CN112182875A - Fatigue design method for test-oriented helicopter rotor wing metal piece - Google Patents

Fatigue design method for test-oriented helicopter rotor wing metal piece Download PDF

Info

Publication number
CN112182875A
CN112182875A CN202011022238.6A CN202011022238A CN112182875A CN 112182875 A CN112182875 A CN 112182875A CN 202011022238 A CN202011022238 A CN 202011022238A CN 112182875 A CN112182875 A CN 112182875A
Authority
CN
China
Prior art keywords
fatigue
load
test
metal piece
rotor
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
CN202011022238.6A
Other languages
Chinese (zh)
Other versions
CN112182875B (en
Inventor
熊欣
刘牧东
赵凤帅
喻溅鉴
崔韦
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
China Helicopter Research and Development Institute
Original Assignee
China Helicopter Research and Development Institute
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by China Helicopter Research and Development Institute filed Critical China Helicopter Research and Development Institute
Priority to CN202011022238.6A priority Critical patent/CN112182875B/en
Publication of CN112182875A publication Critical patent/CN112182875A/en
Application granted granted Critical
Publication of CN112182875B publication Critical patent/CN112182875B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/04—Ageing analysis or optimisation against ageing
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Theoretical Computer Science (AREA)
  • Geometry (AREA)
  • General Physics & Mathematics (AREA)
  • Computer Hardware Design (AREA)
  • General Engineering & Computer Science (AREA)
  • Evolutionary Computation (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Pure & Applied Mathematics (AREA)
  • Mathematical Optimization (AREA)
  • Mathematical Analysis (AREA)
  • Computational Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Investigating Strength Of Materials By Application Of Mechanical Stress (AREA)

Abstract

The invention discloses a fatigue design method for a test-oriented helicopter rotor wing metal piece, which comprises the following steps: determining a proofing fatigue limit corresponding to a target life test, calculating the maximum loading load of the fatigue test carried out under different materials and friction and corrosion modes of the rotor wing metal part, and analyzing under the test load; according to the materials of the rotor moving parts in the inverse calculation fatigue test and the test proof fatigue limit under the abrasion mode state; taking stress ratio effect into consideration to carry out static load correction, calculating the equivalent dynamic load of each characteristic load proofing, and evaluating the fatigue life of the structure; and determining the optimized size of the dangerous part of the rotor wing moving part, guiding the design and meeting the requirement of the fatigue test assessment target. The method has the advantages that the strength, the rigidity and the boundary constraint of the test piece and the matching part under the characteristic test acceleration load can meet the requirements or not, and the early failure caused by unreasonable design of the test scheme is avoided.

Description

Fatigue design method for test-oriented helicopter rotor wing metal piece
Technical Field
The invention belongs to the field of helicopter rotor strength design, and relates to a test-oriented method for evaluating the fatigue life of a helicopter rotor metal part and designing the key size of a structure.
Background
The strength design of the traditional rotor system is that the size is defined by reserving strength margin according to engineering experience on the premise that various failure modes meet the service life. The structural life of a helicopter rotor system part is basically checked by adopting a fatigue characteristic test, and the fatigue characteristic test has the problems of asynchronous life of each failure mode and great acceleration difference, so that the structural fatigue test fails. Such test results do not match the expected failure mode, resulting in wasted test cost and lead time.
Disclosure of Invention
The invention aims to provide a test-oriented fatigue design method for a helicopter rotor wing metal piece, which considers the fatigue failure conditions of different parts and different abrasion states of a part in the fatigue characteristic test process, takes the fatigue test assessment of each part as a target, adopts a test sampling fatigue limit to carry out strength design and guides the structure key size definition, and provides technical support for the development of a helicopter.
In order to realize the task, the invention adopts the following technical scheme:
a fatigue design method for a test-oriented helicopter rotor metal piece comprises the following steps:
establishing a fatigue performance characterization model of a helicopter rotor component based on an accumulated damage theory to obtain a relation between the fatigue limit and the service life of a rotor wing metal piece; determining fatigue limits of different material types and ablation states under the target service life based on the fatigue performance characterization model; carrying out fatigue test verification on the rotor wing metal piece, and determining fatigue test loading amplitudes corresponding to the target service life under different material types and the friction and corrosion states; determining the maximum loading load of carrying out fatigue tests on different materials and under the condition of the fretting corrosion in the rotor wing metal piece based on the fatigue test loading amplitude, substituting the maximum loading load of each characteristic load into a structural strength analysis model of the metal piece and analyzing according to the test design requirements;
according to the reliability of the load spectrum data, taking a load correction coefficient to carry out fatigue limit calculation on the basis of the maximum load;
and selecting proper static load according to the rotor load spectrum, carrying out static load correction on the proper static load to obtain equivalent dynamic load, comparing the equivalent dynamic load with the safe fatigue limit for corresponding metal piece material design, and evaluating whether the key size of the metal piece meets the requirement.
Further, the determining fatigue limits for different material types and ablation states at a target life based on the fatigue performance characterization model comprises:
selecting the material type and the ablation state of the rotor wing part according to actual conditions, and respectively drawing fatigue life curves of different material types and the ablation state according to the fatigue performance characterization model and the load spectrum of the rotor wing part; and determining the fatigue limit of different material types and the friction corrosion state under the target service life through the fatigue life curve, namely determining the target service life fatigue limit of the component.
Further, the fatigue performance characterization model is represented as:
Figure BDA0002700995550000021
in the formula, C and m are the slope and intercept parameters of the fatigue curve of the metal piece material and are determined according to test data; t isa,i(i ═ 1,2, …, p) is a set of dynamic loads for characteristic load T in each flight regime; n isi(i ═ 1,2, …, p) is the frequency of hourly occurrences of the various flight states; l is the high cycle fatigue life of the helicopter structure; t is∞The high cycle fatigue limit of the characteristic load T corresponding to the life L.
Further, the fatigue test loading amplitudes for the different material types and the fretting regime corresponding to the target life are determined by:
Figure BDA0002700995550000022
in the formula, Ci,jAnd mi,j(i is 1,2, …, c, j is 1,2, …, d) is the slope and intercept parameters of the fatigue curve of the metal piece material corresponding to different materials and in the state of abrasion corrosion, and the test frequency is N1，(Ta)i,jFor different material types and for different states of erosion corresponding to a target lifetime LtargetLoading amplitude for fatigue testing.
Further, the proof fatigue limit of each material and the fretting corrosion state of the rotor component obtained by the fatigue test examination can be expressed as:
Figure BDA0002700995550000023
wherein eta is a load correction coefficient and takes a value of 1.0-1.2.
Further, the suitable static load is a maximum static load or a static load corresponding to a dynamic load state.
Further, the equivalent dynamic load T is obtainedSizing LoadSafe fatigue limit S for corresponding metal part material design∞And comparing, and evaluating whether the critical dimension of the metal piece meets the requirement, wherein the specific method comprises the following steps:
i) if S∞≥TSizing LoadThe flight load spectrum is shown at the safe fatigue limit S∞The cumulative damage caused by the lower layer is less than TSizing LoadDamage caused by the rotor structure, so that the high cycle fatigue life of the rotor structure is longer than the target life L corresponding to the dynamic load of the sampletargetAnd can be examined through component tests;
ii) if S∞＜T Sizing LoadThe flight load spectrum is shown at the safe fatigue limit S∞Cumulative damage caused by lower energy at T or aboveSizing LoadDamage caused by the lower part, the high cycle fatigue life of the rotor structure does not meet the target life L corresponding to the proofing dynamic loadtargetFurther optimization is required.
Further, the fatigue design method for the test-oriented helicopter rotor metal piece further comprises the following steps:
and optimally designing the metal piece which does not meet the requirements, acquiring corresponding local maximum stress according to the corresponding part of each equivalent load dynamic load of the metal piece, and determining the optimal size of the dangerous part of the rotor wing component by combining the dangerous section size of the metal piece obtained by measurement before design optimization under the specified load state, loading position, force transmission path and constraint mode.
Further, the optimal size of the rotor component hazard area is expressed as:
Figure BDA0002700995550000031
in the above formula, σmax, optimizationRepresenting the fatigue maximum stress, σ, expected for the metal part performancemaxA is the dangerous section size of the metal piece obtained by measurement before design optimization for the local maximum stress of the corresponding part of each equivalent load dynamic load of the metal piece;
and guiding the design optimization of the dangerous section of the structure according to the calculation result of the formula, so that the fatigue life target requirement can be met.
The invention has the following technical characteristics:
1. the method is based on the characteristic of fatigue life test and assessment of helicopter rotor system components, and fully considers the problem of inconsistent life acceleration conditions of various loads and different failure modes under the same test. The method takes tests as guidance, avoids the conditions that the structure is not fully examined due to different loads and failure mode test service life acceleration differences, the boundary simulation is inaccurate under the condition of test acceleration loads and the like, and improves the strength design, the design iteration efficiency and the test success rate by the method for testing the proof fatigue limit, thereby saving the development period and the test cost.
2. The strength, the rigidity and the boundary constraint of the test piece and the matching part under the characteristic test acceleration load are pre-judged in advance, and early failure caused by unreasonable design of a test scheme is avoided.
Drawings
FIG. 1 is a schematic flow diagram of the process of the present invention.
Detailed Description
Referring to fig. 1, the invention provides a fatigue design method for a test-oriented helicopter rotor metal piece, which realizes fatigue life evaluation and structural design of the metal piece, and comprises the following steps:
step one, analyzing fatigue test loads of different metal piece materials in an abrasion corrosion state
Based on the accumulated damage theory, establishing a fatigue performance characterization model of the helicopter rotor component:
Figure BDA0002700995550000041
in the formula, C and m are the slope and intercept parameters of the fatigue curve of the metal piece material and are determined according to test data; t isa,i(i ═ 1,2, …, p) is a set of dynamic loads for characteristic load T in each flight regime; n isi(i ═ 1,2, …, p) is the frequency of hourly occurrences of the various flight states; l is the high cycle fatigue life of the helicopter structure; t is∞The high cycle fatigue limit of the characteristic load T corresponding to the life L. Formula (1) establishes fatigue limit T of rotor wing metal piece∞The relation between the service life L and the input of a set of fatigue limits T∞The corresponding fatigue life L can be obtained.
According to actual conditions, selecting the material type M of the rotor parti(i ═ 1,2, …, k) and the state of fretting Fj(j ═ 1,2, …, l), the fatigue life T under different material types, rub-out conditions can be plotted according to equation (1) and rotor component load spectrum, respectively∞-an L-curve; through fatigue life T∞An L-curve, from which a target life L can be determinedtargetFatigue limit (T) for different material types and states of fretting∞)i,j(i-1, 2, …, k; j-1, 2, …, l), is the target life fatigue limit for the component.
Carrying out fatigue test verification on rotor wing metal pieces, wherein the test frequency is N1Different material types and states of the ablation correspond toTarget lifetime LtargetFatigue test loading amplitude (T)a)i,jCan be determined by the following formula:
Figure BDA0002700995550000042
in the formula, Ci,jAnd mi,jAnd (i is 1,2, …, c, j is 1,2, …, d) is the slope and intercept parameters of the fatigue curve of the corresponding metal piece material in different materials and in an abrasion state. According to the formula (2), the maximum loading load T of the fatigue test is developed under different materials and friction and corrosion states in the rotor wing metal piecea,maxCan be expressed as:
Ta,max＝max[(Ta)1,1,(Ta)2,1,…,(Ta)c,1,(Ta)1,2,…(Ta)c,1,…(Ta)c,d] (3)
loading the maximum load T of each characteristic loada,maxSubstituting into the structural strength analysis model of the metal piece, and analyzing the test design requirements such as strength, rigidity, boundary conditions and the like. The structural strength analysis model is a known model.
Step two, calculating the fatigue limit of the test guide proofing under different material and friction and corrosion states
The maximum load T of the fatigue test under different rotor wing metal piece materials and friction and corrosion states is given in the formula (3)a,maxConsidering the change of the structure adjustment to the load in the design process, and according to the reliability of the load spectrum data at Ta,maxOn the basis, the load correction coefficient eta (1.0-1.2) is taken to carry out fatigue limit calculation.
Combining formula (2) and formula (3), fatigue test examination of each material of the rotor part and proof fatigue limit (Sizing Load) under the state of abrasioni,jCan be expressed as:
Figure BDA0002700995550000051
the fatigue limit Sizing Load calculated by the formula (4) is a proof fatigue Load corresponding to the target service life and guided by the test, and is used for strength design and guiding structure optimization.
Step three, fatigue life assessment and design optimization
Obtaining the test-oriented proofing load fatigue of different metal piece materials in the friction and corrosion state through a formula (4), selecting proper static load according to a rotor wing load spectrum, and carrying out static load correction on the proper static load to obtain an equivalent dynamic load TSizing Load(ii) a The proper static load is generally the maximum static load or the static load corresponding to the dynamic load state; the method is determined according to actual needs.
Equivalent dynamic load TSizing LoadSafe fatigue limit S for corresponding metal part material design∞And comparing, and evaluating whether the critical dimension of the metal piece meets the requirement, wherein the specific method comprises the following steps:
i) if S∞≥TSizing LoadThe flight load spectrum is shown at the safe fatigue limit S∞The cumulative damage caused by the lower layer is less than TSizing LoadDamage caused by the rotor structure, so that the high cycle fatigue life of the rotor structure is longer than the target life L corresponding to the dynamic load of the sampletargetAnd can be examined through component tests;
ii) if S∞＜T Sizing LoadThe flight load spectrum is shown at the safe fatigue limit S∞Cumulative damage caused by lower energy at T or aboveSizing LoadDamage caused by the lower part, the high cycle fatigue life of the rotor structure does not meet the target life L corresponding to the proofing dynamic loadtargetFurther optimization is needed, and the optimization design process specifically comprises the following steps:
dynamic load T for each equivalent load of metal pieceSizing LoadCorresponding part, obtaining corresponding local maximum stress sigmamaxUnder the appointed load state, loading position, power transmission route and restraint mode, combine the dangerous cross-section size A of metalwork that obtains through the measurement before the design optimization, can confirm the optimum size of rotor part danger position:
Figure BDA0002700995550000052
in the above formula, σmax, optimizationA fatigue maximum stress representing a desired metal part performance; and (4) guiding the design optimization of the dangerous section of the structure according to the calculation result of the formula (5), so that the fatigue life target requirement can be met.
Importantly, the design method taking the test as the guide avoids the condition that after the local part or part of the structure meets the requirement and is firstly damaged, the rest parts or parts are not examined, and saves the test cost and the development period.
The above embodiments are only used for illustrating the technical solutions of the present application, and not for limiting the same; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equally replaced; such modifications and substitutions do not substantially depart from the spirit and scope of the embodiments of the present application, and are intended to be included within the scope of the present application.

Claims (9)

1. A fatigue design method for a test-oriented helicopter rotor metal part is characterized by comprising the following steps:
establishing a fatigue performance characterization model of a helicopter rotor component based on an accumulated damage theory to obtain a relation between the fatigue limit and the service life of a rotor wing metal piece; determining fatigue limits of different material types and ablation states under the target service life based on the fatigue performance characterization model; carrying out fatigue test verification on the rotor wing metal piece, and determining fatigue test loading amplitudes corresponding to the target service life under different material types and the friction and corrosion states; determining the maximum loading load of carrying out fatigue tests on different materials and under the condition of the fretting corrosion in the rotor wing metal piece based on the fatigue test loading amplitude, substituting the maximum loading load of each characteristic load into a structural strength analysis model of the metal piece and analyzing according to the test design requirements;
according to the reliability of the load spectrum data, taking a load correction coefficient to carry out fatigue limit calculation on the basis of the maximum load;
and selecting proper static load according to the rotor load spectrum, carrying out static load correction on the proper static load to obtain equivalent dynamic load, comparing the equivalent dynamic load with the safe fatigue limit for corresponding metal piece material design, and evaluating whether the key size of the metal piece meets the requirement.
2. A test-oriented helicopter rotor metal part fatigue design method according to claim 1 wherein said determining fatigue limits for different material types and rub-out conditions at a target life based on said fatigue performance characterization model comprises:
selecting the material type and the ablation state of the rotor wing part according to actual conditions, and respectively drawing fatigue life curves of different material types and the ablation state according to the fatigue performance characterization model and the load spectrum of the rotor wing part; and determining the fatigue limit of different material types and the friction corrosion state under the target service life through the fatigue life curve, namely determining the target service life fatigue limit of the component.
3. A test-oriented helicopter rotor metal part fatigue design method according to claim 1, characterized in that said fatigue performance characterization model is represented as:
Figure FDA0002700995540000011
in the formula, C and m are the slope and intercept parameters of the fatigue curve of the metal piece material; t isa,i(i ═ 1,2, …, p) is the dynamic load of the characteristic load T in each flight condition; n isi(i ═ 1,2, …, p) is the frequency of hourly occurrences of the various flight states; l is the high cycle fatigue life of the helicopter structure; t is∞The high cycle fatigue limit of the characteristic load T corresponding to the life L.
4. A test-oriented helicopter rotor metal part fatigue design method according to claim 1 wherein the fatigue test loading amplitude for different material types and rub-out conditions corresponding to a target life is determined by the following equation:
Figure FDA0002700995540000021
in the formula, Ci,jAnd mi,j(i is 1,2, …, c, j is 1,2, …, d) is the slope and intercept parameters of the fatigue curve of the metal piece material corresponding to different materials and in the state of abrasion corrosion, and the test frequency is N1，(Ta)i,jFor different material types and for different states of erosion corresponding to a target lifetime LtargetLoading amplitude for fatigue testing.
5. A method of test-oriented helicopter rotor metal fatigue design according to claim 1, wherein the proof fatigue limit for each material and fretting condition of the rotor component assessed by a fatigue test is expressed as:
Figure FDA0002700995540000022
wherein eta is a load correction coefficient and takes a value of 1.0-1.2.
6. A method of trial-oriented helicopter rotor metal fatigue design according to claim 1 wherein said suitable static load is the maximum static load or the static load corresponding to a dynamic loading condition.
7. The trial-oriented helicopter rotor metal part fatigue design method of claim 1, characterized by the equivalent dynamic load TSizing LoadSafe fatigue limit S for corresponding metal part material design∞And comparing, and evaluating whether the critical dimension of the metal piece meets the requirement, wherein the specific method comprises the following steps:
i) if S∞≥TSizing LoadThe flight load spectrum is shown at the safe fatigue limit S∞The cumulative damage caused by the lower layer is less than TSizing LoadDamage caused by the rotor structure, so that the high cycle fatigue life of the rotor structure is longer than the target life L corresponding to the dynamic load of the sampletargetAnd can be examined through component tests;
ii) if S∞＜TSizing LoadThe flight load spectrum is shown at the safe fatigue limit S∞Cumulative damage caused by lower energy at T or aboveSizing LoadDamage caused by the lower part, the high cycle fatigue life of the rotor structure does not meet the target life L corresponding to the proofing dynamic loadtargetFurther optimization is required.
8. The trial-guided helicopter rotor metal part fatigue design method of claim 1, further comprising:
and optimally designing the metal piece which does not meet the requirements, acquiring corresponding local maximum stress according to the corresponding part of each equivalent load dynamic load of the metal piece, and determining the optimal size of the dangerous part of the rotor wing component by combining the dangerous section size of the metal piece obtained by measurement before design optimization under the specified load state, loading position, force transmission path and constraint mode.
9. A test-oriented helicopter rotor metal hardware fatigue design method according to claim 8, wherein the optimal dimensions for the rotor component hazard zone are expressed as:
Figure FDA0002700995540000031
in the above formula, σmax, optimizationRepresenting the fatigue maximum stress, σ, expected for the metal part performancemaxA is the dangerous section size of the metal piece obtained by measurement before design optimization for the local maximum stress of the corresponding part of each equivalent load dynamic load of the metal piece;
and guiding the design optimization of the dangerous section of the structure according to the calculation result of the formula, so that the fatigue life target requirement can be met.
CN202011022238.6A 2020-09-25 2020-09-25 Fatigue design method for test-oriented helicopter rotor wing metal piece Active CN112182875B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202011022238.6A CN112182875B (en) 2020-09-25 2020-09-25 Fatigue design method for test-oriented helicopter rotor wing metal piece

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202011022238.6A CN112182875B (en) 2020-09-25 2020-09-25 Fatigue design method for test-oriented helicopter rotor wing metal piece

Publications (2)

Publication Number Publication Date
CN112182875A true CN112182875A (en) 2021-01-05
CN112182875B CN112182875B (en) 2022-09-06

Family

ID=73944957

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202011022238.6A Active CN112182875B (en) 2020-09-25 2020-09-25 Fatigue design method for test-oriented helicopter rotor wing metal piece

Country Status (1)

Country Link
CN (1) CN112182875B (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113177272A (en) * 2021-04-15 2021-07-27 哈尔滨工业大学（深圳） Method for numerical simulation and parameter analysis of fatigue finite element after corrosion of metal material

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20100235108A1 (en) * 2009-03-16 2010-09-16 Adams David O Usage monitor reliability factor using an advanced fatigue reliability assessment model
US20110057071A1 (en) * 2009-09-10 2011-03-10 Vineet Sahasrabudhe Life improving flight control system
CN105760623A (en) * 2016-03-16 2016-07-13 中国直升机设计研究所 Method for determining allowable defects of helicopter composite main rotor blade
US20170293712A1 (en) * 2016-04-11 2017-10-12 Airbus Helicopters Deutschland GmbH Probabilistic load and damage modeling for fatigue life management
CN108733891A (en) * 2017-04-21 2018-11-02 通用电气公司 System and method for predicting service life period and the consumption of low cycle facigue
CN110789733A (en) * 2019-10-11 2020-02-14 中国直升机设计研究所 Method for evaluating fatigue life of flapping deformation section of tail rotor flexible beam of helicopter
CN110861784A (en) * 2019-12-04 2020-03-06 中国直升机设计研究所 Calculation and monitoring method for fatigue accumulation damage of helicopter flight test load

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20100235108A1 (en) * 2009-03-16 2010-09-16 Adams David O Usage monitor reliability factor using an advanced fatigue reliability assessment model
US20110057071A1 (en) * 2009-09-10 2011-03-10 Vineet Sahasrabudhe Life improving flight control system
CN105760623A (en) * 2016-03-16 2016-07-13 中国直升机设计研究所 Method for determining allowable defects of helicopter composite main rotor blade
US20170293712A1 (en) * 2016-04-11 2017-10-12 Airbus Helicopters Deutschland GmbH Probabilistic load and damage modeling for fatigue life management
CN108733891A (en) * 2017-04-21 2018-11-02 通用电气公司 System and method for predicting service life period and the consumption of low cycle facigue
CN110789733A (en) * 2019-10-11 2020-02-14 中国直升机设计研究所 Method for evaluating fatigue life of flapping deformation section of tail rotor flexible beam of helicopter
CN110861784A (en) * 2019-12-04 2020-03-06 中国直升机设计研究所 Calculation and monitoring method for fatigue accumulation damage of helicopter flight test load

Non-Patent Citations (3)

* Cited by examiner, † Cited by third party
Title
HODGES D H等: ""Finite Element Solution of Nonlinear Intrinsic Equations"", 《JOURNAL OF THE AMERICAN HELICOPTER SOCIETY》 *
梁东亮: ""旋翼机桨叶疲劳特性分析"", 《中国优秀硕士学位论文全文数据库-工程科技II辑》 *
熊欣: ""直升机金属动部件延寿技术研究"", 《直升机技术》 *

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113177272A (en) * 2021-04-15 2021-07-27 哈尔滨工业大学（深圳） Method for numerical simulation and parameter analysis of fatigue finite element after corrosion of metal material

Also Published As

Publication number Publication date
CN112182875B (en) 2022-09-06

Similar Documents

Publication Publication Date Title
US9280620B2 (en) Method and system for probabilistic fatigue crack life estimation
CN104699976B (en) A kind of metal material multiaxis high cycle fatigue failure prediction method influenceed comprising mean stress
El-Aini et al. High cycle fatigue of turbomachinery components-Industry perspective
Klocke et al. Finite element analysis of the roller burnishing process for fatigue resistance increase of engine components
Chen et al. Study on establishment of standardized load spectrum on bogie frames of high-speed trains
CN112182875B (en) Fatigue design method for test-oriented helicopter rotor wing metal piece
Reid Hole cold expansion the fatigue mitigation game changer of the past 50 years
CN111222267A (en) Service life analysis method for hot end part of ramjet
CN113158140A (en) Aircraft structure maintenance project selection and maintenance interval analysis method based on multi-source information fusion
CN112052523B (en) Helicopter rotor moving part strength design method based on load characteristic comparison
Wanhill et al. Fracture mechanics in aircraft failure analysis: Uses and limitations
CN114492074A (en) Probabilistic damage tolerance assessment analysis method
Klimmek et al. Aircraft Loads–A Wide Range of Disciplinary and Process-Related Issues in Simulation and Experiment
Rice et al. Development of MMPDS Handbook aircraft design allowables
Johnson et al. Diagnostics of tool-part interactions during riveting on an aluminum aircraft fuselage
CN110907269A (en) Titanium alloy fatigue life prediction method based on micro scratches
CN113408672B (en) Key parameter identification method for aircraft modal test
Liu et al. Probabilistic properties of fatigue damage and an equivalent damage simplification model for multi-level load spectra
CN114741946B (en) Reliability simulation method for multi-component failure correlation of typical mechanism of aero-engine
Ball et al. Joint strike fighter airframe durability and damage tolerance certification
Boorla et al. Load Variability of a Two‐Bladed Helicopter
US20230118613A1 (en) System, method and apparatus of analytical criteria for composite structure durability and certification
Ai et al. Performance Degradation Analysis and Life Prediction of the Fatigue Damage Process of High Strength Aluminum Alloy Using Acoustic Emission
Dubinsky et al. METHOD FOR DETERMINATION OF ENERGIES FROM ACCIDENTAL IMPACTS ON THE AIRCRAFT STRUCTURE
Xiang et al. Probabilistic fatigue life model of composite laminas under cycle loading

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant