CN113800005A - Emergency breaking test bed and test method for main landing gear of airplane - Google Patents
Emergency breaking test bed and test method for main landing gear of airplane Download PDFInfo
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Abstract
The invention provides an emergency disconnection test bed and an emergency disconnection test method for a main landing gear of an airplane. The test method comprises the steps that a plurality of breaking pins are arranged on the main landing gear according to a main landing gear mounting structure and an emergency breaking requirement; loading a load on an axle of a main undercarriage, wherein the direction and the magnitude of the loaded load are changed according to the actual working condition; when the loaded load reaches the limit load FuKeeping for a certain time and continuously separating; continuously loading the load to break the break pin, recording the break load, and obtaining the break sequence of each break pin; and comparing the obtained breaking sequence and breaking load of each breaking pin with the preset breaking sequence and breaking load, if the obtained breaking sequence and breaking load are consistent, the obtained breaking pin is proved to be reasonable in design, and if the obtained breaking sequence and breaking load are inconsistent, each breaking pin is redesigned and test verification is carried out. The invention can test the emergency breaking sequence and breaking strength of the full-size aircraft main landing gear under different loading working conditions, thereby verifying the reliability of the emergency breaking pin of the main landing gear; in addition, the test bed adopts a modular design.
Description
Technical Field
The invention relates to the technical field of airplane main undercarriage tests, in particular to an airplane main undercarriage emergency disconnection test bed.
Background
With the shortening of the operation period of the airplane and the increase of the taking-off and landing times, higher requirements are put on the safety of the airplane design. In the safety design requirements of civil aircrafts, stable and reliable operation of the aircrafts is ensured, and safety of passengers is guaranteed under extreme conditions of emergency landing, even falling and the like. When the main landing gear structure bears a large load, in order to avoid the damage of the wing body structure connected with the main landing gear system, the occurrence of fire or explosion caused by the tearing of a fuel tank for placing the wings is caused, and the degree of secondary damage to passengers in the emergency falling process of the civil aircraft is reduced, an emergency disconnecting device needs to be designed at the joint of the main landing gear and the wings, so that the main landing gear and the wings are rapidly separated according to a preset disconnecting sequence under the condition that the main landing gear bears overload.
At present, two types of emergency breaking modes are mainly structural member breaking and breaking of a breaking pin in a connecting piece, wherein the breaking of the breaking pin is widely adopted due to simple structure and convenient test. The existing emergency disconnection test bed mainly tests the mechanical property of the emergency disconnection pin so as to provide a test platform and a test method for the structural design and reliability of the disconnection pin. However, the main landing gear is used as a system, and the whole emergency disconnection structure of the main landing gear needs to be tested and verified, so that reference is provided for the design of the main landing gear emergency disconnection device.
Disclosure of Invention
The invention aims to provide an emergency disconnection test bed for a main undercarriage of an airplane, which can be used for carrying out test verification on an integral emergency disconnection structure of the main undercarriage so as to provide reference for the design of an emergency disconnection device of the main undercarriage.
The technical scheme of the invention is as follows: an aircraft main landing gear emergency disconnection test method comprises the following steps:
arranging a plurality of breaking pins on the main landing gear according to the mounting structure of the main landing gear and the emergency breaking requirements; loading a load on an axle of a main undercarriage, wherein the direction and the magnitude of the loaded load are changed according to the actual working condition; when the loaded load reaches the breaking limit load Fu, the load needs to be kept for a certain time and the breaking pin is continuously broken; continuously loading the breaking pin with load until the loaded load exceeds the limit load borne by the breaking pin to break the breaking pin, and recording the breaking load when the breaking pin is broken and the breaking sequence of each breaking pin; and comparing the recorded breaking sequence of each breaking pin with a preset breaking sequence, comparing the recorded breaking load of each breaking pin with the preset breaking load, if the recorded breaking load of each breaking pin is consistent with the preset breaking load, the recorded breaking load is proved to be reasonable in design, and if the recorded breaking load of each breaking pin is inconsistent with the preset breaking load, each breaking pin is redesigned and test verification is carried out.
The direction and the magnitude of the loaded load change according to the actual working condition: the actual working condition is the simulated falling working condition of the civil aircraft. When the airplane falls, the directions and the sizes of loads borne by all parts of the main landing gear are different; in addition, the airplane falling working conditions are different.
Preferably, the loaded load direction comprises a vertical load, a lateral load, a resistance load, a combination of the vertical load and the lateral load, a combination of the lateral load and the resistance load, and a combination of the vertical load, the resistance load and the lateral load; so as to respectively obtain the stress and strain characteristics of each break pin and the load borne by the main landing gear.
Preferably, the plurality of break pins respectively comprise a front lug break pin and a rear lug break pin which are arranged on a main landing gear strut, and a side strut upper break pin, a retracting rotating pin and a side strut lower break pin which are arranged on a side strut;
the breaking sequence of each breaking pin is as follows:
when the main undercarriage mainly bears vertical load and resistance load, firstly, the front lug breaking pin is guaranteed to break, then the main undercarriage bends backwards to cause the rear lug breaking pin to break, and finally, the lower breaking pin of the side stay bar is guaranteed to break; the main landing gear is integrally broken off;
when the main landing gear mainly bears lateral tension or compression load, any breaking pin of the upper breaking pin, the retracting rotating pin and the lower breaking pin of the side stay bar needs to be broken.
Preferably, before loading the load on the axle of the main landing gear, the main landing gear is clamped in an inverted installation mode, so that each loading point is located in the upper space.
Preferably, the ultimate load borne by each breaking pin is obtained in a main landing gear drop test according to the requirements of an airplane main landing gear strength design guideline; or calculated from the main landing gear ground load.
Preferably, the limit load F of each break pinuThe calculation is as follows:
Fu＝S×Fl (1)
in the formula: fuIs the limit load of the break pin; flA limit load for breaking the pin; and S is a safety factor, and the safety factor is 1.5 specified in the pilot flight standard.
Preferably, the minimum breakaway load is determined according to the structural integrity of the main landing gear and the minimum breakaway load failure requirement; determining the maximum breaking load according to the material tolerance, the geometric tolerance, the rigidity tolerance of the main landing gear and the rigidity tolerance of the wings; the breaking load of the breaking pin is obtained by calculation through formulas (2) and (3):
FDmin＝n1×Fu (2)
FDmax＝n2×FDmin (3)
in the formula: fDminIs the minimum break load; fDmaxThe maximum breaking load; n is1、n2Is a breaking load coefficient, n1Taking 1.1, n2Taking 1.03;
when the breaking pin meets the requirement after the axle of the main landing gear is loaded with the limit load, the load is continuously increased until the breaking pin is broken, the breaking load measured in the test is compared with the designed maximum and minimum breaking loads, and if the test breaking load is between the minimum breaking load and the maximum breaking load, the design of the breaking pin is reasonable; otherwise, the breaking pin is redesigned and test verification is carried out.
The invention also provides an emergency disconnection test bed for the main landing gear of the airplane, which comprises a lateral fixing assembly, a lateral loading piece, a vertical fixing assembly, a fixture, a resistance fixing assembly, a resistance loading piece and a main landing gear fixing combined tool;
the fixture comprises an n-shaped gantry structure and side stand rod assemblies arranged beside the gantry structure, a test cavity is formed inside the gantry structure, and the fixture is provided with X, Y, Z three directions of a three-dimensional rectangular coordinate system;
the lateral fixing component is arranged on the vertical edge of the gantry structure and can move in the Z direction; the lateral loading piece is arranged on the lateral fixing component and can move in the Y direction; the lateral loading piece can extend and retract towards the test cavity;
the vertical fixing component is arranged on the horizontal edge of the top of the gantry structure and can move in the X direction; the vertical loading piece is arranged on the vertical fixing component and can move in the Y direction; the vertical loading piece can extend and retract towards the test cavity;
the resistance fixing assembly is arranged on the side stand rod assembly and can move in the Z direction; the resistance loading piece is arranged on the resistance fixing component and can move in the X direction; the resistance loading piece can stretch towards the test cavity;
the main undercarriage fixing combined tool is arranged in the test cavity and is positioned below the vertical fixing component; the main undercarriage fixed combination tool is provided with a plurality of fixed tools which can move towards the Y direction and are used for fixing all parts of the main undercarriage. A plurality of fixed frocks are simultaneously subjected to displacement adjustment, so that the clamping precision is ensured due to the fact that the clamping is carried out in the whole process of adjustment.
Preferably, the resistance fixing assembly comprises a first mounting plate and a second mounting plate, a first slide rail is arranged on one side end face of the first mounting plate, and a second slide rail is arranged on the other side end face of the first mounting plate; a sliding block is arranged on the end face of one side of the second mounting plate; the second sliding rail is connected with the side vertical rod component in a sliding mode in the Z direction; the sliding block is connected with the second sliding rail in a sliding mode in the X direction;
a concave sliding groove is formed in the vertical side face of the first sliding rail, and a sliding plate matched with the sliding groove is arranged on the sliding block in a protruding and extending mode;
the resistance loading member is mounted on the second mounting plate.
Compared with the related technology, the invention has the beneficial effects that:
the method includes the steps that firstly, the emergency breaking sequence and the breaking strength (breaking load) of a main landing gear of a full-size airplane under different loading conditions can be tested, so that the reliability of an emergency breaking pin of the main landing gear is verified, and test reference is provided for the emergency breaking design of the main landing gear;
the emergency disconnection test bed can test the whole structure of the main undercarriage, and is novel in structure;
the three-dimensional space loading device comprises resistance loading along the opposite direction of the course, vertical loading of a vertical machine body and lateral loading along the direction of an axle, and can simulate various space loads borne by the main landing gear emergency disconnecting device;
the three-dimensional coordinate of the test bed can be adjusted, so that the test of the emergency breaking of the main landing gear with different sizes can be tested, and the universality is good;
fifthly, the main undercarriage is clamped on the test bed in an inverted installation mode, so that the load is loaded and arranged in the upper space, and the integral installation, test and test process monitoring of the main undercarriage are facilitated;
and sixthly, the test bed adopts a modular design, and the main landing gear emergency breaking test of different sizes can be carried out through different clamp combinations.
Drawings
FIG. 1 is a schematic illustration of the structure of a main landing gear of an aircraft of the type described;
FIG. 2 is a schematic diagram of the location of the load points of the main landing gear;
FIG. 3 is a schematic view of a main landing gear clamping of an aircraft main landing gear emergency disconnection test bed provided by the invention;
FIG. 4 is an enlarged view of the point A in FIG. 3
FIG. 5 is a schematic structural view of the resistance securing assembly of FIG. 3;
fig. 6 is a schematic structural view of the main landing gear fixing combination tool in fig. 3.
In the drawings: 1. a lateral securing assembly; 2. a lateral loading member; 3. a vertical loading member; 4. a vertical fixing component; 5. forming a frame; 51. a gantry structure; 52. a side stand bar assembly; 6. a resistance fixing assembly; 61. a first mounting plate; 62. a second mounting plate; 63. a first slide rail; 64. a second slide rail, 65 slide blocks, 66 and a slide groove; 67. a slide plate; 68. adjusting bolts, 69 and fixing blocks; 7. a resistance loading member; 8. the main undercarriage fixes the combined tool; 81. the device comprises a flat plate fixing block 82, a movable sliding block 83 and a side stay bar fixing tool; 84. a front lug fixing tool; 85. fixing a tool for a rear lug; 86. adjusting the bolt; 87. a nut;
9. a main landing gear strut; 10. a side stay bar; 11. a front ear break pin; 12. the rear ear break pin; 13. a break pin is arranged on the side stay bar; 14. retracting and releasing the rotating pin; 15. a side stay bar lower break pin; 16. axles, 161, vertical load points; 162. a lateral loading point; 163. a resistance load point.
Detailed Description
The present invention will be described in detail below with reference to the embodiments with reference to the attached drawings. It should be noted that the embodiments and features of the embodiments may be combined with each other without conflict. For convenience of description, the words "upper", "lower", "left" and "right" in the following description are used only to indicate the correspondence between the upper, lower, left and right directions of the drawings themselves, and do not limit the structure.
The invention provides an emergency disconnection test method for a main landing gear of an airplane, which is suitable for the main landing gear of a civil airplane and the positions of emergency disconnection pins of the civil airplane are shown in figure 1. The main landing gear adopts a three-point type and parallel double-wheel strut type structure.
Which comprises the following steps:
and step S1, arranging a plurality of breaking pins on the main landing gear mounting structure according to the emergency breaking requirements. As shown in fig. 1, the plurality of break pins respectively include a front ear break pin 11 and a rear ear break pin 12 provided on the main landing gear strut 9, and further include a side stay upper break pin 13, a retraction rotation pin 14, and a side stay lower break pin 15 provided on the side stay 10.
And step S2, loading the axle of the main landing gear, wherein the direction and the magnitude of the loaded load are changed according to the actual working condition.
As shown in fig. 2, the main landing gear axle 16 is provided with a vertical load point 161, a lateral load point 162 and a drag load point 163. Vertical load point 161, lateral load point 162, and resistive load point 163 are all in three directions X, Y, Z of a three-dimensional rectangular coordinate system, where Z is the vertical direction.
In the test, besides considering the structure of the main landing gear and the requirement of emergency disconnection, the stress characteristic of the main landing gear needs to be considered in combination, wherein the stress characteristic means that the main landing gear bears larger ground load under special working conditions of abnormal landing, obstacle crossing and the like of the airplane, and the stress characteristic usually comprises vertical load, course resistance load and lateral load:
(1) the main landing gear bears larger vertical load when the aircraft lands at a larger landing speed;
(2) when the airplane impacts or passes through an obstacle during the running process, the airplane is subjected to excessive resistance load;
(3) excessive side loads that occur when an aircraft lands due to deviation from a runway or at a high roll or yaw attitude;
in addition, the landing process of the airplane also comprises the combined working condition of the three conditions, and the bearing is complex.
Therefore, the loaded load directions comprise seven loading modes of vertical load, lateral load, resistance load, combination of vertical load and lateral load, combination of lateral load and resistance load and combination of vertical load, resistance load and lateral load, so as to respectively obtain the stress and strain characteristics of each breaking pin and the load borne by the main landing gear, and thus verify the breaking strength (breaking load) of the emergency breaking pin and the breaking sequence of the main landing gear.
When a large ground load is transmitted to the breaking pins at all parts and exceeds the limit load (non-calculated limit load) borne by the breaking pins, the main landing gear needs to be broken off according to the preset breaking direction, so that passengers of the airplane are prevented from secondary damage caused by airplane falling. According to the loading condition, the following requirements are put forward on the emergency disconnection sequence of the main undercarriage:
when the main landing gear mainly bears vertical load and resistance load, the front lug breaking pin 11 is firstly guaranteed to break, then the main landing gear bends backwards and leads to the breaking of the rear lug breaking pin 12, and finally the breaking of the lower breaking pin 15 of the side stay bar is guaranteed; the main landing gear is integrally broken off;
when the main landing gear mainly bears lateral tensile or compressive load, any breaking pin of the upper breaking pin 13, the retracting rotating pin 14 and the lower breaking pin 15 of the side stay bar needs to be broken.
When designing each break pin, it is necessary to consider the break load. The load of each emergency breaking pin can be obtained in two ways, namely, the load is obtained in a main landing gear drop test according to the requirements of an airplane main landing gear strength design guideline; and secondly, the load is calculated according to the ground load of the main landing gear. Limit load F of each break pinuThe calculation is as follows:
Fu＝S×Fl (1)
in the formula: fuIs the limit load of the break pin; flA limit load for breaking the pin; and S is a safety factor, and the safety factor is 1.5 specified in the pilot flight standard.
The breaking load of the breaking pin can be determined according to the limit load and the strength of the supporting structure. Determining the minimum breaking load according to the structural integrity of the main landing gear and the minimum breaking load damage failure requirement; the maximum breakaway load is determined from material tolerances, geometric tolerances, main landing gear stiffness tolerances, and wing stiffness tolerances. The breaking load is calculated as follows:
FDmin＝n1×Fu (2)
FDmax＝n2×FDmin (3)
in the formula: fDminIs the minimum break load; fDmaxThe maximum breaking load; n is1、n2Is a breaking load coefficient, n1Taking 1.1, n2Take 1.03.
The structure and the size of each breaking pin are designed according to the bearing requirements of each emergency breaking pin, finally, a breaking test is carried out through the test bed of the invention to verify the breaking sequence and breaking load of each breaking pin, and the breaking load coefficients of the formulas 2 and 3 are checked to provide test reference for the emergency breaking design of the main landing gear.
The test is carried out on the test bed for the emergency disconnection of the main landing gear of the airplane. As shown in fig. 3, the test bed comprises a lateral fixing assembly 1, a lateral loading part 2, a vertical loading part 3, a vertical fixing assembly 4, a fixture 5, a resistance fixing assembly 6, a resistance loading part 7 and a main landing gear fixing combined tool 8.
The fixture 5 is a bearing part of a test bed and comprises an n-shaped gantry structure 51 and side vertical rod components 52 arranged on the side of the gantry structure, a test cavity is formed inside the gantry structure, and the fixture 5 has X, Y, Z directions.
The lateral fixing component 1 is mounted on the vertical edge of the gantry structure 51, and the lateral fixing component 1 can move in the Z direction. The lateral loader 2 is mounted on the lateral fixing assembly 1, and the lateral loader 2 is movable in the Y direction. The lateral loading member 2 can be extended and retracted towards the test chamber.
The vertical fixing component 4 is arranged on the horizontal edge at the top of the gantry structure 51, and the vertical fixing component 4 can move in the X direction. The vertical loading piece 3 is arranged on the vertical fixing component 4, and the vertical loading piece 3 can move in the Y direction; the vertical loading piece 3 can stretch and retract towards the test cavity.
The resistance fixing assembly 6 is mounted on the side stand bar assembly 52, and the resistance fixing assembly 6 is movable in the Z direction. The resistance loading member 7 is mounted on the resistance fixing assembly 6, and the resistance loading member 7 is movable in the X direction. The resistance loading member 7 is capable of extending and retracting towards the test chamber.
The structure of the resistance fixing component 6 is specifically described as follows: as shown in fig. 3 and 4, the resistance fixing assembly 6 includes a first mounting plate 61 and a second mounting plate 62, a first slide rail 63 is disposed on one side end surface of the first mounting plate 61, and a second slide rail 64 is disposed on the other side end surface. A slide block 65 is arranged on one side end face of the second mounting plate 62.
The side standing rod assembly 52 is vertically provided with a slide rail matched with the second slide rail 64, so that the second slide rail 64 and the side standing rod assembly 52 are connected in a sliding manner in the Z direction. The first mounting plate 61 slides in the Z direction relative to the side stand bar assembly 52, and the sliding adjustment is realized through an adjusting bolt 68 and a fixing block 69, specifically: the fixing block 69 is fixed on the side standing rod assembly 52, a screw hole is formed in the fixing block, the adjusting bolt 68 is in threaded fit with the screw hole, and the tail end of the adjusting bolt 68 is abutted to the first mounting plate 61 so as to push the first mounting plate 61 to slide. In order to fix the sliding position of the first mounting plate 61, an adjusting bolt 68 is provided at both ends in the Z direction, so as to realize abutting fixation. In order to ensure the smoothness of the first mounting plate 61 during sliding, adjusting bolts 68 are provided at both ends on the same side.
The slider 65 is slidably connected to the second slide rail 64 in the X direction. As shown in fig. 2, a recessed sliding groove 66 is provided on a vertical side surface of the first sliding rail 63, and a sliding plate 67 engaged with the sliding groove 66 is protrudingly provided on the sliding block 65. The buckle sliding design structure of spout and slide, when no external force, the auto-lock nature is better.
The resistance loading member 7 is mounted on the second mounting plate 62.
The structure and the adjusting principle of the vertical fixing component 3 and the lateral fixing component 1 are the same as those of the resistance fixing component 6.
Lateral loading piece 2, vertical loading piece 3 and resistance loading piece 7 are the pneumatic cylinder, can stretch out and draw back to the test chamber.
The main landing gear fixing and combining tool 8 is arranged in the test cavity and is positioned below the vertical fixing component 4. And a clamping tool capable of moving towards the Y direction is arranged on the main undercarriage fixed combined tool 8.
As shown in fig. 3 and 5, the main landing gear fixing combination tool 8 includes a flat plate fixing block 81, a moving slider 82, a side stay fixing tool 83, a front lug fixing tool 84, a rear lug fixing tool 85, an adjusting bolt 86, and a nut 87.
The flat plate fixing block 81 is fixed at the bottom below the gantry structure 51, and a T-shaped groove is formed in the upper end face of the flat plate fixing block. The bolt head of the adjusting bolt 86 is inserted into the T-shaped groove, and the screw rod passes through the U-shaped groove on the movable sliding plate 82, and the position of the movable sliding plate 82 is fixed by a multi-nut 87. The movable sliding plate 82 can be adjusted in a movable manner along the T-shaped groove in the Y direction, so that the clamping position of the main landing gear can be adjusted to be optimal.
The movable sliding plate 82 is provided with a side stay bar fixing tool 83, a front lug fixing tool 84 and a rear lug fixing tool 85, and the side stay bar fixing tool, the front lug fixing tool and the rear lug fixing tool are arranged at proper positions according to the structure of the main landing gear.
As shown in fig. 6, during clamping, the position of the main landing gear fixing and combining tool 8 is adjusted to ensure that the tested main landing gear is placed in the test cavity and the working area of each loading oil cylinder. And then installing the main landing gear of the airplane with the emergency disconnection pin on the main landing gear fixing and combining tool 8, finely adjusting the movable sliding plate 82, and simultaneously adjusting each fixing component to ensure that the loaded load acts on each loading point on the wheel shaft 16.
Before testing, strain gauges are attached to the inner diameter of each emergency break pin, and each loading oil cylinder is provided with a force transducer so as to conveniently record break load when the emergency break pin breaks, and verify whether the limit load and the limit load of the emergency break pin meet design requirements.
Step S3, start the test
In the test, seven loading modes of vertical load, lateral load, resistance load, combination of vertical load and lateral load, combination of lateral load and resistance load and combination of vertical load, resistance load and lateral load are respectively carried out.
Loading limit load on an axle of a main landing gear and keeping the limit load for a certain time (CCAR25 airworthiness standard stipulates that a structure must be capable of bearing the limit load for at least 3 seconds without breaking), if the breaking pin does not break, meeting the requirement, and carrying out the next step; if the break pin breaks, the break pin is redesigned and tested.
The next step is that when the breaking pin meets the requirements after the axle of the main landing gear is loaded with the limit load, the load is continuously increased until the breaking pin is broken, the breaking load measured in the test is compared with the designed maximum and minimum breaking loads, and if the test breaking load is between the minimum breaking load and the maximum breaking load, the design of the breaking pin is proved to be reasonable; otherwise, the breaking pin is redesigned and test verification is carried out.
And recording the breaking load when the breaking pin is broken, and obtaining the breaking sequence of each breaking pin. And comparing the breaking sequence of each breaking pin obtained by the test with a preset breaking sequence, comparing the breaking load of each breaking pin obtained by the test with a preset breaking load, if the breaking load is consistent with the preset breaking load, the breaking pins are proved to be reasonable in design, and if the breaking load is inconsistent with the preset breaking load, the breaking pins are redesigned and the test verification is carried out.
In the test process, the stress-strain characteristics of each emergency break pin and the load borne by the main undercarriage are obtained respectively, so that the breaking strength of the emergency break pins and the breaking sequence of the main undercarriage are verified.
The above description is only an embodiment of the present invention, and not intended to limit the scope of the present invention, and all modifications of equivalent structures and equivalent processes, which are made by using the contents of the present specification and the accompanying drawings, or directly or indirectly applied to other related technical fields, are included in the scope of the present invention.
Claims (9)
1. An aircraft main landing gear emergency disconnection test method is characterized by comprising the following steps:
arranging a plurality of breaking pins on the main landing gear according to the mounting structure of the main landing gear and the emergency breaking requirements; loading a load on an axle of a main undercarriage, wherein the direction and the magnitude of the loaded load are changed according to the actual working condition; when the loaded load reaches the breaking limit load FuIn time, the breaking pin is kept for a certain time and is continuously broken; continuously loading the breaking pin with load until the loaded load exceeds the limit load borne by the breaking pin to break the breaking pin, and recording the breaking load when the breaking pin is broken and the breaking sequence of each breaking pin; and comparing the recorded breaking sequence of each breaking pin with a preset breaking sequence, comparing the recorded breaking load of each breaking pin with the preset breaking load, if the recorded breaking load of each breaking pin is consistent with the preset breaking load, the recorded breaking load is proved to be reasonable in design, and if the recorded breaking load of each breaking pin is inconsistent with the preset breaking load, each breaking pin is redesigned and test verification is carried out.
2. The method of claim 1, wherein the loaded load directions include vertical loads, lateral loads, drag loads, combinations of vertical and lateral loads, combinations of lateral and drag loads, and combinations of vertical and drag loads and lateral loads; so as to respectively obtain the stress and strain characteristics of each break pin and the load borne by the main landing gear.
3. The method for testing the emergency disconnection of the main landing gear of the aircraft according to claim 1, wherein the plurality of disconnection pins respectively comprise a front ear disconnection pin and a rear ear disconnection pin which are arranged on a strut of the main landing gear, and an upper side stay disconnection pin, a retracting rotation pin and a lower side stay disconnection pin which are arranged on a side stay;
the breaking sequence of each breaking pin is as follows:
when the main undercarriage mainly bears vertical load and resistance load, firstly, the front lug breaking pin is guaranteed to break, then the main undercarriage bends backwards to cause the rear lug breaking pin to break, and finally, the lower breaking pin of the side stay bar is guaranteed to break; the main landing gear is integrally broken off;
when the main landing gear mainly bears lateral tension or compression load, any breaking pin of the upper breaking pin, the retracting rotating pin and the lower breaking pin of the side stay bar needs to be broken.
4. The method for the emergency breakaway test of the main landing gear of the aircraft according to claim 1, wherein before loading the axle of the main landing gear, the main landing gear is clamped in an inverted installation manner, so that each loading point is located in the upper space.
5. The method for the emergency disconnection test of the main landing gear of the airplane according to claim 1, wherein the ultimate load borne by each disconnection pin is obtained in a main landing gear drop test according to the requirements of the design guidelines for the strength of the main landing gear of the airplane; or calculated from the main landing gear ground load.
6. An aircraft main landing gear emergency breakaway test method according to claim 5, wherein the ultimate load F of each breakaway pinuThe calculation is as follows:
Fu＝S×Fl (1)
in the formula: fuIs the limit load of the break pin; flA limit load for breaking the pin; and S is a safety factor, and the safety factor is 1.5 specified in the pilot flight standard.
7. The method for the emergency breakaway test of an aircraft main landing gear of claim 6, wherein the minimum breakaway load is determined based on the structural integrity of the main landing gear and the minimum breakaway load failure requirement; determining the maximum breaking load according to the material tolerance, the geometric tolerance, the rigidity tolerance of the main landing gear and the rigidity tolerance of the wings; the breaking load of the breaking pin is obtained by calculation through formulas (2) and (3):
FDmin＝n1×Fu (2)
FDmax＝n2×FDmin (3)
in the formula: fDminIs the minimum break load; fDmaxThe maximum breaking load; n is1、n2Is a breaking load coefficient, n1Taking 1.1, n2Taking 1.03;
when the breaking pin meets the requirement after the axle of the main landing gear is loaded with the limit load, the load is continuously increased until the breaking pin is broken, the breaking load measured in the test is compared with the designed maximum and minimum breaking loads, and if the test breaking load is between the minimum breaking load and the maximum breaking load, the design of the breaking pin is reasonable; otherwise, the breaking pin is redesigned and test verification is carried out.
8. An emergency disconnection test bed for a main undercarriage of an airplane is characterized by comprising a lateral fixing component (1), a lateral loading piece (2), a vertical loading piece (3), a vertical fixing component (4), a profile frame (5), a resistance fixing component (6), a resistance loading piece (7) and a main undercarriage fixing combined tool (8);
the fixture (5) comprises an n-shaped gantry structure (51) and side stand rod assemblies (52) arranged beside the gantry structure, a test cavity is formed inside the gantry structure, and the fixture (5) has X, Y, Z three directions of a three-dimensional rectangular coordinate system;
the lateral fixing assembly (1) is arranged on the vertical edge of the gantry structure (51), and the lateral fixing assembly (1) can move in the Z direction; the lateral loading piece (2) is arranged on the lateral fixing component (1), and the lateral loading piece (2) can move in the Y direction; the lateral loading piece (2) can stretch towards the test cavity;
the vertical fixing component (4) is arranged on the horizontal edge of the top of the gantry structure (51), and the vertical fixing component (4) can move in the X direction; the vertical loading piece (3) is arranged on the vertical fixing component (4), and the vertical loading piece (3) can move in the Y direction; the vertical loading piece (3) can stretch towards the test cavity;
the resistance fixing assembly (6) is installed on the side stand rod assembly (52), and the resistance fixing assembly (6) can move in the Z direction; the resistance loading piece (7) is arranged on the resistance fixing component (6), and the resistance loading piece (7) can move in the X direction; the resistance loading piece (7) can stretch towards the test cavity;
the main undercarriage fixing combined tool (8) is arranged in the test cavity and is positioned below the vertical fixing component (4); and the main undercarriage fixed combined tool (8) is provided with a plurality of fixed tools which can move towards the Y direction and are used for fixing all parts of the main undercarriage.
9. The aircraft main landing gear emergency disconnection test bed according to claim 8, wherein the resistance fixing assembly (6) comprises a first mounting plate (61) and a second mounting plate (62), a first sliding rail (63) is arranged on one side end face of the first mounting plate (61), and a second sliding rail (64) is arranged on the other side end face of the first mounting plate; a sliding block (65) is arranged on the end face of one side of the second mounting plate (62); the second sliding rail (64) is connected with the side vertical rod component (52) in a sliding mode in the Z direction; the sliding block (65) is connected with the second sliding rail (64) in a sliding mode in the X direction;
a concave sliding groove (66) is formed in the vertical side face of the first sliding rail (63), and a sliding plate (67) matched with the sliding groove (66) is arranged on the sliding block (65) in a protruding and extending mode;
the resistance loading member (7) is mounted on the second mounting plate (62).
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114778056A (en) * | 2022-06-13 | 2022-07-22 | 中国飞机强度研究所 | Test system and method for dynamic impact mechanical property test of light airplane structure |
| CN117022668A (en) * | 2023-10-08 | 2023-11-10 | 中国飞机强度研究所 | Device and method for continuous cable passing and obstacle crossing test of aircraft landing gear |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4392623A (en) * | 1980-12-22 | 1983-07-12 | The Boeing Company | Fused connection adapted to fail under different overloads acting in different directions |
| CN103604593A (en) * | 2013-11-21 | 2014-02-26 | 南京航空航天大学 | Plane type ejection launch carrier-borne aircraft front undercarriage static force test loading apparatus |
| CN103604619A (en) * | 2013-11-21 | 2014-02-26 | 南京航空航天大学 | Front undercarriage static force test loading device for frame-type catapult-assisted take-off shipboard aircraft |
| CN103630346A (en) * | 2013-11-19 | 2014-03-12 | 北京航空航天大学 | Test table and test method of emergency breaking off pin of undercarriage |
| CN209023161U (en) * | 2018-09-25 | 2019-06-25 | 陕西飞机工业（集团）有限公司 | A kind of aircraft slow test nose-gear loaded member |
| JP2019109175A (en) * | 2017-12-20 | 2019-07-04 | 三菱航空機株式会社 | Strength testing system |
| CN113212795A (en) * | 2021-06-04 | 2021-08-06 | 中国飞机强度研究所 | Loading false wheel for undercarriage strength test |
-
2021
- 2021-11-10 CN CN202111325747.0A patent/CN113800005A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4392623A (en) * | 1980-12-22 | 1983-07-12 | The Boeing Company | Fused connection adapted to fail under different overloads acting in different directions |
| CN103630346A (en) * | 2013-11-19 | 2014-03-12 | 北京航空航天大学 | Test table and test method of emergency breaking off pin of undercarriage |
| CN103604593A (en) * | 2013-11-21 | 2014-02-26 | 南京航空航天大学 | Plane type ejection launch carrier-borne aircraft front undercarriage static force test loading apparatus |
| CN103604619A (en) * | 2013-11-21 | 2014-02-26 | 南京航空航天大学 | Front undercarriage static force test loading device for frame-type catapult-assisted take-off shipboard aircraft |
| JP2019109175A (en) * | 2017-12-20 | 2019-07-04 | 三菱航空機株式会社 | Strength testing system |
| CN209023161U (en) * | 2018-09-25 | 2019-06-25 | 陕西飞机工业（集团）有限公司 | A kind of aircraft slow test nose-gear loaded member |
| CN113212795A (en) * | 2021-06-04 | 2021-08-06 | 中国飞机强度研究所 | Loading false wheel for undercarriage strength test |
Non-Patent Citations (1)
| Title |
|---|
| 杜瑞;许锋;: "大型民用飞机起落架应急断离分析与仿真", 机械设计与制造工程, no. 12, pages 38 - 42 * |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114778056A (en) * | 2022-06-13 | 2022-07-22 | 中国飞机强度研究所 | Test system and method for dynamic impact mechanical property test of light airplane structure |
| CN114778056B (en) * | 2022-06-13 | 2022-09-02 | 中国飞机强度研究所 | Test system and method for testing dynamic impact mechanical property of light airplane structure |
| CN117022668A (en) * | 2023-10-08 | 2023-11-10 | 中国飞机强度研究所 | Device and method for continuous cable passing and obstacle crossing test of aircraft landing gear |
| CN117022668B (en) * | 2023-10-08 | 2024-01-05 | 中国飞机强度研究所 | Device and method for continuous cable passing and obstacle crossing test of aircraft landing gear |
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Application publication date: 20211217 |