Damage tolerance capability
T. Swift
Federal Aviation Administration, Long Beach, CA, USA
Dedication
This paper is dedicated to my friend and colleague Professor Dr Jaap Schijve, required to retire like
many good aircraft components according to a 'safe life' criterion but who is still young in mind,
body and spirit, has not yet reached his unfactored endurance limit, and who could continue subject
to a needed change in philosophy to 'retirement for cause'. Professor Schijve has dedicated his life
to aviation safety through his outstanding continuous research into aircraft fatigue and fracture
phenomena. The Federal Aviation Administration, whose primary goal is aviation safety, is indebted
to Professor Schijve for his continous counsel through many contributions to the literature in his
subject.
The purpose of this paper is to encourage manufacturers of future transport aircraft to retain the
large damage-tolerance capability designed into the first wide-bodied aircraft and to modify their
methodology to establish inspection thresholds for those structures incapable of sustaining large
obviously detectable damage. In the current economic environment there appears to be a general
trend to lower the level of safety built into the original wide-bodied aircraft to reduce assembly costs
and weight. A number of examples are provided highlighting the implications of this general trend.
Most of the wide-bodied aircraft are designed to sustain a two-bay skin crack with a broken central
stiffener at limit load. Some manufacturers would prefer to relax this criterion for wide expanses of
basic structure with a view to saving weight. The implications of this are that extremely sophisticated
NDI will be required over wide areas to satisfy the damage tolerance requirements, thus creating a
considerable burden on the operator. This is illustrated by example. Most of the large transport
aircraft manufacturers establish the threshold for detailed inspection of principal structural elements
through a fatigue evaluations process without considerations of initial manufacturing flaws. These
thresholds are often as long as three quarters of the aircraft design life goal. This practice has been
thought to be satisfactory for fail-safe crack arrest structure. In this case a second line of defence
exists in the event that an initial manufacturing flaw nucleates into a propagating fatigue crack during
the service life of the aircraft. Since crack arrest structure is usually capable of sustaining large
obvious damage it is likely that such large damage would be detectable. However, principal structural
elements do exist that are incapable of sustaining large obvious damage where initial manufacturing
flaws could grow critical prior to the threshold established by fatigue evaluation. This is illustrated
by a typical example. Structures operating beyond the life substantiated by full-scale testing may be
prone to multiple-site damage (MSD). Extremely small in-service undetectable MSD has the potential
for substantially reducing lead crack and discrete source residual strength. This is illustrated by typical
examples. The majority of large transport aircraft developed in the USA have circumferential crackstopper straps attached directly to the fuselage skins to guard against explosive decompression failure
in the event of undetected fatigue damage or discrete source damage. This was thought necessary
after the Comet disasters in 1954. There is a current trend to eliminate these crack stoppers for future
designs and depend only on shear clips for crack arrest capability to save on assembly costs. The
wisdom of this trend is challenged by considering examples and citing a number of secondary effects
that may prevent arrest of fast fracture.
(Keywords: damage toJerl~ce; multiple.site damage; inspection thresholds)
The introduction of wide-bodied commercial transport
aircraft in the late 1960s initiated a fresh look at
damage tolerance capability. Although crack-arresting
concepts had been used on the DC-8 pressurized
fuselage design, which started in 1954 following the
Comet disasters the same year, a disciplined fracture
mechanics analytical approach had not been used until
the development of the DC-10 design, which started
in 1967. Damage tolerance capability, using fracture
mechanics principles, was designed into this aircraft
0142-1123/94/01/0~5-20
© 1994Butterworth-HeinemannLtd
at the outset. Parametric studies were performed using
finite element analysis methods to size crack-arresting
members for the fuselage pressurized cabin. 1,2 The
same methods were used to study crack-arresting
features of the wing design.
For several very good reasons, which will be
discussed in detail later, a goal was established for the
limit load residual strength capability of the basic
structure. This goal included the ability to sustain two
bays of skin cracking with a central broken member. A
Fatigue 1994 Volume 16 Number 1
75
Damage tolerance capability: T. Swift
comprehensive verification component test programme
was conducted to substantiate the residual strength
capability of the structure. Fatigue damage equivalent
to at least two representative design lifetimes was
applied prior to residual strength testing to simulate
the possible presence of widespread fatigue cracking.
A full-scale fatigue test was conducted to two design
lifetimes under representative flight-by-flight spectrum
loading. A structural inspection programme was
developed for this aircraft using the Maintenance
Steering Group 2 (MSG2) approach support by
information on crack propagation rates obtained from
component testing.
The design philosophy in place at the time that the
wide-bodied aircraft were certified included a choice
between fail-safe and safe life. The fail-safe concept
was chosen by the manufacturers of wide-bodied
aircraft. At that time this was interpreted by the
manufacturers to mean that the structure should be
capable of sustaining complete failure or obvious
partial failure of a single principal structural element
at fail-safe load levels.
The wide-bodied commercial transport aircraft were
already in service when the US Air Force introduced
damage tolerance requirements for all military aircraft
in July 1974. The Air Force design philosophy for 20
years prior to this time has been based on a reliability
approach, where safe lives were established and
substantiated by four lifetime full-scale fatigue tests.
The primary lesson learned by the Air Force during
this period was that the safe life approach did not
adequately account for possible initial manufacturing
flaws that may exist in the airframe at delivery. Studies
performed by the Air Force in the early 1970s indicated
that over 50% of fatigue failures nucleated at initial
manufacturing flaws induced during manufacture. This
prompted the Air Force to abandon the safe-life
design approach and convert to a damage tolerance
philosophy.
In the commercial transport arena the fail-safe
concept was thought to be the complete solution
to structural fatigue problems. However, the Civil
Aviation Authority in the United Kingdom were
concerned about loss of fail-safety with age. Their
fears were substantiated by the loss of an A V R O 748
in Argentina on 14 April 1976 due to multiple-site
fatigue damage in the wing. The 748 had been AVRO's
first aircraft designed to fail-safe principles. Soon after
this, a Dan-Air 707 aircraft lost a fail-safe horizontal
stabilizer at Lusaka International Airport in Zambia
on 14 May 1977, because of fatigue. At this point the
commercial transport industry lost faith in the fail-safe
philosophy and introduced damage tolerance principles
by amending FAR 25.571 in December 1978. 3 Since
that time all commercial transport aircraft have been
designed to a damage tolerance philosophy and the
existing ageing fleet have been assessed to the same
principle. The damage tolerance philosophy presumes
that any damage initiated by fatigue, accident or
corrosion will be found before catastrophic failure.
Thus an engineering evaluation considering crack
propagation rates and residual strength limits is made
by the manufacturer and inspections are carried out
by the operator.
Again, in the commercial transport arena the damage
tolerance concept was thought to be the complete
76
Fatigue 1994 Volume 16 Number 1
solution to structural fatigue problems. However, on
28 April 1988 a 737 aircraft operated by Aloha Airlines
suffered a tragic fatal accident due to a pressurized
fuselage fatigue failure. Undetected multisite fatigue
cracking had occurred in the critical rivet row of a
longitudinal skin splice causing skin crack coalescence
resulting in unarrested fast fracture. Now, in the
aftermath of this accident, it becomes apparent that
even the damage tolerance philosophy in itself may
not be completely adequate.
We are now faced with the dilemma that all three
design philosophies - safe life, fail-safe and damage
tolerance - have been shown to be inadequate in
themselves. In fact we need elements of all three
philosophies. We need to build as much redundancy
into the structure as is economically feasible after the
old fail-safe philosophy. We need to establish the life
at the onset of widespread fatigue damage for those
elements prone to multiple-site damage (MSD) and
modify or replace these elements for flight beyond
this point. In essence this is a safe-life approach. We
also need to establish inspections based on crack
growth rates and residual strength limits following the
damage tolerance philosophy.
It is the opinion of this author that the DC-IO widebody structural design incorporates elements of all
three design philosophies. This was the reason for
earlier comments on this aircraft. The basic structure
is redundant, is crack-arrest capable and incorporates
external inspection features in splices. However, as
time passes the reasons for many of these design
features fade. Engineers working on new aircraft are
striving for more cost-effective designs. There appears
to be a strong tendency to sacrifice large damage
capability for reduced cost.
Reflecting on this current trend, and on the past
decade of damage tolerance design, there are a number
of issues that need restating and some that need
initiating. This paper will address some of these issues
outlined as follows.
1. The importance of the two-bay crack design criterion
needs restating.
2. The threshold for detailed inspection of fatiguecritical elements needs close examination, especially
for elements that do not have crack-arrest capability.
3. The effects of multisite damage on residual strength
and discrete source damage capability need to be
addressed, especially for aircraft operating beyond
half their test life.
4. The current trend to eliminate fuselage crack
stoppers should be reconsidered very carefully.
THE TWO-BAY CRACK CRITERION
During the development of wide-bodied aircraft in the
late 1960s considerable effort was expended in design,
analysis and component testing to support large damage
capability, particularly in the pressure cabin. Radial
loading due to cabin pressure, a function of shell
radius, was much greater than for the narrow bodies.
The Comet accidents, caused by pressure cabin fatigue
failures due to lack of crack-arrest capability, w e r e
still fresh in the minds of designers. This author's
experience with the wide bodies was confined to the
DC-IO, so most of the following discussion applies to
Damage tolerance capability: T. Swift
this aircraft. However, the manufacturers of each of
the aircraft were looking very carefully at each other's
designs. Airline engineers and marketing personnel
were also looking closely at each design and commenting on the merits of each.
Studies of in-service cracking problems and considerable parametric analyses were performed in the early
design phases of the aircraft. Critchlow's methods
were used initially to size structural members for
residual strength damage capability. It soon became
apparent that the pressure cabin should be designed
to sustain the damage illustrated by Figure 1. For
longitudinal cracks, propagated by fatigue, it was
decided to consider a two-bay skin crack with a broken
central crack stopper at limit load. This large damage
capability was thought necessary in case fast fracture
occurred from a shorter crack that might have been
missed on inspection. The reason for also considering
a broken central crack stopper was that flat-panel
cyclic testing had indicated that in the event of a skin
crack over a crack stopper a considerable amount of
hoop loading was transferred to the crack stopper,
creating a high cyclic load and eventual fatigue failure
even when the skin crack was still small. In addition
to this criterion, a two-bay longitudinal skin crack was
considered with a broken central frame and crack
stopper at fuselage bending loads equivalent to 1.5 g
plus cabin pressure. This damage was to simulate fast
fracture after discrete source or foreign object damage
as in the case of an engine disintegration. A two-bay
circumferential skin crack with a broken central
stiffener was also considered at limit load.
Figure 2 shows that two-bay damage is logical for
any skin cracking that starts at a circumferential frame.
The cut-out in the frame-to-skin shear clip creates a
considerable stress concentration at the first fastener,
as illustrated in Figure 2. Skin stresses due to frame
bending in certain positions around the frame added
to the stresses due to direct pressure loads, creating
a fatigue hot spot in the skin at the first rivet, as
shown. This crack is likely to propagate into both
adjacent skin bays. Figure 2 also shows a fatiguecritical location at the joint between the fuselage axial
longerons and the circumferential frame. On the
application of internal cabin pressure the skin and
axial stiffeners move outwards in a radial direction.
This radial displacement is resisted by the frame,
causing transfer of radial load from the skin through
the longeron into the frame. The magnitude of this
11rio BAY ~
,B~I O R / O (
WITH CENT1RALBR(XqB~I ORACK I ~ q = ~ R
AT MMIT LOAD
-
-
'
-
•
~
1 ~ O BAY L(2N~IT~K)IN~ gNN ( ~ ( X
PLUB m ~ t ~ d CIWI~RRL C ~ C K IWOIq~R
,~ND FRAME AT 1J ~ PI.U~ IIRES~JRE
Figure 1
F u s e l a g e d a m a g e sizes f o r d e s i g n
1WO BAY
Cl~d~( WWH M q O l ~ l t ~ T I I ~ L
u [ ~ r w l E R AT LIMIT LO~D
load varies with axial load from fuselage bending due
to Poisson's ratio effects but is generally about 200 lbs
(900 N) at nominal cabin pressure. This tension load
causes local high bending stresses in the longeron
flanges and can create cracking as indicated in Figure
2. When the longeron breaks, the skin becomes
overloaded at the first rivet near the break and
eventually causes a skin fatigue crack, which can
propagate in a circumferential direction in both adjacent bays. Thus the two-bay skin-crack scenario makes
sense from a practical viewpoint.
The discrete source damage case for the requirement
of a full two-bay longitudinal skin crack with completely
failed circumferential members has received some
argument from a number of manufacturers. They point
out that harpoon testing to simulate damage from one
third of an engine disc segment does not normally
create a fast fracture situation in 2024-T3 fuselage skin
owing to the high fracture toughness of the material.
Therefore it should not be necessary to consider two
full bays of skin damage for the case.
This is illustrated by Figure 3. The curves shown in
Figure 3 were obtained by finite element analysis
methods described in Ref. 2. The case considered is
for a two-bay skin crack with both centre frame and
crack stopper failed, which is given as case 5 in Table
1, taken from Ref. 2. The skin fracture curve of Figure
3 is obtained from the following equation:
Kc
Residual strength, crR - (~a)l/2 [3S13B
(1)
Fracture toughness for the 2024-T3 sheet was assumed
to be 158 ksi* inj (174 MPa ml). The geometrical
term 13s is the reciprocal of the value Rct listed in
Table 2 of Ref. 2 for case 5. The term 138 is a
geometric effect caused by bulging due to pressure
and shell radius. As all the stiffening material at the
centre of the crack is assumed to have failed, the
bulging effect has been assumed to behave like a onebay crack with the bay equal to two frame spacings.
The term ~B used to develop Figure 3 was based on
Paul Kuhn's unstiffened shell data s together with a
cosine function suggested by Prof. Dr Liider Schwarmann. 6 The resulting term is
5(2a)
Bulge factor, [3a = 1 + R[cos(~ra/P)]
(2)
where: a = half crack length; R = shell radius; and
P = frame spacing, in this case 2 x 20 = 40 in (2 x
0.508 -- 1.016 m).
The outer crack stopper strength curve is determined
by dividing the ultimate strength of the 0.025 in
(0.64 mm) thick crack stopper by the stress concentration factor, GroJ~r, found in Table 2 of Ref. 2 for
case 5. The crack-stopper material is assumed to be
Ti 8-1-1 with ultimate strength 145 ksi (1000 MPa).
The line in Figure 3 intersecting the skin fracture
curve at points B and C represents the principal stress
calculated from 82% of PR/t combined with shear
representing 1.5 g of fuselage down bending. The
harpoon blade shown in Figure 3 is assumed to be 15
in (380 ram) wide and considered representative of
* ksi = 1000 lb in - 2 = 6 . 8 9 5 M P a ; 1 in = 2 5 . 4 r a m ; 1 ksi in ~ =
1.989 MPa m ~
Fatigue 1994 Volume 16 Number 1 77
Damage tolerance capability: T. Swift
SKIN CRACKAT RRST RIVET
NEAR LONGEFION CRACK
PROPAGATES INTO TWO BAYS
N
R
R
LONQERONCRACK
FUSELAGE CIRCUMFERENTIAL
FRAME
,~,,,,,I CRACKAT FIRST RIVET
IN SHEAR ClIP
PROPACaATES INTO "rwo BAYS
Figure 2 Typical fatigue crack locations in fuselage skin
j
40
0.2
0.~1
0.6
Harpoon blade
0.8
1.0
1.2
250
30
200
I
O~
r"
ta
150 a.
:S
20
Frame ' ~
m
t~
100
"O
t~
Crack
stopper
10
50
Circumferential frame
configuration
10
20
30
~0
50
Total crack length 20 (in)
Figure 3 Discrete source damage capability
one third of an engine disc. It can be seen from Figure
3 that the longitudinal damage created by this blade
would not cause fast fracture, as point A is to the left
of the skin fracture curve. This is the reason that
some manufacturers do not consider it essential to
consider two full skin bays of damage.
However, damage to the fusealage may not be
inflicted by the disc segment alone. In order for the
disc to leave the engine it would be necessary to fail
78
Fatigue 1994 Volume 16 Number 1
the engine case. Smaller, fragmentation damage is
likely in addition to disc damage, as indicated in Figure
3. In this event, fast fracture would take place at point
B and the crack would be arrested at point C for the
geometric condition considered.
The high residual strength capability from a skin
fracture standpoint, depicted by point E of Figure 3,
is due primarily to the 3 in (76 mm) wide 0.025 in
(0.64 mm) thick titanium crack stopper riveted to the
Damage tolerance capability: Y. Swift
Table 1 Test results for 24 in (609.6 mm) diameter cylinders 2 (SI
conversions in parentheses)
Crack length at
failure, 2a
(in) (mm)
Hoop stress,
(psi) (MPa)
Shear stress, Principal stress"
T/2At
(psi) (MPa)
(psi) (MPa)
4.00 (102)
4.50 (114)
6.44 (169)
8.50 (216)
4.50 (114)
6.88 (175)
8.44 (214)
12 750 (87.9)
12 880 (88.8)
8 630 (59.5)
6 030 (41.6)
10 250 (70.7)
7 125 (49.1)
5 625 (38.8)
0
(0)
0 (0)
0 (0)
0 (0)
7 240 (49.9)
5 085 (35.1)
4 020 (27.7)
PRIt
1 275
1 288
863
603
1 399
977
771
(8.79)
(8.88)
(5.95)
(4.16)
(9.65)
(6.74)
(5.32)
= Axial stress not included
stress due to fuselage bending. Figure 4, representing
residual strength tests on a n u m b e r of 24 in (0.61 m)
diameter unstiffened cylinders, shows that skin shear
has an effect on residual strength and should be
accounted for. This series of tests was reported over
20 years ago in Ref. 2. The u p p e r curve of Figure 4
is a plot of hoop stress versus crack length with applied
torque providing skin shear. The values of crack
length, hoop stress and shear stress at failure for these
tests are given in Table 1 as a reminder. Full details
of this test p r o g r a m m e can be found in Ref. 2. It is
believed that any residual strength tests, harpoon or
otherwise, p e r f o r m e d to substantiate damage tolerance
capability should include loading from both cabin
pressure and fuselage bending. This is illustrated by
Figure 5. It appears that there is a need to develop
such a fixture to assess the effects of crack propagation
Table 2 Residual strength calculation for wing lower surface configuration (Figure 8) (SI conversions in parentheses)
Strength units
(1)
(2)
(3)
(4)
(5)
(6)
a
(in) (ram)
K/cr
~
~ro=
(ksi) (MPa)
~b
(ksi) (MPa)
cr,(
(ksi) (MPa)
0.18 (4.5)
0.36 (9.1)
0.53 (13.5)
0.71 (18.0)
0.89 (22.6)
1.60 (41)
2.32 (59)
3.03 (77)
3.74 (95)
4.45 (113)
5.17 (131)
5.88 (149)
6.59 (167)
7.30 (185)
1.241 (0.1977)
1.698 (0.2706)
2.000 (0.3187)
2.225 (0.3546)
2.408 (0.3838)
2.955 (0.4709)
3.361 (0.5356)
3.691 (0.5882)
3.959 (0.6309)
4.158 (0.6626)
4.238 (0.6754)
3.985 (0.6351)
3.788 (0.6037)
3.819 (0.6086)
1.6590
1.6050
1.5430
1.4870
1.4400
1.3170
1.2460
1.1970
1.1550
1.1120
1.0520
0.9273
0.8324
0.7973
1.014 (6.99)
1.016 (7.01)
1.019 (7.03)
1.023 (7.05)
1.027 (7.08)
1.054 (7.27)
1.095 (7.55)
1.154 (7.96)
1.238 (8.54)
1.360 (9.38)
1.549 (10.68)
1.853 (12.78)
2.174 (14.99)
2.443 (16.77)
51.91 (357.9)
42.30 (291.6)
37.19 (256.4)
33.87 (233.5)
31.57 (217.7)
30.06 (207.3)
29.47 (203.2)
31.36 (216.2)
32.99 (227.5)
32.73 (225.2)
52.94 (365.0)
44.25 (305.1)
37.71 (260.0)
33.57 (231.3)
=tro, = Stress in outer intact stiffener for unit applied gross stress
b Cr = Residual strength based on skin fracture = Kc/(~/~raf~); Kc/(2) = 137 (125)/(2) [137/(2)]
c ~r,t = Residual strength based on stiffener strength F tu/(4) = 565(82)/(4) [565/(2)]
skin by three rows of rivets. The frame configuration
is seen to the right of the figure. Point D in Figure 3
is the configuration allowable for the two-bay crack.
Any fast fracture below this point would be arrested.
Fast fracture above point D would not be arrested
and failure would be precipitated by crack-stopper
failure. It is the author's opinion that blade impact
tests of the type illustrated by Figure 3 should be
designed to m a k e sure that fast fracture occurs so that
the crack-arresting material can be substantiated. If
the disc size is too small for this to occur then some
fragmentation damage should be simulated. This has
been done by at least one manufacturer of a turbo
prop Part 25 aircraft in recent years.
As mentioned earlier, the horizontal line in Figure
3, intersecting the skin fracture curve at B and C, was
based on a principal stress calculated from hoop
and shear stresses. It has been noticed that some
manufacturers p e r f o r m residual strength tests for the
pressure cabin using nominal cabin pressure only
without compensating for the effects of skin shear
and residual strength in the presence of cabin pressure
and fuselage bending loads.
A design goal for large damage capability in the
lower wing surface was a two-bay chordwise skin crack
with a broken central stiffener at limit load. This is
illustrated by Figure 6. This large damage size was
chosen to allow the opportunity to establish an external
visual inspection at reasonable intervals. The rationale
included the possibility that fast fracture may occur
on a limit load application and be arrested at adjacent
intact stiffeners. This large crack would then be
detected on a walk-around inspection. The central
stiffener was assumed to be broken following the
normal sequence of failure expected and confirmed
by service history and c o m p o n e n t tests. This large
damage scenario provided a balanced design for the
following reasons.
1. The limit stress level chosen for the wing lower
surface was a little lower or about the same as the
residual strength capability for the two-bay crack
Fatigue 1994 Volume 16 Number 1 79
Damage tolerance capability: T. Swift
Torque
100
1
50
lq
I
12
10
(ram)
150
I
Results with z
shear stress
e
r
o
200
I
250
Pressure
80
~
Crack
60
==
==
O.
8
e
gl.
w
6
Results with shear
q0
\
stress applied
4
20
I
I
2
I
q
I
6
8
10
Total crack length 2a(in)
j
Figure 4 24 in (0.61 m) diameter 2024-T3 barrel tests, 0.032 in (0.81 mm) thick
case when proper materials were chosen for skin
and stiffener. Thus additional weight was not
required for the large damage criterion.
2. It allowed visual inspections to be based on crack
growth from a detectable crack size all the way to
the crack-arresting adjacent stiffeners. This was
possible as the maximum spectrum load for normal
usage is only about 60% of limit load and occurs
only once in one tenth of a lifetime.
3. The resulting inspection burden on the operator is
not excessive.
Figare $ Residual strength tests for damage tolerance capability
should include both pressure and fuselage bending
~
g1IFFD~R
_
o
Figure 6 Wing damage tolerance capability
80
Fatigue 1994 V o l u m e
16 N u m b e r
1
It is this author's opinion that design to the two-bay
crack criterion for the lower wing surface should be a
design objective. Unfortunately, some manufacturers
in recent years have allowed their stress levels to
increase to the extent that this is no longer possible.
Other manufacturers are contemplating this also for
future designs. The implications of this from an inservice inspection standpoint can be realized with a
typical example.
Consider the configuration shown in Figure 7. This
represents a typical lower wing surface with riveted Z
section stiffeners of cross-sectional area 0.8933 ine
(576.3 mm z) riveted to 0.3 in (7.6 mm) thick 2024T351 plate. A two-bay skin crack with a broken central
stiffener is assumed. An analysis of this configuration,
based on the displacement compatibility approach, 7
was performed to obtain crack tip stress intensity and
intact stiffener stresses. The skin fracture toughness
for the 2024-T351 material varies with thickness, and
the value assumed was 125 ksi ini (137 MPa m t) for
0.3 in (7.6 mm) thick plate based on Figure 11 of
Ref. 7. The strength of the 7075-T6 extruded intact
outer stiffeners, including the effects of bending, was
based on a static strength Ft, of 82 ksi (565 MPa).
Figure 7 shows the stress intensity factor per unit
applied gross stress and the value of the geometric
Damage tolerance capability: 7". Swift
{mm)
,50
I
4.5
1 O0
I
150
t
200
I
q.0
-
"O
tin
20
3.5
Stiffeners
S t i f f e n e r area
0.8933 in 2 [576.3 mm2}
t-
3.0
-
15
,,in
U
c
2.0
--
/
~(Iq5
ram)""]P~ . . t
~IB~ at°ken
II",~
E
E
2.5
707S-T6
extrusion
Intact
stiffener
stiffenerJ r~ '
10
e-
1.5
~
0.3 in
{7.6 mm}
t.o
e.J
e,
curve/
13
0.5
- S
•
|
,
,
t
•
1.0
2.0
3.0
q.o
5.0
6.0
7.0
Skin
202q-T351
plate
8.0
Half crack length o ( i n )
l~ure 7
Wing lower surface typical two-bay crack configuration unit stress intensity factor and [3 curves
term 13 as a function of half crack length a. The
residual strength of the configuration, as a function of
crack half length, is shown in Table 2 and illustrated
in Figure 8. The allowable stress for the two-bay
crack configuration is given by point A at 33.69 ksi
(232.3 MPa). Any fast fracture higher than this value
would not be arrested and fast fracture below this
point would be arrested. Suppose now the limit stress
was fixed at 33.5 ksi (231.0 MPa). Fast fracture would
take place at point B and be arrested at point C. At
this crack length the intact stiffener is not critical for
the stiffener material used in this example, as indicated
by point D. It is believed that an inspection programme
could be established based on visually detectable
cracking. If fast fracture did occur during a limit load
application in flight then the crack would be arrested
at a large damage size, which would be considered
60
50
(ram)
100
I
I
150
Skin fracture
50
200
l
Intact s t i f f e n e r
strength c r i t e r i o n ' ~ , ~
400
\
300
q0
•~
30
Limit s t r e s s ~
33.5 ksi
"~,,~'~..for
(231.0 MPa)
~
Fast
Allowable
I
crac, karrest
2~.
,,,-r
fracture "~'~ B ' ~ ' C I ~
A
Arrest
33.69 k s i
232.3 MPa)
200
"0
~ 2o
100
10
Two-bay crack
broken central s t i f f e n e r
I
'I
I
I
'
'
'
1.0
2.0
3.0
4.0
5.0
6.0
7.0
8.0
Half crack length o (in)
F~mre g Residual strength diagram for two-bay skin crack with
broken central stiffener
evident on a walk-around inspection. This situation
would not place an undue burden on the operators.
Now suppose it was decided to save weight by
increasing the stress levels by say 10% on subsequent
designs. This may be possible from a static strength
standpoint with some of the newer 2000 series alloys,
developed to replace 2024-T351, which have roughly
a 10% improvement in static strength capability.
Unless there was a corresponding 10% improvement
in fracture properties, which is unlikely, the two-bay
crack design goal would no longer be achievable, as
limit stress would increase to 36.85 ksi (254.1 MPa),
which is above the allowable for crack arrest, as
indicated by Figure 8. This of course assumes the
fracture toughness of the higher strength alloy remains
the same. Let us now consider the implication of this
situation.
An inspection programme is required with inspection
frequency based on crack growth from a detectable
crack size to a critical size at limit load. It will
therefore be necessary to determine the critical crack
size accounting for slow stable growth on a limit load
application.
The stress intensity factor at the onset of slow stable
growth is a function of the crack tip plastic zone size,
which in turn is a function of the applied cyclic stress
level causing crack propagation. Slow stable growth
will not start to occur on a high load cycle until the
plastic zone size created by the cyclic stress that
propagated the crack is equalled on a high load cycle.
This phenomenon is explained in Ref. 8 and has been
verified by a substantial amount of testing. As not all
aircraft in the fleet may experience high gust or
manoeuvre load cycles, it will be conservative to
assume that the crack in question had been propagated
at a moderate cyclic stess. Thus, to be conservative a
crack resistance (R) curve is desired having a low
onset stress intensity factor.
Figure 9 represents such an R curve for 2024-T351
plate, 0.25 in (6.4 mm) thick, taken from Ref. 9 page
Fatigue 1994 Volume 16 Number 1
81
Damage tolerance capability: T. Swift
(ram)
1110
10
20
|
I
c.tlca,
30
q0
|L
*
~f,'~'e
50
|L I
J
, 2 , ks, in - - . - . . - - - - -
I
-
_c 120
100
80
~ 6o
.-~
4o
~
20
~
l
~1
1
14o
Extended by comparison I 120
with other R curves to I00
critical K
"~E
so ~.
/
R curve with low threshold chosen
=E
/
22 ksi in½:(2q.2 MPa m½l
60
/ 2 0 2 1 1 - T 3 5 1 0.25 in (6.4mm) thick
- q0
Ref. Damage Tolerance Design
Handbook MCIC-HB-01R page 7.5-61- 20
I
=
•
=
=
,
•
=
•
•
|
•
•
|
|
,
•
•
=
0.2 0 . 4 ' 0 . 6 " 0 . 8 ' 1 . 0 1.2'1.11 1.6'1.11
Change in crack length /~a (in)
(mm]
50
100
150
200
!
I
60
Intact stiffener
J~ 400
strength criterion ~ _ ~ .
l
Skin fracture ~
/
_ 50
~,w1,, c r i t e r l o n ~
l
.~
~
F~st fracture
N
-4 300
~ _
=/ New limit stressN,
1
110 36.;
AI"~P36.85 k~i (25q.I ~Pa)~. C /
ksi
/~
T'~AIIowable for c r a c k ' ~ 1 3 3 . 6 9 ksi
253, / I_~1
"'-...~.arrest~
~ --'(232.3 MPa)
,0
.p,
I
¢=~'
"O
.0
-
-
-
-
4
2oo
--.~:
~ NUnstable growth
~.
20
~
o:
,
-
r
10
Figure 9 Crack resistance curve
1.0
100
Slow stable tear on
increasing single load
application
i
2.0
I
3.0
I
4.0
i
5.0
I
6.0
i
7.0
8.0
Half crack length a (in)
7.5-61. This curve was extended to reflect a critical
stress intensity factor of 125 ksi in i (137 MPa ml).
Using the R curve of Figure 9, with the unit stress
intensity factor curve of Figure 7, it is possible to
determine the amount of slow stable growth on a
single limit load application. A half crack length of
1.0 in (25 mm) is considered to be the smallest
visually detectable crack at 90% reliability with 95%
confidence. Starting with this length crack the calculation for slow stable growth is shown in Table 3
Figure 10 indicates that the slow growth curve peaks
out at a gross stress of 36.73 ksi (253.2 MPa), slightly
below the limit stress at a half crack length of 1.9 in
(48 mm) illustrated by point A of Figure 10. Beyond
this point the crack would tear slowly at constant load
and fast fracture at point B. Failure of the structure
at point C would be precipitated by outer stiffener
failure as the allowable for crack arrest would be
below the applied stress, as illustrated by point D.
Another way to confirm the instability is to compare
applied stress intensity factor curves, generated at
varying gross stress levels, with the resistance curve.
Instability occurs when the slope of the applied K
curve equals the slope of the resistance curve. This is
illustrated by Figure 11. The slope of a K curve,
generated at a gross stress of 36.73 ksi (253.2 MPa)
becomes tangent to the R curve at a half crack length
of 1.9 in (48 mm). This point is also reflected by point
A of Figure 10.
Table 3 Calculation of slow stable growth on limit load
application, (threshold stress) intensity factor 22 ksi in ~ (24.1
MPa m t) (Figure 9) (SI conversions in parentheses)
(1)
(2)
(3)
(4)
(5)
da=
a
K/(r
Kr
~r Kr/(3)
(ksi) (MPa)
0
0.2
0.4
0.6
0.8
0.9
1.0
(0)
(5.1)
(10.2)
(15.2)
(20.3)
(22.9)
(25.4)
1.0
1.2
1.4
1.6
1.8
1.9
2.0
(25.4)
(30.5)
(35.6)
(40.6)
(45.7)
(48.3)
(50.8)
2.50
2.69
2.82
2.56
3.07
3.13
3.45
(0.398)
(0.429)
(0.449)
(0.408)
(0.489)
(0.499)
(0.550)
22.0 (29.2) 8.80 (60.7)
69.6 (76.5) 25.87 (178.4)
92.4 (101.5) 32.76 (225.9)
104.6 (114.9) 35.34 (246.6)
112.0 (123.1) 36.48 (251.5)
115.0 (126.4) 36.74 (253.3)
117.0 (128.6) 33.91 (233.8)
da is increment in crack length in inches (mm)
Fatigue 1994 Volume 16 Number 1
Figure 10 Residual strength curve for two-bay skin crack configuration with increased limit stress
The implication here is that if stress levels are
increased, and the two-bay crack capability at limit
load is abandoned as a design goal with a view to
saving weight, then the critical crack would be of a
length not considered visually detectable with the
required reliability and confidence. In the example
here the critical half length would be 1.0 in (25 ram),
as this crack would grow to an unstable length on a
limit load application and would not be arrested.
Assuming an external inspection programme were still
desired it would be necessary to detect an extremely
small skin crack requiring non-destructive testing
(NDT). After failure of internal stiffener the skin
crack growth life Ls, shown by Figure 12, would be
small in terms of sophisticated NDT inspection owing
to load transfer out of the broken stiffener into the
skin. The inspection frequency would be Ls/2. This
could be extended to LTOT/2 as indicated by Figure
12, but this would require inspection for stiffener
cracks, which would need internal NDT or external
NDT using low-frequency eddy current. If this noncrack-arrest philosophy were to be used over wide
areas of basic structure, ie, thousands of square feet
(hundreds of square metres), the inspection burden
on the operator would be excessive.
T H R E S H O L D FOR DETAILED INSPECTION
The Federal Aviation Regulations, FAR 25.571,
require that an inspection programme be established
to protect the structure from accidental, corrosion and
fatigue damage. For fatigue-critical elements it has
become customary to establish a threshold and frequency for in-service inspection. The calculation of
the frequency for inspection is comparatively straightforward and is based on the crack growth life between
in-service detectable and critical crack sizes. The
critical crack size at limit load can be calculated and
the detectable crack size can be established depending
on the inspection method to be used. The determination
of the threshold for detailed inspection of fatiguecritical elements is not so straightforward. In fact,
any threshold determined which is longer than the
Damage tolerance capability: T. Swift
-*,= I~,0
".~
120
100
(mm}
20
qo
60
80
,
i
w
i
Unstable half
36.73 (253.2)
crack length ~
~.,,..-.----'-1.9 in)
~ 3 S
(2q1)
(q8 mm)
~
j30
lq0
120
(207)
100
,-M
E
80
£
:E
f
•~
"~o
tJ
qO
20
~
"
~
I
~e
"~
Onset of I
slow, g r o w t h
0.5
60
Appliedgross
\
stress
~
(ksi (MPa))
Applied K curves
~
1.0
R curve
,
i
1.5
2.0
Two-bay crack
broken central stiffener
qO
20
,
,
2.5
3.0
3.5
Half crack length o (in)
Figure 11 Crack resistance R curve with applied K curves
]~
LTOT
i
FLIGHTS
Figure 12 Impfication of not meeting two-bay skin crack criterion; detailed NDT over large area of basic structure
frequency of inspection does not conform to the
damage tolerance philosophy.
When the damage tolerance requirements became
effective in 1978, after amendment 45 to F A R 25.571,
it became apparent that a threshold for detailed
inspection of fatigue-critical elements would be much
more severe than existing thresholds if it was to be
based on the frequency of inspection determined by
crack growth life from in-service detectable cracks to
critical sizes at limit load. Since the aircraft were still
designed to be fatigue-crack-free for their design life
goals it was thought unnecessary to start inspections
earlier than the time when cracks would become
detectable. Thus the true damage tolerance philosophy
became diluted with the safe life approach to some
extent. No standard guidance is provided in the
requirements or advisory material on methods to
determine when detailed inspections should begin, and
consequently a wide variety of methods are used by
different airframe manufacturers.
A number of manufacturers have considered initial
manufacturing flaws when establishing the inspection
threshold. This threshold is based on half the life to
grow the manufacturing flaw to a critical size at limit
load. The manufacturing flaw size chosen (usually a
0.05 in (1.3 ram) crack at a fastener hole) is considered
inspectable in the factory during manufacture. This
approach to establish a threshold is still a damage
tolerance approach undiluted by a safe life philosophy.
Some manufacturers, primarily those producing large
transport aircraft, have established thresholds based
purely on a fatigue life approach without considering
the possibility that initial manufacturing flaws may be
present. Some of these thresholds are as long as three
Fatigue 1994 Volume 16 Number 1 83
Damage tolerance capability: E Swift
was chosen and the details of this analysis are described
in Ref. 10. The results are repeated again here to
further emphasize this problem.
As mentioned in Ref. 10, the wing lower surface
limit stress level for most commercial transport aircraft
is in the vicinity of 35 ksi (240 MPa). Some have
higher and some have lower stresses than this. Figure
14 described what could be expected during a limit
load application assuming a broken spar cap and a
skin crack of half length 0.5 in (12.7 ram), which is
not considered visually detectable with high enough
reliability. The spar cap area is assumed to be 2.788 in 2
(1799 ram2), and the skin is 2024-I"351 plate 0.25 in
(6.4 ram) thick. Figure 14 shows applied stress intensity
factor curves for a number of gross stress values based
on the non-linear displacement compatibility analysis.
These curves are compared with the resistance (R)
curve as indicated in the figure. It can be seen by
point A that slow stable tearing will have already
started at a gross stress of 15 ksi (103 MPa) owing to
the high stress intensity factor caused by load transfer
out of the broken spar cap into the cracked skin. At
20 and 25 ksi (138 and 172 MPa) on the limit load
application the skin crack will have grown to points
B and C respectively. At a gross stress of 27.84 ksi
(192.0 MPa), well below limit stress, the skin crack
would become unstable as indicated by point D and
would result in a fast fracture condition. Instability
occurs when the rate of change of applied stress
intensity equals the rate of change of resistance stress
intensity or when the applied K curve is tangential to
the R curve.
The question to be asked now is: will the fast
fracture at 27.84 ksi (192.0 MPa) be arrested by the
adj acent intact stiffener such that loading could increase
above 27.84 ksi (192.0 MPa) to the limit value of
35 ksi (240 MPa), whereupon the crack would likely
be found on a walk-around inspection? To answer this
question, further displacement compatibility analysis
was performed for the typical configuration. Details
of this analysis are given in Ref. 10. Figure 15 shows
the results. The analysis was performed assuming
quarters of the aircraft life. Under these circumstances,
initial manufacturing flaws could grow to a fast fracture
size before the threshold for detailed inspection. This
situation has been thought to be satisfactory when the
structure is redundant, fail-safe or has crack-arrest
capability. In this case a second line of defence exists
in the event that an initial manufacturing flaw nucleates
into a propagating fatigue crack during the service life
of the aircraft. Since crack arrest structure is usually
capable of sustaining large damage it is likely that
such damage will be readily detectable. This approach
to the manufacturing flaw problem appears to be
reasonable and further reinforces the argument to
design to the two-bay crack capability illustrated by
Figure 1 and discussed in the previous section. However, not all structural elements can be classified as
crack-arrest-capable.
Consider a typical wing rear spar cap as shown in
Figure 13. A threshold for detailed inspection based
on a fatigue life method without considering manufacturing damage for this element would be extremely
long. Any inspections performed before this time
would be merely visual. As can be seen from the
figure, a manufacturing flaw could grow to complete
failure of the cap undetected prior to the threshold
for detailed inspection. The cap cannot be inspected
visually as it is covered by skin and rear spar web and
at rib locations is covered internally by a rib fitting.
Prior to amendment 45 of FAR 25.571 this element
would have been cleared to the fail-safe single element
failure concept without considering any secondary
damage in the skin or web. It is more than likely that
skin or web damage would have developed during
failure of the cap and this should be realistically
accounted for.
A displacement compatibility analysis, considering
the elastic-plastic behaviour of the spar cap to skin
fastening system, was performed to determine whether
a visually undetectable skin crack in the presence of
an undetected broken spar cap could tear in a slow
stable manner to a fast fracture condition on a limit
load application. A typical rear spar cap configuration
[
[~"--L
-1- "L
-L
11
I - L ~
FUBi-i
RIBPHz~IO
j-
"1
8Pgl~ Vk.~.Wd.Y
UNINSPEb~ABLE
Figure 13 Example of the need to consider initial manufacturing flaws when establishing threshold for inspection
84
Fatigue 1994 Volume 16 Number 1
Damage tolerance capability: T. Swift
160
c
lqO
120
U
(mm)
30
40
50
A~plled £~ross S~ress ~ '
I0
20
'
60
il
~
"t 160
~''~7~
lZl0
D 27.8a, (152.0)
" 2 5 (172 "1 120
(ksi (MPa))
Instability
~
Applied K curves . . ~ ~
100
"f/
~
~15 1103}-I 80
Broken
s p a r ~ ~ l ~
~"
60
C
~
qo
Unstable
k
,
•- > I 2o 1 4 - ~ ,
I"
crack length
l!
-____rv.R
cu e
,
0.5
1.0
1.5
Skin
q0
.
,
2O
,
2.0
2.5
Half skin crack length (in)
14 Applied skin stress intensity curves with resistance curve for broken spar cap: 100% spar cap load transferred to skin
5O
q0 I
50
100
7075-T6
202q-T3 ~ j ~
(ram)
150 200
250
300
"'NOp7 Intact stiffener
. X / x strength
300
I n tact
stiffener
Lim,.t ..stre.s.s...' ~ " X "
r-
30;
:3
Instabillt.~. stress N ~
after ;'r~'w'a3bl~"-'~-'r
200 ~.
!
20
100
L,.
o
spar cap
criterion
'
I
~
/ v1,
'
o
r
Skin crack
crack arrest
I
I
I
I
/
2
a,
6
8
10
12
Crack length o (in)
lqgure 15 Residual strength after fast fracture (shows that crack is not arrested)
either 7075-T6 or 2024-T3 stiffener material. It can be
seen that at the instability stress of 27.84 ksi (192.0
MPa), causing fast fracture, failure of the intact 2024T3 stiffener would occur at point A, and at point B
if the stiffener were 7075-T6. Thus the intact stiffener
is inadequate to arrest the fast fracture. In fact, the
allowable for crack arrest indicated by point C, which
is well below the instability stress and considerably
below the limit stress.
The implications of this situation are that if a
threshold for detailed inspection is established without
considering growth of a manufacturing flaw then failure
could occur on a limit load application before cracking
became detectable.
As mentioned earlier, establishment of an inspection
threshold for detailed inspection of fatigue-critical
structure based on a fatigue evaluation without consideration of initial manufacturing flaws can only be
justified if the structure is capable of arresting large
cracks which are then visually detectable. If this is not
the case then initial manufacturing flaws should be
accounted for to establish the threshold, as indicated
by the spar cap example shown here. There may be
other examples falling into this category but probably
not many on large transport aircraft.
THE EFFECTS OF MULTIPLE-SITE D A M A G E
ON R E S I D U A L STRENGTH
Structural safety, maintained through a damage tolerance philosophy, depends upon an inspection programme. The frequency of inspection is based on a
crack growth life evaluation starting with a detectable
crack size and terminating at the critical damage size
under limit load conditions. The critical damage size
can be influenced by the condition of the surrounding
structure. If the structure is young the presence of
multiple-site damage (MSD) may be unlikely to the
extent that the lead crack residual strength would be
affected. However, if the structure is operating beyond
the life substantiated by fatigue testing, there is a
strong possibility that MSD may affect the lead crack
residual strength, as illustrated by Figure 16. This
condition is particularly important in the event of
discrete source damage, as in the case of engine burst.
For example, Figure 3 illustrates residual strength
capability in the event of damage from an engine disc
fragment. The structural configuration described in
the figure has the capability to arrest a fast fracture
in two bays with a broken frame. However, if the
material surrounding the potential arrest point contains
Fatigue 1994 Volume 16 Number 1 85
Damage tolerance capabifity: T. Swift
T
T
T t
.r,_
T
,O-
Figure 16 Primary concern with MSD: effect on load crack residual
strength
fastener holes with undetected MSD present the fast
fracture may not be arrested.
For this reason it has become the opinion of this
author that sufficient fatigue testing followed by
teardown inspection should be required to make sure
that MSD will not influence lead crack residual strength
within the design service life goal. It is not considered
feasible from either a technical or economic standpoint
to live with the potential for MSD within the framework
of a damage tolerance philosophy. The aircraft should
be designed crack-free for the service life goal. Damage
tolerance is intended to take care of inadvertent
damage due to accident, corrosion or early fatigue
cracking, which may occur locally within the design
life owing to poor quality such as a manufacturing
flaw. It is believed that even extremely small undetectable MSD can substantially reduce lead crack residual
strength.
An example analysis, based on an intuitive failure
criterion, can illustrate the effect of MSD on lead
crack residual strength. Figure 17 shows a typical
example of original design capability for a circumferential skin crack on the crown of the fuselage.
Capability usually exists for a two-bay crack with a
broken central stiffener at a limit stress of 34 ksi
(234 MPa) due to fuselage down-bending and pressure.
It is possible to develop a residual strength diagram
for this stiffened structure case through a finite element
analysis, as illustrated by Figure 18. This analysis is
completed in two parts. First, an unstiffened panel is
idealized as shown, by dividing the sheet into bars
and shear panels. Bars carry axial load and panels
carry shear load. Loads applied at the top of the
panel are reacted at the bottom. These reactions are
disconnected sequentially to simulate the propagating
crack. The stress in the last bar still reacted gives the
crack tip stress. The idealization is then modified to
include the stiffening elements idealized as described
in Figure 18. The idealized bars simulating the stiffeners
have areas equivalent to the stiffener area and are
placed to provide the same bending moment of inertia.
Rivet flexibility is simulated by a continuous shear
panel of thickness tse, chosen to provide the same
stiffness as the rivets. Ratios are taken between
stiffened panel and unstiffened panel crack tip stresses
to give the geometric term 13 in the calculation of K =
o'['rra]l/213. The results of this typical analysis are given
in Ref. 2 p.179 Case 15. These results are repeated in
Table 4 and plotted on Figure 19. The panel allowable,
given by the intersection of the stiffener strength curve
and the skin fracture curve is 34.5 ksi (238 MPa). Thus
the configuration has the ability to arrest a fast fracture
at a limit stress of 34 ksi (234 MPa).
Consider now the effect of MSD on the lead crack
residual strength for this typical case. It has been
determined that for equal MSD crack sizes, as shown
in Figure 20, the failure criterion is net section yielding
between crack tips. This appears to occur at the flow
stress or at a net stress approximately equal to (Ftu
+ Fry)/2. It appears conceivable, then, that this failure
criterion could be applied to the ligament between the
lead crack tip and the first MSD crack as illustrated
by Figure 20: that is, when the stress in this ligament
reaches (Ft, + Fry)/2.
Figure 21 illustrates an intuitive link-up criterion
based on the gross stress level that will cause the lead
crack plastic zone to touch the MSD crack plastic
zone, ie:
Plastic zone sizes can be expressed as
(K1/Cry ):
R1 2~r
Kx = 13h13no'0ral )1/2
Therefore
Original design
capability
Two-bay circumferential
skin crack with broken
central stiffener
(~h)2(1311)2o'2"tral + (13s)2(1312)2tr2'~a2
d
20.y2
= P - ~ - al
Therefore
where:
O"R
W
Limit stress 31t ksl {23~1MPa} due to
fuselage down-bending and pressure
lqwBre 17 Residual strength capability of fuselage crown skin
86
(/(2 hry )2
R2 =
2~r
K2 = 13s13i:cr0ra2)l/:
Fatigue 1994 Volume 16 Number 1
Cry
13h
,[
7+
1
(4)
gross stress that causes the plastic zones to
touch;
= flow stress approximately (Ft. + Fty)/2;
= Bowie factor xl for cracks at a hole but
normalized to crack length measured from
----
Damage tolerance capability: T. Swift
APPLIED LON)8
/
QUARTER OF CRACI~D
PANEL
BARS CARRY
AXIAl.LOAD
REACTIONS
y
,
~ X
t m,
RIVET FLEXIBILITY
IDEALIZATION
Figure 18 Finite element analysis of stiffened cracked panel
Table 4 Results of typical finite element analysis for
circumferential crack (SI conversions in parentheses)
138
(1)
(2)
(3)
(4)
(5)
(6)
a
Rot
I~
trr
O'o,,
~
1.5 (38.1) 0.778 1.285 51.98 (358.4) 1.074 (7.41)76.35 (526.4)
2.5 (63.5) 0.820 1.220 42.41 (292.4)1.114 (7.68)73.61 (507.5)
3.5 (88.9) 0.852 1.174 37.25 (256.8)1.176 (8.11)69.73 (480.1)
4.5 (114.3) O.883 1.133 34.04 (234.7) 1.269 (8.75)64.62 (445.5)
5.5 (139.7) 0.918 1.089 32.03 (220.8) 1.414 (9.75)57.99 (399.8)
6.5 (165.1) 0.970 1.031 31.12 (214.6) 1.660 (11.44) 49.40 (341.2)
7.5 (190.5) 1.107 0.903 33.08 (228.1)2.153 (14.84) 38.09 (262.6)
8.5 (215.9) 1.353 0.739 37.97 (261.8) 3.212 (22.15) 25.53 (176.0)
9.5 (241.3) 1.428 0.700 37.92 (261.4)3.875 (26.33)21.16 (145.9)
(3) = 1/(2)
Kc 159 (145 ksi in*) (159 MPa/mi)
(4) = 145/[(7ra)i]
(6) = 82/(5) [565/(5)]
F,o = 565(82 ksi) (565 MPa)
= 13 for the stiffened panel for lead crack
length a 2 , (Table 4);
1312 = [3 for crack 2 tip due to interaction of crack
1, ie this is determined from Ref. 12 Figure
76 based on an equivalent lead crack 2
length ae2 = 13s2a2 and equivalent small
crack 1 length ael = 13h2al •
Consider that MSD cracks in hole 1 of Figure 21
are 0.05 in (1.27 mm) at each side of the hole. The
hole diameter d is 0.19 in (4.83 mm), so that at =
0.145 in (3.68 mm). The rivet spacing P = 1.0 in
(25.4 mm). The stress intensity factor for two cracks
at a hole under uniaxial loading is given as K = (xtL) 1/2
F(L/r) in Ref. 11. At a value of l/r = 0.05/0.095
(1.27/2.41), F(L/r) = 1.8 [by plotting L/r versus F(L/
r)]. The value of 13h can then be obtained as follows:
13h
3.
hole 1 centre, ie al;
= 13 for crack 1 tip due to interaction with lead
crack 2, ie this is determined from Ref. 12
Figure 75 based on an equivalent lead crack
2 length a=2 = 1382a2, ie equivalent crack
length to make K the same as in the
stiffened panel and equivalent small crack 1
length ael -- 13h2al;
= (1.82(0-05)/1/2
\. ~
]
=1.057
The residual strength calculation based on Equation
(3) is given in Table 5. The value try assumed is 50 ksi
(345 MPa) for 2024-T3 material. Assuming the lead
crack tip is continuously influenced by an MSD crack
the residual strength would be reduced below the limit
stress as shown by the lower curve in Figure 22 based
on the calcualtion shown in Table 5. Figure 22 shows
the reduction in residual strength due to MSD ahead
Fatigue 1994 Volume 16 Number 1 117
Damage tolerance capability: T. Swift
Stiffener strength criteria
based on Ftu 82 ksi (565 MPa) 7 0 7 5 - T 6
/
(ram)
5O
J
50
e"
150
200 ,t,,. 250
300
'1
Panel ~
allowable,\
~
I 300
"
~
%,~ ~
Allowable based on
Limit s t r e s s ~
. ~
N critical K
'
."2-
100
'
'
\.~,>"
'
Two-bay crack with
broken central stiffener
30 "Fast f
r a c t ~ ~
Skin fracture criterion X \
20 "based on K 145 ksi in ½ \
c
Arrest
(159 MPa m=) (2024-T3)
10
100
e¢
I
I
I
I
I
2
4
6
8
10
12
Half crack length (in)
~ m r e 19
Residual strength curve for circumferential skin crack with broken central stiffener
UNS'nFFENED PANEL
ttt
t t
CRITERION FOR F N L U R E - WHEN
LIGhMENT B E I W E B ~ G R N ~ TIP8
,u - ~
Figure 20
r~mu.
-O-
-(3-
-(3-
o,mi.~n rPauurm ~,n,, r.r=uN
Stiffened panel skin fracture criterion for lead crack in presence of MSD
tt
tt._ttt
LE~ a ~
ell
Table 5 Residual strength calculation, two-bay circumferential
crack with broken central stiffener, effect of M S D on lead crack
strength, based on Equation (3)
a2
fl,
ae2
a=l/b
ae2/b
/in
[~12
1.5
2.5
3.5
4.5
5.5
6.5
7.5
8.5
9.5
1.285
1.220
1.174
1,133
1.089
1.031
0.903
0,739
0.700
2.477
3.721
4,824
5.777
6.523
6.909
6.116
4.642
4.655
0.179
0.179
0.179
0.179
0.179
0.179
0.179
0.179
0.179
2.737
4.112
5.330
6.383
7.208
7.634
6.758
5.129
5.144
1.50
1.75
1.92
2.05
2.16
2.20
2.10
1.92
1.92
1.0
1.0
1.0
1.0
1.0
1.0
1.0
1.0
1.0
r,
i~m,=~ON F0R I . l ~ - UP
tlq~l I=ld~ 2INto
~
M!~ a~atA0~ ~
RI ÷ It= - lP-cl/S-all
* 0.162/0.905
** a e 2 / 0 . 9 0 5 ,
err = { 2~ry2(p'd/2"al)/(~he~n2al
Figure 21
88
err
Intuitive link-up criterion
Fatigue 1994 Volume 16 N u m b e r 1
erR =
(3800/(0'162~112
+
ae2 = ~s2a2
+ ~,2~1~2a2)}'
13~2~122a2)} '
36.57
30.02
26.48
24.26
22.85
22.22
23.59
26.93
26.90
(252A)
(207,0)
(182.6)
(167.3)
(157.5)
(153.2)
(162,6)
(185,7)
(185,5)
Damage tolerance capability: T. Swift
50
S0
100
I
I
(mm)
150
250
300
I
I
I~
Stiffener strength I
Panel allowable criterion
--I 300
unaffected by ~
B
|
qO
\~SD
Limit~
~
X stress
t-
200
I
~
~
j, 3..s ksi
=
~
" (238 ~Pa)
/
\
30
20 ,Lead crack residual ~
strength affected by
0.05 in
Reduction in panel
allowable due to MSD
(I .3 mm)
100
10
MSD crack
assumed size
I
i
i
I
I
2
q
6
8
10
12
Half crack length (in)
~
22
Lead crack residual strength affected by MSD
of the lead crack tip. Without MSD the allowable for
a two-bay crack with a broken central stiffener is
34.5 ksi (238 MPa) for the example configuration
based on the intersection of the stiffener strength
curve and the skin fracture curve. This point is
identified as point A on Figure 22. Point B is the
peak of the skin facture criterion curve. It can be seen
from the lower curve that fast fracture of the lead
crack would not be arrested at a stress higher than
27.5 ksi (190 MPa) when MSD cracks as small as
0.05 in (1.27 mm) exist ahead of the lead crack tip.
The intuitive link-up criterion used in this evaluation
needs to be verified by carefully controlled unstiffened
panel tests. It appears from this evaluation, however,
that small undetectable MSD can have a substantial
effect on lead crack residual strength.
CRACK STOPPERS
Many of the commercial transport aircraft in scheduled
airline service are fitted with separate crack-stopper
straps in the fuselage to guard against explosive
decompression in the event of longitudinal fast fracture.
Such fast fracture may result from undetected propagating fatigue cracks or discrete source damage created
by debris from a disintegrating engine. Well-known
aircraft fitted with crack stopper straps include the
Boeing 727, 737, 747, 757 and 767. In these aircraft
the crack-stopper straps are aluminium connected
directly to the skin in a circumferential direction either
by bonding or by riveting. Other aircraft, such as the
Douglas DC-8, DC-10 and Lockheed L-1011, are fitted
with titanium crack stoppers connected directly to the
skin. In the aftermath of the British Comet disasters
in 1954, fuselage designers after considerable research
and development, thought that these crack stoppers
were necessary to ensure fuselage structural integrity.
They were correct in these assessments.
Currently, in today's economic environment, there
appears to be a general trend to abandon the use of
crack-stopper straps connected directly to the skin, to
lower assembly costs. Dependence is being placed
solely on the skin-to-frame shear clips to arrest fast
fracture. In the opinion of this author this trend will
lower the level of safety that has existed in the
commercial fleet since the Comet accidents unless
other compensating measures are taken.
On 6 February 1970 a US Air Force C-133 transport
aircraft, cruising fully pressurized over Nebraska, was
lost owing to explosive decompression failure of the
fuselage. The fuselage design did not include crackstopper straps connected directly to the skin. The
frames were connected to the skin by shear clips only.
A fatigue crack had developed in the 7075-T6 fuselage
skin at the end attachment of the shear clip to the
skin in the area of the shear clip cutout, which allowed
axial stiffeners to cross the frame. This location is
identical to the Comet skin crack initiation point in
the vicinity of the automatic-direction-finding windows
on the crown of the fuselage. The C-133 crack started
at a skin countersink dimple and propagated undetected
into two adjacent bays to a total length of 11.0 in
(280 mm), when fast fracture occurred. The crack
progressed rapidly towards the shear clip cutout in the
adjacent frames and was not arrested. Figure 23
illustrates the frame and shear clip configuration in
the area of the failure.
During development of the C-133 aircraft, many
residual strength tests had been conducted on pressurized barrels to assess the crack arrest capability of the
shear clipped frames. In every case, where fast fracture
of the longitudinal skin cracks occurred, the crack was
arrested in a shear clip rivet hole, as indicated by
Figure 24, thus eliminating the stress intensity factor
at the crack tip. The cracks were never propagated
just out of the rivet hole, nor were further loads
applied to determine whether arrest would have
occurred if the crack had missed the rivet hole. This
is what happened on the C-133 lost over Nebraska.
The rapidly moving crack missed the rivet hole. All
of the residual strength testing performed on this
aircraft had created a false sense of security.
This
author perceives this false sence of security being
perpetuated today in the interests of economy. It is
realized and understood that currently used skin
materials have much higher fracture toughness characteristics than the 7075-T6 material used on the C-133
aircraft. Evaluation of residual strength characteristics
Fatigue 1994 Volume 16 Number 1
U
Damage tolerance capability: T. Swift
SKIN CRACK PASSED OVER
CUTOUT IN SHEAR CLIP MISSING RIVET HOLES
AND WAS NOT ARRESTED
8HEAR
(280 MM) AT FAST FRACTURE
SI~N CIaA.~..,KS'I'ARTEO
AT END RIVET NEAR ,
SHEAR CLIP CUTOUT
Figure 23 C-133aircraft fuselage skin crack fast fracture not arrested at shear-clippedframe
CRACK8 ARRESTED IN RIVET I'KXJE8
WERE NOT PROPAQATED BEYOND
RIVET HOLE TO DETEltUNE F
THEY WOULD HAVE BEEN ARRES11ED
IF THE HOLE WAS MISSED
(
Figure 24 C-133residual strength tests indicated that fast fracture was alwaysarrested at a rivet hole
using finite element methods indicates apparent ample
safety margins when considering these materials.
However, there are a number of possible secondary
effects in pressurized fuselage structure that can reduce
these residual strength margins significantly. Some of
these effects will be discussed.
The residual strength for longitudinal cracking in a
fuselage skin is given by Equation (1). The term ~s
provides the geometrical correction to the stress
intensity factor due to stiffening elements and is
90
Fatigue 1994 V o l u m e 16 N u m b e r 1
obtained by finite element analysis similar to that
shown in Figure 18. The term ~s in Equation (1)
represents the effect of skin bulging due to pressure
and radius. For a two-bay longitudinal crack with the
centre frame and crack stopper broken, ~a was
previously expressed as Equation (2). This equation
is written such that crack-tip bulging is completely
eliminated at the adjacent frame locations and this
fact has been verified by curved panel testing for
configurations where the frame crack-arresting material
Damage tolerance capability: T. Swift
includes a titanium crack stopper strap, as shown to
the right of Figure 25. In this case the crack stopper
shear clip combination riveted together provides sufficient stiffness to reduce the crack-tip bulging effect
to zero. However, if the crack stopper strap is
eliminated, as shown to the left in Figure 25, it is not
apparent that the shear clip flange alone will completely
reduce the bulging to zero. Any residual bulging
remaining at the frame location, as shown in the
figure, is probably a function of the stiffness of the
frame flange and possibly the frame area. The author
is not aware of any testing of curved panels or barrels
substantiating the suggestion that bulging is completely
eliminated at a shear-clipped frame without crack
stoppers.
It has been possible in the past to determine the
beneficial effects of crack stoppers compared with
shear clips alone by finite element analysis. Figure 26
shows this comparison for residual strength versus half
crack length. The term [3s was obtained from Case 5
of Ref. 2 for the configuration with crack stoppers.
The frame area was assumed to be 0.5042 in 2
(325.3 mm 2) and the titanium crack-stopper equivalent
aluminium area was 0.1035 in 2 (66.8 ram2). The shear
clip was 0.071 in (1.80 mm) thick. The bulge factor
13a was assumed to be given by Equation (2). The
2024-T3 skin material fracture toughness was assumed
to be 158 ksi in t (174 MPa mt). The residual strength
calculation from a skin fracture viewpoint is performed
in Table 6 and given by the upper curve in Figure 26.
As can be seen by the line representing principal stress
based on 1.5 g plus 82%, PR/t there appears to be
ample margin to arrest a fast fracture.
The curve in
Figure 26 below this curve represents the residual
strength for the same configuration with the crack
stopper removed. The calculation for this case is shown
in Table 7. The curve for this case was developed
from Case 1 of Ref. 2, modified to include a broken
frame. The curve is also based on the assumption that
the bulging effect reduces to zero at the frame. The
benefit of the crack stopper is evident by comparison
of the two cases. However, without the crack-stopper
strap the frame itself acts to reduce the crack tip stress
intensity factor but the load is transferred out of the
skin into the frame through a very flexible shear clip.
Thus the shear-clipped frame is not as effective in
reducing the crack tip stress intensity factor as the
crack stopper connected directly to the skin. There
still appears to be ample margin not accounting for
any other secondary effects. If in fact all the bulging
is not reduced to zero by the shear clip, as illustrated
by the configuration to the left of Figure 25, owing to
shear clip flexibility, then the peak of the second curve
may drop as shown by the dotted line in Figure 26.
Another secondary effect found to influence the
crack-arrest capability is frame bending, illustrated by
Figure 27. Even in a circular fuselage there are areas
in the frame subjected to considerable bending. This
bending is created by floor beam restraint for the
pressure condition plus bending due to transfer of
floor beam loads caused by inertia forces. These
bending moments, when in a direction shown by the
arrows in Figure 27, create additional tension stresses
in the skin locally near the frame as shown in the
figure. This effect reduced the residual strength curve
peak, illustrated in Figure 26, thus reducing the
apparent margin considerably. A very difficult area to
substantiate for this condition is on the lower side of
the fuselage shell at the base of a floor beam support.
Load transfer from the floor beam down the support
into the frame creates high frame bending moments
with tension on the skin side of the frame bending
material. In the author's experience it would be
extremely difficult to show a positive margin in this
area in the event of fast fracture in the skin without
the beneficial effects of a crack-stopper strap. The
peak of the second curve, illustrated in Figure 26,
would be lowered even further, aggravating the residual
bulging effect illustrated by the dotted line. The
message here is that the margins indicated by a residual
strength curve of the type shown in Figure 26 may be
substantially reduced owing to secondary effects such
as skin bulging and frame bending.
Another benefit provided by the additional residual
BULGINGRE~lqlAINED
BY COMBINATIONOF
TIT/~IIUM( ~ / g ~ ' O I = P E R
~
/
CR/g~STOPPER/~ID
FR~EWrrH SH~R
OLIP ONlY
Pi~.re 25
Possible crack tip bulging due to flexibility of shear clip on configuration without stopper
Fatigue 1994 Volume 16 Number 1 91
Damage tolerance capability: T. Swift
Two-bay skin crack
with broken frame
and crack stopper ~
q0
100
200
i
i
q ¢ . _ ~
(mm)
300
J
q00
500
600
I
I
!
~
'250
30i
t~
'~
=
200
Frames with
crack stoppers
,rincipal~ /
1.5 g plus~
, 82tPR/t
~
10l
..... ~°'-<
I
I-
I
I
I ~
8
12
16
20\
(1)
(2)
(3)
(4)
(5)
a
~
~,
Oj~llrap
~
(114.3)
(190.5)
(266.7)
(323.9)
(393.7)
(444.5)
(469.9)
(495.3)
(520.7)
(546.1)
1.2870
1.2136
1.1534
1.1161
1.0482
0.9681
0.8993
0.7133
0.5155
0.5200
1.3560
1.5260
1.6020
1.5800
1.4530
1.2880
1.1840
1.0646
1.0646
1.1840
6.5617
8.9895
10.6124
11.1607
10.6280
9.2455
8.1174
5.9436
4.4042
5.0600
(33.070)
(45.306)
(53.485)
(56.248)
(53.563)
(46.596)
(40.909)
(29.955)
(22.196)
(25.502)
24.08
17.58
14.89
14.16
14.87
17.09
19.46
26.58
35.87
31.23
(166.0)
(121.2)
(102.7)
(97.6)
(102.5)
(117.8)
(134.2)
(183.3)
(247.3)
(215.3)
Material 0.071 in (1.80 mm), Kc assumed 158 ksi ini (174 MPa mt)
trR = 158 (174)(4)
strength margin created by the crack stopper is the
ability to resist some MSD cracking ahead of the lead
crack tip. Figure 22 illustrates how the lead crack
residual strength from a skin fracture standpoint is
reduced from point B to point C in the presence of a
0.05 in (1.27 mm) MSD crack ahead of the lead crack.
This situation also occurs for the longitudinal crack
examples.
Consider the link-up criterion illustrated by Figure
21. For a circumferential crack, the residual strength
was give by Equation (3). Link-up was assumed when
the lead crack plastic zone touched the MSD crack
plastic zone. Equation (3) can be used for longitudinal
cracks by including the additional effect due to bulging.
The residual strength equation then becomes
92
Shear
clip _ ~ .
2q
Possible reduction due to
crack tip bulging induced
by flexible shear clip
Residual strength comparison: frames with and without crack stoppers
Table 6 Residual strength calculation, two-bay longitudinal crack
frames with crack stoppers, centre frame and crack stopper failed
(SI conversions in parentheses)
4.50
7.50
10.50
12.75
15.50
17.50
18.50
19.50
20.50
21.50
Frame
50
q
Half crack length (in)
Figure 2 6
10Q~'~~
Frames without
crack stoppers
\
!02q-T3 skin K~ 158 ksi in ½(17q MPa mj}
Titanium
crack
stopper
I/
~
Crack tip bulging
assumedzero at frames
t'~'~,
=o i',tress \ ' %
I'~
Frame I ' [
"
]
Shear ]"
c l l p
Fatigue 1994 Volume 16 Number 1
7 Residual strength calculation two-bay longitudinal crack,
frames without crack stoppers, centre failed (SI conversions in
parentheses)
Table
(1)
(2)
(3)
(4)
(5)
a
O,
/38
/3J~[~l*
~R
1.4741
1.3779
1.3160
1.2707
1.1962
0.1139
0.0541
0.9589
0.8565
0.8005
1.3560
1.5260
1.6020
1.5800
1.4530
1.2880
1.1840
1.0646
1.0646
1.1840
7.5157 (37.878)
10.2065 (51.439)
12.1084 (61.519)
12.7066 (64.039)
12.1268 (61.117)
10.6379 (53.613)
9.5147 (47.953)
7.9901 (40.269)
7.3212 (36.898)
7.7895 (39.258)
21.02 (144.9)
15.48 (106.7)
13.05 (90.0)
13.09 (90.3)
13.03 (89.8)
14.85 (102.4)
16.61 (114.5)
19.77 (136.3)
21.58 (148.8)
20.28 (139.8)
4.50
7.50
10.50
12.75
15.50
17.50
18.50
19.50
20.50
21.50
(114.3)
(190.5)
(266.7)
(323.9)
(393.7)
(444.5)
(469.9)
(495.3)
(520.7)
(546.1)
Material 0.07 in (1.80 mm) 2024-T3, Kc assumed 158 ksi int (174
MPa mi)
~rR = 158 (174)/(4)
DUETO
Figure 27 Additional skin stress due to frame bending
Damage tolerance capability: E Swift
Table 8
Effect of MSD on residual strength, two-bay longitudinal crack, frames with crack stoppers, centre frame and crack stopper
failed (SI conversions in parentheses)
az
19.5 (495)
20.5 (521)
21.5 (546)
~85
1~
a~2
a¢21b
[311
f312
ca
0.7133
0.5155
0.5200
1.0646
1.0646
1.1840
11.24 (285.5)
6.17 (156.7)
8.15 (207.0)
9.73
5.34
7.06
2.44
2.01
2.12
1.0
1.0
1.0
20.34 (140.2)
27.19 (187.5)
23.85 (164.4)
a I = • h 2 a l = 1.0572 x 0.145 (1.0572 x 3.68) = 0.162 (4.11) = constant
b = 1.25 - 0.095 (31.75 - 2.41) = 1.155 (29.34), a=t/b = 0.162/1.155 = 0.1403
OR = { 2Cy2( p-d/2-a:)/[f32[3n2al + ([3,[3B)2fl~22a2]}'
ae2 = ( [ ~ s [ ~ B ) 2 a 2 ,
CR = (5050/[0.162[3,, + + (~s~a)2~,22a2]} *
(OR = {240070/[4.1113,12 + (~sPa)2p,22a2]} t)
I
o"R =
2
d
performed to substantiate such a configuration should
include:
]1/2
20"y ( P - ~ - a l )
~h 2 ~i12al + ( ~ S ~ B ) 2 ~ I 2 2 3 2
(5)
where fib is given by Equation (2). All the other
terms remain the same as for the circumferential crack
case and are defined after Equation (3).
For the longitudinal crack the rivet spacing P is
assumed to be 1.25 in (31.7 ram) as shown by Figure
28. The MSD crack is assumed to be 0.05 in (1.27 mm)
and the rivet diameter d is 0.19 in (4.83 ram).
Therefore 31 = 0.095 + 0.05 = 0.145 in (2.41 +
1.27 = 3.68 ram). For purposes of crack interaction
the lead crack length is simulated by an effective crack
half length ae2 = (flsfla)2a2, where fls is the geometrical
effect of stiffening, ~B is the bulge factor and a2 is
the actual lead crack length. The effective length of
the MSD-cracked hole is given as ael = ~h2al, which
is the same as for circumferential cracking. The residual
strength calculation based on Equation (4) for the
configuration with crack stoppers is shown in Table 8.
The calculation for the configuration with shear clips
only is shown in Table 9. Crack interaction is again
determined from Ref. 12, Figures 75 and 76. The
effect of the 0.05 in (1.27 mm) MSD crack on the
lead crack residual strength curve in the vicinity of
the crack arresting frames is shown in Figure 28.
It can be seen that ample margin is still available
for the example with crack stoppers but the margin
has been completely eliminated for the configuration
without crack stoppers. This leaves no margin for the
frame bending effect or the potential bulging effect
due to shear clip flexibility. It is this author's opinion
that a configuration without crack-stopper straps may
not provide adequate discrete source damage capability
if a realistic damage scenario such as described by
Figure 3 is encountered along with secondary effects
described herein. Any curved panel or barrel testing
1. effects on cabin pressure plus aerodynamic suction;
2. skin shear due to fuselage down-bending (Figure
5); and
3. frame bending due to payload transfer from floor
beam to frame.
It is believed that great care should be exercised in
making sure that all realistic effects experienced by
the actual aircraft are properly simulated. If fast
fracture of such testing is arrested in a rivet hole the
crack should be propagated out of the rivet hole and
loads reapplied to simulate a crack missing the rivet
hole.
CONCLUSIONS
It has been shown by example that the implications
of not designing for a two-bay crack with a broken
central stiffener for basic airframe structure would be
to impose a significant inspection burden on the
operator. It is hoped that the manufacturers would
seriously consider the implications of lowering the
level of safety created by the dependence on the
resulting sophisticated in-service inspections.
It has been demonstrated by example that certain
elements in typical airframe structures do not have
large visually detectable crack-arrest capability. For
this type of element an initial manufacturing flaw
could grow critical prior to the threshold for detailed
inspection if this threshold were developed by fatigue
analysis alone. The threshold for detailed inspection
of such elements should be based on the growth of
initial likely manufacturing damage.
The substantial loss in the lead crack residual
strength in the presence of undetectable MSD is
demonstrated by analysis for stiffened fuselage structure. An intuitive failure criterion is adopted based
Table 9 Effect of MSD on residual strength, two-bay longitudinal crack, frames with crack stoppers, centre frame failed (SI conversions
in parentheses)
a2
19.5 (495)
20.5 (521)
21.5 (546)
//s
~
a=2
a~2/b
[311
f312
~n
0.9589
0.8565
0.8005
1.0646
1.0646
1.1840
20.32 (516.1)
17.04 (432.8)
19.31 (490.5)
17.59
14.75
16.72
3.31
3.01
3.22
1.0
1.0
1.0
15.12 (104.2)
16.52 (113.9)
15.50 (106.9)
See Table 8 for a=2, a=1, b, a=l/b and trR
Fatigue 1994 Volume 16 Number 1 93
Damage tolerance capability: T. Swift
Two-bay skin crack
a n d c r a c k stopper
crack
~(31.8m7}~.
0.05 in (I .27 rnm)-)H I ~
I
i
(ram)
i
40
100
I
200
I
300
I
qO0
I
500
I
600
t
250
30
Frames with
crack stoppers
C~
r-
20
Principal stress
:.~..
#O
t.s 9 p , u s . t pR,,
r~
10
-
Frames without
crack stoppers
--
100
A
50
Reduction in residual strength
due to MSD ahead of lead crack
I
I
I
I
I
q
8
12
16
20
2q
Half crack length (in)
Figure 28 Effect of MSD on lead crack residual strength: comparison of frames with and without stoppers
on known fracture phenomena, which should be
substantiated by carefully controlled specimen testing.
The examples illustrate the importance of making sure
that MSD does not occur within the design lifetime
by controlled stress levels substantiated by full-scale
fatigue testing followed by teardown inspection.
The importance of circumferential crack stoppers in
the fuselage to guard against explosive decompression
has been emphasized. It is shown that shear-clipped
frames alone may not be able to cope with a
number of secondary effects in arresting discrete source
damage.
3
4
5
6
7
8
ACKNOWLEDGEMENT
Published with permission from the Editors of 'Fatigue
of Aircraft Materials', DUP, The Netherlands, 1992.
10
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2
94
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