502 is arranged on a counterweight frame 501 of the skin loading assembly, one end of a first steel wire rope 503 is connected with the gravity center position of the first group of hydraulic actuating cylinders 205, and the other end of the first steel wire rope 503 bypasses the first fixed pulley 502 to be connected with a first counterweight block 504, so that the effect of the self weight of the skin loading assembly is eliminated. The multistage lever 203 is connected with one end of the skin loading pull rod 202, the other end of the skin loading pull rod 202 is connected with the first group of load cells 204, the first group of load cells 204 are connected with one end of the first group of hydraulic actuating cylinders 205, the other end of the first group of hydraulic actuating cylinders 205 is hinged with lugs of the first group of base adjusting plates 206, and bottom plates of the first group of base adjusting plates 206 are fixed with the first group of bearing beams 207 through bolts. The number of stages of the multi-stage lever 203 is determined according to the number of load points, and the first group of hydraulic cylinders 205 is aligned with the frame B of the test piece P.
As shown in fig. 7 and 8, the frame loading resistance pull rod 301 is sleeved into the sealing hose 114, passes through the frame loading hole 119, and then the frame end is placed between two lugs of the frame loading resistance pull rod 301, aligned with the bolt hole, and fastened by the bolt. One end of the rubber hose 114 is bonded with the frame loading resistance pull rod 301 and is fastened by a hoop, and the other end of the rubber hose 114 is bonded with the inner side of the side wall 101 of the pressure box. The cover 105 is installed on the pressure box bottom plate 104 through bolts to seal the manual inspection opening 118, and a rubber pad 109 is arranged between the sealing groove 106 and the sealing boss 107, so that the sealing effect can be improved. A middle load-bearing beam 304 is mounted on the laboratory ground rail at a middle position directly below the pressure cell 120. And a third group of base adjusting plates 303 are arranged on the middle bearing beam 304, and the bottom plates of the third group of base adjusting plates 303 are fastened with the middle bearing beam 304 through bolts. The base supporting points (bolt holes) of the V-shaped supporting levers 302 are hinged with the lugs of the third group of base adjusting plates 303, and the bottom plates of the third group of base adjusting plates 303 are connected with the middle bearing beam 304. The other end (single lug) of the frame loading resistance pull rod 301 is arranged between the two lugs of the frame supporting point of the V-shaped supporting lever 302, aligned with the bolt hole and fastened through a bolt. One end of the third wire rope 511 is connected to the center of gravity hole of the V-shaped support lever 302, and the other end of the third wire rope 503 is connected to the third counter weight 511 by bypassing the third fixed pulley 509 under the cover 105, so as to eliminate the influence of the self weight of the V-shaped support lever. Install spacing frame 512 on pressure cell braced frame 115, install two spacing roof beams 513 on spacing frame 512, place third balancing weight 511 between two spacing roof beams 513, prevent that third balancing weight 511 from rocking influence experiment in the experiment. A frame loading assembly counterweight frame 505 is built above the second set of hydraulic rams 307 to be installed, and the frame loading assembly counterweight frame 505 is secured to the laboratory ground rail. One end of a second steel wire rope 507 is connected with the lug at the center of gravity of the second group of hydraulic actuating cylinders 307, and the other end of the second steel wire rope 507 is connected with a second balancing weight 508 by bypassing the second fixed pulley 506, so that the influence of the self weight of the frame loading assembly is eliminated. One end of a frame loading resistance pull rod 305 is connected with a loading point of a hydraulic actuator cylinder of the V-shaped support lever 302, the other end of the frame loading resistance pull rod 305 is connected with a second dynamometer 306 and one end of a second group of hydraulic actuator cylinders 307, the other end of the second group of hydraulic actuator cylinders 307 is hinged with a lug of a second group of base adjusting plates 308, a bottom plate of the second group of base adjusting plates 308 is fixed with a second group of bearing beams 309 through bolts, and the second group of bearing beams 309 are installed on a laboratory ground rail.
As shown in fig. 8, a displacement monitoring frame 401 is mounted on the pressure cell support frame 115, and a laser displacement sensor 402 and a displacement alarm 403 are mounted on the displacement monitoring frame 401.
The number of the actuating cylinders in the first group of hydraulic actuating cylinders and the second group of hydraulic actuating cylinders is twice that of the upper frame of the test piece P. During testing, two (left and right) hydraulic actuating cylinders corresponding to the same frame position in the first group of hydraulic actuating cylinders are connected in parallel, so that the load applied to the skin at each frame position is independently adjusted; two (left and right) hydraulic actuating cylinders corresponding to the same frame position in the second group of hydraulic actuating cylinders are connected in parallel, so that the load applied to each frame is independently adjusted; the two hydraulic actuating cylinders (left and right) corresponding to the same frame position are connected in parallel, so that the loads applied to the left side and the right side of the same frame position are equal, and the test piece P is kept in a balanced state.
During the test, the adjustment of the gas pressure in the pressure cell 120 can be realized by adjusting the volume of the gas entering the pressure cell 120 through the gas filling tube 116 and the volume of the gas discharged from the pressure cell 120 through the gas discharging tube 117. Monitoring of the gas within the pressure cell 120 is accomplished by a pressure sensor mounted on the pressure sensing tube 111. The simulation of the pressurization load on the fuselage panels is achieved by adjusting the gas pressure in the pressure cell 120. When the gas pressure in the pressure box 120 exceeds the preset limit value, the pressure relief valve is opened, and the gas is discharged until the pressure is lower than the limit value, so that the test piece P is protected. Adjusting the output force of two hydraulic actuating cylinders at different frame positions in the first group of hydraulic actuating cylinders to realize the loading of skins at different frame positions; and the output forces of the two hydraulic actuating cylinders at different frame positions in the second group of hydraulic actuating cylinders are adjusted to realize the loading of different frames. The laser displacement sensor monitors the displacement of the test piece P in real time, and when the displacement of the test piece P exceeds a preset limit value, the displacement alarm is started, and each loading assembly stops loading.
In civil aircraft operation, the fuselage panel bears not only the pressurization load, but also the skin and the frame bear the hoop load, and the hoop load that the skin and the frame of different frame positions bore is inequality moreover. The fuselage wallboard strength test device disclosed in the prior art can not realize the simulation of the pressurizing load, and simultaneously independently adjust the skin at different frame positions and the circumferential load borne by the frame. The strength test device for the fuselage wall panel of the airplane can simulate the pressurizing load and simultaneously apply different circumferential loads to skins and frames at different frame positions, so that the load borne by the fuselage wall panel can be simulated more truly.
While particular embodiments of the present invention have been described above, it will be understood by those skilled in the art that they are not intended to limit the invention, and that various modifications may be made by those skilled in the art based on the above disclosure without departing from the scope of the invention.
Claims (12)
1. The utility model provides an aircraft fuselage wallboard strength test device, its characterized in that, aircraft fuselage wallboard strength test device is used for carrying out the strength test to the fuselage wallboard under pressurization load and hoop load combined action, aircraft fuselage wallboard strength test device includes: the loading assembly comprises a pressurizing loading assembly, a skin hoop loading assembly and a frame ring, wherein the pressurizing loading assembly is used for applying pressurizing load to the fuselage wall plate, the skin hoop loading assembly is used for applying hoop load to the skin of the fuselage wall plate, the frame hoop loading assembly is used for applying hoop load to the frame of the fuselage wall plate, the pressurizing loading assembly, the skin hoop loading assembly and the frame ring are mutually independent to the loading assembly, and the frame hoop loading assembly is mutually independent to the loading of each frame of the fuselage wall plate.
2. The aircraft fuselage panel strength test apparatus of claim 1, wherein the plenum loading assembly comprises a pressure cell support frame, a pressure cell, wherein the pressure cell is mounted on the pressure cell support frame, and wherein the pressure cell comprises a pressure cell sidewall, a panel support, an inflation tube, an exhaust tube, a pressure measurement tube, a pressure relief tube, a lead tube, a pressure cell floor, and a flap.
3. The aircraft fuselage wall panel strength testing apparatus of claim 2, wherein the pressurization loading assembly further comprises a pressurization protection device, wherein the pressurization protection device comprises a pressure sensor and a pressure relief valve, the pressure sensor is mounted on the pressure measuring pipe, and the pressure relief valve is mounted on the exhaust pipe.
4. The aircraft fuselage panel strength test apparatus of claim 1, wherein the skin hoop loading assembly comprises a skin loading tab, a skin loading pull rod, a first set of load cells, a first set of hydraulic cylinders, a first set of load-bearing beams, and a first set of base adjustment plates, wherein one end of the skin loading tab is connected to a skin loading point, the other end of the skin loading tab is connected to a multi-level lever, the multi-level lever is connected to the skin loading pull rod and connected to the first set of load cells and one end of the first set of hydraulic cylinders, the other end of the first set of hydraulic cylinders is hinged to tabs of the first set of base adjustment plates, and a base plate of the first set of base adjustment plates is connected to the first set of load-bearing beams.
5. The aircraft fuselage panel strength test apparatus of claim 1, wherein the frame hoop loading assembly comprises a frame loading lug, a frame loading drag pull rod, a frame loading power pull rod, a second dynamometer, a second hydraulic cylinder, a second force bearing beam, a second base adjustment plate, a V-shaped support lever, a middle force bearing beam, and a third base adjustment plate, wherein the frame loading lug is connected to a frame loading point, the other end of the frame loading drag pull rod is connected to a frame support point of the V-shaped support lever, a base support point of the V-shaped support lever is hinged to a lug of the third base adjustment plate, a bottom plate of the third base adjustment plate is connected to the middle force bearing beam, a hydraulic cylinder loading point of the V-shaped support lever is connected to one end of the frame loading power pull rod, and the other end of the frame loading power pull rod is connected to one of the second dynamometer and the second hydraulic cylinder And the other end of the second group of hydraulic actuating cylinders is hinged with a lug of the second group of base adjusting plates, and a bottom plate of the second group of base adjusting plates is connected with the second group of bearing beams.
6. An aircraft fuselage panel strength test apparatus as claimed in claim 5, wherein the hydraulic ram load point is located at an intermediate position of the V-shaped support lever, and the ratio of the distance from the hydraulic ram load point to the base support point to the distance from the frame load point to the base support point is 1: 2.
7. the aircraft fuselage panel strength test apparatus of claim 1, further comprising a test piece abnormal displacement monitoring assembly comprising a displacement monitoring frame, a laser displacement sensor, and a displacement alarm, wherein the displacement monitoring frame is mounted on a skin loading assembly counterweight frame, and the laser displacement sensor and the displacement alarm are mounted on the displacement monitoring frame.
8. The aircraft fuselage panel strength test apparatus of claim 1, further comprising a weight assembly comprising a skin loading assembly weight subassembly, a frame loading assembly weight subassembly, a V-shaped support lever weight subassembly, and a test piece weight subassembly.
9. The aircraft fuselage panel strength test apparatus of claim 8, wherein the skin loading assembly counterweight assembly comprises a skin loading assembly counterweight frame, a first fixed pulley, a first steel wire rope, and a first counterweight, wherein the skin loading assembly counterweight frame is built above the first set of hydraulic cylinders, the first fixed pulley is arranged on the skin loading assembly counterweight frame, one end of the first steel wire rope is connected with the center of gravity of the first set of hydraulic cylinders, and the other end of the first steel wire rope bypasses the first fixed pulley and is connected with the first counterweight, so as to realize skin loading assembly counterweight.
10. The aircraft fuselage panel strength test apparatus of claim 8, wherein the frame loading assembly counterweight assembly comprises a frame loading assembly counterweight frame, a second fixed pulley, a second wire rope, and a second counterweight, wherein the frame loading assembly counterweight frame is built above the second set of hydraulic rams, the second fixed pulley is disposed on the frame loading assembly counterweight frame, one end of the second wire rope is connected to a center of gravity position of the second set of hydraulic rams, and the other end of the second wire rope is connected to the second counterweight by bypassing the second fixed pulley, thereby realizing frame loading assembly counterweight.
11. The strength testing device for the aircraft fuselage panel as defined in claim 8, wherein the V-shaped supporting lever counterweight subassembly comprises a third fixed pulley, a third steel wire rope and a third counterweight, wherein two sets of the third fixed pulleys are respectively installed above the V-shaped supporting lever and on the left side and the right side of the lower surface of the opening cover, one end of the third steel wire rope is connected with the gravity center position of the V-shaped supporting lever, and the other end of the third steel wire rope bypasses the third fixed pulley to be connected with the third counterweight, so as to realize the counterweight of the V-shaped supporting lever.
12. The strength testing device for the aircraft fuselage panel as defined in claim 8, wherein the test piece counterweight assembly comprises a test piece counterweight frame, a fourth fixed pulley, a fourth steel wire rope and a fourth counterweight block, wherein the test piece counterweight frame is built above a test piece loading point, the fourth fixed pulley is arranged on the test piece counterweight frame, one end of the fourth steel wire rope is connected with a counterweight hanging ring positioned at an outermost loading point of the test piece, and the other end of the fourth steel wire rope bypasses the fourth fixed pulley and is connected with the fourth counterweight block to realize counterweight of the test piece.
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| CN114778168A (en) * | 2022-06-17 | 2022-07-22 | 中国飞机强度研究所 | Method for determining loading stage number of breaking load in spacecraft cabin section ground strength test |
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