CN114199691B - Aircraft fuselage panel strength test device - Google Patents
Aircraft fuselage panel strength test device Download PDFInfo
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- CN114199691B CN114199691B CN202210065357.2A CN202210065357A CN114199691B CN 114199691 B CN114199691 B CN 114199691B CN 202210065357 A CN202210065357 A CN 202210065357A CN 114199691 B CN114199691 B CN 114199691B
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- 238000012360 testing method Methods 0.000 title claims abstract description 115
- 238000006073 displacement reaction Methods 0.000 claims description 38
- 238000012544 monitoring process Methods 0.000 claims description 19
- 230000005484 gravity Effects 0.000 claims description 13
- 230000002159 abnormal effect Effects 0.000 claims description 6
- 238000009530 blood pressure measurement Methods 0.000 claims 1
- 230000000694 effects Effects 0.000 abstract description 5
- 238000004088 simulation Methods 0.000 abstract description 5
- 230000009286 beneficial effect Effects 0.000 abstract description 2
- 229910000831 Steel Inorganic materials 0.000 description 17
- 239000010959 steel Substances 0.000 description 17
- 238000007789 sealing Methods 0.000 description 13
- 238000007689 inspection Methods 0.000 description 5
- 210000005069 ears Anatomy 0.000 description 3
- 230000001105 regulatory effect Effects 0.000 description 3
- 239000002131 composite material Substances 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 230000003014 reinforcing effect Effects 0.000 description 2
- 230000007547 defect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000000149 penetrating effect Effects 0.000 description 1
- 230000003068 static effect Effects 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
-
- 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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- 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 relates to an aircraft fuselage panel strength test device. This aircraft fuselage wallboard strength test device is used for carrying out strength test to the fuselage wallboard under filling load and hoop load combined action, includes: the frame annular loading assembly is used for applying annular load to frames of the fuselage panel, and the pressurizing loading assembly, the skin annular loading assembly and the frame annular loading assembly are mutually independent, and the frame annular loading assembly is mutually independent for loading each frame of the fuselage panel. According to the technical scheme, the invention has the following beneficial technical effects: the simulation device can simulate the pressurizing load and simultaneously apply different circumferential loads to skins and frames at different frame positions, so that the load born by the wallboard of the airframe can be simulated more truly.
Description
Technical Field
The invention relates to an aircraft fuselage panel strength test device.
Background
The civil aircraft fuselage is approximately a barrel section, the fuselage wall plate is a part of the barrel section of the fuselage, in operation, the fuselage wall plate is subjected to loads and internal pressurizing loads along the circumferential direction of the fuselage from the ground to the air and then to the ground, wherein the pressurizing loads in the whole barrel section of the fuselage are uniform, the circumferential loads at different positions of the fuselage are different, and the relative magnitudes of the circumferential loads respectively born by the frame and the skin are also changed. The quality of static force, fatigue and damage tolerance of the fuselage panel under the combined action of circumferential load and internal pressurization needs to be studied.
In the prior art, various fuselage panel test devices are disclosed at home and abroad, and are typically E-field devices (US 20060101921A1 and US7246527B 2) and D-Box devices of the Boeing company, a fuselage panel test device of Germany IMA and a fuselage panel composite loading test device (CN 104807694A) and a fuselage panel pressurizing test device (CN 105388002A) of the China aircraft strength institute. U.S. Boeing discloses a test apparatus (E-field apparatus) for testing curved wall panels that applies hoop loads via independent rams, but that cannot adjust the relative magnitudes of the hoop loads applied to the frame and skin, respectively; the body panel test device of the D-Box device of boeing, germany IMA, was unable to actively apply hoop loads; the composite loading test device for the fuselage panel of the China aircraft strength institute can apply the circumferential load, but the circumferential load of each frame cannot be respectively adjusted, and the relative sizes of the circumferential loads applied to the frames and the skin cannot be respectively adjusted; a fuselage wallboard pressurization test device of China aircraft strength institute can only simulate pressurization load, and cannot actively apply circumferential load.
Disclosure of Invention
An 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 pressurizing load, so as to more truly simulate the load applied to the fuselage panel.
The above object of the present invention is achieved by an aircraft fuselage skin strength test apparatus for strength testing of a fuselage skin under a combination of a charging load and a hoop load, the apparatus comprising: the frame annular loading assembly is used for applying annular load to frames of the fuselage panel, and the pressurizing loading assembly, the skin annular loading assembly and the frame annular loading assembly are mutually independent, and the frame annular loading assembly is mutually independent for loading each frame of the fuselage panel.
According to the technical scheme, the strength test device for the aircraft fuselage wall plate has the following beneficial technical effects: the simulation device can simulate the pressurizing load and simultaneously apply different circumferential loads to skins and frames at different frame positions, so that the load born by the wallboard of the airframe can be simulated more truly.
Preferably, the pressurizing loading assembly comprises a pressure box supporting frame and a pressure box, wherein the pressure box is installed on the pressure box supporting frame and comprises a pressure box side wall, a wall plate supporting piece, an inflation tube, an exhaust tube, a pressure measuring tube, a pressure relief tube, a guide tube, a pressure box bottom plate and a cover.
Preferably, the pressurizing 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 dynamometers, a first group of hydraulic actuating cylinders, a first group of load 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, the first group of dynamometers and one end of the first group of hydraulic actuating cylinders are connected, the other end of the first group of hydraulic actuating cylinders is hinged with the lugs of the first group of base adjusting plates, and the bottom plate of the first group of base adjusting plates is connected with the first group of load 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 dynamometers, a second group of hydraulic actuating cylinders, a second group of bearing beams, a second group of base adjusting plates, a V-shaped supporting lever, an intermediate 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 supporting lever, a base supporting point of the V-shaped supporting 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 intermediate bearing beams, a hydraulic actuating cylinder loading point of the V-shaped supporting 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 the second group of dynamometers 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 the bottom plate of the second group of base adjusting plates is connected with the second bearing beams.
Preferably, the hydraulic actuator loading point is arranged at the middle position of the V-shaped supporting lever, and the ratio of the distance from the hydraulic actuator loading point to the base supporting point to the distance from the frame loading point to the base supporting point is 1:2.
preferably, the aircraft fuselage panel strength test device 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 arranged on a counterweight frame of the skin loading assembly, and the laser displacement sensor and the displacement alarm are arranged on the displacement monitoring frame.
Preferably, the aircraft fuselage panel strength test apparatus further comprises a weight assembly, wherein the weight assembly comprises 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.
Preferably, the skin loading assembly counterweight subassembly 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 a first group 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 gravity center position of the first group of hydraulic cylinders, and the other end of the first steel wire rope bypasses the first fixed pulley to be connected with the first counterweight, so that the skin loading assembly counterweight is realized.
Preferably, the frame loading assembly counterweight sub-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 actuators, 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 actuators, 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 subassembly comprises a third fixed pulley, a third steel wire rope and a third counterweight block, wherein two groups of the third fixed pulleys are respectively arranged above the V-shaped supporting lever, on the left side and on the right side of the lower surface of the 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 block, so that the V-shaped supporting lever counterweight is realized.
Preferably, the test piece counterweight subassembly 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 lifting ring positioned at the loading point on the outermost side 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 of an aircraft fuselage panel strength test apparatus according to one embodiment of the present invention.
FIG. 2 is a schematic view of a test piece for an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention.
FIG. 3 is a side view of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention.
FIG. 4 is a schematic view of a pressurized loading assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention.
FIG. 5 is another schematic view of a pressurized loading assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention.
FIG. 6 is a schematic illustration of a skin hoop load assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention.
FIG. 7 is a schematic view of a frame hoop load assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention.
FIG. 8 is a schematic view of a counterweight assembly of an aircraft fuselage panel strength test 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 bottom plate;
105: a mouth cover;
106: sealing grooves;
107: sealing the boss;
108: a through hole;
109: a rubber pad;
110: a rubber strip;
111: a pressure measuring tube;
112: a pressure relief tube;
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 port;
119: a frame loading hole;
120: a pressure cell;
200: the skin annular loading assembly;
201: skin loading lugs;
202: skin loading pull rods;
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 set of load beams;
300: a frame hoop loading assembly;
301: a resistance pull rod is loaded on the frame;
302: v-shaped supporting lever;
303: a third set of base adjustment plates;
304: a middle load beam;
305: a frame loading power connecting rod;
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 set of load beams;
400: an abnormal displacement monitoring assembly for 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 weight frame;
502: a first fixed pulley;
503: a first wire rope;
504: a first balancing weight;
505: a frame loading assembly counterweight frame;
506: a second fixed pulley;
507: a second wire rope;
508: a second balancing weight;
509: a third fixed pulley;
510: a third wire rope;
511: a third balancing weight;
512: a limit frame;
513: a limit beam;
514: a test piece weight frame;
515: a fourth wire rope;
516: counterweight hanging rings;
517: a fourth fixed pulley;
518: a fourth balancing weight;
p: a test piece;
b: a frame;
s: a skin;
BP: a frame loading point;
SP: skin loading points.
Detailed Description
In the following, specific embodiments of the present invention will be described, and it should be noted that in the course of the detailed description of these embodiments, it is not possible in the present specification to describe all features of an actual embodiment in detail for the sake of brevity. It should be appreciated that in the actual implementation of any of the implementations, as in any engineering or design project, numerous implementation-specific decisions must be 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 appreciated that while such a development effort might be complex and lengthy, it would nevertheless be a routine undertaking of design, fabrication, or manufacture for those of ordinary skill having the benefit of this disclosure, and thus should not be construed as having the benefit of this disclosure.
Unless defined otherwise, technical or scientific terms used in the claims and specification should be given the ordinary meaning as understood by one of ordinary skill in the art to which this invention belongs. The terms "first," "second," and the like in the description and in the claims, are not used for any order, quantity, or importance, but are used for distinguishing between different elements. 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 "comprising" or "comprises", and the like, is intended to mean that elements or items that are immediately preceding the word "comprising" or "comprising", are included in the word "comprising" or "comprising", and equivalents thereof, without excluding other elements or items. The terms "connected" or "connected," and the like, are not limited to physical or mechanical connections, nor to direct or indirect connections.
FIG. 1 is a general schematic of an aircraft fuselage panel strength test apparatus according to one embodiment of the present invention. FIG. 2 is a schematic view of a test piece for an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention. FIG. 3 is a side view of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention. FIG. 4 is a schematic view of a pressurized loading assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention. FIG. 5 is another schematic view of a pressurized loading assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention. FIG. 6 is a schematic illustration of a skin hoop load assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention. FIG. 7 is a schematic view of a frame hoop load assembly of an aircraft fuselage panel strength test apparatus according to an embodiment of the present invention. FIG. 8 is a schematic view of a counterweight assembly of an aircraft fuselage panel strength test apparatus in accordance with an embodiment of the invention.
In some embodiments, as shown in fig. 1 to 8, an aircraft fuselage panel strength test apparatus for strength testing of a fuselage panel under a combination of a charging load and a hoop load, the aircraft fuselage panel strength test apparatus comprising: the frame hoop loading assembly 300 is used for applying hoop load to the frames B of the fuselage panel, and the pressurizing loading assembly 100, the skin hoop loading assembly 200 and the frame hoop loading assembly 300 are mutually independent, and the frame hoop loading assembly 300 is mutually independent for loading each frame B of the fuselage panel.
According to the invention, 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 the independent actuating cylinders and the V-shaped supporting levers, and the simulation of the pressurizing load in the operation of the civil aircraft is realized by sealing the airframe panel test piece and the pressure box to form a sealed cavity and inflating and deflating the cavity.
In some embodiments, as shown in fig. 4-5, the pressure loading assembly 100 includes a pressure cell support frame 115, a pressure cell 120, and a pressure protection device. The pressure cell support frame 115 includes channel steel beams, load beam connections, and base adjustment plates. The pressure cell side walls 101 and the pressure cell end plates 102 are provided with triangular reinforcing ribs 103 (wall plate supports) to enhance lateral rigidity, and the pressure cell bottom plates 104 are opened with manual inspection openings 118 and are additionally provided with flaps 105. An inflation tube 116 and an exhaust tube 117 are attached to the pressure cell sidewall 101, and a pressure measuring tube 111, a pressure releasing tube 112 and a lead tube 113 are attached to the pressure cell end plate 102. The pressure cell side walls 101, the pressure cell end plates 102 and the pressure cell bottom plate 104 are welded together to form a pressure cell 120. The outer ring of the manual inspection opening 118 is provided with a sealing groove 106, the outer ring of the flap 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 flap 105 and is used for placing a bolt to install the flap 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 adhered to the test piece P and the pressure box 120. The pressure measuring tube 111 on the pressure box end plate 102 is used for installing a pressure sensor, the lead tube 113 is used for leading out strain gauge leads, and the pressure relief tube 112 is used for installing an emergency pressure relief valve. The frame loading hole 119 is formed in the position, corresponding to the end head of the test piece P, of the pressure box side wall 101, the frame loading hole 119 is larger than the section of the frame loading resistance pull rod 301, the frame loading resistance pull rod 301 penetrates through the frame loading hole 119 and is connected with a frame loading point BP (bolt hole) of the end head of the frame, the frame loading resistance pull rod 301 is sleeved into the sealing hose 114, the sealing hose is connected with the frame after penetrating through the frame loading hole 119, one end of the 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 pressure box side wall 101. The pressure cell 120 is mounted on the pressure cell support frame 115, the exhaust pipe 117 is mounted with a relief valve, the pressure measuring pipe 111 is mounted with a pressure sensor, and the relief pipe 112 is mounted with an emergency relief valve.
In some embodiments, as shown in fig. 6, skin hoop load assembly 200 includes skin load tabs 201, skin load tie rods 202, a first set of load cells 204, a first set of hydraulic rams 205, a first set of load beams 207, and a first set of base adjustment plates 206. One end of the 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-stage lever 203, the multi-stage lever 203 is connected with a skin loading pull rod 202 to connect a first group of dynamometers 204 and one end of a first group of hydraulic cylinders 205, the other end of the first group of hydraulic cylinders 205 is hinged with the lug of a first group of base adjusting plates 206, and the bottom plate of the first group of base adjusting plates 206 is connected with a first group of load 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 hoop load assembly 300 includes a frame load tab, a frame load resistance tie rod 301, a frame load power link 305, a second set of load cells 306, a second set of hydraulic rams 307, a second set of load beams 309, a second set of base adjustment plates 308, a V-shaped support lever 302, an intermediate load beam 304, and a third set of base adjustment plates 303. The tab (frame loading tab) of the frame loading resistance stay 301 is connected to the frame loading point BP (bolt hole), and the other end of the frame loading resistance stay 301 is connected to the frame supporting point (bolt hole) of the V-shaped supporting lever 302. The base support points (bolt holes) of the V-shaped support levers 302 are hinged with the lugs of the third set of base adjustment plates 303, and the bottom plates of the third set of base adjustment plates 303 are connected with the middle load beam 304. The hydraulic ram loading point (bolt hole) is disposed at the intermediate position of the V-shaped support lever 302, and the ratio of the distance of the hydraulic ram loading point to the base support point to the distance of the frame loading point to the base support point is 1:2. the hydraulic ram loading point 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 one end of the second group of dynamometers 306 and one end of the second group of hydraulic rams 307, the other end of the second group of hydraulic rams 307 is hinged with the lugs of the second group of base adjusting plates 308, and the bottom plate of the second group of base adjusting plates 308 is connected with the second group of load beams 309. The V-shaped supporting lever 302 is of a truss structure and is designed by adopting an equal-strength method.
In some embodiments, as shown in fig. 8, the aircraft fuselage panel strength test apparatus further includes a test piece abnormal displacement monitoring assembly 400, the test piece abnormal displacement monitoring assembly 400 including a displacement monitoring frame 401, a laser displacement sensor 402, and a displacement alarm 403, wherein the displacement monitoring frame 401 is mounted on the skin loading assembly weight 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 test apparatus further includes a counterweight assembly 500, the counterweight assembly 500 including a skin loading assembly counterweight subassembly, a frame loading assembly counterweight subassembly, a V-support lever counterweight subassembly, and a test piece counterweight subassembly.
In some embodiments, as shown in fig. 7 and 8, the skin loading assembly counterweight subassembly includes a skin loading assembly counterweight frame 501, a first fixed pulley 502, a first wire rope 503, and a first counterweight 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 wire rope 503 is connected to the center of gravity position of the first set of hydraulic rams 205, and the other end of the first wire rope 503 is connected to the first counterweight 504 around the first fixed pulley 502, so as to implement a skin loading assembly counterweight (i.e., counteract the gravity effect of the skin loading assembly during testing).
In some embodiments, as shown in fig. 7 and 8, the frame loading assembly counterweight subassembly includes a frame loading assembly counterweight frame 505, a second fixed pulley 506, a second wire rope 507, and a second counterweight 508, wherein the frame loading assembly counterweight frame 505 is built over the second set of hydraulic rams 307, the second fixed pulley 506 is disposed on the frame loading assembly counterweight frame 505, one end of the second wire rope 507 is connected to the center of gravity of the second set of hydraulic rams 307, and the other end of the second wire rope 507 bypasses the second fixed pulley 506 to connect the second counterweight 508, implementing a frame loading assembly counterweight (i.e., counteracting the gravitational effects of the frame loading assembly during testing).
In some embodiments, as shown in fig. 7 and 8, the V-shaped supporting lever counterweight subassembly includes a third fixed pulley 509, a third wire rope 510, and a third counterweight 511, where two sets of third fixed pulleys 509 are respectively installed above the V-shaped supporting lever 302 and on the left and right sides of the lower surface of the flap 105, one end of the third wire rope 510 is connected to the center of gravity of the V-shaped supporting lever 302, and the other end of the third wire rope 510 bypasses the third fixed pulley 509 to connect the third counterweight 511, so as to implement the V-shaped supporting lever counterweight (i.e., counteract the gravity influence of the V-shaped supporting lever during the test). A limiting frame 512 is mounted on the pressure cell supporting frame 115, two limiting beams 513 are mounted on the limiting frame 512, and the 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 sheave 517, a fourth wire rope 515, and a fourth weight 518, wherein the test piece weight frame 514 is built above the test piece loading point, the fourth fixed sheave 517 is disposed on the test piece weight frame 514, one end of the fourth wire rope 515 is connected to a weight lifting ring 516 located at the outermost loading point of the test piece P, and the other end of the fourth wire rope 515 bypasses the fourth fixed sheave 517 to connect the fourth weight 518, so as to implement the test piece weight (i.e., counteract the gravity influence of the test piece during the test).
As shown in fig. 3 to 8, before the test, the pressure cell support frame 115, the first group of load beams 207, and the second group of load beams 309 were installed on the laboratory ground rails. The pressure cell 120 is mounted on the pressure cell support frame 115. The test piece P is mounted on the pressure box 120, and is sealed by the rubber strip 110 at the contact position of the test piece P and the pressure box 120, and the rubber strip 110 is adhered to the test piece P and the pressure box 120, respectively. The strain 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 has been prefabricated with loading points (bolt holes), the clamping end is placed in the middle of the ears of the skin loading binaural clamp 201, the bolt holes are aligned, and fastening is performed by 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 weight frame 501 is built over the first set of hydraulic rams 205 to be installed, the skin loading assembly weight frame 501 being fixed to the laboratory ground rail. A first fixed pulley 502 is arranged on the 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 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 balancing weight 504, so that the influence of dead 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 dynamometers 204, the first group of dynamometers 204 are connected with one end of the first group of hydraulic cylinders 205, the other end of the first group of hydraulic cylinders 205 is hinged with lugs of the first group of base adjusting plates 206, and the bottom plate of the first group of base adjusting plates 206 is fixed with the first group of bearing beams 207 through bolts. The number of stages of the multi-stage lever 203 is determined by the number of loading points, and the first set of hydraulic rams 205 are aligned with frame B of the test piece P.
As shown in fig. 7 and 8, the frame loading resistance tie 301 is fitted into the sealing hose 114, and after passing through the frame loading hole 119, the frame end is placed between the ears of the frame loading resistance tie 301, aligned with the bolt holes, and fastened by bolts. One end of the rubber hose 114 is bonded to the frame loading resistance stay 301, and is fastened by a clip, and the other end of the rubber hose 114 is bonded to the inside of the pressure cell side wall 101. The cover 105 is mounted on the pressure box bottom plate 104 by bolts, the manual inspection opening 118 is closed, and the rubber gasket 109 is arranged between the sealing groove 106 and the sealing boss 107, so that the sealing effect can be improved. The intermediate load beam 304 is mounted on the laboratory floor rail at an intermediate position directly below the pressure cell 120. The third group of base adjusting plates 303 are mounted on the middle load beam 304, and the bottom plates of the third group of base adjusting plates 303 are fastened with the middle load beam 304 through bolts. The base support points (bolt holes) of the V-shaped support levers 302 are hinged to the lugs of the third set of base adjustment plates 303, and the bottom plates of the third set of base adjustment plates 303 are connected to the intermediate load beam 304. The other end (single lug) of the frame loading resistance pull rod 301 is placed in the middle of the two ears of the frame supporting point of the V-shaped supporting lever 302, aligned with the bolt holes and fastened by bolts. One end of the third wire rope 511 is connected with a hole at the center of gravity of the V-shaped supporting lever 302, and the other end of the third wire rope 503 bypasses the third fixed pulley 509 under the flap 105 to be connected with the third balancing weight 511, so that the influence of the dead weight of the V-shaped supporting lever is eliminated. The pressure box support frame 115 is provided with the limiting frame 512, the limiting frame 512 is provided with the two limiting beams 513, and the third balancing weight 511 is arranged between the two limiting beams 513, so that the third balancing weight 511 is prevented from shaking in the test to influence the test. A frame loading assembly counterweight frame 505 is built over the second set of hydraulic rams 307 to be installed, securing the frame loading assembly counterweight frame 505 to the laboratory floor rail. One end of the second steel wire rope 507 is connected with the lug at the gravity center position of the second group of hydraulic actuating cylinders 307, and the other end of the second steel wire rope 507 bypasses the second fixed pulley 506 to be connected with the second balancing weight 508, so that the influence of the dead weight of the frame loading assembly is eliminated. One end of the frame loading resistance pull rod 305 is connected with a hydraulic actuator cylinder loading point of the V-shaped supporting lever 302, the other end of the frame loading resistance pull rod 305 is connected with one end of a second group of dynamometers 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 lugs 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 force bearing beams 309 through bolts, and the second group of force 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 hydraulic actuating cylinders in the first group of hydraulic actuating cylinders and the second group of hydraulic actuating cylinders is twice the number of the upper frames of the test piece P. During the test, two (left and right) hydraulic actuators corresponding to the same frame position in the first group of hydraulic actuators are connected in parallel, so that the load applied to the skin at each frame position is independently regulated; two hydraulic cylinders (left and right) corresponding to the same frame position in the second group of hydraulic cylinders are connected in parallel, so that the load applied to each frame is independently regulated; the two hydraulic actuating cylinders (left and right) corresponding to the same frame position are connected in parallel, so that the load applied to the left and right sides of the same frame position is equal, and the test piece P is kept in a balanced state.
In the test, the adjustment of the gas pressure in the pressure cell 120 can be achieved by adjusting the volume of gas entering the pressure cell 120 through the gas tube 116 and the volume of gas exiting the pressure cell 120 through the gas exhaust tube 117. Monitoring of the gas in the pressure cell 120 is achieved by a pressure sensor mounted on the pressure tube 111. The adjustment of the gas pressure within the pressure cell 120 allows for the simulation of the charging load imposed by the fuselage wall. When the pressure of the gas in the pressure box 120 exceeds a 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. The output forces of two hydraulic actuators at different frame positions in the first group of hydraulic actuators are regulated, so that the loading of skins at different frame positions is realized; and adjusting the output forces of two hydraulic actuators at different frame positions in the second group of hydraulic actuators to realize 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 panels are subjected to hoop loads in addition to the loading of the fuselage panels, and the skins and frames are subjected to different hoop loads at different frame locations. The fuselage wallboard strength test device disclosed in the prior art can not realize the simulation of the pressurizing load, and simultaneously, the skins at different frame positions and the hoop load born by the frames can be respectively and independently adjusted. The strength test device for the aircraft fuselage wall plate can simulate the pressurizing load, and simultaneously apply different circumferential loads to skins and frames at different frame positions, so that the load born by the fuselage wall plate can be simulated more truly.
While the invention has been described in terms of specific embodiments, those skilled in the art will recognize that the invention is not limited thereto, but that many modifications can be made by those skilled in the art without departing from the scope of the invention.
Claims (9)
1. An aircraft fuselage skin strength test device, characterized in that, the aircraft fuselage skin strength test device is used for carrying out strength test to the fuselage skin under the combined action of charging load and hoop load, the aircraft fuselage skin strength test device includes: the frame annular loading assembly is used for applying annular load to frames of the fuselage panel, and the pressurizing loading assembly, the skin annular loading assembly and the frame annular loading assembly are mutually independent;
the frame annular loading assembly comprises frame loading lugs, frame loading resistance pull rods, frame loading power pull rods, a second group of dynamometers, a second group of hydraulic actuating cylinders, a second group of bearing beams, a second group of base adjusting plates, a V-shaped supporting lever, a middle bearing beam and a third group of base adjusting plates, wherein one ends of the frame loading resistance pull rods are provided with the frame loading lugs, the frame loading lugs are connected with frame loading points, the other ends of the frame loading resistance pull rods are connected with frame supporting points of the V-shaped supporting levers, base supporting points of the V-shaped supporting levers are hinged with lugs of the third group of base adjusting plates, base plates of the third group of base adjusting plates are connected with the middle bearing beams, the loading points of the V-shaped supporting levers are connected with one ends of the frame loading power pull rods, the other ends of the frame loading power pull rods are connected with the second group of the dynamometers and one ends of the second group of the hydraulic actuating cylinders, the other ends of the second group of the hydraulic actuating cylinders are connected with base supporting points of the second group of the second bearing plates, and the base supporting plates of the second group of base adjusting plates are hinged with the base plates of the second group of the base adjusting plates.
2. The aircraft fuselage panel strength test apparatus of claim 1, wherein the pressure loading assembly comprises a pressure cell support frame, a pressure cell, wherein the pressure cell is mounted on the pressure cell support frame, the pressure cell comprising a pressure cell sidewall, a panel support, an inflation tube, an exhaust tube, a pressure measurement tube, a pressure relief tube, a vent tube, a pressure cell floor, a flap.
3. The aircraft fuselage panel strength test apparatus of claim 2, wherein the pressurization loading assembly further comprises a pressurization protection device, wherein the pressurization protection device comprises a pressure sensor mounted on the pressure tube and a pressure relief valve mounted on the exhaust tube.
4. The aircraft fuselage panel strength test apparatus of claim 1, wherein the skin hoop load assembly comprises a skin load tab, a skin load pull rod, a first set of load cells, a first set of hydraulic rams, a first set of load beams, a first set of base adjustment plates, wherein one end of the skin load tab is connected to a skin load point, the other end of the skin load tab is connected to a multi-stage lever, the multi-stage lever is connected to the skin load pull rod to the first set of load cells and to one end of the first set of hydraulic rams, the other end of the first set of hydraulic rams is hinged to the tabs of the first set of base adjustment plates, and the bottom plate of the first set of base adjustment plates is connected to the first set of load beams.
5. The aircraft fuselage panel strength test apparatus of claim 1, wherein the hydraulic ram loading point is disposed in an intermediate position of the V-shaped support lever, and wherein a ratio of a distance from the hydraulic ram loading point to the base support point to a distance from the frame loading point to the base support point is 1:2.
6. 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 the skin loading assembly counterweight frame, and the laser displacement sensor and the displacement alarm are mounted on the displacement monitoring frame.
7. The aircraft fuselage panel strength test apparatus of claim 1, further comprising a weight assembly comprising a skin load assembly weight subassembly, a frame load assembly weight subassembly, a V-support lever weight subassembly, and a test piece weight subassembly.
8. The aircraft fuselage panel strength test apparatus of claim 7, wherein the frame load assembly counterweight sub-assembly includes a frame load assembly counterweight frame, a second fixed pulley, a second wire rope, and a second counterweight, wherein the frame load assembly counterweight frame is built above a second set of hydraulic rams, the second fixed pulley is disposed on the frame load 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 bypasses the second fixed pulley to connect the second counterweight to implement a frame load assembly counterweight.
9. The aircraft fuselage panel strength test apparatus of claim 7, wherein the test piece weight sub-assembly comprises a test piece weight frame, a fourth fixed pulley, a fourth wire rope, and a fourth weight block, wherein the test piece weight frame is built above a test piece loading point, the fourth fixed pulley is arranged on the test piece weight frame, one end of the fourth wire rope is connected with a weight hanging ring positioned at an outermost loading point of the test piece, and the other end of the fourth wire rope bypasses the fourth fixed pulley to connect the fourth weight block, so that the test piece weight is realized.
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| CN114778168B (en) * | 2022-06-17 | 2022-09-02 | 中国飞机强度研究所 | Method for determining loading stage number of breaking load in aerospace plane cabin section ground strength test |
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