Airbus - Mts005b - Fatigue Manual
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Technical Manual MTS 005 Iss. B Outhouse distribution authorised
Fatigue Manual 1 4
5 4
2 1 2
5
Structure Design Manuals
Subject
Fields of application
Tool for calculating the crack initiation life of a metallic structural item of a civil aircraft subjected to cyclic loading. All programs. Structure design speciality.
Computer based tools supporting this manual
Contents
Documentation Manager
Detailed summary Foreword Field of validity Explanations on the method Appendices Bibliography
Logo : A/BTE/CC/CM
Validation
Name : D. CAMPASSENS
Name : JF. IMBERT Function : Assistant to the Department Group Leader Logo : A/BTE/CC/A Date : 27/07/98 Signature
This document belongs to AEROSPATIALE and cannot be given to third parties and/or be copied without prior authorisation from AEROSPATIALE and its contents cannot be disclosed. © AEROSPATIALE - 1998
3page 1
Fatigue Manual - Appendices
Reference documents
Refer to "Bibliography" chapter
Documents to consult
Terminology
Refer to chapter I.1.3 "Rotation"
Table of revisions Revision
Date
Pages modified
A
02/98
All
Reason for changes made New document. Supersedes technical note No. 436.0112/88.
B
© AEROSPATIALE - 1998
07/98
Paragraph 3.3.6
MTS 005 Issue. B
Revisions
3page Ann.
Fatigue Manual - Informations de gestion
DO NOT DISTRIBUTE THIS PAGE
List of approval
Logo
Function
A/BTE/CC/CM
Head of Department
Key words Bibliography
Name/Christian name CAZET G.
Calculation -
Distribution list Logo
Function
Name/Christian name (if necessary)
A/BTE/QN
A/BTE/QN Library
SIBADE Alain
A/BTE/QN
Diderot archives
SIBADE Alain
A/BTE Technical Library
BOUTET Fernand
A/BTE/SM/MG
Distribution list managed in real time by A/POI/D - (Didocost application)
© AEROSPATIALE - 1998
MTS 005 Issue. B
page IG1
CONTENTS
P. 1/4
CONTENTS Revision B (July 1998)
I FOREWORD .......................................................................................................... 1 DEFINITIONS.................................................................................................... 1.1 Reminder: the fatigue phenomenon .................................. Rev. A (Jan. 1998) 1.2 Conventional definition of crack initiation ........................... Rev. A (Jan. 1998) 1.3 Notations ............................................................................ Rev. A (Jan. 1998) 2 GOALS .............................................................................................................. 2.1 Reporting and valorising AS experience ............................ Rev. A (Jan. 1998) 2.2 Multi-user tool..................................................................... Rev. A (Jan. 1998) 3 APPROACH PRINCIPLES ................................................................................ 3.1 Systematic processing of AS tests ..................................... Rev. A (Jan. 1998) 3.2 Global analysis ................................................................... Rev. A (Jan. 1998)
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision B (July1998)
CONTENTS
P. 2/4
II FIELD OF VALIDITY............................................................................................. 1 EXTERNAL PARAMETERS.............................................................................. 1.1 Loading / stress spectrum .................................................. Rev. A (Jan. 1998) 1.1.1 Tension or shear uniaxial loading ....................................................... 1.1.2 "Civil aircraft" type spectra .................................................................. 1.1.3 No frequency effect ............................................................................. 1.2 Environment ....................................................................... Rev. A (Jan. 1998) 1.2.1 No temperature effect ......................................................................... 1.2.2 No fatigue - corrosion interaction ........................................................ 2 INTRINSIC PARAMETERS OF THE PART ...................................................... 2.1 Material .............................................................................. Rev. A (Jan. 1998) 2.2 Surface condition after machining ...................................... Rev. A (Jan. 1998) 2.2.1 Part finishing ....................................................................................... 2.2.2 Roughness .......................................................................................... 2.2.3 Residual stresses................................................................................ 2.3 Technological treatments ................................................... Rev. A (Jan. 1998) 2.3.1 Surface treatments.............................................................................. 2.3.2 Mechanical treatments ........................................................................ 2.3.3 Fastener installation ............................................................................ 3 LIFE................................................................................................................... 3.1 Average / application of a safety factor .............................. Rev. A (Jan. 1998) 3.2 Field of application ............................................................. Rev. A (Jan. 1998)
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision B (July1998)
CONTENTS
P. 3/4
III DESCRIPTION OF THE METHOD 1 GENERAL MATHEMATICAL FORMULATION ................................................. 1.1 Calculation under a monotonic load ................................... Rev. A (Jan. 1998) 1.1.1 Uniaxial tension loading ...................................................................... 1.1.2 Extension to uniaxial shear loading..................................................... 1.2 Spectrum calculation .......................................................... Rev. A (Jan. 1998) 1.2.1 Miner's rule.......................................................................................... 1.2.2 Rain-Flow principle.............................................................................. 1.2.3 Example of application ........................................................................ 1.3 Fundamental properties ..................................................... Rev. A (Jan. 1998) 1.3.1 Monotonic loading equivalent to the spectrum .................................... 1.3.2 Spectrum coefficient (Cs).................................................................... 1.4 Consequences on practical application.............................. Rev. A (Jan. 1998) 1.4.1 Fundamental parameters .................................................................... 1.4.2 Use in sizing........................................................................................ 1.4.3 Use in substantiation........................................................................... 2 DETERMINATION OF THE EQUIVALENT MONOTONIC LOAD..................... 2.1 Basic calculation with the computerised system ................ Rev. A (Jan. 1998) 2.2 High speed calculation using the spectrum calculation ..... Rev. A (Jan. 1998) 3 DETERMINATION OF THE INTRINSIC QUALITY OF THE PART (IQF) ......... 3.1 General IQF law ................................................................. Rev. A (Jan. 1998) 3.2 Effect of material ................................................................ Rev. A (Jan. 1998) 3.3 Influence of scale effect ..................................................... Rev. A (Jan. 1998) 3.4 Effect of technological process .......................................... Rev. A (Jan. 1998)
3.4.1 Surface treatments..............................................................................
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision B (July1998)
CONTENTS
P. 4/4
3.4.2 Mechanical treatments ........................................................................ 3.4.3 Fastener installation ............................................................................ 3.5 Kt in cylindrical shafts......................................................... Rev. A (Jan. 1998) 3.6 Kt in notched plates............................................................ Rev. B (July 1998) 3.7 Kt in drilled plates ............................................................... Rev. A (Jan. 1998) 3.8 Kt in yokes.......................................................................... Rev. A (Jan. 1998) 3.9 Kt in bolted and riveted assemblies.................................... Rev. A (Jan. 1998)
APPENDIX 1 / Substantiation of the general mathematical model .................... A1.1 Demonstration by an elastic-plastic approach ................. Rev. A (Jan. 1998) A1.2 Examples of use .............................................................. Rev. A (Jan. 1998) APPENDIX 2/ Substantiation of the general IQF law ........................................... A2.1 Law on notches................................................................ Rev. A (Jan. 1998) A2.2 Law on yokes ................................................................... Rev. A (Jan. 1998) A2.3 Law on bolted and riveted assemblies............................. Rev. A (Jan. 1998) APPENDIX 3/ Simplified modelling of a fastener................................................. A3.1 Determination of flexibility ................................................ Rev. A (Jan. 1998) A3.2 Determination of the equivalent section........................... Rev. A (Jan. 1998)
BIBLIOGRAPHY ...................................................................................................... ................................................................................................. Rev. A (Jan. 1998)
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision B (July1998)
Ch. I
FOREWORD
P. 1/1
I FOREWORD
This document supersedes the first edition (Ref. 1). Nonetheless, the general approach is globally the same which should facilitate, for potential users, the use of this new version after having used the former version. As far as possible, this new manual takes into account the remarks made by users: - during the 1992-93 audit; - during the proofing phase (July-December 1997); - over the last few years, through the Stress Office Support. This document remains "open-ended" and later may be completed (or modified if the occasion arises) by means of new partial editions (with a new revision index).
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
REMINDER: THE FATIGUE PHENOMENON
Ch. I.1.1
P. 1/2
1 DEFINITIONS 1.1 REMINDER: THE FATIGUE PHENOMENON Damage to a metallic part, under cyclic loading, can be summarised by 3 phases of development of the damage, called "fatigue": - a "crack initiation" period which is generally on the surface for 2 main reasons: . the dislocations in the crystal structure, responsible for material plasticity, are more easily formed on the surface than in the heart of the part and travel more easily; . the surface is exposed to adverse environmental conditions; in general, this entails (when examined under an scanning electron microscope) the following: . initially, formation of surface slide bands, that can be removed by very light polishing; . then the slide bands multiply and persist; . formation of intrusions and extrusions; . lastly, the appearance of micro-cracks along these geometrical defects; one of these defects then becomes more significant than the others; - a stable "growth" period of this crack characterised by a generally smooth and shiny fracture appearance in which lines of arrested growth make it possible to approximately date when crack initiation started; - an abrupt growth entailing a "static fracture" in the part characterised by a rougher and duller appearance. The following figure illustrates this phenomenon:
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
REMINDER: THE FATIGUE PHENOMENON
Ch. I.1.1
P. 2/2
Length of crack a
Static fracture
Initiation
Growth
crack progress
formation of a micro-crack
a
Number of cycles
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
Ch. I.1.2
CONVENTIONAL DEFINITION OF CRACK INITIATION
P. 1/1
1.2 CONVENTIONAL DEFINITION OF CRACK INITIATION Considering the difficulty in measuring the initiation phase of a crack, the corresponding definition can only be conventional. Often, the definition depends on the inspection method implemented to detect the phase. Generally in reference documentation, the phase corresponds to a 0.5 mm long crack. In this "Fatigue Manual" considering that laws are defined using the results from tests on small test specimens until they break, the approximate corresponding length of cracks is, as an average: 3 to 5 mm
Fatigue cracks in test specimen just before abrupt fracture
Also, this corresponds to the length of the crack that can be reasonably detected on aircraft using non-destructive testing techniques, such as eddy currents, ultrasonic, etc.
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
NOTATIONS
Ch. I.1.3
P. 1/2
1.3 NOTATIONS The following notations shall be systematically used to facilitate document understanding: Example of a notched tension test specimen
σmax
Smax
σa
Sa
σaverage
Saverage Smin
σmin
Reference cross-section
S Smax
rated stress (in a reference cross-section) in elastic state maximum rated stress in a loading cycle
Smin
minimum rated stress in a loading cycle
Save
average rated stress in a loading cycle
Sa alternating rated stress in a loading cycle R ratio: Smin / Smax Kt stress concentration coefficient equal to the ratio: local elastic stress at crack initiation point / S
σ local stress (at crack initiation point) in elastic-plastic state σmax maximum local stress in a loading cycle σmin minimum local stress in a loading cycle σave average local stress in a loading cycle σa alternating local stress in a loading cycle
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
NOTATIONS
Ch. I.1.3
P. 2/2
monotonic loading: succession of identical stress cycles
complex or spectrum loading: succession of different stress cycles
N average life until crack initiation, as defined in the previous paragraph and given as: . cycles for a part subject to monotonic loading . generally in flights for a part subject to complex or spectrum loading
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
Ch. I.2.1
REPORTING AND VALORISING AS EXPERIENCE
P. 1/1
2 GOALS
2.1 REPORTING AND VALORISING AS EXPERIENCE The essential goal of the Fatigue Manual, which was edited for the first time in 1988, is to record all AS know-how in the field involved, in the same manner as the other A/BTE/CC manuals: - Design Manual; - Static Manual (Metallic); - Composite Manual. These manuals satisfy the same concerns which can be summarised by the following 4 actions:
"collect" "preserve" "synthesise" "make available"
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
MULTI-USER TOOL
Ch. I.2.2
P. 1/1
2.2 MULTI-USER TOOL The Fatigue Manual has been designed so as to be of a sufficient level of user friendliness for: - designers during the sizing phase; essential goal: minimise the risk of the occurrence of cracks especially in typical areas; this goal can be translated by the following principle:
"prevent rather than cure" - designers in the aircraft certification and follow-up phase; essential goal: substantiation: . of new structures; . of modifications following non-conformities detected during production; . in-service repair following the discovery of a cracked area or an accidentally damaged area.
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
SYSTEMATIC PROCESSING OF AS TESTS
Ch. I.3.1
P. 1/1
3 APPROACH PRINCIPLES
3.1 SYSTEMATIC PROCESSING OF AS TESTS The Fatigue Manual is practically a rule of the thumb method based on overall analysis and synthesis of AS tests conducted over approximately the last 20 years. The approach is summarised in the following graph:
Level 4
Limited tests (generally certification) ==> verification and validation of the manual
Test airframe Level 3
Subassembly Level 2
P
P
P
Structural items ("design" data) Level1
P
P
P
Statistically processable tests ==> direct processing pour for the manual
P
Elementary test specimens (process + material data)
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
Ch. I.3.2
GLOBAL ANALYSIS
P. 1/1
3.2 GLOBAL ANALYSIS The main difficulty consists in finding the right compromise between method simplicity and
reliability. To this end, the following sentences may be considered as key points to the approach: - 1st step: identify the essential parameters, naturally drawing upon available tests and also theoretical analysis, in particular the numerical methods showing certain mechanical behaviours governing crack initiation in metallic structures; - 2nd step: as far as possible, make the parameters independent, one from the other; - 3rd step: quantify the effect of parameters in a friendly manner: . simple analytical formulas; . charts.
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
FIELD OF VALIDITY
Ch. II
P. 1/1
II FIELD OF VALIDITY
The purpose of this chapter is to define as precisely as possible the field of application of this
manual, taking into account that: - this document does not cover the entire metallic material fatigue field which is extremely vast, taking into account a great number of parameters involved; - the laws set forth in this document concern technological processes specific to AS and possibly to its subcontractors or partners. Consequently, reference is made for information purposes to AS standards documents detailing the field of application of these processes.
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
Ch. II.1.1
LOADING / STRESS SPECTRUM
P. 1/2
1 EXTERNAL PARAMETERS
1.1 LOADING / STRESS SPECTRUM 1.1.1 Tension or shear uniaxial loading The local stress state at the crack initiation point (refer to III.1.1) induced by the rated loading of the part involved must be: - in tension direction (or practically), the matrix of stress that can be formulated as follows: æ S 0 0ö
ç 0 0 0÷ çç ÷÷ è 0 0 0ø
- or in shear direction (or practically), the matrix of stress that can be formulated as follows: æ 0 S 0ö
ç S 0 0÷ çç ÷÷ è 0 0 0ø
1.1.2 "Civil aircraft" type spectra The law for calculating damage under complex loading (refer to III.1.2) was built using correlations between calculations / tests conducted under civil aircraft type spectra, i.e. using a known average
mission to which perturbations are randomly added (gusts, manoeuvres, taxiing, etc...). As an example, some statistical stress records are given on the following page.
1.1.3 No frequency effect The effect due to frequency may be considered negligible, knowing that there is (see next paragraph): - no creep; - no corrosion.
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
LOADING / STRESS SPECTRUM
Ch. II.1.1
Upper wing surface
Lower wing surface
Fuselage roof
© AEROSPATIALE 1998
P. 2/2
Engine pylon
FATIGUE MANUAL
Revision A (Jan. 1998)
ENVIRONMENT
Ch. II.1.2
P. 1/1
1.2 ENVIRONMENT 1.2.1 No temperature effect The results of the fatigue tests, carried out in a laboratory at room temperature, are assumed to be
applicable in a temperature field between: - the "low" temperatures, where, generally speaking, an increase to the allowable tensile stress is found; however, on the other hand, a reduction in ductility (embrittlement) is found; - the "high" temperatures, where there is a risk of appearance of the creep phenomenon combined with conventional fatigue; for this reason, the following recommended temperatures must not be exceeded: . 80 to 100°C approx. for aluminium alloys, except 2618 (≈150°C); . 350°C for steels as well as TA6V (200°C for other titanium alloys); . 650°C for nickel alloys.
1.2.2 No fatigue - corrosion interaction It is assumed that there is no corrosion due to an aggressive environment which deteriorates the fatigue strength of the studied parts, i.e.: - either the material is selected correctly: for example, titanium alloys and stainless steels are highly corrosion "resistant"; - a metallic or organic coating is selected, avoiding contact between the part and the aggressive environment (refer to II.2.3.1); - or sealing compounds are used (interlay, added beads, filling of cavities, wet installation or covering of fasteners, filling of countersunk holes). It is assumed that there is no galvanic corrosion due to the bad association of two materials in contact. Consequently, generally speaking: - concerning fastener installation, it is prohibited to use: . aluminium rivets in titanium or steel parts; . untreated fasteners, made of titanium, nickel or steel in aluminium parts; - it is prohibited to install, at the interfaces of parts: . untreated, unpainted, titanium, nickel, steel or composite parts without interlay of sealant with aluminium parts. Refer to the following documents for assembly recommendations: - ASDT 029: "Protection"; - ASDT 072: "Anti-corrosion protection (long-range aircraft)".
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
MATERIAL
Ch. II.2.1
P. 1/5
2 INTRINSIC PARAMETERS OF THE PART
2.1 MATERIAL The following tables list the typical AS qualified materials used and, for reference, purposes their minimum required static characteristics (R: allowable tensile stress / R0,2: allowable tensile yield stress at which permanent strain equals 0,2% / A: elongation at rupture). The fatigue characteristics of the materials underlined are known (therefore, tests have been
carried out on these materials). ALUMINIUM ALLOYS (ASN-B 10000) density around 2,8 Young's modulus E≈72000 MPa (approximately 70% of an aircraft structure) Semi-finished product
R min.
R0,2 min.
A min.
(MPa)
(MPa)
(in %)
T4 / T42
400
255
15
T6 / T62
440
390
7
Extruded bar
T6 / T651
460
415
7
Drawn bar
T6 / T651
450
380
8
Extruded shape
T6 / T 62 / T651
415
370
7
Thin sheet
T4 / T42
385
240
15
T6 / T62
420
345
9
T3
410
290
14
T42
430
265
15
T351
445
290
14
Thick sheet
T351
430
290
12
Structure tube
T3 / T351 / T42
440
290
10
Extruded bar
T3 / T42
440
330
11
Drawn bar
T3 / T351
440
315
12
Extruded shape
T3 / T351
440
330
12
T42
420
280
14
T3
390
260
12
T42
380
230
13
2014
Thin sheet
2014 Pl
2024
Thin sheet
2024 Pl
© AEROSPATIALE 1998
Thin sheet
Heat treatment
FATIGUE MANUAL
Revision A (Jan. 1998)
MATERIAL
Ch. II.2.1
P. 2/5
2124
Thick sheet
T351
440
290
12
2214
Thick sheet
T451
400
250
12
T651
460
410
7
T 62
400
325
7
T8
400
335
7
Thick sheet
T 851
430
385
5
Extruded bar
T 851
415
360
6
Extruded shape
T 62
400
335
7
Thin sheet
T62
390
310
7
T8
395
325
7
T4 / T42
210
110
16
T6 / T62
290
240
10
Thick sheet
T651
290
240
9
Extruded bar
T4 / T42
210
110
14
T6 / T62
270
245
8
Drawn bar
T6
290
245
8
Thin sheet
T6
560
500
7
Thick sheet
T651
570
530
7
T7451
495
430
6
T7651
525
450
5
Extruded shape
T6510
560
510
5
Thick sheet
T7451
510
440
8
T7651
525
455
6
T6
540
470
8
T76
490
410
9
T651
540
460
6
T7351
480
370
7
T7651
490
410
6
Extruded bar
T6
550
480
7
Drawn bar
T6
530
450
8
T73 / T7351
470
385
11
T6
540
480
7
T6510 / T6511
560
490
7
T73511 / T76511
485
420
8
2618A
Thin sheet
2618A Pl
6061
Thin sheet
7010
7050
7075
Thin sheet Thick sheet
Extruded shape
7075 Pl
Thin sheet
T6
505
440
10
7175
Thick sheet
T7351
480
390
7
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
MATERIAL
Ch. II.2.1
7475
7475 Pl
P. 3/5
Thin sheet
T76
490
410
9
Thick sheet
T7351
480
390
8
T7651
490
410
6
T76
470
390
8
R min.
R0,2 min.
A min.
(MPa)
(MPa)
(in %)
Thin sheet
TITANES ET ALLIAGES DE TITANE (ASN-B20000) density around 4,4 Young's modulus E≈110000 MPa (less than 10% of an aircraft structure) Semi-finished product
T40
Heat treatment
Thin sheet
Annealed
390
280
22
Rolled/forged bar
Annealed
390
280
20
Thin sheet
Annealed
570
460
15
Rolled / forged bar
Annealed
540
440
16
Thin sheet
Annealed
540
460
18
Hardened
690
550
10
Rolled/forged bar
Annealed
540
400
16
Thin sheet
Annealed
920
870
8
Thick sheet
Annealed
890
820
8
Rolled / forged bar
Annealed
900
800
10
Hardened
1100
1040
8
R min.
R0,2 min.
A min.
(MPa)
(MPa)
(in %)
T60
T-U2
TA6V
NICKEL ALLOYS (ASN-A 3271/3360/3361) density around 8,2 module de Young E≈200000 MPa (in engine areas of an aircraft) Semi-finished product
Inconel 625
Heat treatment
Thin sheet
Annealed
830
410
30
Inconel 718
Thin sheet /
Hardened
1270
1030
12
(NC19FeNb)
Rolled / forged bar
+ ageing
(NC22DNb)
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
MATERIAL
Ch. II.2.1
P. 4/5
STEELS (ASN-B 01000/05000) density around 7,8 Young's modulus E≈200000 MPa (approximately 10% of an aircraft structure) Semi-finished product
XC18
Heat treatment
R min.
R0,2 min.
A min.
(MPa)
(MPa)
(in %)
Sheet
Annealed
392
235
25
Bar / forged part
WH+Tempered
440
270
21
XC38
Bar / forged part
WH+Tempered
620
400
17
XC65
Bar / forged part
WH+Tempered
900
750
12
15CDV6
Sheet / Structure
AH+Te.>620
980
780
10
tube / Bar / forged part
OH+Te.>600
1080
930
10
Sheet
OH+Tempered
880
690
10
Structure tube
OH+Te.>?
660
470
15
OH+Te.>520
880
690
10
OH/TE+Te.>520
880
690
12
OH/TE+Te.>550
780
590
14
OH/TE+Te.>580
640
470
15
25CD4
Bar / forged part
30CD12
Bar / forged part
OH+Tempered
930
780
14
30CDV13
Bar / forged part
OH+Tempered
1080
880
12
35CD4
Structure tube
OH+Te.>540
1080
960
10
OH+Te.>410
1350
1230
8
40CDV20
Bar / forged part
AH+Tempered
1500
1300
9
12NC12
Bar / forged part
OH+Tempered
930
730
11
16NCD13
Bar / forged part
OH+Tempered
1030
740
11
16NCD17
Bar / forged part
Cem.+Te.
1270
880
8
30NCD16
Bar / forged part
OH+Te.>525
1220
1020
8
OH+Te.>540
1080
880
10
OH+Te.>580
1080
880
10
OH+Te.>550
880
740
14
OH+Te.>580
780
640
15
AH+Te.>200
1760
1420
6
AH+Te.>550
1230
1030
8
AH+Te.>550
1080
880
10
35NC6 35NCD16
Bar / forged part Bar / forged part
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
MATERIAL
Ch. II.2.1
P. 5/5
40CAD6.10
Bar / forged part
OH+Tempered
930
780
12
E-Z1CND12-09
Bar / forged part
AH+Tempered
1300
1200
9
Sheet
Over-hardened
440
180
45
Work hardened
800
700
10
Over-hardened
500
210
40
Work Hardened
800
700
10
Bar / forged part
Over-hardened
440
180
45
Sheet
MS+AC
640
200
40
MS+AC+A+AC
850
550
20
Treated+Temper
1220
1100
6
Treated
1310
1170
10
H 1025
1070
1000
11
OH/WH+Temper
960
660
10
OH+Te.>380
1100
900
14
OH+Te.>580
900
700
16
Over-hardened
490
220
40
forged part
Work Hardened
800
700
10
Z10CNW17
Sheet / Bar / forged part
MS+AC
540
220
35
Z12CN13
Sheet / Bar / forged part
AH/TH+Re.
590
410
16
Z12CN17-07
Sheet
Work hardened
885
600
17
Z12CND16-04
Bar / forged part
Treated
1400
1150
9
Z15CN17-03
Bar / forged part
OH+Te.>300
1350
1050
10
OH+Te.>600
880
690
12
AH/OH+Te.
880
690
10
(Marval X12) Z2CN18-10
Structure tube
E-Z3NCT25 Z6CND15-07
Sheet / Bar / forged part
(PH15.7MO) E-Z6CNU15-05
ed Bar / forged part
(15-5 PH) E-Z6NCT25
Bar / forged part
ed
Z8CND17-04
Bar / forged part
(17.4 PH) Z10CNT18-11
Sheet / Structure tube / Bar /
Z30CN13
Bar / forged part
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
SURFACE CONDITION AFTER MACHINING
Ch. II.2.2
P. 1/1
2.2 SURFACE CONDITION AFTER MACHINING 2.2.1 Part finishing Finishing of metallic parts, excluding bores: prior to any treatment, edges must systematically be deburred to obtain:
0,2 mm ≤ deburring depth (or finish radius) < 0,5 mm. Edges of bores shall be slightly deburred at the top and bottom of the assembly involved to provide a good mating plane for fasteners. The same applies to interfaces if this joint can be disassembled. Refer to the following documents for complementary information on the general directives concerning dimensions: - NSA 2110: "General manufacturing tolerances"; - A/DET 0031: "Finishing of aluminium alloy parts by deburring, breaking sharp edges or radiusing"; - A/DET 0164: "Finishing of edges on hard metallic parts"; - A/DET 0029: "Installation of shear bolts"; - A/DET 0085: "Installation of tension bolts".
2.2.2 Roughness The following rule is mandatory satisfied:
Ra ≤ 1,6 for bores
Ra ≤ 3,2 outside bores.
2.2.3 Residual stresses Inherent residual stresses always remain, they are: - difficult to quantify; - depend on machining conditions and also the material.
The laws proposed in this manual integrate the existence of this type of stress as these laws are built using the results from fatigue tests on test specimens that are generally machined during AS production work
Caution: this is no longer the case if, for example, stress relieving is carried out (in particular for titanium alloys). In this case, the fatigue strength may be considerably modified.
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
TECHNOLOGICAL TREATMENTS
Ch. II.2.3
P. 1/18
2.3 TECHNOLOGICAL TREATMENTS 2.3.1 Surface treatments The following table lists the typical processes used and qualified by AS. Fatigue characteristics are available for the treatments underlined (therefore the treatments for which the test results
are available). ALUMINIUM ALLOYS
Scope of use
Functions
Standard
CAA
Non-conducting layer
Adherence base before
A/DET 0072
(Chromic
No abrasion and wear
painting
Acid
resistance
Good corrosion
Anodising)
(2 to 5 micron layer)
resistance (if sealed)
SAA
Prohibited on fatigue load-
Adherence base before
(Sulphuric
carrying parts, cast, riveted,
painting
Acid
bonded, welded
Good corrosion resistance
Anodising)
(8 to 12 micron layer)
Wear protection
HA
Prohibited on fatigue load-
Adherence base before
(Hard
carrying parts, cast,
painting
Anodising
riveted, bonded, welded
Good corrosion resistance
(30 to 40 micron layer)
Wear protection
ALODINE
Parts where CAA is not
Adherence base before
A/DET 0079
(chromating
possible, touch-up
painting
A/DET 0175
treatment)
No abrasion and wear
Good corrosion resistance
resistance
(if painted)
(< 1 micron layer)
Conducting layer
NICKEL
Prohibited on parts in
Adherence base before
PLATING +
fuel area
painting
CADMIUM
(30 to 50 micron layer)
Good corrosion resistance
PLATING
Conducting layer
WITH SWAB
Protection against galvanic
A/DET 0091
A/DET 0097
A/DET 0147
coupling (carbon composite)
© AEROSPATIALE 1998
FATIGUE MANUAL
Revision A (Jan. 1998)
TECHNOLOGICAL TREATMENTS
Ch. II.2.3
P. 2/18
TITANIUM AND TITANIUM ALLOYS
Scope of use
Functions
Standard
SAA
Parts subject to risks
Adherence base before
A/DET 0084
(Sulphuric
of external aggression
painting
Acid
(< 1 micron layer)
Protection against
Anodising)
galvanic coupling
IVD
Hardware only
Adherence base before
(Ion
(4 to 12 micron layer
painting
Vapour
or 7 to 20 micron layer)
Conducting layer
Deposit)
A/DET 0012
Heat resistant
STEELS
Scope of use
Functions
Standard
Embrittling process
Adherence base before
A/DET 0073
(1 de-embrittlement
painting
A/DET 0167
operation necessary)
Good corrosion resistance
Touch-up
Conducting layer
Non-stainless steels
Protection against
Rm
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