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1 American Institute of Aeronautics and Astronautics
DAMAGE TOLERANT DESIGN AND ANALYSIS OF CURRENT AND FUTURE
AIRCRAFT STRUCTURE
Hans-Jrgen Schmidt
Head of Metal Design Principles and Head of Fatigue and Damage
Tolerance, Airbus, Hamburg, Germany Bianka Schmidt-Brandecker
Metal Design Principles and Fatigue and Damage Tolerance,
Airbus, Hamburg, Germany
Abstract
The primary objective for the aerospace industry is to offer
products that not only meet the operat-ing criteria in terms of
payloads and range but also significantly reduce the direct
operating costs of their customers, the airlines. The struc-ture of
the present civil transport aircraft is de-signed considering the
current and forthcoming airworthiness regulations, the customers
re-quirements and manufacturing aspects. This paper outlines the
implications of the cur-rent airworthiness regulations for fatigue
and damage tolerance (FAR 25.571 Amendment 25-96 and advisory
circular AC25.571-1C), with respect to structural design, analysis
and main-tenance requirements. This includes structure potentially
susceptible to widespread fatigue damage. During the last years
significant improvements have been achieved for fuselage structures
by using new design principles, advanced materials and improved
manufacturing processes. The application of these new technologies
for future fuselage structures requires a new interpretation of the
airworthiness regulations, which were originally defined for
monolithic metallic materi-als and conventionally assembled
structure (e.g. by riveting or bonding). Furthermore the
appli-cation of the new materials and manufacturing processes
requires also further development of the analysis methods to comply
with the regula-tions. Examples of design features using the new
technologies as well as the new aspects of the analysis methods are
presented.
Introduction The continued growth in air traffic has placed an
increasing demand on the aerospace industry to manufacture aircraft
at lower cost, whilst ensur-ing the products are efficient to
operate, friendly to the environment and ensure that the required
level of safety is maintained. Four key airframe drivers are
identified which include the following primary objectives: 1.
Development:
Low weight structure Low non-recurring costs
High performance aircraft Reduced design times
2. Manufacturing Low recurring costs Short flow time Reduced
impact on environment
3. Operation Increased safety and reliability Reduced
inspections and improved repa-
rability Low operating costs Low environmental impact (emissions
and
noise) Increased operational capacity and pas-
senger comfort 4. Disposal
Possibilities of recycling Low environmental impact
To fulfill these targets and to comply with the latest
airworthiness regulations and recommen-dations, the application of
the advanced damage tolerance philosophy, methods and data is
es-sential. The existing analysis and experimental methods as well
as the newest research results have to be taken into account.
Structural criteria and requirements The major structural design
criteria considered during the design development phase are listed
in Table 1. These criteria comprise the basic static strength,
durability and the damage toler-ance aspects, as introduced in 1978
into the regulations as well as the additional major re-quirements
(e.g. discrete source damage, sonic fatigue, wind milling, etc.).
The forthcoming regu-lations must be considered too, which require
a certain structural damage capability (SDC) to provide an
additional design margin to the air-craft. Furthermore other
airworthiness and eco-nomic aspects as corrosion resistance,
repara-bility and inspectability need also to be consid-ered.
Designing for these criteria will provide a structure, which will
meet the certification re-quirements and the customers
expectations. Figure 1 shows in principle the damage types to be
considered during the damage tolerance evaluation. The basic
assumption for all damage tolerance assessments is the local damage
sce-
AIAA/ICAS International Air and Space Symposium and Exposition:
The Next 100 Y14-17 July 2003, Dayton, Ohio
AIAA 2003-2784
Copyright 2003 by the American Institute of Aeronautics and
Astronautics, Inc. All rights reserved.
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nario, i.e. a damage in one or more elements of a principal
structural element (PSE) at a single site, which is not influenced
by damages in adja-cent locations. Furthermore multiple site damage
(MSD) and/or multiple element damage (MED) have to be considered in
structure susceptible to these types of damages. MSD is
characterized by the simultaneous presence of fatigue cracks in the
same structural element and MED occurs simultaneously in similar
adjacent structural elements. MSD or MED may lead to widespread
fatigue damage (WFD), which is reached when
the structure will no longer meet its damage tolerance
requirements, i.e. sufficient residual strength under limit load
condition. The structural damage capability (SDC) will be required
by the forthcoming regulations. It is the characteristic of the
structure which permits it to retain sufficient static load
capability in the presence of damage equivalent to the complete
failure of a load path or partial failure of the load path between
dam-age containment features, i.e. a one- bay-crack-criterion. A
more detailed interpretation of the regulations and requirements is
given by Swift 1.
Table 1: Structural design criteria
Design Criteria Requirements Loads Static strength Undamaged
structure must sustain the loads Ultimate loads Deformation
Deformation of undamaged structure may not inter-
fere with safe operation Limit loads
Durability Damage tolerant structure must meet service life
requirements Safe life components must remain crack free in
service
Operational loads
Residual strength Damaged structure must sustain loads without
catastrophic failure
Limit loads
Crack growth Damage tolerant structure must meet defined
in-spection program
Operational loads
Structural damage capability
Damage tolerant structure must have structural damage
capability
Limit loads
Discrete source damage
Airplane with damaged structure must be able to complete flight
successfully or certain risk level to be meet
Discrete source damage loads get home loads
Sonic fatigue Sonic fatigue cracks leading to catastrophic
failure must be improbable
Operational loads
Further considerations Corrosion resistance, repairability,
inspectability, wind milling, etc.
Figure 1: Damage types (examples) For all locations susceptible
to either local dam-age (LD) or widespread fatigue damage (WFD),
see Figure 2, fatigue and damage tolerance evaluations are
required. These evaluations include the assessment of the fatigue
life (dura-bility), the crack growth between detectable and
critical size and the determination of the residual
strength capability. These results are the neces-sary for the
definition of the structural inspection program.
Current aircraft design and analysis
During the initial design phase of new aircraft types the
application of new materials and pro-duction methods is considered
to reduce the production costs and the structural weight as well as
to comply with the new regulations. The fuselage skins of all
Airbus aircraft certified up to 2001 were made of 2024T3, T42 or
T351. The stringer material was 2024T3 in the upper shell and
7075T73 in the lower shell, which is mainly designed by compression
loads. The first step to apply new materials for the fuse-lage skin
was made for the derivatives of the A340, i.e. for the A340-500 and
600, which are stretched versions of the basic A340-300 and which
have been certified in 2002.
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Figure 2: Local damage versus MSD/MED The dimensioning design
case for the upper fuselage shells is the crack growth behavior
between a damage detectable by general sur-veillance inspection
(walk around, A-, B- or C-check) and the critical crack length
under limit load. To meet the weight target for the A340-500/-600
new materials were selected in many areas, see Figure 3.
Figure 3: Material distribution at Airbus A340-600 fuselage For
the forward and rear fuselage the material 2524T3 has been selected
for the skin in the upper shell, which allows to increase the
allow-able longitudinal skin stresses by approximately 15 percent.
For the side and lower shells the basic 2024 material is kept
except in a small area forward and aft of the center section where
7475T761 was selected due to static reasons. For improvement of the
static strength stringers of high strength material 7349T7 were
selected for the whole fuselage circumference with a few
exceptions. To date pressurized fuselages of commercial transport
airplanes generally consist of a built-up structure where the
skin-to-stringer connection may be riveted or bonded. The other
connec-tions such as skin-clip (shear ties) and clip-frames are
riveted, see Figure 4. The materials used are in general the
aluminum 2000 series (2024, 2524) for all elements. In specific
areas 7000 series alloys (7475, 7075, 7349) are used to increase
the static strength and/or the residual strength. The new
derivative of the Airbus single aisle family, the A318 contains
some panels in
the lower shell where the skin-stringer connec-tion is welded to
reduce the production costs, see Figure 5. Consequently a weldable
material has to be chosen which is 6013 or 6056 for the skin and
6110A or 6056 for the stringers.
Figure 4: Built-up structure
Figure 5: Integral (welded) structure Evaluation of structure
The fatigue and damage tolerance evaluation as required by the
FAR/JAR regulation must be performed by analysis supported by test
evi-dence, i.e. structural tests are performed for certification
purposes to validate analysis meth-ods and design allowables and
finally to proof the structure. Figure 6 shows as an example the
full scale fatigue test of the center fuselage and wing of the
A340-600. Furthermore tests are conducted for development purposes
and to ensure that the in-service airplanes meet or ex-ceed
customers requirements and expectations. Development tests are
accomplished to charac-terize the performance of new materials,
validate new design and manufacturing procedures and
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demonstrate improved durability, safety and maintainability of
the structure. Figure 6: A340-600 full scale fatigue test center
fuselage and wing The analysis of the structure is performed to
justify a sufficient fatigue life of the structure as well as an
adequate damage tolerance behavior, which results in the definition
of an appropriate inspection program. The traditional fatigue life
calculation using the MINER rule is still widely used by the major
manufacturers of civil trans-port aircraft. However, many
investigations have shown that the application of the MINER rule
may lead either to un-conservative results or an under-prediction
of the real fatigue life. There-fore several improvements have been
imple-mented in the fatigue life calculation by the dif-ferent
manufacturers leading to appropriate re-sults. The objective of the
damage tolerance evalua-tion is to provide an inspection program
for each principle structural element (PSE) such that cracking,
initiated by fatigue, accidental damage or corrosion, will not
propagate to catastrophic failure prior to detection. The damage
tolerance analysis consists of fatigue crack growth and residual
strength analysis. The general approach makes use of a basis stress
intensity parameter K, which is a measure of the stress singularity
at the tip of a crack in an infinitely wide panel. This stress
situation is generally characterized by a stress intensity factor.
In addition, correction factors are used for modifying the
influence of the geometry. The crack growth periods are generally
determined using the Forman law. Furthermore a residual strength
analysis is per-formed to determine the critical crack length under
limit loads, which limits the crack growth period for determining
the inspection interval. A new aspect of the damage tolerance
analysis was introduced by the Amendment 25-96 requir-ing the
demonstration that widespread fatigue
damage (WFD) will not occur within the design service goal (DSG)
of the aircraft. There is a general agreement throughout the
literature that MSD and its subsequent phenomenon WFD largely
depend on probabilistic effects. These effects can be derived from
parameters which influence the development of MSD and WFD and which
themselves show a probabilistic char-acter. The major parameters
are the initial de-sign of a structural part, the loading (e.g.
high tension, high induced bending or high load trans-fer), the
manufacturing process, the material properties and to a certain
degree the environ-ment. These parameters obviously have a great
influence on the fatigue life (MSD behavior) of a structure.
Therefore, any approach to assess MSD has to consider the
probabilistic nature of these parameters. In the Airbus approach
this is done by means of a Monte-Carlo simulation. The analysis
model itself consists of two parts, a probabilistic and a
deterministic part. Within the probabilistic algo-rithm the initial
damage scenario is determined, while the subsequent steps, such as
damage accumulation, crack growth and residual strength are
calculated in a deterministic ap-proach. The process is performed
for a pre-defined number of simulations. The AAWG report 2 has
defined the general evaluation process for structure susceptible to
WFD for monolithic aluminum. It is recom-mended to commence the
so-called WFD inspections at 33 percent of the average time to WFD
occurrence. Considering the limited reliability of these
inspections to find small multiple cracks particularly in hidden
areas, it is required to modify, retire or repair the structure at
50 percent of the average time to WFD oc-currence. The threshold
for WFD inspections is defined as Inspection Start Point (ISP) and
the time to repair as Structure Modification Point (SMP).
Figure 7: In-service actions for structure susceptible to WFD
The results of the WFD analysis have to be as-sessed regarding the
repercussions on the aging fleet. An example for service actions as
the re-
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sult of the WFD analysis is given in Figure 7. This example
shows typical values that can be expected for monitoring periods in
fuselage type structure.
Advanced technologies and materials The aircraft industry, as
one of the most innova-tive industries, is always obliged to
introduce new materials and technologies. The aim of this
introduction is the reduction of the manufacturing costs, the
aircraft weight and the direct operating costs (DOCs) as well as
the compliance with the more stringent future airworthiness
regulations. An additional challenge exists for the develop-ment of
very large transport aircraft, e.g. Airbus A380. In theory, when
the size of an aircraft is increased by a certain factor, its
volume and its weight increase with the factor to the third power.
This exponential increase means that weight problems of very large
transport aircraft are quite significant. By improving the
configura-tion of these aircraft types, the effect of this law can
be reduced. Furthermore new materials and technologies play a major
role for very large aircraft. The following chapter describes key
technolo-gies to achieve the goals mentioned above, and their
application to date and/or in future. Figure 8 shows the
distribution of the skin material at the Airbus A380-800.
Figure 8: Material distribution at Airbus A380 Fiber metal
laminate GLARE Fiber metal laminates (FML) were developed at Delft
University of Technology as a family of new hybrid materials
consisting of bonded thin metal sheets and fiber/adhesive layers.
The laminated structure provides materials with excellent fa-tigue,
impact and damage tolerance characteris-tics at low density. The
trademarks are ARALL and GLARE. The prepregs act as barriers
against corrosion and the laminate has an inher-ent high
burn-through resistance as well as good damping and insulation
properties. GLARE provides an attractive weight saving potential of
approximately 10 to 20 percent for
fuselage panels dimensioned by damage toler-ance behavior. The
material provides several improvements such as low density, high
durabil-ity, slow crack growth, high residual strength, high
corrosion resistance and high fire resis-tance. GLARE is a hybrid
material built-up from alter-nating layers of aluminum sheets
(thickness between 0.2 and 0.5 mm, mainly made from 2024T3) and
glass fiber reinforced adhesive unidirectional layers
(FM94-S2-Glass, thickness 0.125 mm). Figure 9 shows the general
defini-tion of GLARE and Table 2 contains the eight standard GLARE
types.
Figure 9: Definition of GLARE Table 2: GLARE types
Standard GLARE
types
Fiber adhe-sive layer (mm)
Fiber/ adhesive
layer build-up
Al alloy
GLARE 1 0.25 0/0 7475T761
GLARE 2A 0.25 0/0 2024T3
GLARE 2B 0.25 90/90 2024T3
GLARE 3 0.25 0/90 2024T3
GLARE 4A 0.375 0/90/0 2024T3
GLARE 4B 0.375 90/0/90 2024T3
GLARE 5 0.50 0/90/90/0 2024T3
GLARE 6 0.25 +45/-45 2024T3 GLARE offers an excellent crack
growth behav-ior for both crack types, i.e. for the so-called
through cracks and part-through cracks. This superior behavior is
the result of the presence of fibers in the laminate, which do not
fail due to fatigue. This enables load transfer over the crack
through the fibers, thus reducing the crack tip opening, the stress
intensity factor and finally the crack growth rate. Figure 10 shows
the crack
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bridging of the fibers and the resulting effect on the crack
growth curves. The GLARE 2 type specimen was loaded in fiber
direction, GLARE 3 includes fibers in both directions, par-allel
and perpendicular to the load direction. Due to less fiber content
perpendicular to the crack GLARE 3 shows a slightly worse crack
growth behavior compared to GLARE 2.
Figure 10: Crack growth behavior in GLARE Most current
regulations and advisory circulars were established when the
aircraft structure was made of monolithic aluminum. Therefore the
present interpretation of the damage tolerance requirements has to
be adapted to the specific characteristics of GLARE material
without changing the overall goals regarding a safe op-eration up
to the end of the service life. Table 3 contains the comparison of
the characteristics between the conventional aluminum and the GLARE
structure. Summarizing Table 3 the GLARE material provides a short
crack initia-tion time, but superior crack growth behavior and
excellent residual strength properties in case of fatigue cracks,
i.e. when the fibers are intact. Figure 11 shows the application of
the damage tolerance philosophy for structure made of GLARE and
monolithic aluminum and suscep-tible to WFD. The curves Al and Gl
indicate the typical crack growth and residual strength be-havior
of monolithic and GLARE structure. As explained above the
Inspection Start Point (ISP) and the Structural Modification Point
(SMP) are defined by applying factors 3 and 2, respectively, on the
WFD average behavior. For the inspection interval a factor 2 is
used on the
crack growth period between detectable and critical MSD damage:
ISPAl = NWFD Al / 3 ISPGl = NWFD Gl / jISP Gl SMPAl = NWFD Al / 2
SMPGl = NWFD Gl / jSMP Gl IWFD Al = NWFD Al / 2 IWFD Gl = NWFD Gl /
jI Gl
Figure 11: Damage tolerance philosophy for WFD in Al and GLARE
structure The scatter factors jISP Gl, jSMP Gl and the crack growth
factor jI Gl will be defined by relevant re-search programs.
However, the probability of failure at SMPGL should not exceed the
probabil-ity of failure at SMPAL, i.e. approximately 510-2. Since
fatigue initiation affects mainly the alumi-num layers in GLARE,
the fatigue initiation process is similar to that of monolithic
aluminum. Therefore a similar stress level in the aluminum will
lead to the same time to crack initiation. The fatigue initiation
in GLARE is calculated in the same way as for monolithic aluminum,
i.e. using the actual stresses in the aluminum layer at the
critical location. The actual stresses in the alu-minum layers in
GLARE consist of stresses due to the curing process, stresses due
to external loads and stresses due to temperature deviating from
the ambient conditions. The actual stresses in the aluminum layers
due to external loads are affected by the different stiffness of
the GLARE components. Due to the lower stiffness of the fibers, the
stresses in the aluminum layers will therefore be higher than the
applied stresses. The total stresses in the aluminum layers are
obtained by superposition of the curing stresses, the stresses due
to external load and the stresses from operational temperatures
(not described here), see Figure 12. The total stress and the
relevant SN curve allow to estimate the fatigue initiation life in
the aluminum layers.
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Table 3: Monolithic Aluminum structure versus GLARE
structure
Aluminum structure GLARE structure
Fatigue and damage tolerance
Long crack free life Moderate growth of short and small cracks
Significant growth of long cracks Significant reduction of residual
strength in the presence of multiple site damage (MSD) Rapid
reduction of residual strength with in-creasing fatigue crack
length Significant reduction of residual strength for the so-called
two bay crack
Shorter crack free life Faster growth of short and small cracks
Slow growth of long cracks Small reduction of residual strength in
the presence of multiple site damage (MSD) Small reduction of
residual strength with in-creasing fatigue crack length (intact
fibers) Similar reduction of residual strength for the so-called
two bay crack caused by foreign object damage
Corrosion (heavy corrosion assumed)
Significant strength reduction Possibility of crack initiation
followed by signifi-cant crack growth (through the thickness
crack-ing)
Limited strength reduction (corrosion is limited to surface
layer) Shorter crack initiation time followed by slow crack growth
in the surface layer (part- through cracking)
Figure 12: Stress cycle in Aluminum layers at
room temperature (example) Since fatigue crack growth occurs in
the Alumi-num layers only, metal methods, i.e. linear frac-ture
fracture mechanics may be used to deter-mine the crack growth
behavior. Different crack cases require different stress intensity
solutions. Surface cracks, for example, may be analyzed according
to Homann 3 using the Paris equation and the following stress
intensity solution:
m2max
nLTmax tF
L24K
=
with: L = material constant = correction factor for loading
direction = correction factor for the number of Al
layers F = finite width correction tm = metal layer
thickness
Figure 13 illustrates the crack growth behavior of a
part-through crack. The crack starts in the surface layer from the
notch. Then cracks are initiated in the subsequent layers.
Figure 13: Part-through crack in GLARE Since GLARE has a low
crack initiation life, early cracking is expected during full scale
fa-tigue test. Consequently future aircraft with GLARE structure
will fly with small undetected cracks in the Al layers of the GLARE
up to the end of the service life. In contrast to monolithic
aluminum structure these cracks are acceptable due to the superior
crack growth and residual strength behavior of the GLARE material.
The requirement of the airworthiness authorities about flyable
crack length allows the operation of an aircraft with known cracks
only, if ultimate load capability exists up to repair. This
philoso-phy is to be applied also to GLARE, i.e. ulti-mate load
capability must exist at the end of the full scale fatigue test,
minimum after demonstra-
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tion of two life times. The procedure shown in Figure 14 has to
be applied.
Figure 14: Procedure for justification of GLARE Laser beam
welding Laser beam welding (LBW) is one of the most promising
welding technologies for aerospace application. The major
motivation of the applica-tion of LBW is the reduction of the
production costs and a slight weight reduction. The LBW technology
is most suitable for welding of T-joints, e.g. skin-to-stringer or
skin-to-clip joints. Weldable aluminium alloys such as 6013 and
6056 have to be used for the time being. Figure 15 shows an Airbus
LBW pilot plant and the LBW tool.
Figure 15: Overview of LBW technology One of the first
applications of LBW on primary structure of a commercial transport
airplane are
the lower and side shells of the Airbus A318 using 6013 and 6056
for skin to stringer welding. Furthermore lower and side shells of
the A380-800 will be welded (skin-stringer joint). However, to date
an application of the welded structure in all areas of the
pressurized fuselage is not ap-propriate due to the limited
residual strength capability of the integral structure. In the
welded areas of the A318 the operational tension stresses (in
stringer direction) are rather low, since the lower and side shells
are dimensioned mainly by compression. Figure 16 shows the fatigue
behavior or the welded structure transverse to the weld line. The
welded joint shows fatigue lives comparable to a Kt = 3.6 specimen.
The actual aircraft stress level is significantly below these SN-
curves.
Figure 16: Fatigue behavior (transverse) of laser
beam welded skin stringer joint The crack growth of longitudinal
cracks in the weld line is shown in Figure 17.
Figure 17: Behavior of cracks in weld line
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If the crack turns into the base material, the be-havior is the
same as for the base material. If the crack remains in the weld
line, the crack growth is faster for stress intensity factors of K
> 28 MPam. Friction stir welding The second promising welding
technology is the friction stir welding (FSW), which is based on
patents developed by the The Welding Institute (TWI) in UK. The
process consists of a rotating tool producing frictional heat so
that plasticized material in kneaded under pressure and there-fore
leading to a tight connection of the sheets. FSW allows joining of
non weldable alloys, e.g. 2000 and 7000 series aluminum alloys.
Fur-thermore different materials may be joined, e.g. different Al
alloys. For series production FSW is today applied in non-aircraft
industry. Examples for application are ship and train manufacturing
as well as aerospace industry (rocket produc-tion). In the aircraft
industry first applications of FSW are envisaged for fuselage
longitudinal joints, wing spanwise joints, wing spars made of
dissimilar alloys and extruded panels, e.g. in center wing box.
Figure 18 shows the excellent fatigue behavior of FSW joints
compared to a riveted joint. The lap joint shown in this figure is
an optimized riv-eted joint with additional doublers in the rivet
area and three rivet rows.
Figure 18: Fatigue behavior of FSW joints Figure 19 contains the
allowable stresses for a three-rivet-row lap joint (same as in
Figure 18) and a FSW joint compared with the behavior of the
baseline material. The allowable maximum fatigue stress (far field
stress) is 54 percent lower for the riveted lap joint compared to
the FSW joint. These figures are valid for specimens with a mean
fatigue life of 250 000 cycles. The application of FSW to joints
instead of rivet-ing offers several advantages:
Reduction of fasteners with - reduced manufacturing costs -
deletion of sealing (less weight, less costs) - no fatigue cracking
initiated at fastener
holes (no MSD) Material utilization by
- reduced by to fly ratio Optimization of performance by
- welding of non weldable alloys and dis-similar alloys
Process automation
Figure 19: Allowable stresses for riveted and
FSW joints On the other hand the FSW process causes additional
features, which need to be consid-ered, e.g. residual stresses
generated by the contraction of the cooling weld nugget which is
impeded by the material on both sides of the weld. These residual
stresses influence both, fatigue and crack growth performances.
Figure 20: Crack growth analysis of FSW joints They depend from
the size and process parame-ters. Dalle Donne and Raimbeaux 4
proposed a fracture mechanics approach based on a crack closure
model with the superposition of external load and internal stresses
(Krs), which can be used to predict the crack growth rate. The
crack opening stress (Kopen) is calculated from empiri-cal
relationships. The da/dn Keff approach
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results in the suppression of the R-ratio effect and the
residual strength effect, see Figure 20. Structural health
monitoring The primary objective for the aerospace industry is to
offer products that not only meet the operat-ing criteria in terms
of payloads and range but also significantly reduce the direct
operating costs of their customers, the airlines. Advanced
structural health monitoring systems may signifi-cantly support
these goals. Table 4 gives an overview of the repercussions of a
health moni-toring system on the structural behavior. Table 4:
Benefits for structural health monitoring Structural criterion
Repercussions Static strength No improvement possibleFatigue
strength (durability)
No improvement possible
Airworthiness Improvements possible, but current structure meets
airworthiness re-quirements
Crack growth periods
Improvements in case of longer cracks due to modified crack
scenarios
Structural damage capability
Improvements in case of fatigue cracks due to modified crack
scenarios,no improvements possi-ble for impact damage due to
accidental dam-age scenario
Several fields of application of structural health monitoring
(SHM) systems are under investiga-tion: Application in laboratory
and full scale tests Monitoring of specific areas of in-service
aircraft Consideration of SHM during the design
phase of new aircraft These applications are briefly discussed
in the following. Application of SHM systems in test specimens will
mainly be performed to gain ex-perience with such systems. The
condition of the structure is well known due to extensive
inspec-tions of the specimens, therefore the SHM re-sults may be
verified. Furthermore the use is possible in a short term approach,
since no qualification process is necessary. In flying aircraft,
there are known hot-spot areas which are sensitive to fatigue
and/or stress cor-rosion or corrosion fatigue problems. A
suitable
SHM system could be installed to monitor these areas. The SHM
application can be very benefi-cial, especially for structural
locations which are difficult to inspect using conventional
inspection methods and/or where access to the structure location is
difficult. The major benefit from SHM systems may be gained, if
considered during the design of new aircraft. As one of the first
possible applications the monitoring of internal stiffeners in wing
or fuselage panels is investigated. The effects of a health
monitoring system on the inspection re-quirements this type of
airframe structure is de-scribed in Figure 21 showing an aircraft
wing or fuselage skin stiffened by stringers. In many cases the
conventional inspection system does not require internal
inspections of the stringers. For these cases it is assumed that
the stringer contains the so-called primary flaw and the skin the
secondary flaw (shorter than the primary). The stringer fails after
a certain number of flights, then the loads are redistributed into
the skin which increases the crack growth rate in the skin. The
inspection interval is based on the crack growth period between the
detectable and the critical crack length in the skin divided by an
appropriate scatter factor. In case of health monitoring of the
stringer a failure of the stringer has not to be assumed (i.e. the
stringer is intact), which reduces the crack growth rate in the
skin significantly.
Figure 21: Effect of SHM on inspections The benefits due to
health monitoring are dis-cussed in Figure 22. One of the major
parame-ters determining the inspection interval is the operational
stress in the structure. The figure shows in principle the interval
versus a reference value of the operational stress (e.g. the
once-per-flight stress) for a structural element for the
conventional inspection system and a monitored structure. Benefit
can be taken of the structural health monitoring: Firstly the
stress level is kept constant. Consequently the inspection interval
may be increased which would lead to a reduc-
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11 American Institute of Aeronautics and Astronautics
tion of the maintenance costs. This results in relatively small
savings for the operators, since also the new inspection intervals
have to be fitted into the scheduled maintenance program, which
depends mainly on the requirements for corrosion inspections and
systems. Secondly a constant inspection interval suitable for the
op-erators is assumed, which allows an increase of the operational
stresses for monitored struc-tures. Increased allowable operational
stresses lead to a reduction of the structural weight in those
aircraft areas, which are dimensioned by crack growth. The overall
weight saving for the aircraft is significantly higher than the
weight saving in the monitored areas due to the so-called snowball
effect. This leads to significant reductions of production costs as
well as main-tenance costs, which improves the efficiency for both
the manufacturers and the operators.
Figure 22: Design and maintenance benefits of
SHM When applying health monitoring systems one specific aspect
has to be taken into account. It has to be assumed that is not
feasible for the operator to repair the structure immediately after
detection of damage by the health monitoring system. An immediate
grounding of the aircraft would lead to significant costs, which
may not be balanced by the benefits gained from the sys-tem.
Therefore in case of a finding operation has to be continued with a
known crack for a certain time. According to the regulations the
structure must be able to sustain the design ultimate loads in case
of a known crack, i.e. the structure should have the same
capability as an intact structure. Therefore the time to repair has
to be based on the crack growth period between the detected crack
length and the critical crack length under ultimate load. It should
be the goal that this period divided by an appropriate scatter
factor is at least one so-called C-check interval, which is usually
a 12 to 18 month period of op-eration. This requirement has to be
taken into account during the design phase, i.e. during definition
of the allowable operational stress.
Conclusions This paper summarizes the major structural cri-teria
and requirements as well as analysis as-pects to be considered
during development, design, certification and operation of civil
trans-port aircraft. During the past few years the de-velopment of
modern transport aircraft has made several important improvements
to cope with the increased expectations of the customers. Ad-vanced
materials and technologies allow signifi-cant reductions in
aircraft weight, production costs and operating costs. These new
technolo-gies and materials are partly introduced in the new Airbus
aircraft A318 and A380. Further ad-vanced developments are planned
for future application. The current and forthcoming certifi-cation
requirements are fully applied to the cur-rent and the advanced
structures.
References 1. Swift, T., Fail-safe design requirements and
features, regulatory requirements. Presented at the
International Air & Space Symposium and Exhibition The Next 100
Years, Day-ton, USA, 2003
2. N.N., Recommendations for Regulatory Ac-
tion to Prevent Widespread Fatigue Damage in the Commercial
Airplane Fleet, Airworthi-ness Assurance Working Group Task
Planning Group, final report, revision A, June 1999
3. Homan, J., Damage Tolerance Analysis in
Glare. TU Delft, Faculty of Aerospace Engi-neering. Presented
during an internal Airbus meeting, 2002
4. Dalle Donne, C. and Raimbeaux, G., Resid-
ual stress effects on fatigue crack propaga-tion in friction
stir welds. German Aerospace Center, Institute for material
Research, Co-logne, Germany.
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