Design of modern aircraft structure and the role of ndi

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

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

Design of Modern Aircraft Structure and the Role of
NDI
Table of Contents ECNDT H.-J. Schmidt, B. Schmidt-Brandecker, G. Tober
'98 Daimler-Benz Aerospace Airbus
Session: Aerospace

TABLE OF CONTENTS
Introduction Introduction
Airworthiness requirements and compliance
Design principles and justification methods
The current generation of civil transport aircraft were designed
Design principle 'safe life'
for at least 20 to 25 years and up to 90 000 flights. These design Design principle 'damage tolerant'
service goals are exceeded by many operators of jets and Example for inspection
turboprops. Future aircraft types are designed for at least the Design of modern aircraft structure
Design criteria
same goals, but structure with higher fatigue life (endurance), Material selection
higher damage tolerance capability and higher corrosion Special NDI application
resistance are required to minimize the maintenance costs and to Aging aircraft issues and activities
comply with the requirements of the operator and the enhanced Aging aircraft initiatives
The aging aircraft issue 'Widespread
airworthiness regulations. Fatigue Damage'
Non destructive inspections (NDI) are still significant means to Repair assessment for aging aircraft
fulfill all the requirements. Further significant applications of ND1 Conclusion
are in the frame of another major aviation issue, the aging aircraft References
issue. Especially the activities regarding widespread fatigue
damage (WFD) and the assessment of existing repairs require the application of newly developed and available
ND1 methods.

Airworthiness requirements and compliance

Due to several structural damages which occurred during service and under consideration of the requirements of the
US american airforce the airworthiness regulations for civil transport aircraft have been developed significantly in
the past 45 years. Especially the introduction of the fatigue and damage tolerance requirements mark the major
steps. 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 Rulemaking
Advisory 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

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


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 assessment
Identification of possible damage locations and extent of damage
Damage tolerance analyses and test
Determination 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 flights
Inspection interval: period between the repeated inspections in flights
Inspection area: detailed description of the area to be inspected including location and access
information of the method to be used, for ND1 methods the detailed description of
Inspection method:
the 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 a
relevant 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 the
structural 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 requirements
Evaluation of the structure by analysis supported tests
Definition 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 single
load path only and the inspectable crack length may be in the range of the critical


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. Nose Landing Gear A320
Fig 3: Design Principle 'safe life'
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 safe
structure. 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 forthcoming
regulations. 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 only
multiple load path - not inspectable for less than one complete load path failure
multiple load path - inspectable for less than one complete load path failure



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 show
sufficient 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 load
paths 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. For
the 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 multiple
load 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 are
performed 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 life
landing 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 Fig 6: Application of ND1
in structural inspection
the pylons. The percentage of ND1 tasks may be higher for widebody aircraft which program of A320-100
have 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



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 visual
inspections, also covered by the zonal program, are described for the upper panel of the Fig 7: External
longitudinal lap joints. Detailed inspections are to be performed of the skin at the inspections of upper and
circumferential joints in the upper area, the surrounding of cut-outs in the upper shell, the skin and theof A320-100
side shells window
center fuselage section
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 lower
panel 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 reach
sufficient 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 Fig 8: Design of aircraft
structures
fuselage of the planned Airbus megaliner A3XX as an example, see Fig. 9. This aircraft
is to be designed for the following goals:
Design service goal 24 000 flights
Inspections goals
- general visual (C-check, zonal program) 24 months
- threshold for detailed inspections / ND1 12 000 flights
- interval for detailed inspections / ND1 6 000 flights Fig 9: Planned Airbus
megaliner A3XX
The design criteria to be met are static strength, residual strength, durability, crack
growth, sonic fatigue strength and the so-called two-bay-crack criterion. This requires
the consideration of corresponding loads as static loads, residual strength loads, discrete
source damage loads, operational loads and sonic fatigue loads. Furthermore the
corrosion resistance, the repairability and the inspectability have to be taken into account. Fig 10: Two-bay-crack
criterion
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 fuselage
with 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 skin
becomes 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 short
range aircraft with the result that the fatigue and damage tolerance criteria have less influence on the design. To limit
the 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 strength
selection of frame material with high static strength
limitation of the allowable frame pitch
adaptation 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 is
considered to reduce the production costs and the weight and to comply with the forthcoming regulations. To
substitute the fuselage material of the current Airbus types, i.e. the 8.aluminium alloy 2024, three different materials
are 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 of
both materials is the higher price. For the GLARE4 material this may be (partly) compensated by a simplified
design 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 be
taken into account besides the structural aspects. The material 6013 leads to similar structural weights as 2524
considering the slightly lower yield strength which is approximately compensated by the lower density. 60 13 can
be 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 A3XX
for 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 stresses
in circumferential direction result from the two-bay-crack criterion, the criterion for the longitudinal stresses is either
the 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
allowable stress in allowable stress in allowable stress in longitudinal
skin material circumferential direction longitudinal direction direction (crack growth / residual
(residual strength) strength)
2024T3 clad 100 % 100 % / 100 %
2524T3 clad 120 % 113 % / 110 %
6013T4/T6
unclad (integral 115 % 104 % / 70 %
stringers)
CLARE4 clad 120 % 120 % / 100 %



The improvements given in table 3 lead to weight reductions in those areas where the
damage tolerance aspects are the dimensioning criteria. Further design cases to be
considered are e.g. the static tension and compression strength and the engine rotor
failure.
Fig. 11 shows the design criteria in the different fuselage areas for an A3XX depending Fig 11: Design criteria for
on the skin material. A3XX fuselage sections

Finite element analyses were carried out for two fuselage sections of a length of 5.3 m and 2.7 m (forward and aft
of the center section) considering the different design cases and the allowable stresses. The resulting structural
weights 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 plus
stringers plus frames) is 12 percent instead of 16 percent for the skin and stringers only.

Table 4: Weights of two fuselage sections
cabin differential weight of two fuselage sections skin and stringer only (frame
skin material
pressure 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 application
The development of a new production technique such as the laser beam welding (LBW) requires a comprehensive
use 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 method
Penetration 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. methods
should be available in the field of production for:

Position of welding gap (pre welding)
Control of process parameters during welding process
Control 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 fuselage
segment, resulted in worldwide activities to increase the safety of the aging aircraft fleet. Further events showed that
the 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 initiatives
The 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) later
renamed as the Airworthiness Assurance Working Group (AAWG) which works under the umbrella of the
Aviation Regulatory Advisory Committee (ARAC). Two other committees were formed which were the Industry


Committee on WFD to study this phenomenon, and the Structural Audit Evaluation Task Group (SAETG) which
was 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 of
Technology tasked with the issue of computational mechanics and the Iowa State University tasked with non
destructive evaluation. Furthermore, rule changes were initiated to require full scale fatigue testing and inspection
threshold 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 the
introduction of FAR 25.57 1 Amendment 45. These models are: Boeing B707, B727, B737, B747, Douglas
DCS, 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 standard
according to FAR 25.57 1 Amendment 45
Assessment 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 influence
each 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 in
the same structural element. The second type is the multiple element damage (MED), which is characterized by the
simultaneous presence of fatigue cracks in similar adjacent structural elements. Both, MSD and MED, are a source
of 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 of
MSD 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 growth
period due to the reduction of the critical crack length.
Fig 12: Effect of multiple
Boeing has made investigations about the effect of MSD on the residual strength of a site damage
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 0.05
inch (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. Fig 13: Effect of MSD on
residual strength of a lead
The Industry Committee on WFD has evaluated the experience of the participating crack
manufacturers 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 and
was then characterized as either multiple element and/or multiple site damage. Fourteen areas were identified as
potentially 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 potentially
susceptible to WFD, circumferential
joints 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 damage
types 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 extend
them 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.


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 has
defined eight tasks to fulfill their charter:

Task 1 - Background: Review actions done
Task 2 - Technology issues: Technology readiness and validation
Task 3 - Model specific issues: Establishment of time frame
FAA recourses if OEM fails to voluntary complete WFD
Task 4 - Regulatory issues:
audit
Task 5 - Management of MSD/MED in fleet: Inspection programs, replacement
Task 6 - Aircraft to be considered in
Define aircraft
recommendation:
Task 7 - March ARAC report issues and items: Issues to be presented to ARAC and AAWG response
Task 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 developments
Establishment of baseline flaw detection
Determination of flaw size that needs to be detected
Determination of additional research and development needs

Repair assessment for aging aircraft
Continuous 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, manufacturers
and authorities. The final draft report of this task group which was issued in December 1996 has recommended a
one 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 certified
prior to introduction of FAR 25.571 amendment 45. Consequently guidelines were developed to assess the
damage tolerance of existing structural repairs which may have been designed without using damage tolerance
criteria.

Based on the general three stage program, which was Fig 15: Airbus repair
developed in a common effort by the major manufacturers and assessment process
operators for categorization of the repairs, the Airbus repair
assessment process was defined, see Fig. 15. Stage 1 (Data
Collection) specifies what should be assessed for repairs. If a
repair is on structure in an area of concern the analysis
continues, otherwise the repair does not require classification
as per this program. Stage 2 (Repair Categorization)
categorizes the repairs regarding maintenance actions to be
applied. The repair categorization contains several steps which


consider the general conditions of the repair, the quality of the
static design, the proximity to other repairs. Stage 3
(Determination of supplementary maintenance requirements)
contains the definition of the necessary maintenance program
for the repair.

For the Airbus A300 aircraft Repair Assessment
Guidelines(RAG) were developed which allow the operators
to determine the inspection threshold and interval for the
category B repairs. Fig. 16 shows a principle sketch of an external skin repair. In principle four fatigue sensitive
Fig 16: External skin repair
locations exist which have to be assessed:

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

The determination of the inspection threshold and interval requires the exact knowledge
about the geometry, materials and fastener data to calculate the correct values for
threshold 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 a
Fig 17: Determination of
repair documentation, they may be taken directly from the aircraft. Some of the data may repair parameters
not easil be measured, but NDI methods have to applied. Fig. 17 shows the application
of NDI methods to determine the cut-out size hidden by the repair doubler, the thickness
of skin and doubler and the rivet material.

The inspection interval for the repair is based on the crack size detectable by NDI Fig 18: Inspection of skin
and external repair
means. Fig. 18 contains the NDI procedures for inspection of the skin and the external doubler
repair doubler. All procedures have been qualified and comply with the defined
inspection requirements that the defect size to be detected is determined with a
probability 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. regarding
widespread fatigue damage and initial flaw concept for threshold determination. Furthermore the general aviation
standard with respect to the two-bay-crack criterion should be reached without special design precautions, such as
crack stoppers, and without disadvantages in weight. Additionally the requirements of the airlines regarding
reduction 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 of
new materials. The development and application of new material is still under investigation to reach the optimum of
material and production costs, weight and maintenance costs. During the development and certification of an
aircraft the NDI plays a major role as shown in this paper. Further significant applications of NDI are within the
frame of the aging aircraft activities where the detection of MSD and MED is an important item during the
assessment of the structure susceptible to widespread fatigue damage.
The Repair Assessment Guidelines which were developed by Airbus also rely on NDI for determination of the
repair 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 ATA
NDT FORUM 1995 in Cromwell (Hartford), Connecticut, USA, September 26 - 28, 1995

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