CN113955145A - Fatigue test state monitoring and troubleshooting method for main control system of airplane - Google Patents
Fatigue test state monitoring and troubleshooting method for main control system of airplane Download PDFInfo
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- CN113955145A CN113955145A CN202111087378.6A CN202111087378A CN113955145A CN 113955145 A CN113955145 A CN 113955145A CN 202111087378 A CN202111087378 A CN 202111087378A CN 113955145 A CN113955145 A CN 113955145A
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Abstract
A fatigue test state monitoring and troubleshooting method for a main control system of an airplane comprises the following steps: determining a key action part in a fatigue test of a main control system of the airplane; marking typical action positions of the actions of the key action parts; setting a monitoring camera, and monitoring whether the action position of the key action part is consistent with the applied manipulation load and displacement coordination loading spectrum or not based on the marked typical action position; if the state of the main control system of the airplane is monitored to be abnormal, faults in the aspect of loading control are checked; after the faults in the aspect of loading control are eliminated, the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated; and after the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated, the faults in the aspect of the structure of the main control system of the airplane are eliminated.
Description
Technical Field
The application belongs to the technical field of fatigue test state monitoring and troubleshooting of an aircraft main control system, and particularly relates to a fatigue test state monitoring and troubleshooting method of the aircraft main control system.
Background
The airplane main control system applies airplane control load and displacement coordination loading spectrum to carry out fatigue control loading on the airplane main control system and carry out fatigue test on the airplane main control system, and the method is an important means for examining the fatigue life and performance indexes of the airplane main control system.
When the fatigue test is carried out on the main control system of the airplane, if the action of the main control system of the airplane does not accord with the applied control load and displacement coordination loading spectrum, the state of the main control system of the airplane is abnormal, the fatigue test on the main control system of the airplane is invalid, if the state of the main control system of the airplane cannot be found in time, the test period can be greatly prolonged, a large amount of manpower and material resources are wasted, and even safety accidents occur.
At present, in a fatigue test of an aircraft main control system, whether the state of the aircraft main control system is abnormal or not is mainly checked through on-site visual inspection or manual inspection, however, under the condition that the aircraft main control system and an airframe are tested in a same way, an aircraft control load and displacement coordinated loading spectrum is applied in a cockpit, a cockpit hatch, an airframe and a wing are all closed, whether the state of the aircraft main control system is abnormal or not can not be checked through on-site visual inspection or manual inspection, and the abnormality of the state of the aircraft main control system cannot be found in time.
In addition, in the fatigue test of the main control system of the airplane, for the condition of checking the state of the main control system of the airplane to be abnormal, an effective means for reliably and efficiently positioning a fault area is lacked, and the test period is prolonged to different degrees.
The present application has been made in view of the above-mentioned technical drawbacks.
It should be noted that the above background disclosure is only for the purpose of assisting understanding of the inventive concept and technical solutions of the present invention, and does not necessarily belong to the prior art of the present patent application, and the above background disclosure should not be used for evaluating the novelty and inventive step of the present application without explicit evidence to suggest that the above content is already disclosed at the filing date of the present application.
Disclosure of Invention
The invention aims to provide a fatigue test state monitoring method for an aircraft main control system and a troubleshooting method thereof, so as to overcome or alleviate the technical defects of at least one aspect known to exist.
The technical scheme of the application is as follows:
in one aspect, a method for monitoring fatigue test state of a main control system of an aircraft is provided, which comprises the following steps:
determining a key action part in a fatigue test of a main control system of the airplane;
marking typical action positions of the actions of the key action parts;
setting a monitoring camera, and monitoring whether the action position of the key action part is consistent with the applied manipulation load and displacement coordination loading spectrum or not based on the marked typical action position;
if the action position of the key action part is consistent with the applied control load and displacement coordination loading spectrum, the state of the main control system of the airplane is normal;
and if the action position of the key action part does not accord with the applied control load and displacement coordination loading spectrum, indicating that the state of the main control system of the airplane is abnormal.
In some optional embodiments, in the method for monitoring the fatigue test state of the main control system of the aircraft, the critical action parts include:
an aileron steering rod of an aileron control system and a simulated aileron booster control dummy piece;
a horizontal tail steering rod of a horizontal tail control system, a simulated horizontal tail booster control dummy piece and a horizontal tail force arm adjusting device;
left and right pedals of a rudder control system, a simulated rudder composite steering engine control dummy piece and a rudder deflection angle limiter.
In some optional embodiments, in the method for monitoring the fatigue test state of the main control system of the aircraft, the typical action positions of the action of the aileron pilot lever are a neutral position, a left deviation limit position and a right deviation limit position of a loading control point;
the typical action position simulating the action of the control dummy piece of the aileron booster is an extension position of a piston rod of the aileron booster corresponding to a neutral position, a left deviation limit position and a right deviation limit position of an aileron steering rod;
typical action positions of the horizontal tail driving rod action are a neutral position of a loading control point of the horizontal tail driving rod, and a forward pushing limit position and a backward pulling limit position corresponding to the conditions of large, medium and small force arms of the horizontal tail;
the typical action position for simulating the action of the horizontal tail booster operating dummy piece is an extension position of a piston rod of which corresponds to a neutral position, a forward pushing limit position and a backward pulling limit position of a horizontal tail driving rod;
typical action positions of the horizontal tail force arm adjusting device are large, medium and small force arm positions;
typical action positions of the left and right foot pedals are a neutral position of a loading control point of the left and right foot pedals and forward limit positions corresponding to large and small speed states of the rudder;
the typical action position of the simulated rudder composite steering engine operating dummy member is an extension position of a piston rod of the simulated rudder composite steering engine corresponding to a neutral position and a forward limit position of a left pedal and a right pedal;
the typical operating position of the rudder deflection angle limiter is a deflection limiting position of the rudder at a large and a small speed of the extending end.
In another aspect, a method for troubleshooting a fatigue test of a main control system of an aircraft is provided, which comprises the following steps:
monitoring the state of the main control system of the airplane based on any one of the fatigue test state monitoring methods for the main control system of the airplane;
if the state of the main control system of the airplane is monitored to be abnormal, faults in the aspect of loading control are checked;
after the faults in the aspect of loading control are eliminated, the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated;
and after the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated, the faults in the aspect of the structure of the main control system of the airplane are eliminated.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, the troubleshooting on the aspect of loading control is specifically as follows:
comparing and analyzing the real-time loading curve data of the control load and the displacement with the normal curve data;
checking whether the neutral position, the left deviation limit position and the right deviation limit position of the aileron steering rod, the neutral position, the forward pushing limit position and the backward pulling limit position of the horizontal tail steering rod, the neutral position and the forward limit position of a left pedal and a right pedal are correspondingly consistent with a zero position for controlling loading and a maximum control displacement;
if the real-time loading curve data of the control load and the displacement are consistent with the normal curve data, and the neutral position, the left deviation limit position, the right deviation limit position, the neutral position, the forward pushing limit position and the backward pulling limit position of the driving rod of the aileron, the neutral position, the forward pushing limit position and the backward pulling limit position of the horizontal tail driving rod, the neutral position and the forward limit position of the left pedal and the right pedal are correspondingly consistent with the zero position and the maximum control displacement for controlling loading, the fault in the aspect of loading control is eliminated;
wherein the content of the first and second substances,
the real-time loading curve data of the control load and the displacement comprises initial value data of an aileron steering column, a horizontal tail steering column and left and right pedal loading control points, the initial value data is collected at the aileron steering column, the horizontal tail steering column and the left and right pedal loading control points, the history curve characteristics of the control load and the displacement coordinated loading of the aileron, the horizontal tail and the rudder can be reflected, and the initial value change ranges of the aileron, the horizontal tail and the rudder can be reflected.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, troubleshooting on finished parts and fake parts of the main control system of the aircraft is specifically as follows:
and troubleshooting is carried out on the simulated aileron booster control dummy piece, the simulated horizontal tail booster control dummy piece, the horizontal tail force arm adjusting device, the simulated rudder composite steering engine control dummy piece and the rudder deflection angle limiter, wherein the action position is not consistent with the control load and displacement coordination loading spectrum.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, a failure in the aspect of a structure of the main control system of the aircraft is specifically:
according to the real-time loading curve data of the control load and the displacement, the area with the abnormal state of the main control system of the airplane is determined by combining visual observation, hand feeling and listening sound on the spot, and whether the parts in the area with the abnormal state of the main control system of the airplane have fatigue damage, abrasion gaps, loose connection and clamping stagnation or not is checked.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, if there are a plurality of areas where the state of the main control system of the aircraft is abnormal, the troubleshooting is performed one by one according to the trend of the main control system of the aircraft.
Drawings
Fig. 1 is a flowchart of a method for monitoring a fatigue test state of a main control system of an aircraft and troubleshooting the fatigue test state of the main control system of the aircraft according to an embodiment of the present application.
For the purpose of better illustrating the embodiments, certain features of the drawings may be omitted, enlarged or reduced, and do not represent the size of an actual product; further, the drawings are for illustrative purposes, and terms describing positional relationships are limited to illustrative illustrations only and are not to be construed as limiting the patent.
Detailed Description
In order to make the technical solutions and advantages of the present application clearer, the technical solutions of the present application will be further clearly and completely described in the following detailed description with reference to the accompanying drawings, and it should be understood that the specific embodiments described herein are only some of the embodiments of the present application, and are only used for explaining the present application, but not limiting the present application. It should be noted that, for convenience of description, only the parts related to the present application are shown in the drawings, other related parts may refer to general designs, and the embodiments and technical features in the embodiments in the present application may be combined with each other to obtain a new embodiment without conflict.
In addition, unless otherwise defined, technical or scientific terms used in the description of the present application shall have the ordinary meaning as understood by one of ordinary skill in the art to which the present application belongs. The terms "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", and the like used in the description of the present application, which indicate orientations, are used only to indicate relative directions or positional relationships, and do not imply that the devices or elements must have a specific orientation, be constructed and operated in a specific orientation, and when the absolute position of the object to be described is changed, the relative positional relationships may be changed accordingly, and thus, should not be construed as limiting the present application. The use of "first," "second," "third," and the like in the description of the present application is for descriptive purposes only to distinguish between different components and is not to be construed as indicating or implying relative importance. The use of the terms "a," "an," or "the" and similar referents in the context of describing the application is not to be construed as an absolute limitation on the number, but rather as the presence of at least one. The word "comprising" or "comprises", and the like, when used in this description, is intended to specify the presence of stated elements or items, but not the exclusion of other elements or items.
Further, it is noted that, unless expressly stated or limited otherwise, the terms "mounted," "connected," and the like are used in the description of the invention in a generic sense, e.g., connected as either a fixed connection or a removable connection or integrally connected; can be mechanically or electrically connected; they may be directly connected or indirectly connected through an intermediate medium, or they may be connected through the inside of two elements, and those skilled in the art can understand their specific meaning in this application according to the specific situation.
The present application is described in further detail below with reference to fig. 1.
In one aspect, a method for monitoring fatigue test state of a main control system of an aircraft is provided, which comprises the following steps:
determining a key action part in a fatigue test of a main control system of the airplane;
marking typical action positions of the actions of the key action parts, specifically marking the typical action positions by marking lines;
setting a monitoring camera, and monitoring whether the action position of the key action part is consistent with the applied manipulation load and displacement coordination loading spectrum or not based on the marked typical action position;
if the action position of the key action part is consistent with the applied control load and displacement coordination loading spectrum, the state of the main control system of the airplane is normal;
and if the action position of the key action part does not accord with the applied control load and displacement coordination loading spectrum, indicating that the state of the main control system of the airplane is abnormal.
For the method for monitoring the fatigue test state of the main control system of the aircraft disclosed in the above embodiment, it can be understood by those skilled in the art that the method selects the key action position of the main control system of the aircraft, marks the typical action position of the action of the key action position, judges whether the action position of the key action position is in accordance with the applied control load and displacement coordination loading spectrum by monitoring the real-time shot image of the camera based on the marked typical action position, judges the state abnormality of the main control system of the aircraft if the action position of the key action position is not in accordance with the applied control load and displacement coordination loading spectrum, has higher timeliness for monitoring the abnormality of the main control system of the aircraft, even if the main control system of the aircraft is in accordance with the fatigue test of the aircraft body, the cockpit hatch cover, the inspection and maintenance of the fuselage and the wing, and the bearing cover are all closed, the abnormity of the state of the main operating system of the airplane can be found in time, so that the test period can be shortened, the manpower and the material resources are saved, and the safety accidents are avoided.
In some optional embodiments, in the method for monitoring the fatigue test state of the main control system of the aircraft, the critical action parts include:
an aileron steering rod of an aileron control system and a simulated aileron booster control dummy piece;
a horizontal tail steering rod of a horizontal tail control system, a simulated horizontal tail booster control dummy piece and a horizontal tail force arm adjusting device;
left and right pedals of a rudder control system, a simulated rudder composite steering engine control dummy piece and a rudder deflection angle limiter.
In some optional embodiments, in the method for monitoring the fatigue test state of the main control system of the aircraft, the typical action positions of the action of the aileron pilot lever are a neutral position, a left deviation limit position and a right deviation limit position of a loading control point;
the typical action position simulating the action of the control dummy piece of the aileron booster is an extension position of a piston rod of the aileron booster corresponding to a neutral position, a left deviation limit position and a right deviation limit position of an aileron steering rod;
typical action positions of the horizontal tail driving rod action are a neutral position of a loading control point of the horizontal tail driving rod, and a forward pushing limit position and a backward pulling limit position corresponding to the conditions of large, medium and small force arms of the horizontal tail;
the typical action position for simulating the action of the horizontal tail booster operating dummy piece is an extension position of a piston rod of which corresponds to a neutral position, a forward pushing limit position and a backward pulling limit position of a horizontal tail driving rod;
typical action positions of the horizontal tail force arm adjusting device are large, medium and small force arm positions;
typical action positions of the left and right foot pedals are a neutral position of a loading control point of the left and right foot pedals and forward limit positions corresponding to large and small speed states of the rudder;
the typical action position of the simulated rudder composite steering engine operating dummy member is an extension position of a piston rod of the simulated rudder composite steering engine corresponding to a neutral position and a forward limit position of a left pedal and a right pedal;
the typical operating position of the rudder deflection angle limiter is a deflection limiting position of the rudder at a large and a small speed of the extending end.
In another aspect, a method for troubleshooting a fatigue test of a main control system of an aircraft is provided, which comprises the following steps:
monitoring the state of the main control system of the airplane based on any one of the fatigue test state monitoring methods for the main control system of the airplane;
if the state of the main control system of the airplane is monitored to be abnormal, faults in the aspect of loading control are checked;
after the faults in the aspect of loading control are eliminated, the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated;
and after the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated, the faults in the aspect of the structure of the main control system of the airplane are eliminated.
For the fatigue test troubleshooting method for the main aircraft control system disclosed in the above embodiment, it can be understood by those skilled in the art that the method monitors the state of the main aircraft control system based on the fatigue test state monitoring method for the main aircraft control system, and when the state of the main aircraft control system is monitored to be abnormal, a troubleshooting method is provided, and the failure is troubleshot from the level of the loading control, the finished product and the dummy of the main aircraft control system, and the main aircraft control system structure, so that the failure region can be quickly and accurately located, and the test region can be shortened.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, the troubleshooting on the aspect of loading control is specifically as follows:
comparing and analyzing the real-time loading curve data of the control load and the displacement with the normal curve data;
checking whether the neutral position, the left deviation limit position and the right deviation limit position of the aileron steering rod, the neutral position, the forward pushing limit position and the backward pulling limit position of the horizontal tail steering rod, the neutral position and the forward limit position of a left pedal and a right pedal are correspondingly consistent with a zero position for controlling loading and a maximum control displacement;
if the real-time loading curve data of the control load and the displacement are consistent with the normal curve data, and the neutral position, the left deviation limit position, the right deviation limit position, the neutral position, the forward pushing limit position and the backward pulling limit position of the driving rod of the aileron, the neutral position, the forward pushing limit position and the backward pulling limit position of the horizontal tail driving rod, the neutral position and the forward limit position of the left pedal and the right pedal are correspondingly consistent with the zero position and the maximum control displacement for controlling loading, the fault in the aspect of loading control is eliminated;
wherein the content of the first and second substances,
the real-time loading curve data of the control load and the displacement comprises initial value data of an aileron steering column, a horizontal tail steering column and left and right pedal loading control points, the initial value data is collected at the aileron steering column, the horizontal tail steering column and the left and right pedal loading control points, the history curve characteristics of the control load and the displacement coordinated loading of the aileron, the horizontal tail and the rudder can be reflected, and the initial value change ranges of the aileron, the horizontal tail and the rudder can be reflected.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, troubleshooting on finished parts and fake parts of the main control system of the aircraft is specifically as follows:
and troubleshooting is carried out on the simulated aileron booster control dummy piece, the simulated horizontal tail booster control dummy piece, the horizontal tail force arm adjusting device, the simulated rudder composite steering engine control dummy piece and the rudder deflection angle limiter, wherein the action position is not consistent with the control load and displacement coordination loading spectrum.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, a failure in the aspect of a structure of the main control system of the aircraft is specifically:
according to the real-time loading curve data of the control load and the displacement, the area with the abnormal state of the main control system of the airplane is determined by combining visual observation, hand feeling and listening sound on the spot, and whether the parts in the area with the abnormal state of the main control system of the airplane have fatigue damage, abrasion gaps, loose connection and clamping stagnation or not is checked.
In some optional embodiments, in the method for troubleshooting a fatigue test of a main control system of an aircraft, if there are a plurality of areas where the main control system of the aircraft is in an abnormal state, the troubleshooting may be performed one by one according to the trend of the main control system of the aircraft, or may be performed one by one from a cockpit to a tail direction.
The embodiments are described in a progressive manner in the specification, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other.
Having thus described the present application in connection with the preferred embodiments illustrated in the accompanying drawings, it will be understood by those skilled in the art that the scope of the present application is not limited to those specific embodiments, and that equivalent modifications or substitutions of related technical features may be made by those skilled in the art without departing from the principle of the present application, and those modifications or substitutions will fall within the scope of the present application.
Claims (8)
1. A method for monitoring fatigue test state of a main control system of an airplane is characterized by comprising the following steps:
determining a key action part in a fatigue test of a main control system of the airplane;
marking typical action positions of the actions of the key action parts;
setting a monitoring camera, and monitoring whether the action position of the key action part is consistent with the applied manipulation load and displacement coordination loading spectrum or not based on the marked typical action position;
if the action position of the key action part is consistent with the applied control load and displacement coordination loading spectrum, the state of the main control system of the airplane is normal;
and if the action position of the key action part does not accord with the applied control load and displacement coordination loading spectrum, indicating that the state of the main control system of the airplane is abnormal.
2. The aircraft main control system fatigue test state monitoring method of claim 1,
the key action parts comprise:
an aileron steering rod of an aileron control system and a simulated aileron booster control dummy piece;
a horizontal tail steering rod of a horizontal tail control system, a simulated horizontal tail booster control dummy piece and a horizontal tail force arm adjusting device;
left and right pedals of a rudder control system, a simulated rudder composite steering engine control dummy piece and a rudder deflection angle limiter.
3. The aircraft main control system fatigue test state monitoring method of claim 2,
typical action positions of the action of the aileron steering rod are a neutral position, a left deviation limit position and a right deviation limit position of a loading control point;
the typical action position simulating the action of the control dummy piece of the aileron booster is an extension position of a piston rod of the aileron booster corresponding to a neutral position, a left deviation limit position and a right deviation limit position of an aileron steering rod;
typical action positions of the horizontal tail driving rod action are a neutral position of a loading control point of the horizontal tail driving rod, and a forward pushing limit position and a backward pulling limit position corresponding to the conditions of large, medium and small force arms of the horizontal tail;
the typical action position for simulating the action of the horizontal tail booster operating dummy piece is an extension position of a piston rod of which corresponds to a neutral position, a forward pushing limit position and a backward pulling limit position of a horizontal tail driving rod;
typical action positions of the horizontal tail force arm adjusting device are large, medium and small force arm positions;
typical action positions of the left and right foot pedals are a neutral position of a loading control point of the left and right foot pedals and forward limit positions corresponding to large and small speed states of the rudder;
the typical action position of the simulated rudder composite steering engine operating dummy member is an extension position of a piston rod of the simulated rudder composite steering engine corresponding to a neutral position and a forward limit position of a left pedal and a right pedal;
the typical operating position of the rudder deflection angle limiter is a deflection limiting position of the rudder at a large and a small speed of the extending end.
4. A fatigue test troubleshooting method for a main control system of an airplane is characterized by comprising the following steps:
monitoring the state of the main control system of the airplane based on the fatigue test state monitoring method of the main control system of the airplane as claimed in any one of claims 1 to 3;
if the state of the main control system of the airplane is monitored to be abnormal, faults in the aspect of loading control are checked;
after the faults in the aspect of loading control are eliminated, the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated;
and after the faults in the aspects of finished parts and fake parts of the main control system of the airplane are eliminated, the faults in the aspect of the structure of the main control system of the airplane are eliminated.
5. The aircraft main control system fatigue test troubleshooting method of claim 4,
the method for checking the loading control faults specifically comprises the following steps:
comparing and analyzing the real-time loading curve data of the control load and the displacement with the normal curve data;
checking whether the neutral position, the left deviation limit position and the right deviation limit position of the aileron steering rod, the neutral position, the forward pushing limit position and the backward pulling limit position of the horizontal tail steering rod, the neutral position and the forward limit position of a left pedal and a right pedal are correspondingly consistent with a zero position for controlling loading and a maximum control displacement;
if the real-time loading curve data of the control load and the displacement are consistent with the normal curve data, and the neutral position, the left deviation limit position, the right deviation limit position, the neutral position, the forward pushing limit position and the backward pulling limit position of the driving rod of the aileron, the neutral position, the forward pushing limit position and the backward pulling limit position of the horizontal tail driving rod, the neutral position and the forward limit position of the left pedal and the right pedal are correspondingly consistent with the zero position and the maximum control displacement for controlling loading, the fault in the aspect of loading control is eliminated.
6. The aircraft main control system fatigue test troubleshooting method of claim 4,
the method is used for checking faults in finished parts and fake parts of the main control system of the airplane, and specifically comprises the following steps:
and troubleshooting is carried out on the simulated aileron booster control dummy piece, the simulated horizontal tail booster control dummy piece, the horizontal tail force arm adjusting device, the simulated rudder composite steering engine control dummy piece and the rudder deflection angle limiter, wherein the action position is not consistent with the control load and displacement coordination loading spectrum.
7. The aircraft main control system fatigue test troubleshooting method of claim 4,
the method is used for troubleshooting faults in the aspect of the structure of the main control system of the airplane, and specifically comprises the following steps:
according to the real-time loading curve data of the control load and the displacement, the area with the abnormal state of the main control system of the airplane is determined by combining visual observation, hand feeling and listening sound on the spot, and whether the parts in the area with the abnormal state of the main control system of the airplane have fatigue damage, abrasion gaps, loose connection and clamping stagnation or not is checked.
8. The aircraft main control system fatigue test troubleshooting method of claim 7,
if the areas with abnormal states of the main control system of the airplane have a plurality of places, the main control system of the airplane is checked one by one according to the trend of the main control system of the airplane.
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Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20050149238A1 (en) * | 2004-01-05 | 2005-07-07 | Arinc Inc. | System and method for monitoring and reporting aircraft quick access recorder data |
| US20100235108A1 (en) * | 2009-03-16 | 2010-09-16 | Adams David O | Usage monitor reliability factor using an advanced fatigue reliability assessment model |
| CN102680221A (en) * | 2012-05-11 | 2012-09-19 | 中国航空工业集团公司西安飞机设计研究所 | Fatigue test method for full-aircraft main operating system of aircraft |
| CN103543029A (en) * | 2013-11-05 | 2014-01-29 | 中国航空工业集团公司西安飞机设计研究所 | Monitoring analysis method for fatigue test for main control system of whole plane |
| CN105928695A (en) * | 2016-05-11 | 2016-09-07 | 中国矿业大学 | Fault diagnosis system and method for mechanical parts of small-size unmanned helicopter |
| US20170293712A1 (en) * | 2016-04-11 | 2017-10-12 | Airbus Helicopters Deutschland GmbH | Probabilistic load and damage modeling for fatigue life management |
| CN110987420A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | Operating force and operating displacement detection method |
| CN110987422A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | Component looseness checking method for aircraft control system dynamic fatigue test |
| CN110987419A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | System state control and implementation method for airplane fatigue test |
| CN110987421A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | Dynamic fatigue test support method for whole-machine main control system |
| CN111122200A (en) * | 2019-12-31 | 2020-05-08 | 中国航空工业集团公司西安飞机设计研究所 | On-site maintenance and inspection method for fatigue test of main control system of whole aircraft |
| CN112685846A (en) * | 2020-12-02 | 2021-04-20 | 徐工集团工程机械有限公司 | Load spectrum acquisition test method and system |
-
2021
- 2021-09-16 CN CN202111087378.6A patent/CN113955145A/en active Pending
Patent Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20050149238A1 (en) * | 2004-01-05 | 2005-07-07 | Arinc Inc. | System and method for monitoring and reporting aircraft quick access recorder data |
| US20100235108A1 (en) * | 2009-03-16 | 2010-09-16 | Adams David O | Usage monitor reliability factor using an advanced fatigue reliability assessment model |
| CN102680221A (en) * | 2012-05-11 | 2012-09-19 | 中国航空工业集团公司西安飞机设计研究所 | Fatigue test method for full-aircraft main operating system of aircraft |
| CN103543029A (en) * | 2013-11-05 | 2014-01-29 | 中国航空工业集团公司西安飞机设计研究所 | Monitoring analysis method for fatigue test for main control system of whole plane |
| US20170293712A1 (en) * | 2016-04-11 | 2017-10-12 | Airbus Helicopters Deutschland GmbH | Probabilistic load and damage modeling for fatigue life management |
| CN105928695A (en) * | 2016-05-11 | 2016-09-07 | 中国矿业大学 | Fault diagnosis system and method for mechanical parts of small-size unmanned helicopter |
| CN110987420A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | Operating force and operating displacement detection method |
| CN110987422A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | Component looseness checking method for aircraft control system dynamic fatigue test |
| CN110987419A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | System state control and implementation method for airplane fatigue test |
| CN110987421A (en) * | 2019-12-25 | 2020-04-10 | 中国航空工业集团公司西安飞机设计研究所 | Dynamic fatigue test support method for whole-machine main control system |
| CN111122200A (en) * | 2019-12-31 | 2020-05-08 | 中国航空工业集团公司西安飞机设计研究所 | On-site maintenance and inspection method for fatigue test of main control system of whole aircraft |
| CN112685846A (en) * | 2020-12-02 | 2021-04-20 | 徐工集团工程机械有限公司 | Load spectrum acquisition test method and system |
Non-Patent Citations (5)
| Title |
|---|
| 卢京明: "飞机全机主操纵系统与机体同时进行疲劳试验的联调测试及试验监控", 飞机设计, vol. 29, no. 3, pages 9 - 13 * |
| 卢京明: "飞机全机主操纵系统疲劳试验故障诊断与分析", 飞机设计, vol. 37, no. 3 * |
| 卢京明等: "飞机主操纵系统疲劳定寿", 飞机设计, vol. 40, no. 2 * |
| 卢京明等: "飞机主操纵系统疲劳试验载荷监控方法研究——应变数据分析法", 飞机设计, no. 2 * |
| 卢京明等: "飞机全机主操纵系统疲劳试验的零组件拆检及分析", 飞机设计, vol. 38, no. 1, pages 44 - 48 * |
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