CN103543029A - Monitoring analysis method for fatigue test for main control system of whole plane - Google Patents
Monitoring analysis method for fatigue test for main control system of whole plane Download PDFInfo
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Abstract
The invention belongs to the technical field of plane main control system fatigue tests and fatigue life tests, and relates to a monitoring analysis method for a fatigue test for a main control system of a whole plane. According to the method, in general, the system state and fatigue loading can be effectively monitored in a fatigue testing process of the main control system of the whole plane, the process of the fatigue test for the main control system of the whole plane is quickened, test quality is guaranteed, and the phenomenon that it is found out that the fatigue test is in vain afterwards and needs to be done for supplement or a fatigue test cycle of the system is repeated or a block spectrum test cycle of the system is repeated is avoided.
Description
Technical field
The invention belongs to flight control system torture test and tired determine longevity technical field, relating to a kind of monitoring analysis method for the torture test of aircraft primary flight control system.
Background technology
Because test condition and experimental scale are limit, the torture test of existing aircraft primary flight control system, all do not relate to and use control loads and the displacement spectra of primary flight control system, in addition, do not consider Aircraft Main control system state issues, do not consider with airframe tired with may well ask topic yet, therefore, after the test check of its corresponding flight control system torture test is generally the proof cycle requiring at test mission book, or complete after regulation test period piece spectrum, or there is obvious testing failure and stop after examination to carry out, test monitoring content in its torture test process and form are only loading curve tolerance or the degree of closeness aspect of torture test loading and test load value of feedback.Visible, existing flight control system torture test inspection or monitoring means can not meet the monitoring requirement of the primary flight control system torture test of implementing primary flight control system control loads and displacement spectra, examination control system spare part and supporter fatigue behaviour thereof and control system static performance index.
Summary of the invention
Order of the present invention: a kind of analytical approach that can carry out effective monitoring for the aircraft primary flight control system torture test (system state and fatigue loading) of implementing control loads and displacement spectra is provided, thereby in the system state that relates to Aircraft Main control system, in passenger cabin jociey stick and left and right pedal fatigue loading and with the tired situation with examination of airframe under, improve aircraft primary flight control system torture test progress and guarantee test mass, reach comprehensively and actual examination aircraft primary flight control system spare part and supporter fatigue behaviour thereof, examine the object of Aircraft Main maneuverability static system performance index simultaneously, avoid doing for supplement because of invalidate the test the phenomenon of tired spectrum piece test afterwards.
Technical solution of the present invention:
(1) the system state of aircraft primary flight control system and fatigue loading debugging is complete and primary flight control system torture test stable normal after, for control loads and the displacement spectra of aircraft primary flight control system, gather respectively the control loads of aircraft aileron, horizontal tail and rudder control system and the real-time loading curve data of displacement;
(2) in real-time loading curve data, choose jociey stick and left and right pedal Loading Control point initial value curve and data, typical real-time loading curve maximal value and real-time loading curve shape feature as typical case or effective real-time loading curve data, meanwhile, by the real-time loading curve data of collection, according to reference mark manipulation displacement, form, the reference mark control loads with displacement controlled quentity controlled variable provides with the form of load monitoring amount;
(3) in the full machine system of aircraft torture test control loads and displacement spectra, arrange and determine " 15+2 " individual typical monitoring/measurement point, wherein 15 typical monitoring/measurement points are 15 typical ground maintenances and ground actual measurement load condition: the ground actual measurement aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick right avertence limit, before ground actual measurement horizontal tail jociey stick under the large arm of force state of horizontal tail, push away the limit, the ground actual measurement horizontal tail jociey stick post-tensioning limit, before ground maintenance horizontal tail jociey stick, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit, before ground maintenance horizontal tail jociey stick under the little arm of force state of horizontal tail, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit, ground maintenance yaw rudder left foot under the large speed state of yaw rudder is the limit forward, ground maintenance yaw rudder right crus of diaphragm is the limit forward, ground actual measurement yaw rudder left foot under the little speed state of yaw rudder is the limit forward, ground actual measurement yaw rudder right crus of diaphragm is the limit forward, ground maintenance yaw rudder left foot is the limit forward, ground maintenance yaw rudder right crus of diaphragm is the limit forward, 2 typical monitoring/measurement points in " 15+2 " are 2 typical ground maintenance brake load situations: the little brake situation of ground maintenance under the large speed state of rudder control system and the large brake situation of ground maintenance under the little speed state of rudder control system.Respectively " 15+2 " individual typical case monitoring/measurement point is verified to load test, to determine that the real-time loading curve data of the individual typical case of correspondence " 15+2 " monitoring/measurement point is Loading Control point initial value curve and data, typical control loads and the peaked range of stability of displacement real-time loading curve data, and the loading curve shape facility that characterizes intrinsic shape feature and the loading linear degree of control loads/displacement real-time loading curve, thereby form the system state of aircraft primary flight control system torture test and the template of fatigue loading monitoring analysis;
(4) in aircraft primary flight control system torture test process, real-time loading curve data and its template data of the collection of corresponding typical case's monitoring/measurement point are contrasted respectively, in time primary flight control system state and fatigue loading are implemented to effective monitoring; If data are not in the range of stability of template, at torture test scene, the control system of relevant range and support position thereof are checked, if data in the range of stability of template, continue the aircraft primary flight control system torture test of current test period or carry out the aircraft primary flight control system torture test of next test period.
The advantage that the present invention has and beneficial effect: compared with prior art, the present invention can carry out effective monitoring to its system state and fatigue loading in the primary flight control system torture test process of implementing aircraft handling load and displacement spectra, can find in time and process the relevant system state that occurs in torture test and all kinds of testing failures and the problem of fatigue loading aspect, avoid finding to do for supplement or repeat because of invalidate the test the phenomenon of full machine system cycle or the test of piece spectrum afterwards, guaranteed test mass, meanwhile, greatly improved test progress.
The monitoring analysis method of aircraft primary flight control system fatigue test system state provided by the invention and fatigue loading successful Application in the normal operating situation 14000 pilot time torture test of aircraft primary flight control system and prove its torture test method for supervising practicality and effect very obvious.The primary flight control system torture test initial stage before implementing the present invention causes having doed for supplement this periodic test because of system state mistake afterwards, but after enforcement the present invention, in the primary flight control system torture test of nearly 14000 pilot time (for up to for many years), owing to can effectively testing monitoring, find in time and processed testing failure or problem that various relevant system states or test load, do not occur again causing its primary flight control system torture test cycle or composing the phenomenon that piece invalidate the test is had a penalty heat afterwards because of the mistake of system state or fatigue loading aspect.
The present invention can be for aircraft primary flight control system routine tests and primary flight control system and two kinds of situations of the same examination of organism fatigue, wherein, control system loading spectrum used is primary flight control system control loads and the displacement spectra of comprehensively dispatch from the factory tune-up data, ground maintenance data and control system measured data; Aircraft horizontal tail control system can divide greatly/in large/in/little four kinds of arm of force states; Can divide greatly/little two kinds of speed states of rudder control system.Visible, the scope of application of the present invention is large, can meet comprehensive assessment control system spare part and supporter fatigue behaviour thereof and Aircraft Main control system static performance index, embody the monitoring analysis requirement of general aircraft primary flight control system torture test that flight control system is the system fatigue behaviour of structure You Shi mechanism.
In addition, the present invention is directed to the aircraft primary flight control system torture test of the rigid nonreversible control of common mechanical, and its primary flight control system fatigue load spectrum form is control loads and displacement spectra, the monitoring analysis of applicable all kinds of Aircraft Main control system torture test, every general aircraft/mechanical backup/fax aircraft control stick (dish) that relates to is to full machine, parts or the spare part torture test monitoring analysis of primary control circuit between load simulated device, can therefrom use for reference or directly quote, therefore, the present invention has larger application and promotional value.
Accompanying drawing explanation
Fig. 1 is aircraft primary flight control system fatigue test system state and fatigue loading monitoring analysis process flow diagram.
Embodiment
The first step gathers the real-time loading curve data of control loads/displacement
The system state of aircraft primary flight control system and fatigue loading debugging is complete and the torture test of full machine system stable normal after, for control loads and the displacement spectra of aircraft primary flight control system, gather respectively the control loads of aircraft aileron, horizontal tail and rudder control system and the real-time loading curve data of displacement;
The object of this step is to gather the real-time loading curve data that represents control loads/displacement in aircraft primary flight control system torture test process, because in the aircraft primary flight control system torture test that relates to control loads and displacement spectra, manipulation displacement and control loads apply simultaneously, and dynamic fatigue loads.
Second step is chosen effective real-time loading curve data
In above-mentioned real-time loading curve data, choose jociey stick and left and right pedal Loading Control point initial value curve and data, typical real-time loading curve maximal value and real-time loading curve shape feature as typical case or effective real-time loading curve data of primary flight control system torture test monitoring analysis, meanwhile, by the real-time loading curve data of collection, according to reference mark manipulation displacement, form, the reference mark control loads with displacement controlled quentity controlled variable provides with the form of load monitoring amount;
The present invention that this step is clear and definite is to the effective typical real-time loading curve data of primary flight control system torture test monitoring analysis (being called for short real-time loading curve data)---jociey stick or left and right pedal Loading Control point initial value curve and data, typical control loads/displacement real-time loading curve and maximal value, simultaneously clear and definite real-time loading curve data is gathered to requirement, reference mark is handled displacement and is provided with the form of load monitoring amount with form, the reference mark control loads of displacement controlled quentity controlled variable; Should be noted that loading curve shape facility refers to intrinsic shape feature and the loading linear degree that characterizes relevant control loads/displacement real-time loading curve.At primary flight control system state and fatigue loading commissioning test and normal stable in the situation that, the shape facility of control loads/displacement real-time loading curve of jociey stick or left and right pedal Loading Control point fatigue loading is intrinsic.
The 3rd step arranges typical case's monitoring/measurement point and formulates corresponding real-time loading curve data template
In the full machine system of aircraft torture test control loads and displacement spectra, arrange and determine " 15+2 " individual typical monitoring/measurement point, wherein 15 typical monitoring/measurement points are 15 typical ground maintenances and ground actual measurement load condition: the ground actual measurement aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick right avertence limit, before ground actual measurement horizontal tail jociey stick under the large arm of force state of horizontal tail, push away the limit, the ground actual measurement horizontal tail jociey stick post-tensioning limit, before ground maintenance horizontal tail jociey stick, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit, before ground maintenance horizontal tail jociey stick under the little arm of force state of horizontal tail, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit, ground maintenance yaw rudder left foot under the large speed state of yaw rudder is the limit forward, ground maintenance yaw rudder right crus of diaphragm is the limit forward, ground actual measurement yaw rudder left foot under the little speed state of yaw rudder is the limit forward, ground actual measurement yaw rudder right crus of diaphragm is the limit forward, ground maintenance yaw rudder left foot is the limit forward, ground maintenance yaw rudder right crus of diaphragm is the limit forward, 2 typical monitoring/measurement points in " 15+2 " are 2 typical ground maintenance brake load situations: the little brake situation of ground maintenance under the large speed state of rudder control system and the large brake situation of ground maintenance under the little speed state of rudder control system.Relevant " 15+2 " individual typical monitoring/measurement point refers to shown in table 1.
Respectively " 15+2 " individual typical case monitoring/measurement point is verified to load test, to determine that the real-time loading curve data of the individual typical case of correspondence " 15+2 " monitoring/measurement point is Loading Control point initial value curve and data, typical control loads and the peaked range of stability of displacement real-time loading curve data, and the loading curve shape facility that characterizes intrinsic shape feature and the loading linear degree of control loads/displacement real-time loading curve, form the system state of aircraft primary flight control system torture test and the template of fatigue loading monitoring analysis.
Typical monitoring/the measurement point of table 1 " 15+2 "
Load condition explanation in table 1: ground actual measurement or ground maintenance refer to that the spectrum in aircraft primary flight control system torture test control loads and displacement spectra carries situation; Aileron jociey stick or horizontal tail jociey stick refer to that promoting jociey stick handles aileron or handle horizontal tail, and yaw rudder left foot or yaw rudder right crus of diaphragm refer to left foot or right crus of diaphragm steering rudder.System state explanation in table 1: the large arm of force refers to that horizontal tail control system is large arm of force state; The little arm of force refers to that horizontal tail control system is little arm of force state; Large speed refers to that rudder control system is large speed state or carries greatly state; Little speed refer to rudder control system be little speed state or in carry state.
The 4th step is carried out primary flight control system torture test monitoring analysis
In aircraft primary flight control system torture test process, real-time loading curve data and its template data of the collection of corresponding typical case's monitoring/measurement point are contrasted respectively, in time the system state of primary flight control system and fatigue loading are implemented to effective monitoring; If data are not in the range of stability of template, at torture test scene, the control system of relevant range and support position thereof are checked, if data in the range of stability of template, continue the aircraft primary flight control system torture test of current test period or carry out the aircraft primary flight control system torture test of next test period.
Embodiment mono-(the independent torture test of aircraft primary flight control system of take is carried out monitoring analysis as example)
One, gather the real-time loading curve data of control loads/displacement: before aircraft primary flight control system torture test official test, when the debugging of primary flight control system state and fatigue loading is complete and full machine system torture test commissioning test stable normal after, for control loads and the displacement spectra of aircraft primary flight control system, respectively the real-time loading curve data of the control loads/displacement of aircraft aileron, horizontal tail and rudder control system is gathered.
Two, choose effective real-time loading curve data: the control loads/displacement real-time loading curve data gathering is analyzed to judgement, determine the validity of its real-time loading curve data to primary flight control system fatigue test system state and fatigue loading monitoring analysis, finally choose jociey stick and left and right pedal Loading Control point initial value curve and data, typical real-time loading curve maximal value and real-time loading curve shape feature as typical case or effective real-time loading curve data; Meanwhile, by the real-time loading curve data of collection, according to reference mark manipulation displacement, form, the reference mark control loads with displacement controlled quentity controlled variable provides with the form of load monitoring amount.
Three, typical case's monitoring/measurement point be set and formulate corresponding real-time loading curve data template: in aircraft primary flight control system torture test control loads and displacement spectra, arrange and determine " 15+2 " individual typical monitoring/measurement point, referring to table 1.Respectively " 15+2 " individual typical case monitoring/measurement point is verified to load test, to determine that the real-time loading curve data of the individual typical case of correspondence " 15+2 " monitoring/measurement point is Loading Control point initial value curve and data, typical control loads and the peaked range of stability of displacement real-time loading curve data, and the loading curve shape facility that characterizes intrinsic shape feature and the loading linear degree of control loads/displacement real-time loading curve, thereby form the system state of aircraft primary flight control system torture test and the template of fatigue loading monitoring analysis.
Four, carry out aircraft primary flight control system torture test monitoring analysis: in aircraft primary flight control system torture test process, real-time loading curve data and its template data of the collection of corresponding typical case's monitoring/measurement point are contrasted respectively, in time full machine system state and fatigue loading are implemented to effective monitoring; If data are not in the range of stability of template, at torture test scene, the control system of relevant range and support position thereof are checked, if data in the range of stability of template, continue the aircraft primary flight control system torture test of current test period or carry out the aircraft primary flight control system torture test of next test period.
Embodiment bis-(take primary flight control system and organism fatigue with trying to carry out monitoring analysis as example)
One, gather and determine effective real-time loading curve data: in aircraft primary flight control system fatigue test system state and fatigue loading monitoring analysis process about the real-time loading curve data that gathers control loads/displacement is with to choose effective real-time loading curve data step identical with the step 1～step 2 of above-described embodiment one.
Two, typical case's monitoring/measurement point be set and formulate corresponding real-time loading curve data template: the setting of typical case's monitoring/measurement point is identical with the step 3 of above-described embodiment one, in primary flight control system control loads and displacement spectra, " 15+2 " individual typical monitoring/measurement point is set and formulates the system state of corresponding aircraft primary flight control system torture test and the template of fatigue loading monitoring analysis.
Three, carry out aircraft primary flight control system torture test monitoring analysis: the monitoring analysis process of aircraft primary flight control system torture test is identical with the step 4 of above-described embodiment one, in aircraft primary flight control system torture test process, real-time loading curve data and its template data of the collection of corresponding typical case's monitoring/measurement point are contrasted respectively, in time full machine system state and fatigue loading are implemented to effective monitoring; If data are not in the range of stability of template, at torture test scene, the control system of relevant range and support position thereof are checked, if data in the range of stability of template, continue the aircraft primary flight control system torture test of current test period or carry out the aircraft primary flight control system torture test of next test period.
Should be noted that, problem or fault handling when the primary flight control system torture test problem of trying together from organism fatigue or fault handling and primary flight control system torture test are carried out are separately slightly different, when data are not in the range of stability in template or when abnormal, should suspend airframe torture test, guaranteeing safe in the situation that, the control system of relevant range and support position thereof are checked, find in time, get rid of test problem or fault, continue the primary flight control system torture test of current period or piece spectrum, until complete aircraft primary flight control system fatigue life test.In addition, because the torture test of aircraft primary flight control system and airframe torture test are carried out on same airplane fatigue tester, organism fatigue loads has certain " fluctuation " impact to Aircraft Main control system, may when monitoring analysis, there is once in a while real-time loading curve data slightly to exceed the phenomenon of template (range of stability), this belongs to normal phenomenon, and when the torture test of aircraft primary flight control system is carried out separately, the template data of formulation is still effective.
Claims (3)
1. for a monitoring analysis method for aircraft primary flight control system torture test, it is characterized in that, torture test monitoring analysis method is following steps:
(1) the system state of aircraft primary flight control system and fatigue loading debugging is complete and primary flight control system torture test stable normal after, for control loads and the displacement spectra of aircraft primary flight control system, gather respectively the control loads of aircraft aileron, horizontal tail and rudder control system and the real-time loading curve data of displacement;
(2) in real-time loading curve data, choose jociey stick and left and right pedal Loading Control point initial value curve and data, typical real-time loading curve maximal value and real-time loading curve shape feature, meanwhile, by the real-time loading curve data of collection, according to reference mark manipulation displacement, form, the reference mark control loads with displacement controlled quentity controlled variable provides with the form of load monitoring amount;
(3) in aircraft primary flight control system torture test control loads and displacement spectra, arrange and determine " 15+2 " individual typical monitoring/measurement point, wherein 15 typical monitoring/measurement points are 15 typical ground actual measurements and ground maintenance load condition: the ground actual measurement aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick right avertence limit, before ground actual measurement horizontal tail jociey stick under the large arm of force state of horizontal tail control system, push away the limit, the ground actual measurement horizontal tail jociey stick post-tensioning limit, before ground maintenance horizontal tail jociey stick, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit, before ground maintenance horizontal tail jociey stick under the little arm of force state of horizontal tail control system, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit, ground maintenance yaw rudder left foot under the large speed state of rudder control system is pedaled the limit forward, ground maintenance yaw rudder right pedal is the limit forward, ground actual measurement yaw rudder left foot under the little speed state of rudder control system is pedaled the limit forward, ground actual measurement yaw rudder right pedal is the limit forward, ground maintenance yaw rudder left foot is pedaled the limit forward, ground maintenance yaw rudder right pedal is the limit forward, 2 typical monitoring/measurement points in " 15+2 " are 2 typical ground maintenance brake load situations: the little brake situation of ground maintenance under the large speed state of rudder control system and the large brake situation of ground maintenance under the little speed state of rudder control system.Respectively " 15+2 " individual typical case monitoring/measurement point is verified to load test, to determine that the real-time loading curve data of the individual typical case of correspondence " 15+2 " monitoring/measurement point is Loading Control point initial value curve and data, typical control loads and the peaked range of stability of displacement real-time loading curve data, and characterize the intrinsic shape of real-time loading curve and the loading curve shape facility of loading linear degree, form the template to the system state of aircraft primary flight control system torture test and fatigue loading monitoring analysis;
(4) in aircraft primary flight control system torture test process, real-time loading curve data and its template data of the collection of corresponding typical case's monitoring/measurement point are contrasted respectively, in time primary flight control system state and fatigue loading are implemented to effective monitoring; If data are not in the range of stability of template, at torture test scene, the control system of relevant range and support position thereof are checked, if data in the range of stability of template, continue the aircraft primary flight control system torture test of current test period or carry out the aircraft primary flight control system torture test of next test period.
2. a kind of monitoring analysis method for the torture test of aircraft primary flight control system according to claim 1, it is characterized in that, described " 15+2 " individual typical monitoring/measurement point that arranges in the full machine system of aircraft torture test control loads and displacement spectra and determine, for the debugging of dispatching from the factory based on Aircraft Main control system, ground maintenance, and general aircraft primary flight control system control loads and the displacement spectra of the data organization such as Aircraft Main control system control loads and displacement actual measurement arrange, monitoring analysis for general primary flight control system torture test, monitoring analysis for the Aircraft Main control system torture test of component-level or component level, by the setting of aircraft primary flight control system torture test control loads and displacement spectra and " 15+2 " individual typical case monitoring/measurement point, or the part typical case monitoring/measurement point of choosing wherein combines and is transformed into the control system on described parts or assembly, being about to jociey stick and pedal control loads and displacement spectra is control loads and the displacement spectra of described parts or assembly control system by control system gear ratio relationship conversion, the corresponding load condition of corresponding all or part of typical case's monitoring/measurement point is constant, monitoring analysis method to described component-level or the torture test of component level control system is identical with the monitoring analysis method of aircraft primary flight control system torture test.
3. a kind of monitoring analysis method for the torture test of aircraft primary flight control system according to claim 1, it is characterized in that, described " 15+2 " individual typical monitoring/measurement point that arranges in the full machine system of aircraft torture test control loads and displacement spectra and determine, typical case's monitoring/measurement point is wherein chosen or combined, for the monitoring analysis of each unidirectional primary control circuit torture test of aileron, horizontal tail or yaw rudder:
(1) the typical case's monitoring/measurement point for the monitoring analysis of Aileron control system torture test is, the ground actual measurement aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick left avertence limit, the ground maintenance aileron jociey stick right avertence limit;
(2) the typical case's monitoring/measurement point for the monitoring analysis of horizontal tail control system torture test is, before pushing away the limit, the ground actual measurement horizontal tail jociey stick post-tensioning limit, ground maintenance horizontal tail jociey stick before ground actual measurement horizontal tail jociey stick under large arm of force state, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit, before the ground maintenance horizontal tail jociey stick under the little arm of force state of horizontal tail, push away the limit, the ground maintenance horizontal tail jociey stick post-tensioning limit;
(3) the typical case's monitoring/measurement point for the monitoring analysis of rudder control system torture test is, ground maintenance yaw rudder left foot under the large speed state of yaw rudder is the limit forward, ground maintenance yaw rudder right crus of diaphragm is the limit forward, ground actual measurement yaw rudder left foot under the little speed state of yaw rudder is the limit forward, ground actual measurement yaw rudder right crus of diaphragm is the limit forward, ground maintenance yaw rudder left foot is the limit forward, ground maintenance yaw rudder right crus of diaphragm is the limit forward, the little brake situation of ground maintenance under the large speed state of yaw rudder and the large brake situation of ground maintenance under the little speed state of yaw rudder.
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Cited By (10)
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|---|---|---|---|---|
| CN105571839A (en) * | 2014-10-11 | 2016-05-11 | 中国航空工业集团公司西安飞机设计研究所 | Continuous variable-angle cyclic loading 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 |
| 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 |
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| CN113942659A (en) * | 2021-09-16 | 2022-01-18 | 中国航空工业集团公司西安飞机设计研究所 | Airplane left and right pedal double-foot brake fatigue test and monitoring method thereof |
| CN113955145A (en) * | 2021-09-16 | 2022-01-21 | 中国航空工业集团公司西安飞机设计研究所 | Fatigue test state monitoring and troubleshooting method for main control system of airplane |
| CN115639805A (en) * | 2022-12-24 | 2023-01-24 | 中国航空工业集团公司西安飞机设计研究所 | Airborne PHM system state monitoring function test method |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS55119038A (en) * | 1979-03-08 | 1980-09-12 | Sundstrand Corp | Tester for motive power machine |
| JP2008269455A (en) * | 2007-04-24 | 2008-11-06 | Ihi Corp | Vibration control-system machine evaluation device and method |
| CN102279101A (en) * | 2011-07-13 | 2011-12-14 | 北京航空航天大学 | Six-dimension force high-frequency fatigue testing machine and method for using same |
| 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 |
| CN103043225A (en) * | 2012-12-10 | 2013-04-17 | 中国飞机强度研究所 | Load system and method combining force control and position control and used for airplane structural strength tests |
-
2013
- 2013-11-05 CN CN201310543028.5A patent/CN103543029B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS55119038A (en) * | 1979-03-08 | 1980-09-12 | Sundstrand Corp | Tester for motive power machine |
| JP2008269455A (en) * | 2007-04-24 | 2008-11-06 | Ihi Corp | Vibration control-system machine evaluation device and method |
| CN102279101A (en) * | 2011-07-13 | 2011-12-14 | 北京航空航天大学 | Six-dimension force high-frequency fatigue testing machine and method for using same |
| 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 |
| CN103043225A (en) * | 2012-12-10 | 2013-04-17 | 中国飞机强度研究所 | Load system and method combining force control and position control and used for airplane structural strength tests |
Non-Patent Citations (1)
| Title |
|---|
| 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》 * |
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