design of modern aircraft structure and the role of ndi

6/1/12 Design of Modern Aircraft Structure and the Role of NDI

1/13www.ndt.net/article/ecndt98/aero/001/001.htm

TABLE OF CONTENTS

IntroductionAirworthiness requirements and complianceDesign principles and justification methods

Design principle 'safe life'Design principle 'damage tolerant'Example for inspection

Design of modern aircraft structureDesign criteriaMaterial selectionSpecial NDI application

Aging aircraft issues and activitiesAging aircraft initiativesThe aging aircraft issue 'WidespreadFatigue Damage'Repair assessment for aging aircraft

ConclusionReferences

NDT.net - June 1999, Vol. 4 No. 6

Table of Contents ECNDT'98Session: Aerospace

Design of Modern Aircraft Structure and the Role ofNDI

H.-J. Schmidt, B. Schmidt-Brandecker, G. ToberDaimler-Benz Aerospace Airbus

Introduction

The current generation of civil transport aircraft were designed

for at least 20 to 25 years and up to 90 000 flights. These designservice goals are exceeded by many operators of jets and

turboprops. Future aircraft types are designed for at least the

same goals, but structure with higher fatigue life (endurance),higher damage tolerance capability and higher corrosionresistance are required to minimize the maintenance costs and tocomply with the requirements of the operator and the enhancedairworthiness regulations.

Non destructive inspections (NDI) are still significant means tofulfill all the requirements. Further significant applications of ND1are in the frame of another major aviation issue, the aging aircraftissue. Especially the activities regarding widespread fatiguedamage (WFD) and the assessment of existing repairs require the application of newly developed and availableND1 methods.

Airworthiness requirements and compliance

Due to several structural damages which occurred during service and under consideration of the requirements of theUS american airforce the airworthiness regulations for civil transport aircraft have been developed significantly inthe past 45 years. Especially the introduction of the fatigue and damage tolerance requirements mark the majorsteps. Table 1 shows an overview of the regulations developed in the USA.

Table 1: Development of airworthiness regulations in the USA

1953 - CAR4b: no special regulations regarding fatigue

1956 - CAR4b Amendment 3: regulations regarding 'safe life' and 'fail-safe'.

1962 - CAR4b Amendment 12: regulations regarding fatigue for landing gears

1966 - FAR25 Amendment 10: sonic fatigue

1978 - FAR25 Amendment 45: introduction of 'damage tolerance' regulations

1981 - FAR25 Amendment 54: further airworthiness regulations for aircraft certified prior to amendment 45

To guarantee an equivalent standard of regulations in the USA and Europe harmonization meetings were held

between the airworthiness authorities and the manufacturers under the umbrella of the Aviation RulemakingAdvisory Committee (ARAC). Furthermore the new aspects regarding widespread fatigue damage (WFD) were

considered. The harmonized forthcoming regulation and advisory circular require: 'An evaluation of the strength,detailed design, and fabrication must show that a catastrophic failure due to fatigue, corrosion, or accidental

damage, will be avoided throughout the operational life of the airplane.' and 'The ultimate purpose of the damage



tolerance evaluation is the development of a recommended structural inspection program considering probable

damage locations, crack initiation mechanisms, crack growth time histories and crack detectability.'

The major requirements of the damage tolerance evaluation are:

Widespread fatigue damage assessmentIdentification of possible damage locations and extent of damage

Damage tolerance analyses and testDetermination of inspection threshold and intervals

The major differences compared with the current regulations are the requirements that:

Sufficient fullscale testing must be accomplished to ensure that widespread fatigue damage will not occur

within the design service goal of the airplane.The inspection threshold for certain types of structure has to be established based on crack growth analysis

and/or tests.

The development of the structural inspection program is shown in Fig. 1. For each structural element to be

inspected the following information has to be provided which are comprised in the Maintenance Review Board(MRB) report:

Inspection threshold: time of first inspection in flightsInspection interval: period between the repeated inspections in flightsInspection area: detailed description of the area to be inspected including location and access

Inspection method:information of the method to be used, for ND1 methods the detailed description ofthe method is given in a special handbook

Fig 1: Development of structure inspection program



In general the inspection threshold is determined by the fatigue life to crack initiation under consideration of arelevant scatter factor. For specific structure the threshold is to be based on crack growth analysis.

The inspection interval is determined from the crack growth period between the detectable crack length for thestructural detail and the critical crack length under limit load divided by a scatter factor, see Fig. 2.

Fig 2: Principle of damage tolerance investigation

The damage tolerance requirements lead to three major tasks for the aircraft manufacturer:

Structural design according to fatigue and damage tolerance requirementsEvaluation of the structure by analysis supported testsDefinition of a structural inspection program

Design principles and justification methods

Due to the complexity of the structural elements, their function and location,

several design principles are used to design a damage tolerant structure. In

addition to this the safe life principle is still applied for specific cases.

Design principle 'safe life'

The safe life design principle was applied in aircraft design prior to 1960.According to JAR/FAR 25.57 1 a safe life design is now allowed for the landing

gear and its attachments only.

An example is given in Fig. 3. A structure designed as safe life contains a singleload path only and the inspectable crack length may be in the range of the critical



Nose Landing Gear A320 Fig 3: Design Principle 'safe life'

crack length. Consequently inspection intervals to monitor the structure cannot be

defined. A failure of one of the structural elements leads to the complete failure of

the safe life structure and possibly to significant consequences for the aircraft.

A fatigue resistant design of safe life structure is based on fatigue life calculations

for all structural elements during the design phase and is justified by full scale fatigue test with the complete safestructure. The fatigue life calculations are performed using the linear damage accumulation according to Palmgren-

Miner considering relevant load spectra and material (S-N) data. The calculated fatigue life as well as the achieved

test life are divided by relevant scatter factors.

Design principle 'damage tolerant'The damage tolerance design principle comprises two categories which are 'single load path' and 'multiple load

path' structure.

Fig. 4 shows a single load path design where the justification is based on the following analyses. Fatigue life

calculations are performed to justify the reliability during service and to determine the inspection threshold. For

future projects the inspection threshold has to be based on crack growth analysis according to the forthcomingregulations. The inspection interval is determined from the crack growth period between the detectable and the

critical crack length divided by a scatter factor. The calculation of the crack growth is based on the Forman

equation or equivalent.

Example:

Fig 4: Design principle 'damage tolerant - single load path'

The 'multiple load path' category is sub-divided into three groups:

multiple load path - externally inspectable onlymultiple load path - not inspectable for less than one complete load path failure

multiple load path - inspectable for less than one complete load path failure



Fig 6: Application of ND1in structural inspectionprogram of A320-100

Only the latter group is described here, see Fig. 5. For structures 'damage tolerant - multiple load path -

inspectable for less than one complete load path failure' again fatigue life calculations are performed to showsufficient reliability during service and to determine the inspection threshold, which is derived from the structural

element with the lowest fatigue life. The inspection interval is based on the crack growth behavior of both loadpaths were in the primary load path an initial flaw of 1.27 mm is assumed and in the secondary load path an initial

flaw of 0.127 mm. The interval is determined by the crack growth period between the detectable crack length in

the primary load path and the critical crack length in the secondary load path divided by an appropriate factor. Forthe crack growth calculations the same method as for single load path structure is applied.

Example:

Fig 5: Design principle 'damage tolerant - multiple load path -inspectable for less than one complete load path failure'

The current, and forthcoming, regulations allow both damage tolerance categories, i.e. single load path and multipleload path. The multiple load path design, however, is highly recommended in the interpretation of the regulations

(advisory circular AC/ACJ 25.571). The recommended multiple load path design leads to additional safety, but

causes, in exceptional cases, significant costs during design and production.

Examples for inspections

The structural inspection program comprises three categories or inspection levels which are:

General visual inspection (GVI):

a visual examination to detect obvious unsatisfactory conditions and discrepancies. The inspections areperformed in frame of the so called zonal inspection program where the complete aircraft, divided in zones,

is inspected in regular time intervals.

Detailed visual inspection (DET):an intensive visual examination of a specified detail or assembly searching for evidence of irregularity.

Special detailed inspection (SDET):an intensive examination of a specific location similar to the detailed inspection but requiring special

techniques, mostly NDI.

Fig. 6 shows the distribution of the inspection levels for the structural significant items

(SSI's) of the major aircraft components using the standard body Airbus A320400 as an

example. Several SSI's comprise more than one inspection task. Except for the safe lifelanding gears the 5.percentages of the ND1 tasks are 6 percent for the stabilizer (mainly

composite), 11 percent for fuselage and doors, 18 percent for wing and 19 percent for

the pylons. The percentage of ND1 tasks may be higher for widebody aircraft whichhave in general higher stress levels in most of the structural details leading to faster crack

propagation and lower critical crack length. Therefore sometimes an ND1 method is



Fig 7: Externalinspections of upper andside shells of A320-100center fuselage section

Fig 8: Design of aircraftstructures

Fig 9: Planned Airbusmegaliner A3XX

Fig 10: Two-bay-crackcriterion

chosen to reach a sufficient inspection interval.

The external inspections of the upper and side shells of the A320-100 are given in Fig. 7.

Besides a general visual inspection of the complete shells, special tasks of general visualinspections, also covered by the zonal program, are described for the upper panel of the

longitudinal lap joints. Detailed inspections are to be performed of the skin at thecircumferential joints in the upper area, the surrounding of cut-outs in the upper shell, the skin and the window

frames and the cut-out comers of the emergency exits. ND1 methods are used for the strap at the circumferential

joints (upper area) and, offered as an alternative to a detailed inspection of externally visible cracks, for the lowerpanel of the longitudinal lap joint in the upper shell. In principle these external inspections are typical examples for

the fuselage upper and side shells at standard body and wide-body Airbus aircraft. The only exception are the cut-

out comers of the doors where on widebody aircraft mostly ND1 are applied due to the higher stress level.

Design of modern aircraft structure

Design criteria

During the design of aircraft structures several aspects have to be considered to reachsufficient static strength as well as sufficient fatigue and damage tolerance behavior, see

Fig. 8. The result of iterative calculations is an optimized design regarding weight, costs

and aircraft performance.

Several aspects of the design of modern aircraft structure are described here using the

fuselage of the planned Airbus megaliner A3XX as an example, see Fig. 9. This aircraftis to be designed for the following goals:

Design service goal 24 000 flightsInspections goals

- general visual (C-check, zonal program)

- threshold for detailed inspections / ND1- interval for detailed inspections / ND1

24 months

12 000 flights6 000 flights

The design criteria to be met are static strength, residual strength, durability, crackgrowth, sonic fatigue strength and the so-called two-bay-crack criterion. This requiresthe consideration of corresponding loads as static loads, residual strength loads, discrete

source damage loads, operational loads and sonic fatigue loads. Furthermore thecorrosion resistance, the repairability and the inspectability have to be taken into account.

One of the major criteria which an aircraft has to fulfill to reach the safety standard of the competitors is the two-

bay-crack criterion, see Fig. 10. It has to be shown, that a longitudinal crack in the skin of the pressurized fuselagewith a length of two frame bays above a broken center frame does not lead to a complete failure of the structure.The load case to be considered is 1.15 of the onerational cabin differential nressure at cruise altitude without

consideration of external loads.

The structure of a pressurized fuselage which fulfills this criterion has to guarantee that neither the crack in the skinbecomes unstable nor that the stiffeners perpendicular to the crack (i.e. the frames) fail statically. The two-bay-crack criterion is the designing criterion for large areas in the upper and side shells of the pressurized fuselage of

medium and long range aircraft. These aircraft types have lower design service goals in flights compared with shortrange aircraft with the result that the fatigue and damage tolerance criteria have less influence on the design. To limitthe implications on the weight due to the compliance with the two-bay-crack requirement following precautions are

possible:



selection of skin material with high residual strengthselection of frame material with high static strength

limitation of the allowable frame pitchadaptation of the stress level to the two-bay-crack criterion.

Material selection

During the initial design phase of the Airbus A3XX the application of new materials and production methods isconsidered to reduce the production costs and the weight and to comply with the forthcoming regulations. Tosubstitute the fuselage material of the current Airbus types, i.e. the 8.aluminium alloy 2024, three different materialsare under consideration; these are 2524,60 13 and GLARE, see table 2.

Table 2: Materials for fuselage skin

material data 2024T3 clad 2524T3 clad 6013T4/T unclad GLARE4 (LT/TL) unclad

Rm (in %) 100 100 ~75 190 / 120

Rp0.2 (in %) 100 100 -94 ll0 / 80

blunt notch (in %) 100 100 not tested l43 / 100

young's modulus(tension) (in %) 100 100 ~95 79 / 70

KC (in %) 100 -120 ~115 ~120 / -110

(in %) 100 100 97 87

corrosion resistance basis equal equal / less higher

The materials 2524 and GLARE4 show significantly higher fracture toughness compared with 2024 which results in

significant weight reductions in those areas which are designed by the two-bay-crack criterion. The disadvantage ofboth materials is the higher price. For the GLARE4 material this may be (partly) compensated by a simplifieddesign and production, GLARE4 has additionally advantages with respect to the static strength, the yield strength

and the corrosion resistance. Furthermore GLARE4 shows a very good bum through behavior which should betaken into account besides the structural aspects. The material 6013 leads to similar structural weights as 2524considering the slightly lower yield strength which is approximately compensated by the lower density. 60 13 canbe welded which allows to substitute the bonding or riveting of the stringers to the skin by welding. This new

production method is very promising with respect to the reduction of the production costs.

The different material data allow an increase of the allowable circumferential stresses in the fuselage of the A3XXfor all of the three new materials. An increase of the allowable longitudinal stress in the fuselage is possible when

using 2524T3. Table 3 contains the allowable skin stresses for a the frame pitch of 656 mm. The allowable stressesin circumferential direction result from the two-bay-crack criterion, the criterion for the longitudinal stresses is eitherthe crack growth,i.e. the inspection interval, or the two-bay-crack criterion depending on the ratio of static and

fatigue loads.

Table 3: Allowable stresses for fuselage skin

skin material

allowable stress in allowable stress in

circumferential direction longitudinal direction(residual strength)

allowable stress in longitudinal

direction (crack growth / residualstrength)

2024T3 clad 100 % 100 % / 100 %

2524T3 clad 120 % 113 % / 110 %

6013T4/T6unclad (integral

stringers)115 % 104 % / 70 %

CLARE4 clad 120 % 120 % / 100 %



Fig 11: Design criteria forA3XX fuselage sections

The improvements given in table 3 lead to weight reductions in those areas where thedamage tolerance aspects are the dimensioning criteria. Further design cases to be

considered are e.g. the static tension and compression strength and the engine rotorfailure.Fig. 11 shows the design criteria in the different fuselage areas for an A3XX dependingon the skin material.

Finite element analyses were carried out for two fuselage sections of a length of 5.3 m and 2.7 m (forward and aftof the center section) considering the different design cases and the allowable stresses. The resulting structuralweights for the skin and the stringers were determined, see table 4. If the weight of the frame is taken into account

in addition the total weight reductions are less, e.g. for GLARE4 the weight reduction of the fuselage shell (skin plusstringers plus frames) is 12 percent instead of 16 percent for the skin and stringers only.

Table 4: Weights of two fuselage sections

skin materialcabin differential

pressureweight of two fuselage sections skin and stringer only (frame

pitch 656 mm)

2024T3 clad 605 hPa 100%

2524T3 clad 605 hPa 94%

6013T4/T6 unclad 605 hPa 103%

GLARE4 clad 605 hPa 84%

Special ND1 applicationThe development of a new production technique such as the laser beam welding (LBW) requires a comprehensiveuse of sophisticated inspection methods, especially the ND1 techniques. During the development of the LBW

technique for connection of the stringers to the fuselage skin the following standard ND1 methods are used:

High frequency ultra sonic test methodPenetration test method

Eddy current test method

The overall target is to provide an online ND1 method for valuation of the welding beam quality, i.e. methodsshould be available in the field of production for:

Position of welding gap (pre welding)

Control of process parameters during welding processControl of welding area (post welding)

Aging aircraft issues and activities

The well known Aloha accident near Hawaii in April 1988 which led to the loss of an upper forward fuselagesegment, resulted in worldwide activities to increase the safety of the aging aircraft fleet. Further events showed thatthe damage mechanism which led to the Aloha accident was not a single case and that the issue of widespread

fatigue damage (WFD) was not sufficiently covered by the current regulations.

Aging aircraft initiativesThe Aloha accident prompted considerable aviation community activity related to aging air frames. Manufacturers,operators and authorities got together to initiate changes to the system for safety improvement. A number of

industry committees were formed and the first was the Air worthiness Assurance Task Force (AATF) laterrenamed as the Airworthiness Assurance Working Group (AAWG) which works under the umbrella of theAviation Regulatory Advisory Committee (ARAC). Two other committees were formed which were the Industry



Fig 12: Effect of multiplesite damage

Fig 13: Effect of MSD onresidual strength of a leadcrack

Committee on WFD to study this phenomenon, and the Structural Audit Evaluation Task Group (SAETG) whichwas charged to develop guidelines to establish the beginning of WFD.

The FAA organized a number of conferences on aging aircraft and structural integrity which were supported by

NASA. They created centers of excellence by providing funding; two examples are the Georgia Institute ofTechnology tasked with the issue of computational mechanics and the Iowa State University tasked with nondestructive evaluation. Furthermore, rule changes were initiated to require full scale fatigue testing and inspectionthreshold determination for new aircraft as described in chapter 2.

Early in all these activities an interim solution was defined for eleven aircraft types which were defined prior to theintroduction of FAR 25.57 1 Amendment 45. These models are: Boeing B707, B727, B737, B747, DouglasDCS, DC9, DClO, Lockheed LlOll, BAe BAC 111, Fokker F28 and Airbus A300.

For these aircraft types the following activities were defined:

Periodical review of the inservice experience regarding structural damage (review of service bulletins)Introduction of a Corrosion Prevention and Control Program (CPCP)Assessment of the fatigue life of structural repairs

Establishment of an Supplement Structural Inspection Program (SSIP) to reach the safety standardaccording to FAR 25.57 1 Amendment 45Assessment of the structure regarding WFD.

The aging aircraft issue 'Widespread Fatigue Damage' The main issue of the aging aircraft fleet is the occurrence of multiple damages at adjacent locations which influenceeach other. Two types of multiple damages are known. The sketch on the upper righthand side of Fig. 12 shows an

example of multiple site damage (MSD), which is characterized by the simultaneous presence of fatigue cracks inthe same structural element. The second type is the multiple element damage (MED), which is characterized by thesimultaneous presence of fatigue cracks in similar adjacent structural elements. Both, MSD and MED, are a sourceof WFD which is reached when the MSD or MED cracks are of sufficient size and density that the structure will

not longer meet its damage tolerance requirement.

The effect of MSD is shown in Fig. 12. The lefthand diagram describes the effect ofMSD on a single lead crack used to establish the inspection program. In the presence of

MSD adjacent to the lead crack the critical crack or the residual strength, respectively,are reduced drastically. The righthand diagram shows the reduction of the crack growthperiod due to the reduction of the critical crack length.

Boeing has made investigations about the effect of MSD on the residual strength of a

lead crack which are published in /l/, see Fig. 13. The residual strength load of a 14 inch(356 mm) long lead crack is reduced in the presence of adjacent MSD cracks of 0.05inch (1.27 mm) by 30 percent. This demonstrates the dramatic effect even of small MSD

cracks which are uninspectable by state of the art techniques.

The Industry Committee on WFD has evaluated the experience of the participatingmanufacturers based on the results of large component and full scale fatigue tests as well

as in service experience in order to identify the locations potentially susceptible to WED.From this compilation of data each area was assessed for its susceptible to WFD andwas then characterized as either multiple element and/or multiple site damage. Fourteen areas were identified aspotentially susceptible to WFD:

Fuselage:



Longitudinal skin joints, frames and tear straps (MSD, MED)Circumferential joints and stringers (MSD, MED)Fuselage frames (MED)Aft pressure dome outer ring and dome web splices (MSD, MED)

Other pressure bulkhead attachment to skin-web attachment to stiffener and pressure decks (MSD, MED)Stringer to frame attachment (MED)Window surround structure (MSD, MED)

Over wing fuselage attachments (MED)Latches and hinges of nonplug doors (MSD, MED)Skin at runout of large doubler (MSD)

Wing and empennage:

Skin at runout of large doubler (MSD)Chordwise splices (MSD, MED)Rib to skin attachments (MSD, MED)

Stringer runout at tank end ribs (MED9 MSD)

Fig 14: Example of area potentiallysusceptible to WFD, circumferentialjoints and stringers

For each of these fourteen areas a typical design was given and the type and possible location of MSD/MED was

defined. An example is given in Fig. 14 showing circumferential joints and stringers. In detail the following damagetypes were defined:

MSD - circumferential joint

without outer doubler:

- splice plate - between and/or at the inner two rivet rows- skin - forward and aft rivet row of splice plate- skin - at first fastener of stringer coupling

with outer doubler:

- skin - outer rivet rows- splice plate/outer doubler - inner rivet rows

MED - stringer/stringer coupling- stringer - at first fastener of stringer coupling

-stringer coupling - in splice plate area

In August 1997 the FAA has tasked the ARAC to continue the activities on the WFD assessment and to extendthem to all transport category jets and turboprops with maximum gross weights greater than 75000 lbs. The ARAC

then chartered a new group in frame of the AAWG called Task Planning Group (TPG) with the following activities:

(1)

Review capability of analytical methods and their validation relative to the detection of WFD.Review evidence of WFD occurring in the fleet.

Recommend means of collection of inservice data where data missing.Determine extent of WFD in fleet.



Fig 15: Airbus repairassessment process

Extent AAWG 1993 report for all large transport aircraft > 75000 lb GW.

(2)

Establish time standards for the initiation and completion of model specific programs for prediction,verification and rectification of WFD.Recommend actions for the authorities, if a program for certain model airplanes is not performed prior to the

time standard.

The AAWG-TPG started their work in autumn 1997 in order to complete it within 18 months. The TPG hasdefined eight tasks to fulfill their charter:

Task 1 - Background: Review actions done

Task 2 - Technology issues: Technology readiness and validationTask 3 - Model specific issues: Establishment of time frame

Task 4 - Regulatory issues:FAA recourses if OEM fails to voluntary complete WFD

auditTask 5 - Management of MSD/MED in fleet: Inspection programs, replacementTask 6 - Aircraft to be considered in

recommendation:Define aircraft

Task 7 - March ARAC report issues and items: Issues to be presented to ARAC and AAWG responseTask 8 - Final report: Results of tasks 1 to 5

One major item of task 2 deals with the readiness of the ND1 technology. In frame of this subtask four actions

were defined to push the development of the methods needed:

Review of recent developmentsEstablishment of baseline flaw detectionDetermination of flaw size that needs to be detected

Determination of additional research and development needs

Repair assessment for aging aircraftContinuous airworthiness assessment of existiong repairs was identified as one of the five significant concerns by the

AAWG which formed a Repair Assessment Task Group (RATG) with participation of operators, manufacturersand authorities. The final draft report of this task group which was issued in December 1996 has recommended aone time structural repair assessment task for the external fuselage pressure boundary (skin and bulkhead webs) to

assure the continued airworthiness. This recommendation is again applicable to the eleven aircraft models certifiedprior to introduction of FAR 25.571 amendment 45. Consequently guidelines were developed to assess thedamage tolerance of existing structural repairs which may have been designed without using damage tolerance

criteria.

Based on the general three stage program, which wasdeveloped in a common effort by the major manufacturers andoperators for categorization of the repairs, the Airbus repair

assessment process was defined, see Fig. 15. Stage 1 (DataCollection) specifies what should be assessed for repairs. If arepair is on structure in an area of concern the analysis

continues, otherwise the repair does not require classificationas per this program. Stage 2 (Repair Categorization)categorizes the repairs regarding maintenance actions to be

applied. The repair categorization contains several steps which



Fig 16: External skin repair

Fig 17: Determination ofrepair parameters

Fig 18: Inspection of skinand external repairdoubler

consider the general conditions of the repair, the quality of thestatic design, the proximity to other repairs. Stage 3

(Determination of supplementary maintenance requirements)contains the definition of the necessary maintenance programfor the repair.

For the Airbus A300 aircraft Repair AssessmentGuidelines(RAG) were developed which allow the operatorsto determine the inspection threshold and interval for thecategory B repairs. Fig. 16 shows a principle sketch of an external skin repair. In principle four fatigue sensitivelocations exist which have to be assessed:

skin, longitudinal rivet row at doubler run-outskin, circumferential rivet row at doubler run-outdoubler, longitudinal rivet row adjacent to cut-outdoubler, circumferential rivet row adjacent to cut-out

The determination of the inspection threshold and interval requires the exact knowledgeabout the geometry, materials and fastener data to calculate the correct values forthreshold and interval. For dat not known conservative assumptions are to be made

which would lead to a worse threshold and / or interval. If the data are not available in arepair documentation, they may be taken directly from the aircraft. Some of the data maynot easil be measured, but NDI methods have to applied. Fig. 17 shows the applicationof NDI methods to determine the cut-out size hidden by the repair doubler, the thicknessof skin and doubler and the rivet material.

The inspection interval for the repair is based on the crack size detectable by NDImeans. Fig. 18 contains the NDI procedures for inspection of the skin and the external

repair doubler. All procedures have been qualified and comply with the definedinspection requirements that the defect size to be detected is determined with aprobability of detection (POD) of 90 percent at a confidence level of 95 percent.

Conclusion

The next aircraft generation has to comply with the forthcoming more stringent regulations, e.g. regardingwidespread fatigue damage and initial flaw concept for threshold determination. Furthermore the general aviationstandard with respect to the two-bay-crack criterion should be reached without special design precautions, such ascrack stoppers, and without disadvantages in weight. Additionally the requirements of the airlines regardingreduction of the maintenance costs have to be considered, i.e. among others the inspection intervals have to be

increased by decreasing the crack growth. These goals may be reached for fuselage structures by application ofnew materials. The development and application of new material is still under investigation to reach the optimum ofmaterial and production costs, weight and maintenance costs. During the development and certification of anaircraft the NDI plays a major role as shown in this paper. Further significant applications of NDI are within theframe of the aging aircraft activities where the detection of MSD and MED is an important item during theassessment of the structure susceptible to widespread fatigue damage.

The Repair Assessment Guidelines which were developed by Airbus also rely on NDI for determination of therepair parameters and the inspections of the repair.

References



1. T. Swift: Aging Aircraft From The Viewpoint of FAA, Presentation at Daimler-Benz Aerospace Airbus GmbH,Hamburg, Germany, September 17, 1997

2. D. Schiller, G. Tober, H.- J. Schmidt: NDT Technology for Fuselage Repair Assessment, Presentation at ATANDT FORUM 1995 in Cromwell (Hartford), Connecticut, USA, September 26 - 28, 1995

|Top|

NDT.net

Copyright © NDT.net, [email protected]/DB:Article /SO:ECNDT /AU:Schmidt_H-J /AU:Schmidt-Brandecker /AU:Tober_G /CN:DE /CT:NDT /CT:aerospace /ED:1999-06

design of modern aircraft structure and the role of ndi

Business

damage tolerance requirements

widespread fatigue damage

damage tolerance evaluation

accidental damage

inspection turboprops

sonic fatigue

current regulations

special regulations