AIRCRAFT DYNAMICS AND AGEING - fpz.hrfiles.fpz.hr/Djelatnici/zmarusic/Aircraft-dynamics-and...Ž....

UNIVERSITY OF ZAGREB

Faculty of Traffic and Transport Sciences

Aeronautics Study Programme

AIRCRAFT DYNAMICS AND

AGEING

A U T H O R I S E D L E C T U R E S

Željko Marušić, PhD

Zagreb, 2014

TABLE OF CONTENT:

1 INTRODUCTION .............................................................................................................................................. 1

1.1 WHY DO WE USE OLD AIRCRAFT? .................................................................................................................. 1

1.2 MODELS AND STATUS OF OLD AIRCRAFT ....................................................................................................... 2

1.3 EXAMPLES OF OLD AIRCRAFT IN ACTIVE SERVICE ......................................................................................... 3

1.4 HISTORY OF AGEING AIRCRAFT TESTING PROGRAMMES ................................................................................ 5

1.5 EVOLUTION OF TECHNOLOGY ....................................................................................................................... 5

1.6 AATF I AAWG ............................................................................................................................................ 6

1.6.1 Structure Task Groups ............................................................................................................................. 7

1.6.2 Supplemental Structural Inspection Programme ..................................................................................... 8

1.6.3 Corrosion Prevention and Control Programme ...................................................................................... 8

1.6.4 Service Bulletin ........................................................................................................................................ 9

1.6.5 Repair Assessment Programme ............................................................................................................... 9

1.7 DISTRIBUTION OF FATIGUE DAMAGE ........................................................................................................... 10

1.8 SUPPLEMENT TO MAINTENANCE PROGRAMME ............................................................................................ 10

1.9 ACTIONS TO SATISFY THE INTEGRITY OF THE STRUCTURE ........................................................................... 11

LIST OF QUESTIONS .................................................................................................................................................. 12

2 ACCIDENTS CAUSED BY AIRCRAFT STRUCTURE MATERIAL FATIGUE .................................... 12

2.1 1954 – FLIGHT 781 AND FLIGHT 201 – DE HAVILLAND COMET 1 ............................................................... 13

2.1.1 Flight 781 .............................................................................................................................................. 13

2.1.2 Flight 201 .............................................................................................................................................. 13

2.1.3 De Havilland DH 106 Comet design and characteristics ...................................................................... 14

2.2 1985 – FLIGHT 123 – BOEING 747 ............................................................................................................... 15

2.2.1 Flight 123 .............................................................................................................................................. 15

2.2.2 Boeing 747 ............................................................................................................................................. 17

2.3 1988 – FLIGHT 243 – BOEING 737 ............................................................................................................... 18

2.3.1 Flight 243 .............................................................................................................................................. 18

2.3.2 Boeing 737 ............................................................................................................................................. 19

2.4 1992 – FLIGHT 1862 – BOEING 747 ............................................................................................................. 20

2.5 2002 – FLIGHT 611 – BOEING 747 ............................................................................................................... 22

2.6 OTHER MATERIAL FATIGUE RELATED ACCIDENTS ....................................................................................... 22


3 BASIC THEORY OF VIBRATIONS ............................................................................................................. 25

3.1 FUNDAMENTALS OF MECHANICAL VIBRATIONS .......................................................................................... 25

3.2 VIBRATIONS IN AIRCRAFT STRUCTURES ...................................................................................................... 26

3.3 ANALYTICAL DESCRIPTION OF FREE AND FORCED VIBRATIONS ................................................................... 27

3.3.1 Free vibrations ....................................................................................................................................... 28

3.2.2 Forced vibration .................................................................................................................................... 31

3.4 VIBRATIONS WITH DAMPING ....................................................................................................................... 33

3.4.1 Free vibration with damping ................................................................................................................. 33

3.4.2 Forced vibration with damping ............................................................................................................. 36

3.5 ANALYSIS OF JITTER AND RESONANCE IN TRANSPORT VEHICLES ................................................................ 37

3.5.1 Analysis of jitter in transport vehicles ................................................................................................... 37

3.5.2 Analysis of resonance in transport vehicles ........................................................................................... 38

3.6 INFLUENCE OF VIBRATIONS ON MATERIAL FATIGUE OF AIRCRAFT STRUCTURES ......................................... 38

3.7 SENSOR SYSTEMS FOR MEASURING VIBRATION CHARACTERISTICS ............................................................. 39

3.8 ADVERSE EFFECTS CAUSED BY VIBRATIONS: FLUTTER, BUZZ, DIVERGENCE ................................................ 40

3.8.1 Flutter .................................................................................................................................................... 40

3.8.2 Buzz ........................................................................................................................................................ 40

3.8.3 Divergence ............................................................................................................................................. 40


4 MECHANICAL LOADS ON AIRCRAFT STRUCTURE ........................................................................... 42

4.1 STATIC LOADS............................................................................................................................................. 42

4.2 DYNAMIC LOADS ........................................................................................................................................ 42

4.3 COMBINED LOADS ....................................................................................................................................... 42

4.4 THERMAL LOADS ........................................................................................................................................ 42

4.5 AIRCRAFT STRUCTURAL LOADS .................................................................................................................. 42


5 INFLUENCE OF AERODYNAMIC LOADS ON AIRCRAFT STRUCTURE ......................................... 44

5.1 STRUCTURAL LOADS OF AIRCRAFT STRUCTURES ........................................................................................ 44

5.1.1 Loads and fatigue .................................................................................................................................. 44

5.1.2 Determination of design loads ............................................................................................................... 45

5.1.3 Structural design criteria ....................................................................................................................... 46

5.2 LOADS MONITORING ................................................................................................................................... 48

5.2.1 Monitoring of aircraft loads .................................................................................................................. 48

5.2.2 Flight parameter envelopes ................................................................................................................... 51

5.2.3 Loads model ........................................................................................................................................... 52

5.2.4 Aircraft component loads and design cases ........................................................................................... 52

5.3 IMPACT OF CHANGES ON COMPONENT LOADS ............................................................................................. 54

5.3.1 Qualification of loads via static and dynamic tests ............................................................................... 55

5.3.2 Static load conditions and fatigue spectrum .......................................................................................... 56

5.3.3 Conversion of external loads into structural airframe loads ................................................................. 58


6 LOADS ON AIRCRAFT ELEMENTS ........................................................................................................... 62

6.1 LOADS ON FUSELAGE .................................................................................................................................. 62

6.2 LOADS ON WINGS ........................................................................................................................................ 62

6.3 LOADS ON FLIGHT CONTROLS ..................................................................................................................... 63

6.4 LOADS ON PROPULSION SYSTEM ................................................................................................................. 64

6.5 LOADS ON UNDERCARRIAGE ....................................................................................................................... 65


7 DEGRADATION OF AIRCRAFT STRUCTURE ........................................................................................ 67

7.1 ADHESION ................................................................................................................................................... 67

7.2 ABRASION ................................................................................................................................................... 68

7.3 EROSION ..................................................................................................................................................... 69

7.4 CORROSION ................................................................................................................................................. 70

7.5 CAVITATION ............................................................................................................................................... 70

7.6 FATIGUE ..................................................................................................................................................... 70


8 CORROSION .................................................................................................................................................... 72

8.1 CORROSION IN AIRCRAFT STRUCTURES ....................................................................................................... 72

8.2 CHEMICAL CORROSION ............................................................................................................................... 72

8.3 ELECTROCHEMICAL CORROSION ................................................................................................................. 73

8.5 LOCAL CORROSION ..................................................................................................................................... 74

8.6 SELECTIVE CORROSION ............................................................................................................................... 75

8.7 INFLUENCE OF CORROSION ON AIRCRAFT AGEING ....................................................................................... 75

8.8 CORROSION PROTECTION METHODS ............................................................................................................ 75


9 MATERIAL FATIGUE ................................................................................................................................... 77

9.1 FATIGUE CHARACTERISTICS ........................................................................................................................ 77

9.2 MATERIAL FATIGUE AND CRACKS ............................................................................................................... 77


10 BASICS OF FRACTURE MECHANICS ....................................................................................................... 89

10.1 HIGH CYCLIC CRACKS ................................................................................................................................. 89

10.2 INFLUENCE OF STRESS CONCENTRATION ON MATERIAL FATIGUE ................................................................ 90

10.2.1 Concentration factor for cracks ........................................................................................................ 91

10.2.2 Concentration factor calculation....................................................................................................... 91

10.2.3 Stress concentration prevention ........................................................................................................ 91

10.3 EXAMPLES OF STRESS CONCENTRATION ..................................................................................................... 92

10.4 FRACTURE CRITERION ................................................................................................................................. 92

10.5 BASIC CONCEPT OF CRACK PROPAGATION DUE TO MATERIAL FATIGUE ....................................................... 95

10.6 CONSEQUENCES OF FRACTURE IN AIRCRAFT STRUCTURES .......................................................................... 97

LIST OF QUESTIONS ................................................................................................................................................ 107

11 METHODS TO DETECT AND PREVENT MATERIAL FATIGUE ...................................................... 108

11.1 INFINITE LIFETIME CONCEPT METHODS ..................................................................................................... 109

11.2 FINITE LIFETIME CONCEPT METHODS ........................................................................................................ 109

11.3 DAMAGE TOLERANT DESIGN METHODS OR METHODS OF NON-DESTRUCTIVE TESTING .............................. 109

11.3.1 Visual inspection .................................................................................................................................... 110

11.3.2 Penetrant testing .................................................................................................................................... 111

11.3.3 Ultrasonic testing ................................................................................................................................... 111

11.3.4 Magnetic particle inspection .................................................................................................................. 113

11.3.5 Radiographic testing .............................................................................................................................. 114

11.3.6 Eddy current testing ............................................................................................................................... 114

11.3.7 Low coherence interferometry ............................................................................................................... 115

11.4 COLD EXPANSION METHODS ..................................................................................................................... 116

11.5 BITE CONCEPT ......................................................................................................................................... 123

11.6 SHM CONCEPT .......................................................................................................................................... 124

11.6.1 Wireless Sensors for Structural Health Monitoring ........................................................................ 125

11.6.2 Future of sensors technology ........................................................................................................... 126


12 REPARATION OF AIRCRAFT STRUCTURE .......................................................................................... 128

12.1 TYPES OF AIRCRAFT STRUCTURES ............................................................................................................. 128

12.1.1 Riveted steel structures .................................................................................................................... 128

12.1.2 Aluminium alloy structures .............................................................................................................. 128

12.2 METHODS TO REPAIR AIRCRAFT STRUCTURES ........................................................................................... 128

12.2.1 Splicing of tubes .............................................................................................................................. 128

12.2.2 Repairs to aluminum alloy members ............................................................................................... 129

12.2.3 Reinforcement of wing and tail surface ribs .................................................................................... 129

12.2.4 Reinforced splices of trailing and leading edges and tip strips ....................................................... 129

12.2.5 Repair of damaged skin ................................................................................................................... 129

12.2.6 Patching of small holes ................................................................................................................... 129

12.2.7 Splicing of sheets ............................................................................................................................. 129

12.2.8 Straightening of stringers or intermediate frames ........................................................................... 129

12.2.9 Local heating ................................................................................................................................... 129

12.2.10 Splicing of stringers and flanges ..................................................................................................... 130

12.2.11 Drill small holes .............................................................................................................................. 130

12.2.12 Adding reinforcement of joints ........................................................................................................ 130

12.2.13 Steel fittings ..................................................................................................................................... 130

12.2.14 Aluminium and aluminium alloy fittings ......................................................................................... 130

12.2.15 Selective plating in aircraft maintenance ........................................................................................ 130


LIST OF TABLES ................................................................................................................................................... 133

LIST OF FIGURES ................................................................................................................................................. 134

CONCLUSION ........................................................................................................................................................ 138

REFERENCES ........................................................................................................................................................ 139

Ž. Marušić Aircraft Dynamics and Ageing

1

1 Introduction

Intensive development of air traffic in early 1930s was followed by the regulations on

resources for aircraft engines. The first maintenance systems based on the previous experiences

were built according to the model of preventive maintenance with pre-defined resources (Hard

Time Limit).

Because of this kind of maintenance, unnecessary dismantling of engines and assemblies

often took place, and also there was considerable properly operating remaining resource left

behind. It is known that the intensity of ageing depends on the manner and conditions of

exploitation which are highly variable. This way of maintenance increased maintenance costs

and decreased overall efficiency of the airline operators.

After the World War Second a new system of maintenance (On Condition) is developed.

It was also a type of preventative maintenance. The systems are tested in a given time in order to

assess the situation and, if necessary, take corrective actions to prevent the occurrence of

function failures during operation.

In the early sixties maintenance systems are developed and based on reliability programs

(Reliability Centred Maintenance). It is a methodology (set of rules) to prevent failures. It results

in the identification of the required type (or policy) of maintenance that needs to be applied to

individual failures. It also enables the development of sensor technology and continuous

monitoring of system status (Condition Monitoring).

Accidents caused by fatigue are presented in next Chapter. De Havilland Comet 1 aircraft

accidents indicated the catastrophic effects of material fatigue. Hence, different methods are

developed to detect and prevent fatigue of material in aircraft structures. The most important and

most commonly used methods are infinite lifetime concept methods, finite lifetime concept

methods, non-destructive testing methods, cold expansion methods, methods using built-in

testing equipment and structural health monitoring methods.

The most commonly used methods of non-destructive testing are: visual, magnetic,

penetrant, radiographic, ultrasonic, eddy current methods and low-coherence interferometry.

Cold hole expansion induces a zone of residual compressive stress around and through a hole.

All modern transport aircraft today have some form of permanent monitoring of technical

condition which is integrated on the aircraft. These systems allow detection of defects during

operation and are called built-in test equipment. The process of damage detection in aircraft

structures via sensors is known as structural health monitoring.

1.1 Why do we use old aircraft?

First generation of transport jet aircraft from the 50s and 60s had engines with high SFC,

which was not a problem until the 70’s when fuel prices suddenly increased. Manufacturers have

predicted replacement of high consumption aircraft with new ones (Boeing expected order of

250-300 replacement aircraft, which reduced to only 60 in 1988) but there was a delay due to the

replacement of the fleet for several reasons.

Higher prices of new aircraft, which were homologated by strict and rigorous standards

for structure (ICAO Annex 16 Chapter 3 – Noise Level); price inflation, which made

exploitation of new aircraft more expensive; increasing travel demand, increase of traffic; lack of

new aircraft in the market due to sudden increase of demand with a two-year delivery deadline;


2

increased competition with fares, free market (lower profits); etc. “Catch-22” is known as

dilemma or a circumstance from which there is no escape because of mutually conflicting or

dependent conditions which was the case here.

Management was faced with a dilemma: more flying hours, and no new aircraft because

of a reduced profit, more expensive loans (from banks), long delivery times, etc. Management

decision was “flying with old aircraft until they fall apart”. The result of such decision pointed

out previously led to moving further away from the concept of “Safety: Profit = Accidents”.

Corrective legislation was introduced and adopted stricter standards of the approval and

maintenance, globalization framework, the introduction of quality control and assurance (ISO

9000).

1.2 Models and status of old aircraft

The first systematic studies of old aircraft began in 1988 and were focused on e eleven

aircraft types (“AATF Eleven”), A300, BAC 111, B 707/720, B 727, B 737, B 747, DC-8, DC-9,

DC-10, F-28, L-1011;

At the end of 1995, there were 5671 jet (4.65%) and 2036 turbo-prop aircraft older than

15 years, of which there were 2889 jet aircraft older than 20 years, from total of 122,000

commercial jet aircraft in service.

Most of these old aircraft does not meet stricter noise standards (Chapter III or FAR/JAR

36) and are commonly sold to airlines “in developing countries” (the new owners generally do

not have neither the technical nor the financial resources to maintain airworthiness), or air

carriers move Operations to countries with more lenient national aviation regulations (lease

agreements, joint venture companies, etc.).

It is most likely that “young engineers” with little experience will be faced with such a

problem in countries in transition and developing countries.


3

1.3 Examples of old aircraft in active service

Table 1 Old aircraft in active use

McDonell Douglas Model Number of aircraft in

active fleet

Life time of aircraft

structures (economical

aspect)

Number of aircraft that

exceed their life time

100%

DC-8 342 25,000 cycles

50,000 hours

20 years

49

203

350

DC-9 924 40,000 cycles

30,000 hours

20 years

504

636

184

MD-80 496 50,000 cycles

50,000 hours

20 years

0

0

0

DC-10 595 42,000 cycles

60,000 hours

20 years

0

0

0

Boeing Model Number of aircraft in

active fleet



aspect)



75% 100%

B 707 342 20,000 cycles

60,000 hours

20 years

266

236

201

109

68

165

B 727 1,760 60,000 cycles

50,000 hours

20 years

262

334

341

13

446

721

B 737 1,547 75,000 cycles

51,000 hours

20 years

99

257

297

1

85

209

B 747 595 20,000 cycles

60,000 hours

20 years

112

205

202

11

105

4


4

Table 2 European manufacturers

Airbus Model Number of aircraft in

active fleet



aspect)



25% 50%

A300-B4 237 36,000 cycles

60,000 hours

178 34

A300-600 58 30,000 cycles

67,500 hours

10 0

A310 149 35,000 cycles

60,000 hours

14 0

A320 23 48,000 cycles

60,000 hours

0 0

A330 - 40,000 cycles

60,000 hours

- -

A340 - 20,000 cycles

30,000 hours

- -

Other aircraft Number of aircraft in

active fleet


structure (hours)

Successfully realized

Flight hours

Tristar 97 115,000 72,202

BAC-111 81 85,000 62,050

Tu-154 79 30,000 38,131

F-28 59 60,000 64,644

Il-62 28 30,000 39,026

Concorde 4 45,000 18,778


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1.4 History of ageing aircraft testing programmes

In the last 30 years of integrity and durability of aircraft structure testing programmes are

studied and developed. Accident in Lusaka in 1977 pointed out the problem of ageing aircraft

structure and initiated the introduction of fatigue testing and concept of construction fail-safe

technology.

Accident in Ramsden in 1981 revived the issue, and it was realized that 15 years old

aircraft with 50,000 flight hours very “large” numbers in aircraft exploitation. FAA convened a

conference on old aircraft in Washington DC in 1984, and parameters of 20 years and 60,000

flight hours for the “old” aircraft were adopted. Since 1986, Boeing is investigating “ageing” of

jet aircraft, with the goal to create empirical findings about the state of structure and systems and

the effectiveness of anti-corrosion methods.

A strong impact to take concrete measures made the accident of Aloha Airlines in 1988

when a 5.5 m of hull separated from the cockpit at an altitude of 24,000 ft. and the crew and the

passengers found themselves in a decompressed convertible.

In 1988 airworthiness of old aircraft became questionable and compelling activities have

been initiated with goals such as: mandatory modification of structure that will reduce

dependence on check-ups; development of programmes to prevent and control corrosion; update

the maintenance programme for the old aircraft; development of new improved test for the

assessment of fatigue damage, and development of methods and procedures to repair hidden

cracks and damage.

The problem of ageing is primarily a problem of airworthiness extending and control of

activities related to the area of exploitation and maintenance.

When solving problems, the role and share of producers, i.e. TCH (Type Certificate

Holder's) inevitably imposes itself, or to say, where it all started.

When the exploitation time limit of the particular aircraft is approaching the end, further

action in maintenance procedures should be discussed (in order to preserve the integrity of an

airframe (ASIP) and to extent airworthiness), between manufacturers (TCH), operators and the

relevant regulatory authority.

Old aircraft, due to persistent safety threats, were and still are an international safety

issue. In the last 10 years, led by world organization Assurance Airworthiness Working Group

(AAWG, ex AATF – Aging Aircraft Task Force), the implementation of the organizational tasks

as the solution of stated problems is continually examined.

1.5 Evolution of technology

The first generation of transport aircraft was designed in the 1950s and 1960s. They are

characterized by the principle of construction structures with emphasis on static strength and

safety in the case of failure (fail-safe structure), for example Boeing 707, 727, 737 and 747. This

concept is based on the redundancy, which allows satisfactory detection capabilities of damage

and “ageing” (deterioration) caused by corrosion.

Experience has shown that fail-safe concept served well in commercial transport and that

it has credibility, but there are imperfections in safety aspects (safety records). Due to detecting

and preventing fatigue, especially in the old aircraft that have numerous defects, such

redundancy is not always effective.


6

Additional programmes for detecting fatigue are developed in 1970s, which allow

satisfactory detection of fatigue damage. Period of structure testing or check-ups for most of

structures is based on the detection and monitoring of fatigue damage, using estimates for the

propagation of individual structure components with hidden details and limited ability to

“capture” a crack. Currently, procedures are being developed to calculate the period of structure

testing in terms of widespread fatigue damage.

Second generation of transport aircraft introduced models B 757 and 767 and modified B

737 and 747 models. Additionally to previous demands of strength and fail-safe technology,

these models are characterized by “the ability of tolerance”, which allows uncompromising

structure safety until the moment of damage detection with planned testing programme. In

addition, the standards of durability and damage tolerance are incorporated in these models.

Exploitation experience shows significant improvements in maintenance requirements

due to corrosion and fatigue-induced structure damage.

Third generation of transport aircraft represents a model B 777 and next generation of

modified model B 737. Those aircraft compared to the second generation have incorporated

additional improvements to prevent and control corrosion. Full-scale fatigue testing of model

777 shows significant improvements compared to the 757/767 high-performance testing in the

previous decade.

1.6 AATF i AAWG

Many commercial transport aircraft are homologated indefinitely in terms of structure

airworthiness and Airframe Service Life is not strictly defined.

Maintenance programmes must ensure that the structure meets all requirements for safe

and effective exploitation throughout its lifespan.

To achieve this, in 1988 Air Transport Association of America (ATA), the Aerospace

Industries Association of America (AIA) and the FAA initiated the establishment of an

international group of experts i.e. Aging Aircraft Task Force (AATF), which is known as “AATF

Eleven” (for 11 types of aircraft).

In 1992, group was renamed to Airworthiness Assurance Working Group (AAWG) and

is currently working under the supervision of the Aviation Rulemaking Advisory Committee,

and is composed of operators, manufacturers and legislators.

European Airworthiness Assurance Group is currently studying existing materials in

effort to regulate the instructions requirements for extending airworthiness.

AAWG research resulted in the establishment of five core programmes that articulate the

following elements:

1. Structure Task Groups (STG)

2. Supplemental Structural Inspection Programme (SSIP);

3. Corrosion Prevention and Control Programme (CPCP);

4. Service Bulletin – Modification Programme (SB);

5. Repair Assessment Programme (RAP).


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1.6.1 Structure Task Groups

Manufacturer and JAA/FAA should:

Determine the anticipated service life (Design Service Goal – DSG);

Determine possible extended service life (Extended Service Goal – ESG);

Ensure that DSG fits into other documents (MRB Report, SSID, etc.);

Apply additional programmes to test the structure (Supplementary Structural

Inspection Program – SSIP). The programme should not start after the aircraft reaches

one half of DSG;

Ensure the application of Corrosion Prevention and Control Programme – CPCP,

which is also implemented before the aircraft reaches one half of DSG;

Ensure the application of Service Bulletin in regular maintenance (SB);

Implement assessment and evaluation programme, repair and maintenance

programme, before the aircraft reaches 3/4 of DSG (Repair Assessment – RAT);

Determine the damage caused by the expansion of fatigue (Widespread Fatigue

Damage – WFD); fit actions to prevent WFD into the maintenance programme before

the end of DSG;

Provide Maintenance Records of maintenance, SSIP, CPCP, and derived SB.

Figure 1 Structures Task Group [author]


8

1.6.2 Supplemental Structural Inspection Programme

Supplemental Structural Inspection Programme or SSIP is made on the basis of detailed

studies and assessments of aircraft structure characteristics. It is designed by TCH

(manufacturer) and certified by the authority. It is designed to assess propagating cracks

according to fail tolerant philosophy. For aircraft certificated by FAR 25.571 with Amendment

45, the standards for “damage reports” are introduced, and additional examinations are

prescribed with AC91-56 and AWN 89.

Norms related to fatigue testing were added in FAR 25.571 and Amendments 45/55, and

are part of the MRB Report, in accordance with requirements of MSG 3.

The purpose of assessment and adjustment of the maintenance programme is an extension

of airworthiness while ensuring the safety of the aircraft type. Assessment of the structure

integrity is based on evidence (maintenance tests and data or findings) and is a continuous

process. In accordance with findings, next testing is determined.

It is important to identify Principal Structure Elements (PSE) that contribute to safety; i.e.

in the event of cancellation PSE, structural integrity is compromised. Testing and evaluation is

carried out with effective NDT methods, applying the partial dismantling of certain areas of the

hull.

1.6.3 Corrosion Prevention and Control Programme

Corrosion Prevention and Control Programme or CPCP provides a basis for assessment

and control of the structure weakening due to corrosion. The level of allowable corrosion is

usually regulated with AD. If exceeding of level of corrosion is detected in the area observed in

the aircraft or fleet, inspection intervals will be reduced and rehabilitation measures will be

taken.

CPCP must be part of the MRB Report which is periodically used for structure checks. If

the manufacturer has not developed CPCP program, it must be developed by an operator or

group of operators. Measuring the level of corrosion and determining the efficiency of

prevention is defined with:

1. Level 1: corrosion occurred between two inspections, and can be repaired without

limitations given by TCH (does not require replacement of components or reinforcing

structure);

2. Level 2: corrosion occurred between the two inspection and requires repair in

accordance with the instructions of TCH, amplification, etc., but it is not urgent in

safety aspect;

3. Level 3: corrosion was found at the first examination and urgent repair is required.

Using the CPCP corrosion above the first level should never be found or change the

programme is needed.

First level is recorded and filed by an operator, second level by a manufacturer and third

level by the Management and by a manufacturer.


9

1.6.4 Service Bulletin

Each component of the aircraft, which requires frequent examinations or review is hardly

feasible (hidden if not disassembled) should be modified or replaced. Modifications are carried

out in accordance to the manufacturer instructions in the form of a Service Bulletin (SB)

Modification Programme.

Operator should report to the manufacturer about conducted modifications or report

Modification status. Manufacturer monitors the entire fleet, and based on of such general

overview of the fleet status, decides to issue the SB. Choosing mandatory SBs and

recommendations is determined by Structural Task Group (STG) of the manufacturer, and such

recommendations are based on:

The possible impact on the structure airworthiness;

Credibility of overviews;

Frequency of incidence;

Damage done to the adjacent structure.

1.6.5 Repair Assessment Programme

Even the best exploited and properly maintained aircraft will gather a certain amount of

repairs (or “patches”) with time. However, the increased number of flight hours and age years

does not reduce the safety requirements and standards for damage tolerance. Practice had shown

that aircraft repairs do not return the aircraft to the “original state” or back to the initial level of

safety.

Additional commitments are prescribed for operators when developing assessment

programme of repair quality (Repair Assessment Programme or RAP), in a period that does not

exceed 3/4of DSG, and which should:

Define the repair time that does not exceed ¾ of DSG;

Establish a review by zones (Baseline Zone Inspection – BZI);

Perfect repair with an emphasis on the hull repair (pressurization);

Review the application and all repairs by SB.

The Repair Assessment Programme consists of three levels:

1. Level 1: data collecting (creating files);

2. Level 2: categorization of repairs (regardless of the repairs that need to be done

immediately, the data collected in the first level are used to categorize repairs), and

the categories are:

1) Category A: final repair to ensure airworthiness is equal to the initial state of the

structure;

2) Category B: final repairs that require additional examinations in order to extend

airworthiness (under supervision);

3) Category C: temporary repairs that require additional inspections in order to

extend airworthiness (under supervision);

3. Level 3: definition of additional maintenance (frequency of inspections, new or

additional technologies, etc.). In this level, additional examinations of categories B


10

and C are defined; and inspections are performed by the operator and are defined by

the manufacturer.

1.7 Distribution of fatigue damage

High number of flight hours and age of a particular structure area (or structure zone) can

develop cracks that can significantly reduce the strength of the structure. WFD is characterized

as crack development in the carrier part; of sufficient size and density that structure can no

longer take predicted damage (damage exceeds tolerance).

It is necessary to define the limit of the resulting fatigue damage or widespread fatigue

damage (WFD), over which the aircraft could no longer fly. Repairs to extend airworthiness are

only partial reconstruction, and are regulated by the Advisory Circular AC 91-56B.

Table 3 Examples of “life-prolonging” Boeing fleet

Model of aircraft Total number

Aircraft

DSG flights DSG flights

hours

Maximum

flights

Maximum flight

hours

B 707 728 20,000 60,000 40,700 89,600

B720 153 30,000 60,000 45,000 69,300

B727 1,819 60,000 50,000 72,000 78,400

B737 2,706 75,000 51,000 90,100 81,100

B747 1,051 20,000* 60,000 32,800 98,600

B757 686 50,000* 50,000 23,100 45,000

B767 582 50,000* 50,000 28,200 52,300

*some derivatives have special design objectives or DSG

1.8 Supplement to maintenance programme

As a response to the aircraft accidents, which were caused by compromised structural

integrity of the aircraft, the FAA first established National Aging Aircraft Research Program

(NAARP), which considers three elements when checking the integrity of the aircraft structure:

1. Methodology for predicting the expansion of fatigue damage;

2. Control inspections of passenger aircraft structure;

3. Repair assessment of aircraft structure.

The first supplementary documents for structure inspection (Supplemental Structural

Inspection Documents – SSID) were issued between 1979 and 1983 for old Boeing models, with

the intention to detect any damage caused by fatigue in given periods of time. Procedures have

been developed over time based on empirical standards, and Boeing issued a series of SSID and

CPCP documents which are the basis for AD issued by the FAA.

European regulations and the manufacturers followed the example of the United States,

and although they were faced with such a large number of old aircraft, the first A300 Task Unit

was created with task to: - re-examine SBs (modification status); develop CPCP; assess the

quality of repair – RAT; develop guidelines for maintaining the structure and check the SSID.

These five programmes (SSID, CPCP, SB, RAT and WSD) are the basis for updates of

existing aircraft maintenance programmes. For new models, MRB Report predicts immediately


11

and includes the continuous development of these programs as the aircraft keeps records of flight

hours, years i.e. age and cycles. Existing NDT methods like ultrasonic, eddy current, visual with

borescope, D sight and edge of light (EOL), and X-rays are used according to the applicability

and reliability for detecting damage to the structure and continuous monitoring the situation and

demonstrating the fulfilment of conditions for airworthiness. The development of multi-modal

NDT Analysis software increases the possibilities of the inspection.

Implementation of ISO 9000 (ISO 9001 for suppliers, ISO 9002 for manufacturer and

ISO 9003 for final inspection) regulates the introduction of quality control system to control

quality of the manufacturing and maintaining procedures as an independent system of continuous

control of quality or airworthiness.

Pursuant to the issuance of the AD and the obligation to introduce quality systems, it is

regulated by management of the operator that safety must be priority before profit. At the same

time, stricter standards for homologation when designing new models are imposed on the

manufacturer, in the way that new technologies must provide the possibility of continuous

surveillance of structure integrity and quality, and fulfilment of the requirements for

airworthiness in accordance with global regulations.

1.9 Actions to satisfy the integrity of the structure

Continuity of tracking the aircraft structure integrity must not stop after the construction

and delivery of the aircraft (quality control system), but it continues also during aircraft

exploitation through constant data collecting and updating.


12

List of questions

1. What are the reasons to use old aircraft instead of new ones?

2. What is AATF eleven? Explain. Specity several old aircraft in active service.

3. Explain important events in history that initiated studies and development of ageing

aircraft testing programmes.

4. Explain development of testing programmes by generations of transport aircraft.

5. Explain tasks of AAWG. Specify five AAWG programmes.

6. Explain tasks of STG.

7. Explain tasks of SSIP.

8. Explain tasks of CPCP.

9. Explain tasks of SB.

10. Explain tasks of RAP.


13

2 Accidents caused by aircraft structure material fatigue

2.1 1954 – Flight 781 and Flight 201 – De Havilland Comet 1

Two accidents occurred in 1954 which involved De Havilland Comet 1 aircraft. First

accident involving Comet 1 occurred in January and it was British Overseas Airways

Corporation Flight 781. The cause of accident was material fatigue failure which led to explosive

decompression in flight, breakup and consequently crashing into the Mediterranean Sea.

Second accident occurred three months later in the same year also involving Comet 1

(South African Airways Flight 201). The cause of this accident was also material fatigue failure

which led to explosive decompression in flight and breakup. All passengers and crew were killed

in both accidents. These two accidents indicated the catastrophic effects of material fatigue.

2.1.1 Flight 781

Flight 781 of British Overseas Airways Corporation (BOAC) crashed into the

Mediterranean Sea between the islands Elba and Montecristo on January 10th 1954. Everyone

was killed in this accident (29 passengers and 6 crew members). The main cause of the accident

was material fatigue failure which led to explosive decompression in flight and breakup.

Flight 781 started in Kallang Airport in Singapore, with last stopover in Ciampino

Airport in Rome, Italy where BOAC Comet 1 took off and headed to its destination, which was

London Heathrow Airport in England. Comet 1 took off at 09:34 GMT from Ciampino Airport

and at about 10:51 GMT a group of fishermen witnessed the aircraft crash into the sea.

After fishermen first started the search for survivors, navy ships Barhill and Gambia, and civilian

ship Sea Salvor searched for the survivors and for the remains of the aircraft. This was a start of

a very long and difficult investigation. Aircraft wreckage was found on the bottom of the sea and

transported to the Royal Aircraft Establishment for the puposes of further investigation.

There were several conclusions of what it could have happened like bomb on board,

engine turbine explosion, etc. Eventually, it was clear that the aircraft have broken up in the air,

but the true cause was recognized after second accident occurred three months later with Comet

1 of South African Airways.

2.1.2 Flight 201

South African Airways Flight 201 (Comet 1) crashed in the vicinity of Naples, on its way

from Ciampino Airport in Rome to Cairo in Egypt on April 8th 1954. Comet 1 took off from

Ciampino Airport at 18:32 UTC, and at 19:07 UTC, after reporting passing a beam of Naples and

contacting Cairo, crashed into the sea. All 21 passengers and crew were killed in this crash. After

many unsuccessful attempts to contact the aircraft, search has begun for survivors and the

aircraft.

Wreckage and bodies were found 110 km east from Naples and 48 km south-east of

Stromboli. After finding the wreckage, the investigation begun. Investigators also thought that

the cause could be an engine turbine failure. This investigation begun while the investigation of

Flight 781 still hasn’t finished, and Comets were already grounded in order to find the cause of

first accident. The similarities were huge in these two cases, hence the joint investigation was


14

conducted after this accident. After several years of investigation, the cause was found and it was

manufacturer design defects and material fatigue which caused the explosive decompression and

the breakup of both aircraft.

2.1.3 De Havilland DH 106 Comet design and characteristics

The first commercial jet aircraft was de Havilland DH 106 Comet. It was developed and

manufactured by de Havilland at Hatfield, Hertfordshire in United Kingdom. On July 27th in

1949 Comet 1 prototype flew for the very first time. Comet 1 had four de Havilland Ghost

turbojet engines located in the wings, a pressurised fuselage with silent and comfortable

passenger cabin, square windows, and also it was the first aircraft that showed a commercial

success back in 1952.

Comets began to have problems, one year after starting active service, especially with

two of them breaking up in flight in accidents described above. The cause was material fatigue in

the airframes, which was not well understood at the time. After the accidents occurred, Comets

were withdrawn from service and tested in order to discover the cause.

As already mentioned, main cause was fatigue, but there were many other reasons which

helped in process of aircraft breaking up like design defects which included dangerous stresses at

the corners of the square windows, and installation methodology also. It was revealed that

breaking up initiated with cracks around square window of the aircraft and then spread onto the

entire aircraft, tearing it into million pieces. Comets were re-designed with oval windows,

structural reinforcement and many other changes.

Figure 2 Comet 1A at Le Bourget Airport [30]


15

Comet 1A, which was an updated version of Comet 1, could take more weight, it had

larger fuel capacity and water methanol injection. Two Comets 1A were re-designed with

heavier gauge skins, reinforced fuselage structure and were renamed to Comet 1X.

2.2 1985 – Flight 123 – Boeing 747

2.2.1 Flight 123

Flight 123 of Japan Airlines (Boeing 747) crashed near Mount Osutaka on August 12th in

1985, when the aircraft lost vertical stabilizer because the maintenance done on rear bulkhead

wasn’t done correctly. This accident is known as the second deadliest aircraft accident in the

history of aviation. 520 people (505 passengers and 15 crew members) were killed in this

accident (of 524 in total). Only 4 people survived. Flight 123 took off at 6:12 of local time from

Tokyo International Airport to its destination, Osaka International Airport. 32 minutes later it

crashed into Mount Takamagahara, approximately 100 km from Tokyo.

Figure 3 Japan Airlines Flight 123; Boeing 747 at Osaka International Airport in 1982 [29]


16

Figure 4 Route of JAL123 and sequence of events [29]

The cause of the accident was the lost of vertical stabilizer due to incorrect repairs of rear

bulkhead. After the rear bulkhead gave way, the result was explosive decompression which tear

down the lines of all four hydraulic systems, ejected the vertical stabilizer, pressurized air rushed

out of the cabin and brought down the ceiling around the rear lavatories.

This lost made the aircraft uncontrollable. Seven years earlier, the same aircraft suffered

a tail strike incident at Osaka International Airport as Flight 115, in which the aircraft's rear

pressure bulkhead was initially damaged. The repair of the rear bulkhead didn’t comply with

approved methods. Back then, technicians repairing the aircraft used two separate doubler plates,

one with two rows of rivets and one with only one row when the procedure dictated one

continuous doubler plate with three rows of rivets to reinforce the damaged bulkhead.

The investigation of the Flight 123 accident revealed that the faulty repair reduced the

resistance to metal fatigue of that specific part by 70%. Also, it has been calculated during the

accident investigation that this incorrect method of installation would have failed after

approximately 10,000 pressurizations and the aircraft accomplished 12,318 successful flights

before the crash occurred.

Figure 5 This photograph shows the plane as it looked after explosive decompression; the vertical stabilizer

is missing (circled in red) [29]


17

Figure 6 Diagram of correct and incorrect repairs of the bulkhead [29]

Figure 7 Diagram of the aft pressure bulkhead [29]

2.2.2 Boeing 747

Boeing 747SR-46 (registered JA8119) was the Flight 123 accident aircraft. On January

28th 1974 it made its first flight. It had 18,835 cycles before the accident (one cycle is one take-

off and landing).

The Boeing 747 is a wide-body commercial jet and cargo transport aircraft, also called

Jumbo Jet. It is the most recognizable aircraft in the world and the first wide-body aircraft

produced. Boeing 747 was manufactured by Boeing's Commercial Airplane in the United States.

Boeing 747 main characteristics are four engines, double deck configuration, passenger,

freighter and other versions, possibility to convert passenger into cargo aircraft by removing the

seats and installing cargo door, etc.


18

Figure 8 Boeing 747SR-46 [30]

2.3 1988 – Flight 243 – Boeing 737

2.3.1 Flight 243

On April 28th 1988 Boeing of Aloha Airlines (Flight 243 from Hilo to Honolulu in

Hawaii) was severely damaged because of an explosive decompression in flight, but managed to

land safely at Kahului Airport on Maui. One person was killed and most of other people were

injured. This accident represents an important event in the history of aviation because it affected

safety policies and procedures.


19

Figure 9 Aloha Airlines Flight 243 [29]

The main cause of the accident was explosive decompression as a result of fatigue

failure. Flight 243 took off from Hilo at 13:25 of local time. After the aircraft reached its flight

altitude of 24,000 ft. about 43 km south-east of Kahului, small part on the left side of the roof

ruptured which made captain realize that the aircraft is rolling left and right, as well as the

controls were loose.

The result of this small rupture was explosive decompression which tore off most of the

roof. Captain steered the aircraft to the closest airport in Kahului. NTSB of the USA conducted

an investigation. The conclusion was that the crash was caused by material fatigue which was the

result of corrosion as the aircraft flew over seas for 19 years and was repeatedly exposed to salt

and humidity.

2.3.2 Boeing 737

The main characteristics of Boeing 737 are: it is a short-to-medium range, narrow-body

aircraft with two engines. This aircraft is Boeing's only narrow-body aircraft ever produced.

From 1964 many variants of B 737 were produced, such as: 737-100, 737-200, 737-300, 737-

400, and 737-500 models (these models are also called Boeing 737 Classic series).

From 1990 more variants of B 737 family were produced, such as: 737-600, 737-700,

737-800, and 737-900ER models (these models are also called Boeing 737 Next Generation.

Variants of Boeing 737 Next Generation Business Jet were produced as well. Boeing 737 series

are the best selling aircraft in the history of aviation.


20

Figure 10 Boeing 737-297 [30]

2.4 1992 – Flight 1862 – Boeing 747

On October 4th 1992, cargo Boeing 747 of airline El Al servicing Flight 1862 crashed

into the two high building complexes in Amsterdam neighbourhood called Bijlmermeer in

Netherlands. The aircraft lost engines on its right wing because of the fatigue failure in the pylon

of the engines. 43 people were killed and many more were injured. This accident is known to be

the deadliest aircraft accident ever to occur in Netherlands.

Boeing 747-258F took-off at 18:22 of local time from J.F.K. International Airport in New

York and headed to its destination which was Ben Gurion International Airport, with a stopover

at Amsterdam Schiphol Airport. Five minutes later, above the lake Gooimeer, a loud noise was

heard. That noise was made by two fuse pins (attaching engine number three) failing because of

material fatigue cracks. As the pins failed, engine number three detached from the wing, did

huge damage to the right wing flaps, struck and tore off engine number four, and scraped off 9 m

part of the leading edge of the wing.

Because of these damages done to the aircraft, the captain circled back to the aircraft to

make the emergency landing, without knowing that two engines were completely lost as well as

the right wing flaps. Hence, the emergency landing was conducted but unsuccessfully because

the aircraft was uncontrollable and at 18:35 of local time it ended up crashing into the buildings.

The investigation of this accident revealed the cause of the accident: failure of fuse pins in the


21

pylon of the engine due to material fatigue which led to engine detachment, damage of wing

flaps, tearing off another engine and part of leading edge, and the crash of the aircraft.

Figure 11 El Al Flight 1862 [29]

Figure 12 Fatal El Al flight of Boeing 747-200 [29]

Figure 13 A map of Amsterdam showing the aircraft's flight path (marked in green) [29]


22

2.5 2002 – Flight 611 – Boeing 747

On 25th May 2002 Boeing 747-209B of China Airlines (Flight 611) departed from

Taiwan Taoyuan International Airport and headed to Hong Kong International Airport. The

flight departed at 15:08 of local time. At 15:28 the aircraft broke up in flight near Penghu

Islands. Everyone on board was killed (all 225 passengers and crew).

Figure 14 China Airlines Flight 611 [29]

The main cause of accident was material fatigue which was result of incorrect

maintenance performed after the previous incident. In 1980, the aircraft was in tain strike

incident while landing in Hong Kong. The part of aircraft tail was scraped off along the runway.

The repairs that were conducted afterwards didn’t comply with Boeing Structural Repair

Manual.

Damaged skin was not trimmed and new installed doubler plate was inadequate and did

not cover the entire damaged area in order to restore required structural strength. After 22 years

of repeated cycles of loading, depressurization and pressurization, due to fatigue, hull of the

aircraft finally gave away, and the aircraft disintegrated in the air.

2.6 Other material fatigue related accidents

Accidents caused by fatigue and similar causes are presented below.

In 1945 (January 31st) Stinson Tokana operated by Australian National Airways crashes,

killing all 10 on board. The accident was caused by a fatigue crack in a wing-spar.

In 1948 (August 29th) Flight 421 of Northwest Airlines crashed due to fatigue failure in a

wing-spar root. All 37 passengers and crew were killed.


23

In 1957 (March 17th) Douglas C-47 airplane named “Mt. Pinatubo”, a presidential plane

of Philippine President Ramon Magsaysay crashed due to engine failure caused by metal fatigue.

25 people were killed.

In 1968 (August 14th) Los Angeles Airways Flight 417 lost one of its main rotor blades

due to fatigue failure. All 21 people were killed.

In 1968 (December 31st) MacRobertson Miller Airlines Flight 1750 lost a wing due to

improper maintenance what led to fatigue failure and killed all 26 passengers and crew on board.

In 1968 (July 3rd), BKS Air Transport Heathrow crash, an Airspeed Ambassador freight

aircraft experiences metal fatigue and crashes while landing, striking two unoccupied British

European Airways airliners.

In 1971 (October 2nd) British European Airways Flight 706, a Vickers Vanguard, breaks

up in mid-air due to corrosion and crashes near Aarsele, Belgium, killing all 63 passengers and

crew on board.

In 1976 (April 14th) Hawker Siddeley 748 crashed in Argentina due to undetected metal

fatigue as starboard wing failed outboard of engine.

In 1977 (May 14th) Dan-Air Boeing 707 crashed, caused by fatigue failure resulting in the

loss of the right horizontal stabilizer, killing all 6 on board.

In 1978 (June 26th) Helicopter Service Flight 165, a Sikorsky S-61, crashes into the North

Sea while en route to Statfjord oil field due to fatigue failure of a rotor, killing all 18 on board.

In 1980 (December 22nd) Saudia Flight 162, a Lockheed L-1011 Tristar, suffers an

explosive decompression over Qatar, killing two passengers who are sucked out of the aircraft;

the cause is traced to a fatigue failure of a main landing gear wheel flange.

In 1980 (March 14th) LOT Flight 7 crashed due to fatigue in an engine turbine shaft

resulting in engine disintegration leading to loss of control, killing all 87 on board.

In 1981 (August 22nd) Far Eastern Air Transport Flight 103, a Boeing 737, disintegrates

during flight and crashes near Taipei, Taiwan; severe corrosion in the fuselage structure leads to

explosive decompression and disintegration at high altitude; all 110 on board are killed.In 1989

(July 19th) United Airlines Flight 232 lost its tail engine due to fatigue failure in a fan disk hub

and 111 people were killed.

In 1997 (June 26th) Helicopter Service Flight 451, Eurocopter AS 332L1 Super Puma

crashed in Norway (Norwegian Sea). The accident was caused by a fatigue crack in the spline,

which ultimately caused the power transmission shaft to fail. The helicopter crashed into the sea

and killed all 12 people on board.

In 2005 (December 19th) Chalk's Ocean Airways Flight 101 lost its right wing due to

fatigue failure caused by inadequate maintenance practices, killing all 20 people on board.


24

List of questions

1. Flight 781 and Flight 201. Describe accidents and explain cause of accidents.

2. Flight 123. Describe accident and explain cause of accident.




6. Specify ten fatigue related accidents.


25

3 Basic theory of vibrations

3.1 Fundamentals of mechanical vibrations

Vibration is defined as an oscillation about a reference point. It is an everyday

phenomenon and it is usually very destructive to materials.

Dynamic forces in machine elements generate vibrations and once vibrations are

generated they can cause material fatigue. Three basic elements of every mechanical system are

spring, damper, and mass.

Figure 15 Mechanical parameters and components [4]


26

Figure 16 Simplest form of vibrating system – mass and spring [4]

After the system of a mass and a spring is put in motion it will oscillate with constant

frequency and amplitude. This system oscillates according to a sinusoidal function.

The sine curve obtained from a system which oscillates can be interpreted by its

amplitude (D) and its period (T). The number of cycles per second is called frequency (f) and it is

equal to the reciprocal of the period (T). The angular frequency is obtained when the frequency

(f) is multiplied by 2, which is then proportional to the square root of spring constant k divided

by mass m.

The natural frequency fn is defined as the frequency of oscillation. The sine function can

be expressed by the formula tDd nsin , where d is instantaneous displacement and D is peak

displacement.

3.2 Vibrations in aircraft structures

Vibration is any motion that repeats itself in a certain time interval, and it is also called an

oscillation. The motion of the pendulum is a typical example of vibration. Vibrations in the

system include the transfer of potential energy to kinetic and vice versa. If the system is muted,

the energy is lost in each cycle.

Most human activities involving vibration are in one of its forms. For example: we hear

because the vibrations are transmitted through our eardrums, breathing is associated with the

vibration of our lungs, including walking (periodic) fluctuations in our arms and legs and talking

due to vibration of our vocal cords.

Most motor vehicles or means of transport have vibrational problems due to imbalances

operating system. Examples:


27

Because the aircraft with piston engines should have a much more robust

construction;

Because they are much more burdened vibration than those with turbojet engines;

Imbalance diesel engine can produce seismic waves strong enough to be a nuisance in

urban areas;

Locomotive wheels at high speeds because of imbalances can be separated more than

an inch of track;

Turbine vibrations caused by spectacular mechanical failures. In general, the

vibrations resulting in faster wear and failures of engine components such as girders

and wheels, and also create and noisy.

Whenever the natural frequency of vibration of the engine coincides with the frequency

of the external stimulus, resulting phenomenon called resonance occurs. It leads to exceptional

deviations and failures. Literature is full of examples of the devastating effects of resonance.

The vibrations have catastrophic effect on structures, and because of that fact, vibration

testing is now a standard procedure in development of many different systems. Vibrations in the

field of aircraft dynamics and ageing are studied for two basic reasons:

The impact of vibration on the aircraft structure stability and durability, and the

occurrence of material fatigue in vital parts of the airframe;

Analysis of the vibrational characteristics of an airframe vital parts and assemblies, in

order to assess the validity and to detect cracks, wear out, ageing, etc.

Types of vibration are the following:

1. Free;

2. Free with damping;

3. Forced;

4. Forced with damping.


28

3.3 Analytical description of free and forced vibrations

3.3.1 Free vibrations

Figure 17 Free vibrations [4]

If there is a mass m hanging on a spring with constant k, the equation of motion of the

mass m is given by Newton's second law as kxma . The acceleration is derivative of velocity

and it is ''xa . Therefore, the equation reduces to: kxmx '' .

The same equation can be obtained if the law of energy conservation. By solving the

equation: 0'' kxmx , and thus obtaining a function that describes the motion of the mass

hanging from a spring. The equation is written in the form: 0'' 2 xx , where m

k . The

general solution of this equation is given in the form )sin()cos()( tBtAtx .

The period of this motion is

2T . T depends only on the mass and spring constant. If

we have the initial position of mass 0)0( xx and initial speed 0')0(' xx , it is easy to calculate

the constants A and B. Substituting 0 in )sin()cos()( tBtAtx we get 0xA . If we

derive )sin()cos()( tBtAtx , and then insert 0 in the resulting equation, we obtain

0'xB .

In this way we obtain a solution in the form )sin('

)cos()( 0

0 tx

txtx

.


29

If we take A and B to be cosA and sinB , for some and, and we obtain

next equations: 2

2

00

222 )'(

xxBA and

B

A

x

xtg

0

0'

. In this case, the solution

can be written in the form )cos()'(

)(2

2

00

2

tx

xtx . For example, the graph of such a

solution looks like in Figure 18.

Figure 18 Free vibrations [4]

Figure 19 Free vibration without damping [29]


30

Figure 20 The spring – simple harmonic motion of the mass–spring system [29]

The solution of stated equation that expresses the motion of mass is: )2cos()( tfAtx n .

It states that system will oscillate in harmonic motion with amplitude A and frequency fn.

The undamped natural frequency is marked with number fn . Also, fn is defined as: m

kf n

2

1 .

If free undamped system is set into oscillation, the extra energy is constant, but it

switches from kinetic form to potential form during the oscillation.

The velocity and the kinetic energy is 0, while the potential energy is equal to ½kD2. The

potential energy is 0 and the kinetic energy reaches its maximum at ½mV2, at the equilibrium

point. The sinusoidal function: tDd nsin can find the velocity by differentiating

tVtDdt

tDdV nnn

n

coscos)sin(

and obtain DfnV 2 . By applying laws of energy

conservation the natural resonance frequency can be obtained: m

kf n

2

1 .


31

Figure 21 Increase of mass [4]

An increase in the mass of a vibrating system causes an increase in period i.e. a decrease

in frequency.

3.2.2 Forced vibration

Forced vibrations occur when force f acts on the mass m that hangs from a spring with

spring constant k. In this case the equation for the mass m according to Newton's second law is:

)()()('' tftkxtmx . Also we have )cos()( tFtf . Now we want to express the equation

)()()('' tftkxtmx in the form )()()( txtxtx ph where )cos()( 0 ttxh (where angular

frequency of free oscillations is m

k0 ) is the solution of homogeneous equation

0)()('' tkxtmx . We want to obtain the particular solution in the form )cos()( tAFtxp .

Substituting particular solution in )()()('' tftkxtmx we obtain

2

0

2

1

1

k

FA . A problem

arises for 0 which is the case that will be considered separately.


32

Hence, the solution to the problem )()()('' tftkxtmx is given as

)cos(

1

1)cos()(

2

0

20 tk

Fttx

.

Here we also have a superposition of two harmonic vibrations. For a case 0 we

seek for particular solution in the form )sin()( 00 ttAtxp . Substituting particular solution into

the equation we obtain k

FA

2 . Hence, in this case, the general solution in the form:

)sin()cos()( 000 ttk

Fttx . This solution blows up in time, i.e. the amplitude is

increasing rapidly. Graph of such solutions is shown in a Figure 22 below.

Figure 22 Forced vibration [4]

Figure 23 Forced vibration [29]


33

It is important to note that vibration without damping, which are in the real world, affect

the system to vibrate with constant amplitude, i.e. with maximum shift values.

3.4 Vibrations with damping

There are free and forced vibrations with damping. When vibration damping forces are acting the

amplitude decreases cyclically depending on the damping factor k.

3.4.1 Free vibration with damping

Figure 24 Mass-spring-damper model [29]

Sum of the forces is expressed in the differential equation: 0''' kxcxmx . The solution

to this equation depends on the amount of damping. To solve this equation solve the

corresponding characteristic equation must be solved first, 02 kcm . We get

solutionsm

k

m

c

m

c4

2

1

2

2

2,1

. The general solution of equation 0''' kxcxmx is

then given as ttBeAetx 21)(

. There are three different outcomes.

In first case we have 04

2

m

k

m

cand kmc 2 . This leads to critical damping.

Solution is given in the form: tt

BteAetx 21)(

. Graph of such solution is shown in Figure 25.


34

Figure 25 Constant damping [4]

In second case we have 04

2

m

k

m

cand kmc 2 . In this case, viscous damping is

greater than critical. The solution is given in the form ttBeAetx 21)(

, but because such c this

solution numbs again. Such case is called great damping. Graph of such solution is shown in

Figure 26.

Figure 26 Viscous damping [4]

In third case we have 04

2

m

k

m

cand kmc 2 . In this case viscous damping is less

critical. As we obtain complex solutions of characteristic equation, the solution of initial

equation is given in the form ))sin()cos(()( 212 tAtAetx

tm

c

, where

24

2

1

m

c

m

k . If

we introduce substitute m

k0 we obtain

km

c

41

2

0 . In this case we have 0 . Such

case is called small damping. Graph of such solution is shown in Figure 27.


35

Figure 27 Small damping [4]

Figure 28 Free vibration with 0.1 and 0.3 damping ratio [29]


36

Figure 29 Mass, spring and damper [4]

When a damper is added to the system it results in a decrease in amplitude with time.

3.4.2 Forced vibration with damping

Equation of harmonic force is )2sin(0 ftFF . Sum of the forces is expressed in differential

equation: )2sin(''' 0 ftFkxcxmx . The solution can also be expressed as:

)2sin()( ftXtx . The result indicates that, at the same frequency f, the mass will oscillate

with a phase shift . The following formula defines the amplitude of the vibration X:

222

0

)2()1(

1

rrk

FX

where r is the ratio of the harmonic force frequency and undamped

natural frequency, nf

fr . The phase shift is defined by the following formula:

21

2arctan

r

r .


37

Figure 30 Phase shift [29]

3.5 Analysis of jitter and resonance in transport vehicles

3.5.1 Analysis of jitter in transport vehicles

Jitter is defined as an important and but undesired factor in all communications

connections and links. Jitter period is defined as the period between two times of maximum

effect or as the period between two times of minimum effect of a signal.

Its reverse is called jitter frequency. Jitter frequency, the more commonly quoted figure,

is its inverse. ITU defines jitter as frequencies at or above 10 Hz. Electromagnetic interference

may sometimes cause jitter.

Figure 31 Normal distribution to determine or predict jitter [29]

There are several types of jitter: random, deterministic, and total. Random jitter, also

known as Gaussian jitter, is defined as unpredictable electronic noise. Deterministic jitter (also

called non-Gaussian) is defined as clock timing jitter which is possible to predict. Total jitter (T)

represents random jitter (R) and deterministic jitter (D) as the combination: T = Dpeak-to-peak +

2× n×Rrms.


38

The standards for jitter measurement include jitter tolerance, jitter transfer function and

jitter generation. Devices that prevent jitter are: anti-jitter circuits, jitter buffers or de-jitter

buffers, a dejitterizer, a jitter filter, etc.

3.5.2 Analysis of resonance in transport vehicles

Simple system with elements of a weight and a spring has the natural frequency expressed as:

m

kf

2

1 , where k is the spring constant and m is the mass.

Example of a resonant system is a swing set which is a form of pendulum. If we push the

system with a period between each push that equals to the inverse of the natural frequency of the

pendulum, the swing will move higher and higher, but if the frequency is different, it won’t

hardly move. Pendulum resonance frequency, is obtained by the followingequation:

l

gf

2

1 , where g is the gravity acceleration (9.8 m/s2), and l is the length starting from the

pivot point to the mass centre.

Figure 32 Mechanical resonance in a mechanical oscillatory system [29]

3.6 Influence of vibrations on material fatigue of aircraft structures

Effect of vibration on fatigue and ageing is analyzed in the process of designing aircraft.

Vibrations lead to cyclic stresses, which are analyzed in Smith diagram and Wöhler curve.


39

Figure 33 Wöhler curve [29]

3.7 Sensor systems for measuring vibration characteristics

The most commonly used device for measuring vibration characteristics is an

accelerometer, and it measures proper acceleration of vibrations. Those accelerations are

measured in terms of gravity force. An accelerometer works as a damped mass on a spring.

When it detects acceleration, the mass is moved to the point where spring can accelerate the

mass. The displacement is used to measure the acceleration. Accelerometers are used to measure

the motion and vibrations in the structure which is subjected to dynamic loads.

There are several types of accelerometer such as: bulk micromachined capacitive

accelerometer, bulk micromachined piezoelectric resistive accelerometer, capacitive spring mass

base accelerometer, DC response accelerometer, electromechanical servo accelerometer, high

gravity accelerometer, high temperature accelerometer, laser accelerometer, low frequency

accelerometer, magnetic induction accelerometer, modally tuned impact hammers, null-balance

accelerometer, optical accelerometer, pendulous integrating gyroscopic accelerometer,

piezoelectric accelerometer, quantum (rubidium atom cloud, laser cooled) accelerometer,

resonance accelerometer, seat pad accelerometers, shear mode accelerometer, strain gauge

accelerometer, surface acoustic wave accelerometer, surface micro-machined capacitive

(MEMS), thermal (submicrometre CMOS process) accelerometer, triaxial accelerometer,

vacuum diode with flexible anode accelerometer, etc.


40

3.8 Adverse effects caused by vibrations: flutter, buzz, divergence

3.8.1 Flutter

Dynamic instability of an elastic structure in a fluid flow is called flutter. There are two

types of flutter: hard flutter, where damping decreases rapidly; and soft flutter, where damping

decreases gradually. Structures subjected to aerodynamic forces such as wings or aerofoils, are

designed carefully with aim to avoid flutter. For example, changing of the aircraft mass

distribution can generate flutter. Automatic control systems prevent vibration related to flutter.

3.8.2 Buzz

The shock wave in front or inside a duct may be unstable. This is called the buzz. The

shock waves can oscillate upstream and downstream. Buzz can cause severe structural damage or

failure.

3.8.3 Divergence

Divergence is a phenomenon which occurs when a surface deflects under aerodynamic

load in order to move the load. If the structure reaches the diverge point, the structure deflects

under the load.


41

List of questions

1. Define and explain vibrations.

2. Explain the simplest form of vibrating system.

3. Explain vibrations in aircraft structures. Specify types of vibrations.

4. Explain free vibrations.

5. What is natural frequency? Explain.

6. Explain forced vibrations.

7. Explain free vibrations with damping.

8. Explain forced vibrations with damping.

9. What is phase shift? Why is it important? Explain.

10. What is jitter? Explain.

11. What is resonance? Explain.

12. Explain how vibrations affect aircraft structures.

13. What are accelerometers? Give examples.

14. Explain flutter as adverse effect caused by vibrations.

15. Explain buzz as adverse effect caused by vibrations.

16. Explain divergence as adverse effect caused by vibrations.


42

4 Mechanical loads on aircraft structure

All mechanical structures including aircraft, satellites, ships, and others, have particular

structural loads. Structural loads are evaluated based on regulations, contracts, or specifications.

For acceptance testing and inspection, the accepted technical standards are used. Types of

mechanical loads are: static loads, dynamic loads, thermal loads and combined loads.

4.1 Static loads

Mechanical load is called static load when it refers to loads that are constant. An example

of a static load is weight of aircraft structure.

4.2 Dynamic loads

Mechanical load which is a dynamic load is temporary and short in duration. It is the load

that move or change when it is acting on a structure. An example of a dynamic load is force of

the wind. Dynamic loads also consider impact, momentum, vibrations, dynamics of fluids and

material fatigue.

4.3 Combined loads

Combined load represents a combination of several loads acting on the structure. An

example for combined load is combination of static and dynamic load.

4.4 Thermal loads

Thermal load is the temperature that effects on objects, like air temperature, solar

radiation, underground temperature, and some heat source equipment inside the object. The

change of the temperature causes thermal stress in the structures.

4.5 Aircraft structural loads

There are into two main categories of aircraft structural loads: limit loads and ultimate

loads. Limit loads represent flight loads and there are two types as well: maneuvering loads and

gust loads. Ultimate loads represent crash loads. Other loads important loads are pressure loads

and ground loads. Ground loads can come from adverse braking or taxiing. Aircraft loads impact

on fatigue because aircraft are constantly subjected to repeated loading and that can initiate

cracks.


43

List of questions

1. Describe and specify loads on aircraft structures.

2. Explain static loads.

3. Explain dynamic loads.

4. Explain aircraft structural loads. What are two main categories?


44

5 Influence of aerodynamic loads on aircraft structure

This Chapter describes loads, analysis of loads during the phases of the structure life.

Also, it explains how structural loads influence the structure through three stages:

1. Structural loads during design and qualification of aircraft structures;

2. Loads monitoring during usage;

3. Impacts due to aircraft modification.

5.1 Structural loads of aircraft structures

All loading scenarios in the early stage of design process of a new aircraft must be

considered and included, and they have to ensure that these loads can be handled throughout all

specific tasks. Figure 34 shows how loads are initially generated and how are they used

throughout the design, qualification and usage process.

Figure 34 Loads main tasks [9]

5.1.1 Loads and fatigue

The main prerequisite for successful design and safe operation of any new aircraft is

determination of loads and qualification for static strength and fatigue which is conducted

through tests of all important elements of the structure.


45

Ageing aircraft is not just the aircraft in terms of flight hours and flight cycles, it also

means that some of the reference data have changed during time, such as:

1. Airframe and equipment mass growth;

2. Enhancement of systems performance, especially engine thrust;

3. New configurations;

4. Update of flight control systems (electronically or hardware changes like added slats

or enlarged ailerons);

5. Mission profiles and additional/changed roles;

6. Actual usage spectrum.

5.1.2 Determination of design loads

Design of Loads, or Initial Design of Loads is one of the most important steps and

influences greatly the design of a component (such as wing or fuselage structure).

Figure 35 shows a typical “loads loop” is which loads are usually repeated several times

in the different phases of the aircraft design.

Figure 35 Loads loop [9]


46

5.1.3 Structural design criteria

Specification document Structural Design Criteria determines aircraft loads in accordance

with all collected regulations and requirements. Figure 36 shows V-n diagram which defines the

regime of speeds in combination with max/min allowable load factor Nz including gust

conditions.

Figure 36 V(Ma)-n diagram in altitude [9]

Figure 37 shows a typical flight envelope for the Tornado aircraft.


47

Figure 37 Altitude-Mach number envelopes [9]

Figure 38 indicates what part of the flight envelope is of importance for the investigation

of loads and shows points in the Mach-Altitude range for which loads are calculated.


48

Figure 38 Mach-Altitude points of loads model [9]

Influence on structural loads also include:

System pressures;

Cabin and fuselage bay pressures;

Temperatures and noise levels;

Local accelerations for qualification of equipment;

Vibration levels.

5.2 Loads monitoring

5.2.1 Monitoring of aircraft loads

Loads classification is:

1. Quasi-static loads:

1) Flight Loads:

Symmetric manoeuvres;

Asymmetric manoeuvres;

Deep and flat spin;

Gust loads;

2) Ground Handling:

Take off;

Landing;

Repaired runway;


49

Taxiing (asymmetric braking, turning etc.);

Towing, pivoting etc.;

3) Local and Internal Loads:

Max./min. aerodynamic pressures (outer surfaces);

Local accelerations;

System pressures;

Bay pressures (pressurised areas);

Hydrostatic pressures (fuel tanks);

Intake duct pressures (steady state);

Engine thrust;

4) Dynamic Loads:

Buffet (outer wing, vertical fin buffet, etc.);

Dynamic gust;

Vibrations;

Acoustic noise;

Limit cycle oscillation;

Shimmy (undercarriage);

Engine hammershock conditions (duct);

2. Fatigue Loads. [9]

Figure 39 shows measured buffet on a vertical fin in flight.


50

Figure 39 Fin buffet at high angle of attack (flight test results) [9]


51

5.2.2 Flight parameter envelopes

Loads depend on many flight parameters, the most important are:

1. Incidence or angle of attack (AOA):

Sideslip (for design the significant factor is Q, the product of sideslip and dynamic

pressure);

Control surface deflection angles (aileron, rudder, tailplane, etc.);

Lateral load factor Ny;

Vertical load factor Nz;

Roll rate /Roll acceleration;

Pitch acceleration;

Yaw acceleration;

2. Usually less important for load derivation:

Longitudinal load factor Nx;

Pitch rate;

Yaw rate. [9]

Figure 40 shows flight parameters during a typical pitch manoeuvre in relation to time.

Figure 40 MIL-SPEC pitch manoeuvre [9]

Figure 41 shows typical envelopes used in the early stages of design process.


52

Figure 41 Flight parameter envelopes for structural design [9]

5.2.3 Loads model

The Loads Model is the main tool for testing and calculating loads. It is a computer model of the

total aircraft, integrating the physics of motion, the aerodynamic dataset, structural design

criteria etc. and has interfaces to other disciplines.

5.2.4 Aircraft component loads and design cases

Loads can be calculated in three different ways:

Interface or component loads;

Load distributions;

Nodal point loads for Finite Element Analysis.

An example can be seen in Figure 42, showing the aircraft components:

Wing;

Wing spoiler;

Front fuselage transport joint;


53

Forward front fuselage;

Radom;

Rear fuselage transport joint;

Taileron;

Fin;

Rudder;

Airbrake. [9]

Figure 42 Load monitoring stations [9]

Figure 43 shows the load envelopes concept for the front fuselage and the wing root.


54

Figure 43 Major aircraft component loads envelopes [9]

Practical sequence of steps to calculate a flight load at a specific structural component

could be as follows:

1. Define mass and c.g.;

2. Define point in Mach-Altitude range;

3. Define sort of manoeuvre;

4. Simulate manoeuvre and calculate response parameters;

5. Calculate external net loads on component from aerodynamic pressures;

6. Convert external load distribution to nodal point loads on FE grid;

7. Analyse structure and determine local stresses.

5.3 Impact of changes on component loads

Table 4 Forces acting on an aircraft caused by various effects

Load Dependent on (list not complete)

Aerodynamic loads Incidence, sideslip, control angles, Mach, altitude, etc.

Inertia loads Nx, Ny, Nz angular ratesand accelerations, etc.

Engine thrust Mach, altitude, thrust, idle, etc.

Internal loads (e.g. cabin pressure) Specs, local accelerations

Actuator forces for control surfaces Hinge moment = f (Mach, altitude)

Hydrostatic pressure Local accelerations

Figure 44 shows how will an increase in the front fuselage result in higher front fuselage load.


55

Figure 44 Front fuselage transport joint critical load conditions; an increasing front fuselage mass will lead

to higher front fuselage loading [9]

Figure 45 shows the influence of wing loading conditions on wing loads. [9]

Figure 45 Influence of wing loading conditions on wing loads; adding mass to the wing (e.g. carriage of

stores) leads to reduced wing loads [9]

5.3.1 Qualification of loads via static and dynamic tests

Critical static and dynamic loads for the structure are checked during the early stages of

aircraft operational flight test but also through ground tests as required by the certification

procedures for the individual aircraft type. A typical layout of pressure measurement locations

for flight test is shown on Figure 46.


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Figure 46 Prototype pressure plotting for flight load survey [9]

5.3.2 Static load conditions and fatigue spectrum

Figure 47 shows a typical static loads criteria for a „care free handling“ flight control

system equipped aircraft.


57

Figure 47 Static loads design criteria for airframes [8]

Figure 48 shows excedance curves that are generated for combat aircraft.

Figure 48 Typical excedance curves for combat aircraft [8]


58

5.3.3 Conversion of external loads into structural airframe loads

Figure 49 shows a typical „coarse mesh“ finite element model of a wing structure with

wing box and flaps, where 40-50 „design loadcases“ were identified from the loads database of

500 load conditions and used for subsequent strength analysis. [9]

Figure 49 Coarse mesh FE model of wing structure [9]

Figure 50 shows model of a center fuselage for a fighter aircraft.

Figure 50 FE half-model of center fuselage structure [9]


59

Figure 51 EF2000 Global Model for Unified Analysis [9]

Figure 52 and 53 show an examples for a center fuselage bulkhead.

Figure 52 Coarse mesh FE model of center fuselage frame [9]


60

Figure 53 Fine mesh FE model for detail analysis [9]


61

List of questions

1. Explain influence of aircraft structural loads.

2. How are loads initially generated and how are they used throughout the design,

qualification and usage process?

3. Explain Loads Loop.

4. What is structural design criteria? Explain.

5. Specify all types of aircraft loads which are monitored during design.

6. Define fatigue loads. Explain.

7. What is Loads Model? Explain.

8. Specify load monitoring stations. Explain.


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6 Loads on aircraft elements

6.1 Loads on fuselage

Fuselage is a main section of the aircraft. It carries crew and the passengers (Figure 54).

Fuselage supports structure for the wings and tail, the cockpit, and structure to carry cargo,

passengers, and equipment.

Fuselage has to be able to resists bending moments (which are results of weight and lift

from the tail), torsional load (which is caused by fin and rudder) and cabin pressurization.

Aircraft fuselage also plays an important role in stability and maneuverability of the aircraft.

The loads that will be imposed on the aircraft have to be identified before the structure is

designed. Fuselage has to support two types of basic loads: ground loads and air loads. Fuselage

loads are also distributed and concentrated loads. Those loads transfer from the fixing bolts on

the wing, tail stabilizers, and the landing gear.

Figure 54 Fuselage structure [31]

6.2 Loads on wings

Wing loading represents the loaded weight of the aircraft divided by the area of the wing.

The more lift is produced by each unit area of wing as faster an aircraft flies.


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Figure 55 Loads acting along half of wing [32]

6.3 Loads on flight controls

Aircraft flight control system contains flight control surfaces, cockpit controls, connecting links,

engine controls, and required operating mechanisms to control the direction of an aircraft. Main

flight control surfaces are: ailerons, elevators and rudder. Secondary control surfaces are

spoilers, flaps, slats, air brakes, control trimming surfaces, etc.

Figure 56 Flight control surfaces [33]


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Figure 57 Loads on ailerons [34]

6.4 Loads on propulsion system

Propulsion system of an aircraft consists of an aircraft engine and thrust generators, like a

propeller or a propulsive nozzle. There are several types of engines: piston, turboprop, turbojet,

turbofan, etc.

Figure 58 Piston engine [30]


65

Figure 59 Turbofan jet engine propulsion system [35]

6.5 Loads on undercarriage

Landing gear of an aircraft supports the aircraft when it is on the ground, when taking off,

landing and taxiing. The main wheel tires are loaded by static loading.

Figure 60 Landing gear of Airbus A340-300 [30]


66

List of questions

1. Explain loads on fuselage.

2. Explain loads on wings.

3. Explain loads on flight controls.

4. Explain loads on propulsion system.

5. Explain loads on landing gear.


67

7 Degradation of aircraft structure

Whenever surfaces move at one another, a wear and tear occurs, i.e. damage to one or

both surfaces, which typically involves a progressive loss of material. In most cases, the wear has

adverse effect; causes increased clearance between the moving components, unwanted freedom

of movement, loss of precision, often vibrations, increased mechanical loading and faster

wearing, and sometimes fatigue.

The loss of a relatively small amount of material may be sufficient cause of operation

failure of large and complex machines. However, sometimes (as in friction) high wear is

desirable, for example, grinding, milling and polishing using the friction for fast removal of

material in a controlled environment, a low level of friction is sometimes desirable in some

break-in processes of mechanisms.

Wear can be slip wear, which occurs without the presence of solid particles, and abrasive

wear, which occurs with their presence. Under certain conditions, slip wear can produce debris

that causes further abrasion. This is why it should be borne in mind that the boundaries between

different types of wear often can not be uniquely determined.

7.1 Adhesion

Adhesion wear occurs as a consequence of intermolecular forces acting at the contact

points of the body and manifests itself through “redistribution” of material from one surface to

another.

This often leads to rupture and cold welding of working parts. Adhesion wear is reduced

by using different materials and hard surfaces resistant to this type of wear. This is most often the

wear case that leads to ageing of aircraft structures. Examples of wear in different elements:

1. Brake linings and discs of packet disc brakes;

2. Landing gear tires, especially in moments of touchdown while landing and braking in

the process of landing;

3. Gear parts of engines;

4. Joint mechanisms of landing gear, control surfaces, etc.;

5. Hydraulic mechanism.

Figure 61 Adhesion [author]


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7.2 Abrasion

Abrasion is the most common wear in the industry. It occurs as a result of harder material

peaks penetrate into the surface layers of softer material with rutting at the reciprocal motion of

the body.

Abrasive medium can be: rock, glass, ceramics, etc. Wear therefore occurs during the

excavation of ore, crushing, extraction and shipment. Machine parts are then subjected to high

stresses or so-called abrasion between the two bodies (Two-Body Abrasion). The fight against

this kind of wear requires very hard, dense and resistant materials for surface protection.

Abrasion between the three bodies (Three-Body Abrasion) occurs in devices such as

pumps or valves when abrasive medium is stuck between the friction surfaces (e.g. bearings).

Even here, due to the high stress, strong and durable materials must be used.

The main difference between these two types of abrasion is that abrasion between the two

bodies occurs exclusively because of hard bumps on the surfaces in contact, while in three-body

abrasion there are two surfaces and hard abrasive particles in between that move freely and cause

damage.

Some of the ways to reduce abrasion are:

1. Selection of pairs of materials resistant to abrasive wear;

2. Appropriate treatment of the surface layers;

3. Separation of surfaces with layer of fluid lubricant.

An extreme form of abrasive wear is when the aircraft flies into a cloud of volcanic ash,

and fine chipped stones and other debris cause wear of engines’ bearings and rotor blades. It also

occurs on the surface of the aircraft, due to the impact of solid particles on the surface of the

aircraft as fuselage skin, glass, etc.

Abrasive wear occurs when bird flies into the engine and crushed bone fragments act as

an abrasive medium; and also when the engine intakes larger chunks of ice.

Figure 62 Abrasion [author]


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Figure 63 Abrasion between two objects [author]

Figure 64 Abrasion between three objects [author]

7.3 Erosion

Erosion is a type of wear generated by the action of fluid particles (with or without hard

particles of fluid) that struck the surface of the body with great velocity. The severity of erosion

depends largely on the velocity and angle of impact of particles, and their hardness. There are

two main forms of erosion:

1. Erosion with obtuse angle is a process where most of the energy is consumed on

surface deformation; preventing this form of erosion requires a flexible protective

layer, usually an elastomer;

2. Erosion with acute angle is a process that is similar to abrasion and cutting; to reduce

the rate of wear, a great hardness of the surface is requred.

Some of the ways to reduce erosion are:

1. Elimination of solid particles from the fluid;

2. Changing fluid angle of attack on the surface;

3. Decrease of the relative velocity of fluid;

4. Selection of suitable material;

5. Additional changes to the surface material in order to improve its characteristics.


70

Figure 65 Erosion [author]

7.4 Corrosion

Corrosion occurs in cases when an adhesive or abrasive wear occurs in combination with

corrosive environment. The rate of loss of material can be very high. The reason for this lies in

the fact that the coatings for corrosion protection remove easily because of wear, leaving the

unprotected metal to corrode rapidly. Stable oxide layer that prevents the progress of corrosion is

lost due to adhesion or abrasion wear.

Corrosion itself is electrolytic process that involves changing the electrons and ions. It

can occur between different metals or between different parts of the same metal or alloy where

there is a difference in electrochemical potential. The difference arises also because of the

presence of oxides, various impurities, the alloy phases, etc. For corrosion to occur there is a

conductive electrolyte (moisture, salt water, etc.) to establish an electrical circuit.

Corrosion severely threatens all exterior parts of the aircraft, particularly in areas of high

pollution (creating acidic moisture) or near the sea.

7.5 Cavitation

Cavitation wear is actually subspecie of corrosion wear. It occurs when the pressure in

the fluid lowers the value of the vapor pressure and can lead to the vapor bubbles. They drift in

the area of higher pressure where they implode (return to the liquid phase).

If the implosion of vapor bubbles occurs near a solid wall, it leads to its damage. The

phenomenon is accompanied by vibrations and noise. Cavitation wear often occurs in hydraulic

system (phenomenon of cold fermentation of hydraulic fluid), causing the system to

overpressure.

7.6 Fatigue

The weakening of a material which is caused by repeatedly applied loads is called

fatigue. It represents the fast-growing and localized damage to the structure which happens when

a material is subjected to repeated loading and unloading. When loads are bigger than the level

the material can hold, cracks start to form, grow, reach their critical size, propagate rapidly and

ultimately the whole structure will fracture.


71

List of questions

1. What is adhesion? Explain.

2. What is abrasion? Explain.

3. What is erosion? Explain.

4. What is corrosion? Explain.

5. What is cavitation? Explain.

6. What is fatigue? Explain.


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8 Corrosion

The gradual tearing and destruction of materials by chemical reaction with environment

is known and called corrosion. This refers to electrochemical oxidation of metal materials in

reaction with an oxidant like oxygen. Example of electrochemical corrosion is rusting. Corrosion

also occurs in other materials like ceramics or polymers.

8.1 Corrosion in aircraft structures

Corrosive processes or corrosion, which are more or less present in all engineering

constructions, convert a large number of useful metal compounds into useless, even harmful,

corrosive products. Most often it is another expression for metal oxidation. That means the great

importance of high-quality corrosion protection, and research shows that a quarter of corrosion

damage can be prevented by applying modern technology protection means.

Corrosion processes affect the classification of metals. Material’s resistance to corrosion

divides them into classes: noble (gold and platinum), which do not react with oxygen forming

oxides; non-noble (aluminum, copper, zinc, etc.) which react with oxygen and other substances,

forming layers of oxides and other compounds that are compact and protect the metal from

further degradation; other non-noble (iron, steel, etc.), which create a layer of oxide on the

surface, which is not stable and does not prevent new oxygen contact with the metal and

corrosion can progress without control.

Corrosive processes can be divided according to the mechanism of the corrosion process

with regard to manifestation of corrosion. Corrosion occurs in metallic and non-metallic

construction materials; hence division of corrosion with regard to metals and non-metals is also

used. In today’s engineering construction industry, including the aerospace industry, metals are

still basic materials, which brings out special attention to corrosion.

Corrosion is classified:

1. According to the mechanism of the process: chemical and electrochemical corrosion;

2. According to the appearance: general, local, and selective corrosion.

8.2 Chemical corrosion

Chemical corrosion of metals occurs in nonelectrolytes, i.e. in the media which do not

conduct electricity, wherein the compounds of metal with non-metal elements are created

(usually oxides and sulphides). The most important nonelectrolytes that in practice cause

chemical corrosion of metals are certainly hot gases and organic liquids. Chemical corrosion of

metals is formed by reaction of the crystal lattice metal atoms with the molecules of an element

or compound from the environment, in which the compound molecules are formed directly to the

corrosion product.

In order to prevent chemical corrosion, the construction of the aircraft that is to be used

must be made of materials resistant to chemical corrosion. Hence, aluminum is extensively used

for aerospace structures, because of its resistance to corrosion and action of most acids. The air

resistance of aluminum is based on the creation of a thin oxide layer which is structurally

attached to the metal surface, which does not peel and thus protects the metal from further


73

oxidation. Thickness of oxide layer, which is only a few thousandths of a millimeter, is so dense

that moisture and air cannot get through to the aluminum (4Al + 3O2 -> 2Al2O3).

For example, thoroughly washing the aircraft Canadair CL-215 after the daily service of

firefighting and taking fluid from the sea is the most important action to prevent chemical

corrosion. If necessary, organic and inorganic coatings are used.

8.3 Electrochemical corrosion

Electrochemical corrosion of metals occurs in the electrolyte, i.e. in the media with ionic

conductivity. It is a redox process in which metal atoms are oxidized as a reductant (electron

donor) in the free cation with simultaneous reduction of an oxidant, so-called depolarizer

(electron acceptor). An anode, cathode, and electrolyte are necessary for electrochemical

corrosion to occur. Each metal has its own electrochemical potential, which is listed below in the

Table 5.

If only one metal is included in the process, it is then electrochemical corrosion, and if

the process includes two or more metals, it is then a special type of electrochemical corrosion or

galvanic corrosion.

Table 5 Electrochemical potential of metals

Magnesium and magnesium alloys -1.60 to -1.63 V

Zinc -0.98 to -1.03 V

Aluminium and aluminium alloys -0.76 to -1.00 V

Soft and high-alloy steels -0.60 to -0.71 V

Noncorrosive Steel -0.42 to -0.58 V

Aluminium bronze -0.31 to -0.42 V

Tin -0.31 to -0.32 V

Copper -0.30 to -0.57 V

Stainless steel shaft -0.25 to +0.06 V

Bronze -0.24 to -0.31 V

Lead -0.19 to -0.25 V

Nickel -0.10 to -0.20 V

Titan -0.05 to -0.06 V

Gold +0.10 to +0.22 V

Platinum +0.19 to +0.25 V

Graphite +0.20 to +0.30 V

If the system has different metals, the one that has a higher negative voltage will be the

anode, while one with smaller negative voltage or positive voltage will be the cathode and will

create a galvanic part. In terms of corrosion, especially in the outer parts, which are in contact

with atmospheric conditions (therefore in the possible electrolyte), it is important to use the same

as or similar materials, so anode and cathode could not be formed. Thus, aluminium airframe

shall be bonded by aluminium rivets, and not steel, just as the copper roof should only use

copper. Galvanic corrosion normally occurs in aircraft operations, and sometimes

electrochemical corrosion occurs.

Electrochemical corrosion occurs in natural and non-natural water, in water solutions of

acids, alkalis, salts and similar substances, in moist soil, in fluids of biological origin (reason to


74

count on the occurrence of corrosion in the area below the toilet), in molten salts, oxides and

hydroxides, and in the atmosphere.

The most important is atmospheric corrosion for aircraft operation, and it occurs with

precipitation, or in the water absorbate or condensate, which arises due to the humidity on the

surface of metal and has the character of an electrolyte. It is particularly high in the marine

environment.

There are ways to prevent galvanic corrosion of the aircraft. The basic prerequisite for the

anti-corrosion construction is optimized by reducing the electrical potential of metals in contact

or in vicinity. It is important that this principle stays alive in the aircraft maintenance also. In

aircraft exploitation the most important is prevention the creation of electrolytes, avoidance of

entering the area with the possibility of creating an electrolyte, i.e. regular cleaning of the

aircraft.

8.4 General corrosion

General corrosion affects the entire surface of the material, and it can be uniform or

uneven. Uniform general corrosion is technically the least dangerous because the process can be

easily monitored and predicted when a particular part needs to be repaired or replaced with new.

Uneven general corrosion is much more dangerous. General corrosion occurs when the entire

surface of the material is exposed to aggressive environment under approximately the same

conditions with regard to internal and external corrosion factors. It takes up to approximately 33

percent of the aircraft structure corrosive processes.

8.5 Local corrosion

Local corrosion attacks only some parts of the exposed surface, and is also the most

widespread manifestation of corrosion. Local corrosion can be divided into sunspots, pitting,

subsurface, contact, galvanic, stress, and intergranular corrosion. It takes up to approximately 37

percent of the aircraft structure corrosive processes, with the exception of sunspots corrosion.

Types of local corrosion of aircraft are: sunspots corrosion, pitting corrosion, subsurface

corrosion, contact corrosion, galvanic corrosion, stress corrosion, and intergranular corrosion.

Pitting corrosion is closely localized form of corrosion that occurs when a medium that

causes corrosion attacks material and causes the formation of small holes. It usually occurs in

places where the protective coating is broken due to mechanical damage or chemical

degradation. Pitting corrosion is one of the most dangerous forms of corrosion because it is very

difficult to predict and prevent, and relatively difficult to detect, it happens very quickly and

penetrates the metal without causing visible loss of weight. It mainly occurs in steel materials.

Subsurface corrosion occurs when the focal pitting corrosion widens in depth of

material; and delaminates it. The most widespread subsurface corrosion occurs in rolled metals

in contact with sea water and acids. Bubbles appear in the surface of the material because solid

corrosion products are accumulating in its interior, and those products have volume greater than

the volume of completely destroyed materials.

Contact corrosion can be divided into galvanic contact corrosion that occurs when

contact of two different metals occurs and crack contact corrosion that occurs when the contact

of two parts of the same metal occurs or when the contact of metals and non-metals occurs.


75

Galvanic corrosion occurs when two metals with different electrical potentials

electrically connect, either by physical contact or through a medium that conducts electricity. A

system that meets these criteria will form the electrochemical cell that will conduct electricity.

Induced electricity then withdraws electrons away from one metal, which acts as the anode, and

the other metal acts as the cathode and receives electrons. Galvanic corrosion is greatest near the

surface where two metals are in contact. Reducing of galvanic corrosion preferences is achieved

by selection of materials.

Stress corrosion occurs when some part is simultaneously exposed to aggressive media

and tensile stress. Therefore, it occurs on parts of aircraft such as the fuselage and wings, as a

result of pressurisation and aerodynamic forces. Due to the action of tensile stress, area of the

metal expands, creating a small crack, and current cracks increase. This increases the surface

area exposed to corrosion, and slots keeps means (media) that causes corrosion. It is the most

dangerous form of corrosion in aircraft, because it connects (superimposes) with the influence of

chemical and electrochemical corrosion.

Intergranular corrosion destroys the material at the grain boundaries and consequently

spreads in depth. This type of corrosion occurs mostly in alloys. Intergranular corrosion is one of

the most dangerous forms of corrosion because it can stay unnoticed for long time and rapidly

reduce the strength and toughness of the material. The final result of intergranular corrosion is

fracture or even dissolution of the material in grains. The most common affecting materials are

non-corrosive steels, nickel-based alloys and aluminium-based alloys.

8.6 Selective corrosion

Selective corrosion is a rare form of corrosion in which it attacks one element of metal

alloy, and it results in modified structure. The most common form of selective corrosion is

dezincification where zinc is separated from the brass alloy. The rule is that this type of

corrosion does not participate in the corrosion processes of aircraft structure.

8.7 Influence of corrosion on aircraft ageing

Corrosion effects on aircraft ageing in the way of reducing the bearing section and

increasing the negative impact of stress concentration (corrosion often occurs in areas of

narrowing, slots, etc., hence in the places of increasing corrosion impact, and corrosion activity

creates additionally very dangerous stress concentrators (Aloha accident, Boeing 737)).

8.8 Corrosion protection methods

Methods to protect the aircraft from corrosion are:

1. Selecting corrosion resistant materials (aluminium is more resistant to corrosion than

steel, and composites are more resistant than aluminium);

2. Designing and introducing construction measures (avoiding the slots that gather

excessive moisture, more aggressive media are used, sealing is improved, etc.);

3. Installing protective coatings (galvanizing, chroming, browning, etc. or colouring);

4. Using electrochemical methods of protection (anodic and cathodic protection).


76

List of questions

1. Explain corrosion and its effect on aircraft structures.

2. Explain chemical corrosion.

3. Explain electrochemical corrosion.

4. Explain importance of galvanic corrosion effect on aircraft structures.

5. Explain general corrosion.

6. Explain local corrosion. Specify types of galvanic corrosion.

7. Explain influence of corrosion on aircraft ageing and give example.

8. What are the methods to protect aircraft from adverse effects of corrosion?


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9 Material fatigue

Repeatedly applied loads cause the weakening of a material. This phenomenon is called

fatigue. It represents the fast-growing and localized damage to the structure which happens when

a material is subjected to repeated loading and unloading. When loads are bigger than the level

the material can hold, cracks start to form, grow, reach their critical size, propagate rapidly and

ultimately the whole structure will fracture.

9.1 Fatigue characteristics

Fatigue strength (fatigue limit, endurance limit) is defined as value for stress amplitude

below which the material will not fail for any number of cycles. The fatigue strength of the

structure can be increased with round holes and oval corners, as well as it can improve fatigue

life. Fatigue life, Nf, is defined by American Society for Testing and Materials (ASTM) as the

number of stress cycles that material sustains before failure occurs.

9.2 Material fatigue and cracks

Fatigue is a phenomenon of gradual damage of materials due to long-term periodic

fluctuating loads (stresses). Fatigue occurs mainly in dynamically strained constructions. During

dynamic loading, the material will break under much less stress than the maximum tensile

strength (σM), and even the yield point (σ0.2). Material resistance to dynamic or cyclic (flickering)

stress is called fatigue strength of materials.

The greatest stress, in absolute value, which material may be submitted to at any number

of cycles for a given asymmetry coefficient and load form of the sample, is called a permanent

dynamic strength of the material. Permanent dynamic strength is also called dynamic endurance

or permanent oscillating strength.

Fracture that occurs due to fatigue can be recognized when there are two clearly

separated surfaces: wavy-smooth and dark surface, which indicates a slowly propagating crack

and grainy and silvery-light surface, wich is a result of the current residual refractive supporting

part. A wavy surface has “wave peaks” which indicate one phase of cracking, as tree rings

indicate tree age. In next phase load cycles smooth its fibers, and the material under the influence

of air, moisture, and corrosion darkens. This is followed by a new cycle of cracking, stabilizing,

smoothing, and darkening.

While investigating the causes of the crash, the crash site is being closely monitored,

especially every broken part of structure, according to the criterion of fracture surface

appearance, as it is already described. If any broken part of the structure points to fatigue, the

investigation is continued in that direction.


78

Figure 66 Example of screw fracture due to material fatigue [29]

Figure 67 Example of fracture of hexagonal profile due to material fatigue; it is clear that the dark area is a

result of gradual, propagatory material fracture and light part indicates fast fracture [29]


79

Figure 68 Example of handle fracture due to material fatigue; besides the fact that fracture surfaces have

different shapes, the area of slow and prolonged propagating crack has a darker colour, because it is oxidized

(corroded), in contrary to the area of current fast fracture which is grainy and light [29]

Figure 69 Example of piston rod fracture due material fatigue; figure clearly shows the area of slow and fast

(current) fracture; initial area is also apparent, where beginning of cracking took place, due to stress

concentration, inclusions, local weakening, etc. [29]


80

Figure 70 Example of screw fracture due to material fatigue on the surface, caused by irregularities in the

structure, i.e. by damage to the edge (initial crack), cyclic loads due to incorrect torque, overload, corrosion,

etc., (not necessarily all conditions and not necessarily in that order) [29]

Figure 71 Example of fracture of aircraft structure of de Havilland DH 106 Comet after the accident in 1954,

due to fatigue caused by cyclic pressurisation, stress concentration, lack of advenced material technology and

errors in technology production [29]

Although at first glance it does not look like it, the share of material fatigue in all

fractures is 80% to 90%. The impact of fluctuating loads can be monitored in high or low

temperatures, and with the simultaneous action of aggressive media (e.g. sea water).

Initial cracks that occur due to fatigue are the sharpest natural cracks that can be very

difficult to detect before the fracture occurs. Fracture due to fatigue begins at the site of the

largest concentration of stresses what then leads to fatigue caused by fluctuating loads, although


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the maximum load, which occurs in cycles, is well below the elastic limit, as shown in the

diagram in Figure 72.

Figure 72 Fracture limits [29]

An explanation of how damaging material in fatigue is very complex and it is closely

associated with the behavior of defects in material (dislocations). During the action of long-term

dynamic stresses, glide lines are produced in the material. This is an indication that there has

been a local plastic deformation in the center of maximum stress (peak of initial crack in the

material). A hardening of material occurs in these places and, with further periodic loads, initial

microcracks generate, which expands until the load-bearing section is reduced so that the

maximum value of the variable stress can cause immediate fracture.

Very small cracks (cracks) are regular occurrence in structures and their parts and as such

are harmless for construction. This shows that microscopic defects are always present in the form

of irregularities in the crystal lattice of the metal, and are due to a number of imperfections in the

process of developing. Fatigue in this case represents growth and fusion of these irregularities,

followed by the formation of cracks, and its expansion to the final rupture. The process of fatigue

damage is usually divided into three steps:

1. Initiation or initial cracking;

2. Expansion or growth of cracks;

3. Violent fracture.

In addition to the above stated, generally accepted definition of material fatigue

occurrence, there is a simpler and more accurate interpretation of the material fatigue occurrence,

although the peak loads are more or less below the elastic limit.

Reference diagram “stress-strain”, which is based on the above discussion, applies to the

average value of stress in relation to deformation.

The average value of the strain in relation to the deformation is valid for 99.9 to 99.99

percent of the material.

There are localities due to irregularities, defects and inclusions which reduce the local

load-bearing section and create stress concentrations.

At these locations, in terms of average loads, crossing the border of elasticity may happen

and enter the plastic area. After unloading the material becomes stiffer and loses its ability to

elastic deformation. In the following phase or one of the phases after that an initial crack may

occur.


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At certain places specified load, which is by reference “stress-strain” diagram, below the

elastic limit, can instantly turn the value of fracture limit and the initial crack can occur. Next

what follows is crack propagation and violent fracture. The causes of material fatigue are:

1. Variable load;

2. Structural stress concentration;

3. Stress concentration in the material of the structure;

4. Decrease in toughness at low temperatures.

Variable load is stress that changes size and direction due to time intervals. If time

periods are regular time periods, and values are equal to the minimum and maximum values,

then it is a cyclic stress, and if time periods are irregular, then it is a stochastic variable load.

In the case of cyclic loading, it may be symmetrical and unsymmetrical.

Load changes in time and that is the main difference when compared to the original

approach in the dimensioning of parts using Hook’s law. Therefore, it is necessary to classify the

following types of load: constant, single variable, clean single variable, alternating variable, and

clean alternating variable. The mere approximation of dynamic changes with sinusoid law is

“engineering error” which is considered to be negligible for collegiate level. A detailed review of

certain types of loads in respective to time periods are given in Figure 73.

Figure 73 Stresses versus loading [29]


83

In consideration of cyclic stresses there is Wohler’s curve which emphisizes the

relationship between the number of stress cycles to failure N and nominal stress range Δσ: Δσm x

N = C, where m and C are constants dependent on the material, type of weld, type of load and

environmental conditions. Nominal stress range Δσ is defined as the difference of the maximum

upper stress cycle σmax and the the maximum lower stress cycle σmin: Δσ = σmax – σmin.

Wohler’s curves are determined with the experiments on samples (tubes), which are

subjected to a constant stress with alternating amplitude until they fracture, and the service life

can be learned from such experiments (the life of the compound). Stresses can be: constant,

single variable, clean single variable, alternating variable, and clean alternating variable.

Figure 74 Wohler’s curve [29]

One of the earliest and best-known approaches in the proper analysis of dynamic loads is

given by August Wohler in 1871. According Wohler's diagram durability of the component

increases with decrease in stress σa and vice versa.

However, their mutual dependence is not linear. Wohler's curve asymptotically

approaches the value of the stress which is called “dynamic durability”. Less the σa is, higher

number of cycles to failure the structure could withstand. The dynamic durability Rd is the

biggest variable (dynamic) stress that the material can withstand with a virtually infinite number

of cycles without causing failure.


84

Figure 75 Wohler’s diagram [29]

Every body is made of elementary particles or atoms which are connected via molecules

in one single compact unit. To ensure that set of molecules could survive together,

intermolecular forces must exist between individual molecules that will not allow their

dissipation, so the body remains “firm”. If we take into account that the body is composed of an

enormous number of molecules that are piled up in certain layers, one above the other, then it is

possible to understand that “power lines” or stress lines are being created, which stem out from

series of intermolecular forces that hold the body together.

Stresses are concentrated in areas where the continuity of the shape is interrupted

(crossing of small section into the larger one), where the continuity of sufaces is interrupted

(cuts, sectional vents) or where there is interruption of continuity in the structure of the material

(shrinkages, bladders, graphite nests). In these places of the structure, due to variable loads,

cracks in the material are developed. When maximum voltage exceeds the highest yield strength

impact load occurs, and the material on these sites is plastically deformed. If there is more

frequent incidence of such overloads, the material yields over time in such areas and fatigue

starts to develop.


85

Figure 76 Stress concentration [29]

Particularly sensitive sites for stress concentration to occur in the aircraft are:

Skin elements, particularly in the areas of rivet holes (Figure 75) and window frames;

Lugs and bolts for engine attachment;

Lugs and calibrated screws hold empennage;

Vital parts of landing gear;

Rotating parts of jet engines.

Stress concentration in the material structure is the result of imperfections, defects,

inclusions and defects in the material, which cause local overload.

Irregularities in the structure of the material can be:

Localized (dotted);

Lined;

Flatted;

Spacial.

All metals are realistically imperfect and imperfections are related to their crystal lattice,

hence, theoretically there are 1 percent of positive properties, including the proper arrangement

of atoms and intermolecular forces.


86

Figure 77 Point defects of crystals [29]

Local or dotted irregularities are caused by following occurrences of irregularities:

Lined as the result of accumulation and multiplication of dotted ones;

Flatted as the result of accumulation and multiplication of lined ones;

Spatial as the result of accumulation and multiplication flatted ones.

Hence, it all occurs according to propagation principle: point, line, surface, volume.

To understand the irregularities it is sufficient to analyze localized irregularities:

Void (vacancy) occurs when there is a missing atom in the grid.

A grid can have more atoms (interstitial impurity), when another atom, which

represents an impurity, is placed in the full grid, or between two full grids;

Replacement impurity (substitution impurity) occurs when in the place of one missing

atom comes another atom, which then represents impurity;

Frenkel’s defect is when an atom leaves its place and moves into the space between

the second grid, or two other grids;

Half interstitial irregularities occur when “local” atom comes in between the metal

grids, and that it is not lacking in any of other grids.

As long as there is a local problem or local problems, there is practically no problem at

all. The problem occurs when local problems start to build up in lined ones, flatted and spatial

resulting in spiral of problems and occurrence of fracture.

Even spatial imperfections are not a problem that can be observed with non-destructive

testing. They are the only precondition or the cause of the first microcracks. Spatial irregularities


87

in terms of dynamic and cyclic loads cause emergence of the first microcracks and that's the

moment when phenomena occur that can be observed and detected using NDT methods.

Toughness, as the most important feature of construction materials is a property that

helps material to resist dynamic loading of all types. Toughness is also “fracture tendency” and

has a unit in joules (J) and shows how much energy should be used to disintegrate concrete

material or how much it takes for it to break or to be destroyed.

Each metal with decrease in temperature loses toughness. The problem is the gradient of

toughness decrease. While decrease in toughness is low, essentially there is no problem. When

an accelerated decrease in toughness starts, brittle fracture or fracture without previous

deformation occurs (without previous warning). Prevention of brittle fracture is a priority for all

aircraft structures, which operate during flight in the field of low temperatures, which make them

susceptible to brittle fractures.

Brittle fracture occurs when toughness, due to low temperature, drops to 50% of its

original value. Below this temperature, metal breaks like glass. This is reason why structure

should be designed in the way that each metal in the construction of the aircraft has a critical

temperature for a minimum of 30 degrees Celsius lower than the lowest exploitative temperature.

This can be achieved with appropriate technology and production materials.

The worst thing that can happen if defects occur in the sequence, as it was the case of de

Havilland DH 106 Comet (Figure 77) and accidents which have occurred in 1954. It is following

sequence of irregularities:

1. Macro stress concentrations (rectangular and oversized windows openings);

2. Stress concentrations (countersunk holes for rivets made with poor technology that

caused the initial cracks);

3. Micro stress concentrations (inclusions in aluminum weight);

4. The sensitivity of the metal to brittle fracture due to a critical temperature close to the

lowest exploitative temperature;

5. Improper dimensioning (too small thickness of the aircraft skin).


88

Figure 78 De Havilland DH 106 Comet [29]

List of questions

1. Explain material fatigue and its characteristics.

2. Specify and explain the process of fatigue damage.

3. Specify causes of material fatigue and explain stress concentration.

4. What is Wohler curve? What is Wohler diagram? Explain.


89

10 Basics of fracture mechanics

Fracture mechanics is studying the propagation of cracks in materials. It is the field of

mechanics that uses methods to calculate the force on a crack and to characterize the resistance

of the material to fracture. Fracture mechanics develops methods to improve the performance of

mechanical components. Crack growth prediction is at the nucleus of the damage tolerance

discipline. If we want to enable a crack propagation, three ways of applying force can be

conducted:

1. Mode I fracture – Opening mode;

2. Mode II fracture – Sliding mode; and

3. Mode III fracture – Tearing mode.

10.1 High cyclic cracks

The situations that require more than 10,000 cycles to reach the failure or the situations where

stress is low and deformation is elastic, have the most attention.

Materials performance is usually described by an S-N curve, which is also called

a Wöhler curve. Figure 79 shows the graph of the cyclic stress (S) in relation to the scale of

cycles to failure (N).

Figure 79 S-N curve (Wohler curve) [29]


90

Many factors like corrosion, temperature, residual stresses, and other can influence the S-

N curve progression. The method to assess the influence of the mean stress to fatigue strength is

called the Goodman’s line (as shown in Figure 80). Goodman’s relation is quantifies the

interaction of mean and alternating stresses on the material fatigue life. A Goodman’s diagram is

a graph of mean stress in relation to alternating stress. The area below the curve on Figure 80

indicates that the material should not fail given the stresses. The area above the curve on Figure

79 represents likely failure of the material.

The Goodman’s relation is expressed as:

ts

m

fata x

1 where a is the alternating

stress, m is the mean stress, fat is the fatigue limit, and ts is the ultimate tensile stress of the

material.

Figure 80 Mathematical description of Goodman’s line [29]

10.2 Influence of stress concentration on material fatigue

A location in an object where stress is concentrated represents a stress concentration.

When force is equally distributed over its area, an object is the strongest. Crack can create

reduction in area which can result in an increase in stress, and material can fracture, through

a propagating crack, when a concentrated stress exceeds the theoretical strength of the material.


91

Figure 81 Internal force lines are denser near the hole [29]

10.2.1 Concentration factor for cracks

The maximum stress occurs in the area of lowest radius of curvature. For example, in an

elliptical crack of length 2a and width 2b, under an applied external stress , the stress at the

ends is expressed by:

a

b

a2121max where is the radius of curvature of the

crack tip. The ratio of the highest stress (max) in relation to a reference stress () is called the

stress concentration factor. The maximum stress approaches infinity as the radius of curvature

approaches 0.

10.2.2 Concentration factor calculation

Methods for measuring stress concentration factors are photoelastic stress analysis, brittle

coatings or strain gauges. In the design phase, there are many ways to assess stress concentration

factors. Today, most commonly used methods are finite element methods. All methods have their

advantages and disadvantages. Engineering judgment must be used in data selection.

10.2.3 Stress concentration prevention

One method to reduce stress concentrations i.e. crack, is to drill a large hole at the end of

the crack. That drilled hole which has large diameter, causes smaller stress concentration than the


92

sharp end of a crack. Another method to reduce the stress concentration is creating the fillet at

the sharp edges which gives smooth flow of stress streamlines.

10.3 Examples of stress concentration

Example of metal failure as a result of stress concentrations are corners of the windows

of the De Havilland Comet 1 aircraft as it is shown in Figure 82.

Figure 82 The sharp corners at the windows of Comet 1 G-ALYP [29]

10.4 Fracture criterion

The law of energy conservation dictates that the energy (F) spent on the deformation of

the structure is equal to the accumulated internal energy of deformation (U), i.e. 0U-F .

Two different cases can be considered: first case is when shift does not change the

appearance with length da when fracture occurs and second case when the load is constant with

length da when fracture occurs. In the first case dF=0, since there is no shift, while in the second

case energy is equal to shift difference before and after the occurrence of the crack. In both

cases, the following equation is obtained: da

dW

da

dU .

Left side is called strain energy release rate, while the right side is called fracture energy

or fracture resistance. As the energy of deformation is under the influence of the crack, it can be

written as: U=Uwithout crack+Udue to crack. For very large panel of unit thickness (with central crack

of length 2a) final expression is obtained: E

aπσLW

2E

σU

22

.


93

If the expression is derived by a and by taking into account that it is a crack with two

crack tips, the following expression is obtained: E

aπσ

da

dU 2

. Thus the equation for fracture

criterion becomes: E

aπσ

da

dW 2

.

In that case, dW/da ratio is called fracture resistance and is often marked with R, while

dU/da, i.e. change of the energy of deformation is indicated as G, and hence it is: G=R.

Last equation shows that fracture occurs when the product πσ2a reaches the value ER,

where πσ2a represents the squared coefficient of stress intensity, K. Therefore, it can be

concluded that the fracture will occur when K=(ER)1/2, wherein (ER)1/2 represents the toughness

Kc, and the fracture resistance is then equal to R=Kc2/E. We conclude that the fracture criterion is

derived from the law of energy conservation and it is also identical to fracture criterion derived

from the stress at cracks tips.

Figure 83 Increasing R curves: (a) stable fracture from Gi to G3; (b) Stable fracture for cracks with different

lengths [24]

Instability of the fracture occurs when the G line is tangent to the R curve (Figure 83), i.e.

when they both have the same slope. The condition of instability is as follows: G=R,

respectivelyda

dR

da

dG .

Regardless of considering the material in the elastic or plastic area, the law of energy

conservation must be valid. Previously derived fracture criterion is: da

dW

da

dU or G=R.


94

For elastic behavior the following expression is obtained: da

dW

E

aπσβ 22

or R=G, where

R represents fracture energy and G represents the change of deformation energy.

The engineering analysis it is useful to express the previous equation in the term of stress:

πaβ

K

aπβ

ERσ c

2fr .

In the case of plastic deformation geometric factor β is changed, but it is still

dimensionless and is denoted by H. In this case, the expression for the change of deformation

energy is as follows: RHσσε .

Furthermore, in the case of non-linear materials, G is denoted with J, while R is indicated

as JR in the case of non-linear material. With this equations obtained for the elastic behavior of

the material are converted into the equations for the plastic behavior of materials: J=JR and

JHσσε .

It must be noted that these are the same equation derived from the law of energy

conservation, but with few marks changed. Although Hooke's law is used for linear material

behavior (linear relationship between stress and deformation), same equations can be applied to

the plastic area if there is a non-linear function between ε and σ, such that it is corresponding

with behavior of the material.

The most suitable (empirical) function of the non-linear relationship between stress and

deformation is the exponential function, known as the Ramberg-Osgood equation: F

σ

E

σε

2

or plel εεε .

Since the Ramberg-Osgood equation can be used to evaluate the fracture criterion and

how the first addend of the equation represents elastic (linear) part of stress-deformation curve, it

is necessary to consider the effects of the plastic part (second addend) of Ramberg-Osgood

equation: F

σε

n

pl .

It is obvious that for the case n=1 (F=E) the above expression reduces to Hooke's law,

while with the inclusion of the equations J=JR and JHσσε into the equation F

σε

n

pl , we

obtain: R

1n

JF

aHσ

.

For n=1 (F=E) above expression implies that H=πβ2. Fracture criterion, after having

plastic parts defined, can now be written as follows: da

dW

F

aHσ

E

aσπβ 1n22

or in abbreviated

notation G+R=JR.

Analysis shows that the use of elastic-plastic fracture mechanics in relation to the linear-

elastic fracture mechanics is nothing more complicated and/or more complex. The biggest

difference is that in the elastic-plastic fracture mechanics iterative method for solving equation


95

da

dW

F

aHσ

E

aσπβ 1n22

should be applied and that H does not depend only on the geometry

but on n also, and is therefore more difficult to H than β.

10.5 Basic concept of crack propagation due to material fatigue

Fatigue is technically the most important mechanism of crack growth. Although in many

structuras fatigue is very difficult to prevent or it can not be prevented, it certainly should be

controlled. Damage due to cyclic loading has four levels: crack initiation, small crack

propagation, large crack propagation and final fracture, as illustrated in Figure 84.

Crack length of 1mm is usually taken as the boundary between the stages of propagation

of small and large cracks. This value is also accepted as the threshold value of crack length

which can be determined by non-destructive testing methods to structure in service. It was

generally found that the component has work life of about 80% of its life span in a state of small

crack propagation. Once the presence of a crack is detected, it is important to know how its

going to progress to efficiently repair or replace the defected part.

Figure 84 Level of damage for different crack lenghts as a function of number of cycles (loads) [24]

In order to fully describe the cyclic stresses due to loads with constant or variable

amplitude it is sufficient to know one combination of two different parameters, as shown in

Figure 79: Δσ and R, σmin and R, σmax and R, σa and R, and σm and R. σmin denotes the least stress,

while σmax denotes the maximum stress in the cycle. Δσ is the stress range, defined as Δσ = σmax-

σmin. R denotes the ratio of stress or stress ratio, defined as R = σmin/σmax, while σa and σm are

stress amplitude and mean stress. The life span of a crack or crack growth life is expressed as the

number of cycles required for fatigue crack to grow to a certain size, wherein the number of

cycles is denoted as N.

Figure 85 (a) shows the mechanism of crack progression as geometric result of blunting

the crack tip with each load cycle, while re-sharpening of the crack tip during unloading causes

an increase of the crack during next load cycle. It can be concluded that the crack growth Δa per


96

cycle will be higher if the maximum stress per cycle is higher (greater opening of the crack tip)

and if the minimum stress per cycle is less (greater sharpening of the crack tip). Localized

stresses at the tip of the crack can be described with the stress intensity factor K, where

πaβσK , wherein σ is applied nominal stress.

Figure 85 Parametres of crack propagation due to material fatigue: (a) un-sharpening and re-sharpening of

the crack tip; (b) Δσ, ΔK [24]

As the applied stress varies within the cycle i.e. Δσ between σmax and σmin, the local

stresses will also vary in accordance with the following equations: πaβσK minmin ,

πaβσK maxmax , and πaσβK . From the stated equations it follows that the crack

growth per cycle will be higher if Kmax is higher and/or if ΔK is higher. According to the

equations πaβσK minmin , πaβσK maxmax , and πaσβK , it follows that for any size


97

of crack ratio of stress values is max

max

max

min

max

min

K

ΔK-K

K

K

σ

σR . It may be further noted that the

crack grows when the ΔK and/or R is larger, the function of the crack growth speed can generally

be written in mathematical form: R)f(Δ(ΔdN

da .

10.6 Consequences of fracture in aircraft structures

In ribed panels of aircraft structures damage due to cracking can occur. These defects

occur due to cyclic workloads. In aircraft structures basic cyclic loading is a cabin pressurisation.

In the experiments and simulations it is demonstrated that in addition to the membrane stresses in

ribed panels it may occur additional component stress due to bending.

Component of bending on the side of the panel which has a positive value, additionally

increases the value of the stress intensity factor, which increases due to relative crack growth

progression and thereby reduces the lifetime of the panel, and hence, aircraft structure.

Catastrophic damage of aircraft structure caused by defects due to fatigue had influenced

the development of practical knowledge about the prevention of similar accidents later. In the

20th century there were several catastrophic failures of aircraft structures as a result of fatigue.

Aircraft De Havilland Comet, which entered the service in the 1952, was the first passenger jet

airplane (Figure 86). With pressurized cabin and silence during flight, Comet has shortened the

journey from New York to London for 4 hours.


In January 1954, Comet G-ALYP, and in April of the same year, Comet G-ALYY, broke

up in the air and took many lives. Tests and studies of the remains of the structure of the first

crashed aircraft (Figure 87) showed that the cracks developed due to metal fatigue near the ADF

window located at the front of the cabin ceiling (Figure 88 and Figure 89). These cracks

eventually spreaded to the window and developed into a crevasse of the great length and caused

the collapse of the local construction and breakdown of the entire structure.


98

Figure 87 Reconstruction of the fuselage and empennage from the Comet G-ALYP reckage [24]



99

Figure 89 a) Crack in the De Havilland DH 106 Comet G-ALPY; b) detailed illustration of the initial crack

on the Comet G-ALYP [24]


100

In order to determine the cause of failure of Comet G-ALYP and G-ALYY extensive

tests were performed, including complete repeated pressurisation in the aircraft Comet G-ALYU

which was grounded. Tests were conducted under water, to minimize the damage caused by the

breakup of the structure (Figure 90).

Before testing, aircraft had 1231 cycles of pressurisation. After 1825 tested cycles of

pressurisation, there was a collapse of the structure. Tests have shown that the cracks initiated

mainly at the corners of windows and emergency exits and then spreaded almost axially along

the fuselage (Figure 91 and Figure 92).

Figure 90 Comet G-ALYU in testing pool [24]


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Figure 91 Fuselage break-up of tested model of Comet G-ALYU [24]


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Figure 92 Probable cause of Comet G-ALYU fuselage break-up; stress dissemination calculated during

repairs [24]

Much has been learned from these investigations after accidents and Comet was

redesigned into more robust structure (Figure 93). However, in the four years that it took Comet

to obtain permission for the flight, Boeing released its 707 series of aircraft that has taken

primacy in passenger jets.


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Figure 93 Advances in construction design of Comet’s windows [24]

The accident that involved Aloha Airlines Boeing 737 also attracted much attention. At

an altitude of 7300 meters aircraft lost most of its fuselage (Figure 94). Although aircraft

suffered large damage, it continued the flight to the airport.

Figure 94 Damage to Aloha Airlines Boeing 737 [24]


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The investigation showed that the loss of fuselage was caused by the merger of the

fuselage cracks sequently in a very short time. This type of fatigue is called multiple cracking

damage (Figure 95), where cracks initiated on the sharp edges of rivets’ holes (Figure 96).

Figure 95 Multiple crack damage [24]

Figure 96 Illustration of rivet and its fatigue crack on Aloha Airlines Boeing 737 [24]


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The aircraft was 19 years old and had 35,496 flight hours and 89,680 landings and often

flew on short routes. Environment in which it flew was warm, the air was often a marine air,

moist and with salt particles which led to corrosion of the rivet joints accompanied by separation

of doubler.

In August 1985, Japan Airlines Boeing 747SR flew from Tokyo to Osaka. While

climbing to an altitude of 24,000 ft the rear pressure bulkhead yielded due to which there has

been an explosive decompression that caused the loss of hydraulics and the aircraft became

unmanageable. After 30 minutes, the plane crashed into the mountains. It was then the largest

airline disaster, 524 passengers were killed, and only four of them survived.

The aircraft was previously damaged in June of 1978 when the tail struck the runway

partly causing the damage to the lower part of the rear pressure bulkhead. The bulkhead was

made of a thin sheet of aluminum alloy and was of hemispherical shape. Panels were linked with

rivets, and among them was a doubler for increased bond strength. In repair of damage from

1978, new lower part of the bulkhead was nailed to the upper half.

However, the two halves were not well connected. On the upper half of the panel was

doubler and stiffener on the inner side of the bulkhead (Figure 97). On the lower half of the

compound there was a doubler, but it was separated from the upper doubler, so there was a rift

between the two doublers so essentially only panel was carrying all loads. Furthermore, the

center of the cross section which transferred the load was shifted toward the inner side of the

bulkhead. Therefore, the load of the panel that was overarching gap, was not only consisted of

tensile stress, but also from bending.

Every time the cabin was pressurized, there would be an increase in stress in the panel

that overarched the gap. As a result of the increase of the stress, there were fatigue cracks at rivet

holes in the lower part of the bulkhead below the gap. These fatigue cracks were eventually

linked to one large crack that eventually led to explosive decompression.


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Figure 97 Back pressure obstacle on JAL Boeing 747-SR [24]

It can be concluded that many accidents happened because of flaws in the design of

structures and selecting materials with relatively low resistance to fatigue. On the other hand,

some of the accidents could have been avoided if the flight simulation tests were conducted with

realistic workloads. Undoubtedly, much can be learned from the investigation of accidents, but it

is certainly hard way of learning.


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List of questions

1. Define fracture mechanics. What are the most important tasks in fracture mechanics?

2. Explain stress concentration factor.

3. Explain method to prevent stress concentrations.

4. Explain fracture criterion.

5. Explain basic concept of crack propagation due to material fatigue.

6. Explain consequences of fracture in aircraft structures.


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11 Methods to detect and prevent material fatigue

Different methods are developed to detect and prevent fatigue of material in aircraft

structures. The most important and most commonly used methods are discussed in this paper.

Those methods are infinite lifetime concept methods, finite lifetime concept methods, non-

destructive testing methods, cold expansion methods, methods using built-in testing equipment

and structural health monitoring methods.

Dependable design against fatigue-failure requires thorough education and supervised

experience in structural engineering, mechanical engineering, or materials science. There are

four principal approaches to life assurance for mechanical parts that display increasing degrees

of sophistication. Design to keep stress below threshold of fatigue limit (infinite lifetime

concept):

1. Infitite lifetime (fail-safe, graceful degradation and fault-tolerant) design: instruct the

user to replace parts when they fail. Design in such a way that there is no single point

of failure, and so that when any one part completely fails, it does not lead to

catastrophic failure of the entire system.

2. Safe-life design: design (conservatively) for a fixed life after which the user is

instructed to replace the part with a new one (a so-called lifed part, finite lifetime

concept, or "safe-life" design practice); planned obsolescence and disposable product

are variants that design for a fixed life after which the user is instructed to replace the

entire device;

3. Damage tolerant design: instruct the user to inspect the part periodically for cracks

and to replace the part once a crack exceeds a critical length. This approach usually

uses the technologies of nondestructive testing and requires an accurate prediction of

the rate of crack-growth between inspections. The designer sets some aircraft

maintenance checks schedule frequent enough that parts are replaced while the crack

is still in the "slow growth" phase. This is often referred to as damage tolerant design

or "retirement-for-cause".

Fatigue cracks that have begun to propagate can sometimes be stopped by drilling holes,

called drill stops, in the path of the fatigue crack. This is not recommended as a general practice

because the hole represents a stress concentration factor which depends on the size of the hole

and geometry, though the hole is typically less of a stress concentration than the removed tip of

the crack. The possibility remains of a new crack starting in the side of the hole. It is always far

better to replace the cracked part entirely.

Changes in the materials used in parts can also improve fatigue life. For example, parts

can be made from better fatigue rated metals. Complete replacement and redesign of parts can

also reduce if not eliminate fatigue problems. Thus helicopter rotor blades and propellers in

metal are being replaced by composite equivalents. They are not only lighter, but also much

more resistant to fatigue. They are more expensive, but the extra cost is amply repaid by their

greater integrity, since loss of a rotor blade usually leads to total loss of the aircraft. A similar

argument has been made for replacement of metal fuselages, wings and tails of aircraft.

The durability and life of dynamically loaded, welded steel structures are determined

often by the welds, particular by the weld transitions. By selective treatment of weld transitions

with the High Frequency Mechanical Impact (HFMI) treatment method, the durability of many


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designs can be increased significantly. This method is universally applicable, requires only

technical equipment and offers high reproducibility and a high grade of quality control.

11.1 Infinite lifetime concept methods

Infinite lifetime concept is design to keep stress below threshold of fatigue limit. Fatigue

limit, endurance limit, and fatigue strength are all expressions used to describe a property of

materials. It is the amplitude or range of cyclic stress that can be applied to the material without

causing fatigue failure.

Fail-safe design is design in such a way that there is no single point of failure, and so that

when any one part completely fails, it does not lead to catastrophic failure of the entire system.

A fail-safe or fail-secure devices are ones that, in the event of failure, responds in a way

that cause no harm, or at least a minimum of harm, to other devices or danger to personnel. For

example, aircraft landing on an aircraft carrier increases the throttle to full power at touchdown,

and if the arresting wires fail to capture the plane, it will be able to take off again.

Graceful degradation is the property that enables a system to continue operating properly

in the event of the failure of (or one or more defects within) some of its components. If its

operating quality decreases at all, the decrease is proportional to the severity of the failure.

Fault tolerant design enables a system to continue its intended operation, possibly at a

reduced level, rather than failing completely, when some part of the system fails. The term is

most commonly used to describe computer systems designed to continue more or less fully

operational with, perhaps, a reduction in throughput or an increase in response time in the event

of some partial failure.

11.2 Finite lifetime concept methods

Safe-life design (finite lifetime concept or safe-life design practice) is design for a fixed

life after which the user is instructed to replace the part with a new one. Planned obsolescence or

built-in obsolescence in industrial design is a policy of planning or designing a product with a

limited useful life, so it will become obsolete, that is, unfashionable or no longer functional after

a certain period of time.

A disposable product is a product designed for a single use after which it is recycled or is

disposed as solid waste. The term often implies cheapness and short-term convenience rather

than medium to long-term durability.

11.3 Damage tolerant design methods or methods of non-destructive testing

Damage tolerant design instructs the user to inspect the part periodically for cracks and to

replace the part once a crack exceeds a critical length. This approach usually uses the

technologies of non-destructive testing and requires an accurate prediction of the rate of crack-

growth between inspections. The designer sets some aircraft maintenance checks schedule

frequent enough that parts are replaced while the crack is still in the “slow growth” phase.

Methods for non-destructive testing (Non-Destructive Testing – NDT) is a set of methods

for finding defects in material and in such a way that materials or devices after testing remain

intact, and if there are not defects detected, they can be placed in a normal exploitation.


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NDT methods are focused on the detection of micro cracks on the surface or below the

surface of the element, and to detect defects in inaccessible places when regular visual inspection

without tools is conducted. NDT methods are applied to the aircraft itself and individual

elements of the aircraft.

The most commonly used methods of NDT are: visual, magnetic, penetrant, radiographic,

ultrasonic and eddy current methods.

11.3.1 Visual inspection

Visual testing is the oldest and most common form of inspection. It consists of an

overview using human eye, a magnifying glass, a light source or special optical devices. The

reliability of this method depends on the ability and experience of personnel who must know

how to look for structural defects and how to identify an area where such defects are found.

A visual examination with the help of optical devices can be considered an extension of

the human eye. From inaccessible areas, an image on the screen is formed that can be

electronically captured, increased, analysed and recorded for future directions.

Basic equipment for the visual inspection is endoscopes. Endoscopes allow the technician

to view the interior of the equipment, components or structures that have closed or hidden areas

not accessible to ordinary visual inspection.

Endoscopes (flexible borescopes and inflexible fiberscopes) are optical devices with

optical probe which can penetrate through small openings to the work area to be tested. These

devices have the possibility of strong internal illumination of dark areas to make good visual

inspection, photographing or video recording.

Their optical system transmits a clear picture of high resolution to the eyepiece or video

monitor which is distant, but connected to the device. Eyepiece is used for direct visual

inspection or it can be connected to a photographic or video camera.

Figure 98 Video borescope [36]

Videoscope (Figure 98) is an electronic version of the optical fiberscope. Instead of the

lens at the end of a flexible tube and coil which transmits the image, videoscope has a small

video camera and a lens at the end. The camera, which is based on a compact chip technology,

sends back a video image to the unit where it is shown on a video monitor.


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The basic advantages of this method using endoscopes are: it is cheap, it is portable, it

gives instant results, and it takes a minimum of special skills and a minimum of preparation.

Disadvantages are that it is only suitable for areas that can be visually detected and it

reveals only greater defects.

11.3.2 Penetrant testing

Penetrant method (penetrant testing) is used to reveal discontinuities opened towards the

surface of the parts made of material that is not porous. The method depends on the ability of

liquid to penetrate the discontinuity of the material on which it is applied (Figure 99).

Ordinary penetration method is used for the detection of small cracks or gaps that go up

to the surface and that may not be visible under normal visual inspection. Penetration methods

can be used in most parts of the structures that are available for its application.

Figure 99 1) section of material with a surface-breaking crack that is not visible to the naked eye, 2)

penetrant is applied to the surface, 3) excess penetrant is removed, 4) developer is applied, which renders the

crack visible [29]

The process of penetrant testing involves a few basic steps. Part to be tested is first

thoroughly cleaned and then a liquid penetrant is applied to the surface. It is left for some time to

sit on the surface while the penetrant penetrates the surface discontinuity, if there is any. After

that, penetrant remained on the surface is removed with water, cloth or thinner. Such removal

cleans the surface of objects, but allows the penetrant to remain in the discontinuity.

Then a developer is applied that works as a blotter and draws penetrant on the surface of

objects and creates indication. Subject is inspected in order to detect indications that are visible

due to contrast dye between a drawn penetrant and a background surface. After testing the

surface is cleaned of residual penetrant and the protective layer is returned.

The advantages of this method are: low cost, portable equipment, it has a high degree of

sensitivity; it gives instant results and requires minimum skills for performance and

interpretation.

The disadvantages are that it requires a high degree of purity, it can detect only those

defects that have a connection with the surface and there are no permanent written test results.

11.3.3 Ultrasonic testing

Ultrasonic testing method (ultrasonic inspection) is suitable for the examination of most

metals, plastics and ceramics and defects on the surface or below the surface. Ultrasound


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examination requires that at least one part of the surface near surface to be tested is available.

Examination of aircraft structures can be achieved by inducing ultrasonic waves on the object

with the contact probe and receiving of reflected waves from that point. Reflection of ultrasonic

waves is projected electronically into the tube of oscilloscope and is used to indicate defects

(Figure 100).

Figure 100 An example of Ultrasonic Testing (UT) on blade roots of a V2500 IAE aircraft engine (step 1: the

UT probe is placed on the root of the blades to be inspected with the help of a special borescope tool (video

probe), step 2: instrument settings are input, step 3: the probe is scanned over the blade root, in this case, an

indication (peak in the data) through the red line (or gate) indicates a good blade; an indication to the left of

that range indicates a crack) [2]

Ultrasonic testing method is a method of non-destructive testing that uses high-frequency

sound energy (above 20 kHz) for testing the structural integrity of the material.

Ultrasonic waves can spread only in the medium. Precisely this fact is used to detect

defects in the tested object. Ultrasonic waves at the medium border, as well as all other types of

waves follow the laws of wave motion. For this reason, at the medium border, whether it is the

wall of the tested object or irregularity, the reflection of ultrasonic waves and/or fracture,

diffraction or other interaction between means and transmitted ultrasound energy will occur.

Proper interpretation of ultrasound energy obtained from material testing can assess the

condition of material and parameters of detected irregularities. Ultrasound can be generated in

several ways. The most commonly used is piezoelectric effect for production of ultrasonic

energy.

If the ultrasonic wave passes through the material, it encounters a defect, it can

(depending on the size, location and type of defect) be reflected from surface of defect, or, due to

the different angles of reflection, be reduced when returning back to the surface from which it

was transmitted. The measurement of these phenomena is the basis of ultrasonic defectoscopy.

Transmitting and receiving ultrasound from and into materials is conducted using a

transmitter and receiver or ultrasonic probes, which generate ultrasound waves and convert

received ultrasound into an electronic impulse.

Ultrasonic device must allow activation of probes with required electronic impulses and

receive electronic impulses from the probe.

Advantages of this method are the following: area thickness of the tested material is very

large, tested materials can have very large thickness of the material; devices and accessories are

very simple, easily portable and do not take up much space, the risk of device failure is reduced

by using modern electronics, which also reduced the weight; handling of devices and equipment

poses no danger to the operator; method is very sensitive and detecting defects is fairly simple;

and it is possible to test the object from only one side of the tested object.


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Disadvantages are: generally cannot leave a permanent document; assessment of defect

detected by ultrasound is subject to operator subjectivity and his level of knowledge how to test

and properties of the test material; because of the way of testing and assessment results at the

point of control, personnel handling devices must, besides greater knowledge, have long practice

with the possibility of comparative determination of the character of the detected defects, which

is why the introduction of this method takes much longer than any other; ability to detect

subsurface defects and defects in objects that have irregular and complicated shapes is relatively

weak.

11.3.4 Magnetic particle inspection

Magnetic testing method (magnetic particle inspection) is a method used to detect surface

and subsurface discontinuities in ferromagnetic materials. Testing is done by inducing a

magnetic field in observed part and applying the particles of a dry powder or a liquid suspension

of iron oxide. If there is a discontinuity in the material (in the form of cracks, nicks, or

inclusions), it leads to increased resistance in magnetic field at the site of the cracks.

If, for example, particles of iron oxide are applied on the surface of material to be tested

and if there is a crack perpendicular to the direction of the passage of forces, particles will gather

around the cracks. Particles can be seen by colour contrast or fluorescent light under the UV

lamp. To obtain the best results of indications, magnetic field forces should be at right angles to

the crack, i.e. maximum length of the crack.

Cracks are detected by looking at shapes that are formed on the surface of the particles.

Fine discontinuity will diffuse visible indication when formed with a fluorescent magnetic

particle. The dry particles are best used on rough surfaces and the supporting equipment.

The simplest devices for magnetic tests are ferroflux devices (Figure 101). These are the

generators of DC or AC currents of high strength and low voltage.

The basic advantages of magnetic method are: simplicity of principles and low cost of

devices, unless they are automated; easy and simple handling devices; good sensitivity to detect

surface and subsurface defects, particularly cracks; interpretation of the results does not require

special expertise; the ability to apply on products of complex shapes and with large differences

in wall thicknesses (preferred over methods of radiography and ultrasound).

Disadvantages of magnetic method are: it establishes the existence of defects, but not

their geometry; depth defects are difficult to detect, reliable detection depth is only a few

millimetres and most reliably reveals the defects that have contact with the surface; generally

quite slow; interpretation of results depends a lot on the state of the surface of the tested object.

Figure 101 A horizontal MPI machine with a 36 in (910 mm) coil [29]


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11.3.5 Radiographic testing

The radiographic testing of material, material irregularities are obtained in the way that

the object of testing is aired with appropriate ionizing radiation. Radiographic testing will

therefore show internal and external structural details of all portions of the material. It is a

method that is used to check aircraft structure that is unavailable or unsuitable for the use of any

other method. As a source of radiation X or gamma rays are to be used to detect objects' interior

which is shown in the film.

Radiographic methods differ with respect to: source of ionizing radiation, means for

registering the radiation, and way of conducting and/or registering results (Figure 102).

Advantages of film radiography: good visibility of surface, subsurface and internal

defects; obtaining of a permanent document (film), which may subsequently determine the level

of quality of inspected items; recording and interpretation are widely separated and the review

conduct does not require a large number of highly specialized personnel as recording and

interpretation can be separated in time; minimal preparation work.

Disadvantages of film radiography: devices to conduct control measures are relatively

expensive, and accessories for recording and reviewing films require constant new investments;

unsuitable for control of products whose walls show large differences in thickness, certain types

of small thickness defects such as cracks transversely to the direction of radiation are very

difficult to detect by this method; it may be necessary to drain the fuel from the tank; because of

X-ray other staff should be out of the area designated as harmful.

Figure 102 Radiographic testing [37]

11.3.6 Eddy current testing

Eddy current testing is used to detect fractures on the surface or near the surface in most

metals. It can be used for aircraft parts or assemblies where the damaged area is accessible to

contact probe. Examination is done by inducing eddy currents in the part to be tested and

monitoring electronically variations in the induced field.

Eddy current testing method (Figure 103) is a method of non-destructive testing that can

detect discontinuities in parts that are electrically conductive. Eddy currents are circular electric

currents that are induced in conductive material by an alternating magnetic field, and exhibit

properties such as the compressible medium (e.g. air) when encountering an obstacle, such as


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crack, they are traveling around it, being compressed and getting weak in a similar manner as the

air does. Character of observed changes in the field is interpreted as a defect in the material.

Figure 103 Eddy current testing [29]

Advantages of this method are the following: tests can be quickly conducted; it is

possible to detect defects under a layer of coating (paint, etc.) from 1.5 to 2 mm and thus save

money and time for removal of the same; it is possible to measure the thickness of coatings

(paint, etc.) and determine the metallurgical differences; in addition to cracks, in the interior of

the tube sediment build-up can be detected; there is a good possibility to determine the size

(depth) of damage; it is not necessary to use a contact means.

Disadvantages of this method are: small possibility of finding shallow defects in

materials with little wall thickness, and relatively high cost of equipment and probes.

11.3.7 Low coherence interferometry

New materials such as polymer composites and ceramics have found a wide variety of

applications in modern industries including energy, aviation and infrastructure. This is due to

their superior properties as compared to traditional materials. For process enhancement, quality

control and health monitoring where these types of materials are used, sophisticated techniques

are needed to non-destructively inspect the complex geometries inside the structures produced.

Low-coherence interferometry has been developed as a powerful tool for the cross-sectional

imaging of microstructures.

Interferometry uses the principle of superposition to combine waves in a way that will

cause the result of their combination to have some meaningful property that is diagnostic of the

original state of the waves (Figures 104 and 105). Most interferometers use light or some other

form of electromagnetic wave.


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Figure 104 The schematic setup of the low-coherence interferometer [19]

Figure 105 The cross-sectional image for composite material sample [20]

11.4 Cold expansion methods

Over the past 40 years, cold expansion processes have impacted fatigue crack mitigation,

structural integrity and airworthiness. In conjunction with improving materials, reducing effect

of stress concentration in hole and mitigation of fatigue cracks at holes, was focus in 1960s.

Term “cold working” or as it is later called “cold expansion” usually applies to processes

that improve strength and performance via cold work hardening.The probability of fatigue

increases where stresses are concentrated, like sharp bends and fastener holes.

Holes typically magnify the applied stresses by three times. Holes make the metal

especially susceptible to fatigue because the hole alters the stress flow and concentrates stress

due to cyclic loads. Many systems have been conceived to offset or mitigate this stress

concentration problem in fastener holes.

Methods to reduce fatigue problems are following: thicken the structure locally to reduce

stress levels with a structural weight penalty; install interference fit fasteners; reduce amplitude

of the applied strain; fastener preload or clamp up or bridge hole throught interface friction; and

induce a compressive hoop pre-stress associated with a cold working process. [22]

The sleeve cold working or pre-stressing system is the one that offers advantages in terms

of simplicity, flexibility, performance, cost, and minimized skill requirements. From the early

advent of metal structures, surface treatment methods had been devised to improve ageing and

durability of holes and surfaces.

From 1940s to 1950s used methods are ballizing, mandrelizing (mandrel only) and

broaching, in 1960s there are improved materials, surface treatments (shot peening), 1960s

interference fit fasteners, in 1970s high interference cold working, split sleeve cold expansion,

stress coining (ring and pad), in 1980s derivative cold expansion processes, and today ForceMate

methods of derivative cold expansion processes.


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Figure 106 Surface treatment cold working methods [22]

Figure 107 High interference expansion methods [22]

Basic split sleeve cold expansion process is: the split sleeve is slipped onto the mandrel,

attached to the hydraulic puller unit; the mandrel and sleeve are inserted into the hole with the

nose-cap held firmly against the work-piece; and when the puller is activated, the mandrel is

drawn through the sleeve radially expanding the hole.

Figure 108 Split Sleeve Cold Expansion [22]

Advantages (mandrel only) are: it is a one sided process; higher applied expansion; radial

expansion; uses lower pull force; there is no damage to hole; and minimal surface upset. The

split sleeve, though simple in design, makes the process possible and effective.


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Figure 109 Induced residual stress distribution [22]

Cold expansion induces a zone of residual compressive stress around and through a hole,

typically extends radially at least one radius around hole. Hole is effectively “shielded”, reducing

effective stress intensity factor and therefore the propagation of fatigue cracks.

It improves fatigue life, durability and damage tolerance of structure.

Characteristics are: hole sizes from 0.070” to over 5.5”; typically applied expansion levels

ranging from: 3% to 5% aluminium and 5% to 6% Ti & HS steels; effective in most aerospace

materials; and it is one sided process.

Figure 110 Effect of applied expansion on fatigue life improvement [22]


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Figure 111 Fatigue life improvement, weight savings and design optimization [22]

Figure 112 Fatigue life improvement in aluminium [22]

Figure 113 Fatigue life improvement in aluminium [22]


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Figure 114 Fatigue life extension in titanium [22]

Figure 115 Effect of fastener interference [22]

Cold expansion can completely stop the growth of small cracks without increasing the

thickness (or weight) of the structure and it can also increase the damage tolerance life. Reducing

crack growth and improving damage tolerance of structure can be achieved by permanent

compressive stresses.


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Figure 116 Effect of cold expansion on crack tip stress intensity [22]

Figure 117 Fatigue crack abatement [22]

Table 6 Crack growth investigations

Investigator Spectrum Material Initial flaw Life improvement

Northrop TAC T-38 7075-T6

7075-T73

0.002 -> 0.01 5:1

Boeing KC-135 7178-T6 0.07 -> 0.08 4:1 -> 15:1

Toor

(Lockheed GA)

Constant

Amplitude

7075-T6 0.05 3.5:1

Tiffany, Stewart &

Moore (USAF)

C-5A 7075-T6 0.05 -> 0.09 4:1

No cold-work failures

Kobler, Huth, &

Schutz (LBF

Germany)

Constant

Amplitude

2024-T3 0.05 -> 0.23 3:1 -> 6:1

Hoosen &

Eidenhoff

(Grumman)

F-14 Ti-6AL-6V-2SN 0.035 -> 0.040 10:1

Ozelton & Coyle

(Northrop)

Constant

Amplitude

7050-T7451

Ti-6AL-6V-2SN

0.020 -> 0.120 8.6:1 -> 1.6:1

>100:1 -> 3:1

Petrak & Stewart Constant

Amplitude

7075-T6 0.030 -> 0.100 >100:1 -> 2:1

FTI (Fatigue 7075-T6 0.05 -> 0.1 >100:1 -> 2:1


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Technology) 2024-T3 6:1 -> 4:1

Figure 118 Extended inspection interval with cold expansion (with initial 0.005” flaw) [22]

Figure 119 Damage tolerance benefits (fighter aircraft study) [22]

Cold expansion can: increase the fatigue life of a hole or joint; provide a solution

independent of fastener fit or hole quality; reduce or arrest crack growth; enhance damage

tolerance; extend inspection thresholds; provide terminating repair action; mitigate design or

repair consequences; facilitate structural weight reduction; be effective in tensile and


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compressive load environments; do all mentioned above without changing weight, stiffness or

dynamic response of the structure; and present the foundation of so many derivative products or

processes.

11.5 BITE concept

All modern transport aircraft today have some form of permanent monitoring of technical

condition which is integrated on the aircraft (On-board Maintenance Systems). For mechanical

components it means that there are sensors placed that continuously measure a parameter of the

system by which it can be evaluated the technical condition. Examples are sensors that measure

pressure, temperature, vibration, shifts, etc. These systems allow the detection of defects during

operation and are called BITE (Built In Test Equipment).

The BITE is characterized primarily as a passive error management and diagnosis built

into airborne systems to support the maintenance process. Built-in test equipment refers to multi-

meters, oscilloscopes, discharge probes, and frequency generators that are provided as part of the

system to enable testing and perform diagnostics.

The term BITE often includes: detection of defects, placement of the defect (how the

system actively responds to the defect) and annunciation or logging of the defect to warn of

possible effects and/or aid in troubleshooting the affected equipment.

Its functionalities are: analysis of failure monitoring results, reporting and memorization

of failures and management of tests (Figure 120).

Figure 120 BITE concept on the Airbus 320 [2]


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11.6 SHM concept

The process of implementing a damage detection and characterization strategy for

engineering structures is referred to as Structural Health Monitoring (SHM). Here damage is

defined as changes to the material and/or geometric properties of a structural system, including

changes to the boundary conditions and system connectivity, which adversely affect the system’s

performance.

The SHM process involves the observation of a system over time using periodically

sampled dynamic response measurements from an array of sensors, the extraction of damage-

sensitive features from these measurements, and the statistical analysis of these features to

determine the current state of system health.

For long term SHM, the output of this process is periodically updated information

regarding the ability of the structure to perform its intended function in light of the inevitable

aging and degradation resulting from operational environments. After extreme events, such as

earthquakes or blast loading, SHM is used for rapid condition screening and aims to provide, in

near real time, reliable information regarding the integrity of the structure.

Monitoring of the structure status (Structural Health Monitoring) is a concept of system

for continuous monitoring the technical condition of the aircraft structure. SHM will be a key

technology for maintaining the integrity of the aircraft structure in the future.

The basic principle lies in creating a reliable non-destructive technology that would be an

integral part of the aircraft structure. Sensors mounted on the plane would detect the defect in

time and further take action. SHM sensors will be able to detect: fatigue, hidden cracks,

weakened joints, erosion, corrosion, cracks in the structure, etc. The integrated network, except

for structural elements, will oversee even other systems on the aircraft (hydraulic, electronic, and

electric).

Positive SHM features in comparison to conventional NDT: sensors are placed on the

plane to oversee the entire structure; physical access to equipment or human is not needed to

access areas to be tested; inspection of hazardous areas shall be carried out in a safe manner;

inspection is fully automated; at the same time it examines several areas; and the results of the

testing have no impact of the human factor (Figure 121).


125

Figure 121 Structural health monitoring (SHM) systems can be arrayed in similar fashion to the human

nervous system, with sensors concentrated in key areas where loads are highest [38]

11.6.1 Wireless Sensors for Structural Health Monitoring

MicroStrain is the company which produces wireless sensors. MicroStrain wireless

sensor nodes are low-power sensors enabling monitoring with one greatest advantage and that is

no need for installing or maintaining wires. Wireless nodes support a wide range of sensor types

including on-board accelerometers, external devices for supplying data on strain, temperature,

load, pressure, humidity, vibrations and other.

MicroStrain has also developed wireless health monitoring system for rotorcraft which

provide data about technical condition of all aircraft components. Main products are: G-Link-

RGD, SG-Link-RGD, and WSDA-RGD.

Wireless sensors enable the autonomous reporting and tracking of pitch link strain,

landing gear load, structural loads and fatigue, bearing health monitoring, swashplate

temperature and vibration, drive train torque, gearbox vibration, and flight regime recognition.

Figure 122 shows an example of rotorcraft sensing system.


126

Figure 122 Rotorcraft Wireless Health Monitoring [39]

11.6.2 Future of sensors technology

Wireless sensors will become deeply embedded within structures and machines in the

future. Sensed information will be automatically compressed and forwarded for condition-based

maintenance. Sensors wil be able to: harvest and store energy from strain, vibration, light, and

motion; use power management to balance the energy “checkbook”; and use embedded

processors to compress data and compute fatigue life.


127

List of questions

1. Specify what methods are developed to detect and prevent material fatigue in aircraft

structures. Specify designs against fatigue related failures.

2. Explain infinite lifetime concept to keep stress below threshold of fatigue limit.

3. Explain finite lifetime concept methods.

4. Explain damage tolerant design. Explain and specify NDT methods.

5. Explain NDT visual inspection method.

6. Explain NDT penetrant method.

7. Explain NDT ultrasonic testing method.

8. Explain NDT magnetic particle inspection method.

9. Explain NDT radiographic testing.

10. Explain NDT eddy current testing.

11. Explain NDT low-coherence interferometry method.

12. Explain cold expansion methods. Specify cold expansion methods.

13. How does cold expanision improve fatigue life? Explain.

14. What are important benefits and improvements that cold expansion methods bring?

15. What id BITE? Explain.

16. What is SHM? Explain.

17. Explain importance of SHM wireless sensors.


128

12 Reparation of aircraft structure

Fixed-wing aircraft airframe is generally considered to consist of five main elements: the

fuselage, wings, stabilizers, flight control surfaces, and landing gear. Aircraft main structural

elements and joints are designed to carry loads by distributing them as stresses. The elements and

joints are strong enough to resist these stresses, and must remain so after any repairs.

Any repair made on an aircraft structure must allow all of the stresses to enter, sustain

these stresses, and then allow them to return into the structure. The repair must be equal to the

original structure, but not stronger or stiffer, which will cause stress concentrations or alter the

resonant frequency of the structure. All-metal aircraft are made of very thin sheet metal, and it is

possible to restore the strength of the skin without restoring its rigidity.

All repairs should be made using the same type and thickness of material that was used in

the original structure. If the original skin had corrugations or flanges for rigidity, these must be

preserved and strengthened. If a flange or corrugation is dented or cracked, the material loses

much of its rigidity; and it must be repaired in such a way that will restore its rigidity, stiffness,

and strength. [23]

12.1 Types of aircraft structures

12.1.1 Riveted steel structures

The general principles on repairs of riveted or bolted structures must be followed

according to relevant aircraft repair manuals and manufacturer’s instructions.

12.1.2 Aluminium alloy structures

Repairs to damaged skin on monocoque types of aluminum alloy structures should be

conducted in accordance with specific manufacturer’s instructions or other approved sources.

12.2 Methods to repair aircraft structures

Methods to repair aircraft structure include careful examination of all adjacent rivets

outside of the repair area to ascertain that they have not been harmed by operations in adjacent

areas; drilling rivet holes round, straight, and free from cracks; deburring the hole with an

oversize drill or deburring tool. Information on special methods of riveting, such as flush

riveting, usually may be obtained from manufacturer’s service manuals.

12.2.1 Splicing of tubes

Round or streamline aluminum alloy tubular members may be repaired by splicing

(Figure 123). Splices in struts that overlap fittings are not acceptable.

Rivets which are loaded in shear should be hammered only enough to form a small head

and no attempt made to form the standard roundhead. Correct and incorrect examples of this type

of rivet application are shown in Figure 123. [23]


129

12.2.2 Repairs to aluminum alloy members

Aluminum alloy members’ repairs must be done with the same material or with adequate

material of higher strength. Particular attention should be paid to design of parts to avoid

notches, small radii, and large or rapid changes in cross-sectional areas, in order to take

advantage of strength characteristics.

Caution must be taken to avoid processing and handling defects, like machine marks,

nicks, dents, burrs, scratches, and forming cracks.

12.2.3 Reinforcement of wing and tail surface ribs

Damaged aluminum alloy ribs may be fixed by the addition of adequate reinforcement. In

small and medium size aircraft, there are specific types of ribs that this method can deal with.

Also, schemes of repair can be acceptable if they are developed by the aircraft manufacturer.

12.2.4 Reinforced splices of trailing and leading edges and tip strips

Repairs to wing, control surface trailing edges, leading edges, and tip strips should be made by

properly executed and reinforced splices. [23]

12.2.5 Repair of damaged skin

In the places where metal skin is severely damaged, repair should be done by replacing a

whole sheet panel from one structural member to the next.

12.2.6 Patching of small holes

Undamaged small holes in skin panels could be patched by covering the hole with a patch

plate. Also, in stressed-skin type construction, the flush patches may be installed.

12.2.7 Splicing of sheets

The splice must be designed properly when the sheet has cut-outs, or doubler plates at an

edge seam, or when other members transmit loads into the sheet.

12.2.8 Straightening of stringers or intermediate frames

Slightly bent members could be cold straightened and examined for cracks or ruptures in

the material with a magnifying glass.

12.2.9 Local heating

Local heating must not be applied to facilitate bending, swaging, flattening, or expanding

operations of heat-treated aluminum alloy members, as it is difficult to control the temperatures


130

closely enough to prevent possible damage to the metal, and it may impair its corrosion

resistance. [23]

12.2.10 Splicing of stringers and flanges

All splices should be made in accordance with the manufacturer’s recommendation.

Splices should be designed in such way to carry both tension and compression.

12.2.11 Drill small holes

At the extreme ends of the cracks, drilling small holes, is done to minimize the possibility

of further spreading.

12.2.12 Adding reinforcement of joints

Adding reinforcement is done to carry the stresses across the damaged area and to

strengthen the joints.

12.2.13 Steel fittings

Steel fitting are inspected for defects, in careful examination of the fitting with a medium

power magnifying glass. Inspection is performed according to the manufacturer’s instruction

manual.

12.2.14 Aluminium and aluminium alloy fittings

Damaged fittings should be replaced with new parts that have the same material

specifications. Repairs may be made in accordance with data furnished by the aircraft

manufacturer, or data substantiating the method of repair.

12.2.15 Selective plating in aircraft maintenance

Selective plating is a method of depositing metal from an electrolyte to the selected area.


131

Figure 123 Typical repair method for tubular members of aluminum alloy [23]


132

List of questions

1. Explain the principals and considerations in aircraft structure reparation.

2. Specify methods to repair aircraft structure. Explain one of them.


133

List of Tables

Table 1 Old aircraft in active use ................................................................................................... 3 Table 2 European manufacturers ................................................................................................... 4 Table 3 Examples of “life-prolonging” Boeing fleet ................................................................... 10 Table 4 Forces acting on an aircraft caused by various effects ................................................... 54

Table 5 Electrochemical potential of metals ................................................................................ 73 Table 6 Crack growth investigations ......................................................................................... 121


134

List of Figures

Figure 1 Structures Task Group [author] ....................................................................................... 7 Figure 2 Comet 1A at Le Bourget Airport [30] ........................................................................... 14 Figure 3 Japan Airlines Flight 123; Boeing 747 at Osaka International Airport in 1982 [29] .... 15 Figure 4 Route of JAL123 and sequence of events [29] .............................................................. 16

Figure 5 This photograph shows the plane as it looked after explosive decompression; the

vertical stabilizer is missing (circled in red) [29] ......................................................................... 16 Figure 6 Diagram of correct and incorrect repairs of the bulkhead [29] ..................................... 17 Figure 7 Diagram of the aft pressure bulkhead [29] .................................................................... 17 Figure 8 Boeing 747SR-46 [30] ................................................................................................... 18

Figure 9 Aloha Airlines Flight 243 [29] ...................................................................................... 19

Figure 10 Boeing 737-297 [30].................................................................................................... 20

Figure 11 El Al Flight 1862 [29] ................................................................................................. 21 Figure 12 Fatal El Al flight of Boeing 747-200 [29] ................................................................... 21 Figure 13 A map of Amsterdam showing the aircraft's flight path (marked in green) [29] ........ 21 Figure 14 China Airlines Flight 611 [29] .................................................................................... 22

Figure 15 Mechanical parameters and components [4] ............................................................... 25 Figure 16 Simplest form of vibrating system – mass and spring [4] ........................................... 26 Figure 17 Free vibrations [4] ....................................................................................................... 28

Figure 18 Free vibrations [4] ....................................................................................................... 29 Figure 19 Free vibration without damping [29] ........................................................................... 29

Figure 20 The spring – simple harmonic motion of the mass–spring system [29] ...................... 30 Figure 21 Increase of mass [4] ..................................................................................................... 31

Figure 22 Forced vibration [4] ..................................................................................................... 32 Figure 23 Forced vibration [29] ................................................................................................... 32

Figure 24 Mass-spring-damper model [29] ................................................................................. 33 Figure 25 Constant damping [4] .................................................................................................. 34 Figure 26 Viscous damping [4].................................................................................................... 34

Figure 27 Small damping [4] ....................................................................................................... 35 Figure 28 Free vibration with 0.1 and 0.3 damping ratio [29] ..................................................... 35

Figure 29 Mass, spring and damper [4] ....................................................................................... 36 Figure 30 Phase shift [29] ............................................................................................................ 37 Figure 31 Normal distribution to determine or predict jitter [29] ................................................ 37 Figure 32 Mechanical resonance in a mechanical oscillatory system [29] .................................. 38

Figure 33 Wöhler curve [29]........................................................................................................ 39 Figure 34 Loads main tasks [9] .................................................................................................... 44

Figure 35 Loads loop [9].............................................................................................................. 45 Figure 36 V(Ma)-n diagram in altitude [9] .................................................................................. 46 Figure 37 Altitude-Mach number envelopes [9] .......................................................................... 47 Figure 38 Mach-Altitude points of loads model [9] .................................................................... 48 Figure 39 Fin buffet at high angle of attack (flight test results) [9] ............................................. 50

Figure 40 MIL-SPEC pitch manoeuvre [9] ................................................................................. 51 Figure 41 Flight parameter envelopes for structural design [9] ................................................... 52 Figure 42 Load monitoring stations [9] ....................................................................................... 53


135

Figure 43 Major aircraft component loads envelopes [9] ............................................................ 54 Figure 44 Front fuselage transport joint critical load conditions; an increasing front fuselage

mass will lead to higher front fuselage loading [9]....................................................................... 55 Figure 45 Influence of wing loading conditions on wing loads; adding mass to the wing (e.g.

carriage of stores) leads to reduced wing loads [9] ....................................................................... 55 Figure 46 Prototype pressure plotting for flight load survey [9] ................................................. 56 Figure 47 Static loads design criteria for airframes [8] ............................................................... 57 Figure 48 Typical excedance curves for combat aircraft [8] ....................................................... 57 Figure 49 Coarse mesh FE model of wing structure [9] .............................................................. 58

Figure 50 FE half-model of center fuselage structure [9] ............................................................ 58 Figure 51 EF2000 Global Model for Unified Analysis [9] .......................................................... 59 Figure 52 Coarse mesh FE model of center fuselage frame [9] .................................................. 59

Figure 53 Fine mesh FE model for detail analysis [9] ................................................................. 60 Figure 54 Fuselage structure [31] ................................................................................................ 62 Figure 55 Loads acting along half of wing [32] .......................................................................... 63

Figure 56 Flight control surfaces [33] ......................................................................................... 63 Figure 57 Loads on ailerons [34] ................................................................................................. 64

Figure 58 Piston engine [30] ........................................................................................................ 64 Figure 59 Turbofan jet engine propulsion system [35] ................................................................ 65 Figure 60 Landing gear of Airbus A340-300 [30] ....................................................................... 65

Figure 61 Adhesion [author] ........................................................................................................ 67 Figure 62 Abrasion [author]......................................................................................................... 68

Figure 63 Abrasion between two objects [author] ....................................................................... 69 Figure 64 Abrasion between three objects [author] ..................................................................... 69 Figure 65 Erosion [author] ........................................................................................................... 70

Figure 66 Example of screw fracture due to material fatigue [29] .............................................. 78

Figure 67 Example of fracture of hexagonal profile due to material fatigue; it is clear that the

dark area is a result of gradual, propagatory material fracture and light part indicates fast fracture

[29] ................................................................................................................................................ 78

Figure 68 Example of handle fracture due to material fatigue; besides the fact that fracture

surfaces have different shapes, the area of slow and prolonged propagating crack has a darker

colour, because it is oxidized (corroded), in contrary to the area of current fast fracture which is

grainy and light [29]...................................................................................................................... 79

Figure 69 Example of piston rod fracture due material fatigue; figure clearly shows the area of

slow and fast (current) fracture; initial area is also apparent, where beginning of cracking took

place, due to stress concentration, inclusions, local weakening, etc. [29] .................................... 79 Figure 70 Example of screw fracture due to material fatigue on the surface, caused by

irregularities in the structure, i.e. by damage to the edge (initial crack), cyclic loads due to

incorrect torque, overload, corrosion, etc., (not necessarily all conditions and not necessarily in

that order) [29] .............................................................................................................................. 80

Figure 71 Example of fracture of aircraft structure of de Havilland DH 106 Comet after the

accident in 1954, due to fatigue caused by cyclic pressurisation, stress concentration, lack of

advenced material technology and errors in technology production [29]..................................... 80 Figure 72 Fracture limits [29] ...................................................................................................... 81 Figure 73 Stresses versus loading [29] ........................................................................................ 82


136

Figure 74 Wohler’s curve [29] ..................................................................................................... 83 Figure 75 Wohler’s diagram [29] ................................................................................................ 84 Figure 76 Stress concentration [29] ............................................................................................. 85 Figure 77 Point defects of crystals [29] ....................................................................................... 86

Figure 78 De Havilland DH 106 Comet [29]............................................................................... 88 Figure 79 S-N curve (Wohler curve) [29] .................................................................................... 89 Figure 80 Mathematical description of Goodman’s line [29] ...................................................... 90 Figure 81 Internal force lines are denser near the hole [29] ........................................................ 91 Figure 82 The sharp corners at the windows of Comet 1 G-ALYP [29] ..................................... 92

Figure 83 Increasing R curves: (a) stable fracture from Gi to G3; (b) Stable fracture for cracks

with different lengths [24] ............................................................................................................ 93 Figure 84 Level of damage for different crack lenghts as a function of number of cycles (loads)

[24] ................................................................................................................................................ 95 Figure 85 Parametres of crack propagation due to material fatigue: (a) un-sharpening and re-

sharpening of the crack tip; (b) Δσ, ΔK [24] ................................................................................. 96

Figure 86 De Havilland DH 106 Comet [24]............................................................................... 97 Figure 87 Reconstruction of the fuselage and empennage from the Comet G-ALYP reckage [24]

....................................................................................................................................................... 98 Figure 88 De Havilland DH 106 Comet [24]............................................................................... 98 Figure 89 a) Crack in the De Havilland DH 106 Comet G-ALPY; b) detailed illustration of the

initial crack on the Comet G-ALYP [24] ...................................................................................... 99 Figure 90 Comet G-ALYU in testing pool [24]......................................................................... 100

Figure 91 Fuselage break-up of tested model of Comet G-ALYU [24] .................................... 101 Figure 92 Probable cause of Comet G-ALYU fuselage break-up; stress dissemination calculated

during repairs [24]....................................................................................................................... 102

Figure 93 Advances in construction design of Comet’s windows [24] ..................................... 103

Figure 94 Damage to Aloha Airlines Boeing 737 [24] .............................................................. 103 Figure 95 Multiple crack damage [24]....................................................................................... 104 Figure 96 Illustration of rivet and its fatigue crack on Aloha Airlines Boeing 737 [24] ........... 104

Figure 97 Back pressure obstacle on JAL Boeing 747-SR [24] ................................................ 106 Figure 98 Video borescope [36] ................................................................................................ 110

Figure 99 1) section of material with a surface-breaking crack that is not visible to the naked

eye, 2) penetrant is applied to the surface, 3) excess penetrant is removed, 4) developer is

applied, which renders the crack visible [29] ............................................................................. 111 Figure 100 An example of Ultrasonic Testing (UT) on blade roots of a V2500 IAE aircraft

engine (step 1: the UT probe is placed on the root of the blades to be inspected with the help of a

special borescope tool (video probe), step 2: instrument settings are input, step 3: the probe is

scanned over the blade root, in this case, an indication (peak in the data) through the red line (or

gate) indicates a good blade; an indication to the left of that range indicates a crack) [2] ......... 112 Figure 101 A horizontal MPI machine with a 36 in (910 mm) coil [29] ................................... 113

Figure 102 Radiographic testing [37] ........................................................................................ 114 Figure 103 Eddy current testing [29] ......................................................................................... 115 Figure 104 The schematic setup of the low-coherence interferometer [19] .............................. 116 Figure 105 The cross-sectional image for composite material sample [20] .............................. 116 Figure 106 Surface treatment cold working methods [22] ........................................................ 117


137

Figure 107 High interference expansion methods [22].............................................................. 117 Figure 108 Split Sleeve Cold Expansion [22]............................................................................ 117 Figure 109 Induced residual stress distribution [22].................................................................. 118 Figure 110 Effect of applied expansion on fatigue life improvement [22] ................................ 118

Figure 111 Fatigue life improvement, weight savings and design optimization [22] ............... 119 Figure 112 Fatigue life improvement in aluminium [22] .......................................................... 119 Figure 113 Fatigue life improvement in aluminium [22] .......................................................... 119 Figure 114 Fatigue life extension in titanium [22] .................................................................... 120 Figure 115 Effect of fastener interference [22] .......................................................................... 120

Figure 116 Effect of cold expansion on crack tip stress intensity [22] ...................................... 121 Figure 117 Fatigue crack abatement [22] .................................................................................. 121 Figure 118 Extended inspection interval with cold expansion (with initial 0.005” flaw) [22] . 122

Figure 119 Damage tolerance benefits (fighter aircraft study) [22] .......................................... 122 Figure 120 BITE concept on the Airbus 320 [2] ....................................................................... 123 Figure 121 Structural health monitoring (SHM) systems can be arrayed in similar fashion to the

human nervous system, with sensors concentrated in key areas where loads are highest [38] .. 125 Figure 122 Rotorcraft Wireless Health Monitoring [39] ........................................................... 126

Figure 123 Typical repair method for tubular members of aluminum alloy [23] ...................... 131


138

Conclusion

An aircraft begins to age during its first flight and various effects of ageing begin to occur

almost immediately. However, the term “ageing” is usually applied to the issues which begin to

arise when it gets significant. Damage tolerance and safe-life design philosophies are applied

nowadays and appropriate inspection methods are developed to detect the effects of accidental,

environmental or fatigue damage. It is also common that control and inspection programmes are

established for fatigue and corrosion related issues.

Keeping older aircraft in an airworthy condition has been found to present special

difficulties which have not all been addressed by prescribed maintenance. The serious continuing

airworthiness issues which have arisen in many ageing aircraft have often been a direct

consequence of the gap between current and former practices required for Aircraft Type

Certificate issue and Maintenance Programme approval.

Until quite recently, some significant issues arising from aircraft age had not been

recognised and addressed until after fatal accidents had occurred. Also, the general principles of

system deterioration that affect all older aircraft started to receive renewed attention.

The United States, which has seen most examples of accidents attributed to ageing

aircraft problems, has had for several years now a joint civil-military organisation called the

Joint Council on Ageing Aircraft (JCAA) to coordinate the development of risk management

solutions for the various types of ageing aircraft problems, especially addressing ageing aircraft

structures problems.

Awareness of these safety issues in the other leading airworthiness jurisdictions of

design, production and maintenance regulation is now similarly high and preventive methods and

interventions are being developed. The maintenance issues which have particularly arisen from

ageing aircraft structural failures, also generally arised from fatigue or corrosion effects, with

corrosion sometimes initiating fatigue effects.

For existing aircraft, improved inspections, including the use of non-destructive testing

(NDT), and the management of any corrosion found through effective repair techniques,

mapping technologies, and recording are the main option and solution.


139

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