M I N I S T R Y OF A V I A T I O N

R. & M. No. 3248;

AERONAUTICAL RESEARCH COUNCIL

REPORTS AND MEMORANDA

Behaviour of Skin Fatigue Cracks at the Corners of Windows in a Comet I Fuselage

By R. J. ATKINSON, W. J. WINKWORTH and G. M. NORRIS

LONDON: HER MAJESTY'S STATIONERY OFFICE

1962

F O U R T E E N S H I L L I N G S N E T

Behaviour of Skin Fatigue Cracks at the Corners of Windows in a Comet Fuselage

B y R . J . A T k I N S O N , W . J . W I N K W O R T H a n d G . M . N O R R I S

COMMUNICATED BY THE DEPUTY CONTROLLER AIRCRAFT fRESEARCH AND DEVELOPMENT),

MINISTRY OF AV.IATION

Reports and Memoranda No. 3248* June, ±960.

Summary. Fatigue tests on a Comet I pressure cabin subjected to operational pressure cycles are described. Cracks at window corners are the main subject of investigation. Results are compared with earlier experiments on other Comet I pressure cabins. Conclusions are reached that appear to have some general significance.

1. Introduction. Fatigue investigations made on one of a number of Comet I fuselages that had

been specially provided for research purposes are described in this Report. In these investigations

the pressure cabin was subjected to pressure loading cycles only, and attention was directed mainly

to the initiation and development of cracksat the window corners. Special care was taken to avoid

unnecessary destruction, and cracks were repaired as necessary to prevent catastrophic failure.

Nine cracks out of a total of sixteen actually reached the stage where sudden extension appeared

imminent. Results are examined below in the light of earlier fatigue work on a Comet pressure cabin 1, 2

subjected to wing loads as well as pressure. In the case of strain measurements, use is also made of the results from a static test to destruction 3 made on a third fuselage under pressure alone.

2. The Test Specimen. The fatigue tests were made on the fuselage of Comet I G-ALYR, (Fig. 1). This aircraft was built in 1952 and had made 747 pressurized flights. (N.B. All numbers of

pressure cycles are totals, i.e., service plus test cycles.) The diameter of the fuselage was 10 ft 3 in.

and the pressure cabin was 70 ft long.

2.1. Basic Structure. The basic structure consisted of circumferential f rames 21 inches apart, stringers at approximately 5.5 in. pitch,, and the skin covering. The frames were of zed section,

2.75 in. deep and notched to take the witch-hat stringers which were bonded to the skin (Fig. 2).

The skin was generally 22 s.w.g. (0. 028 in.) except along the sides of the fuselage, where 20 s.w.g. (0. 036 in.) skin contained the windows (Fig. 3). Skin material was D.T.D.546 (Appendix).

Attachment between the skin-stringer panels and the frames was usually by 2 B.A. countersunk-

head bolts at the stringer flanges only. In the centre section, however, the skin was additionally

riveted to the frames (Fig. 4), in the region of the windows.

* Previously issued as R.A.E. Report No. Structures 257--A.R.C. 22,270.

2.2. Local Structure at Windows and Escape Hatches. With the exceptions of the two forward

escape hatches, which interrupted a circumferential frame (Fig. 6), the windows and escape hatches

were positioned between the frames. The windows and escape hatches were rectangular, their relative sizes being:

Window: 16.6 in. wide x 14 in. high, corner radii 3 in.

Escape hatch: 19.0 in. x 21.5 in. high, corner radii 4 in. (see Fig. 5).

The apertures were reinforced by peripheral members of zed section bonded to the skin with

Redux adhesive and additionally riveted by ~ in. countersunk-head rivets at the corners (Figs. 7

to 11).

3. Method of Test. Supported on its wing centre section in a tank, the fuselage was filled with

water and completely submerged in water, internal pressure being applied by pumping in more

water. The cycling action was controlled by pressure switches. No loads other than those due to internal pressure were applied. The cycles were repeated

pressure cycles of the form given in Fig. 12; the peak pressure was 8.25 p.s.i, and the loading cycle

took about 65 seconds. Detailed visual inspections were made at frequent intervals. When a fatigue crack was found its

subsequent development was observed continuously with an inverted periscope. When it was judged that a fatigue crack would soon develop catastrophically the affected aperture was repaired

to prevent excessive dan:age to the specimen (Fig. 13). By means of resistance-wire strain gauges, strains were measured at selected window and escape-

hatch corners before the fatigue test was started.

4. Results. 4.1. Measured Strains. Readings from strain gauges positioned at the corners of the

third starboard window and the forward port escape hatch were taken at increments of pressure.

Stresses were deduced for a pressure of 8.25 p.s.i. (Figs. 14 to 19). The highest stresses are as

given in Table 1.

4.2. Fatigue Test. A total of 11,319 pressure cycles of 0 to 8.25 p.s.i, to 0 were applied to the

fuselage. Fatigue cracks occurred in the skin at the corners of nine windows and two escape hatches,

sixteen corners being affected (Table 2 and Fig. 20). No fatigue cracks occurred at the A.D.F. aerial hatches (at which the 22 s.w.g, reinforcing plates were subsequently removed for examination

of the skin underneath), the crew and passenger doors, or at the freight hatches.

The first crack was seen at 5,248 cycles at the third window on the port side, i.e., the window just

forward of the rear spar frame, and by 8,941 cycles all six windows in the centre section had fatigue

cracks at one or more of their corners. Nine fatigue cracks were observed continuously throughout their growth; six at windows in the

centre section, one at a window in the aft section, and two at the port forward escape hatch. The fatigue cracks originated at the rivet holes at the aperture corners, not at the aperture edges, and when first seen were usually about 0.25 in. long. Development away from the aperture was initially about 1 in. in 500 pressure cycles, and, as all the windows were located between frames, the growing crack invariably had to cross a frame when approximately 4.5 in. long. When the cracks had spread 2 in. or so past the frames, i.e., were about 6.5 in. long, they were judged to be critical in that the

application of a few more cycles would cause a catastrophic failure.

2

Differences of behaviour occurred during growth towards the frames and across the frames.

During the first stage, several cracks became critical when about 3 in. long, i.e., between the aperture

and the frame. This condition was noted at most of the windows in the centre section where the

skin was also riveted to the frames. The presence of this extra attachment appeared to have a strong

influence in delaying crack growth across the frame, and was clearly demonstrated in the case

where a crack extended a distance of 2 in. in one pressure cycle (Fig. 22). An exception however,

occurred at the port forward escape hatch (where the riveted frame was a partial frame only--Fig. 6)

when the crack at the bottom forward corner caused a catastrophic failure when 2-75 in. long (Fig. 6).

The growth of one crack was observed in detail at a window in the aft portion of the fuselage

where the skin was not riveted to the adjacent frame. No critical stage occurred at 3 in., the crack

grew uninterruptedly across the frame to a critical length of 7.1 in. This behaviour was also shown in a crack at a window in the same section of G-ALYU a (Fig. 30).

In crossing the frames the cracks behaved in various ways:

(1) The frame with the normal bolted attachment appeared to have no influence whatever on crack growth.

-(2) Where the skin was riveted to the frames:

(a) Cracks passing between rivet holes were slowed down, but not stopped altogether.

(b) Cracks entering rivet holes were stopped temporarily; e.g., one crack was contained for more than 1,800 cycles.

Development beyond the frames progressed for about 2 in. when it was evident that catastrophic failure was imminent. Generally it was possible to stop the test before the fast-running stage, but

four catastrophic failures did occur either because of misjudgment of rates of growth or of the difficulty involved in observing more than two cracks at the same time.

Table 4 summarises the data on critical crack lengths. Curves of crack growth are plotted in Figs. 21 to 29; photographs of typical cracks are given in Figs. 31 to 36 (Table 3).

5. Discussion. 5.1. Origins of the Fatigue Cracks. All the fatigue cracks originated at the counter-

sunk rivet holes in the skin at the window and escape hatch corners. Those cracks which eventually

became catastrophic started at outer-row rivet holes. The few cracks that originated at holes in the

inner row grew inwards to the edge of the aperture and did not become catastrophic. No cracks originated at the edges of the apertures.

As indicated by the strain measurements, the stress at the corner of an ~ aperture attained its peak

value at the edge; at the outer row of rivet holes it was about 20,000 p.s.i, or perhaps half the stress

at the edge. The presence of a sharp-edged (countersunk) rivet hole in a high stress field might,

however, increase the stress locally, perhaps by a factor of 3, and, in addition, there would be a

certain amount of fretting action, so it is reasonable to expect fatigue cracks to be initiated at the rivet holes.

5.2. Locations of the Fatigue Cracks. The test on G-ALYR showed that fatigue cracks were initiated earliest and most numerously at the windows in the centre section, and though the first

failure in G-ALYU was at a forward escape hatch 4, fatigue cracks had also formed at several windows in the centre section by the time of this failure (Fig. 37).

That the fatigue cracks should occur first at the corners of the apertures is easily explainable in that the general level of stress there is some two to three times that found elsewhere in the fuselage.

The tendency for cracks to occur first in the centre section may be explained by a combination of

three reasons. First, the average stress in the skin at the corners of the windows would appear to be some 20 per cent greater than at the corners of the escape hatches, as is shown by the strain measurements made on three different fuselages (Table 5); in this connection it should be noted that

the radius at the corner of a window is smaller, 3 in. compared with 4 in. for the escape hatch.

Second, it is possible that there were aggravating distortions in the centre section of the cabin due

to the reaction of the internal pressure by the floor instead of by a complete cylinder as elsewhere.

Third, the effect of previous service use should not be forgotten, in that the shear stresses from

the usual flight and ground loads are highest in this part of the fuselage.

5.3. Propagation of the Fatigue Cracks. Many of the cracks when first observed were about

0.25 in. long. This is relatively short, but it is pointed out that the conditions for observing cracks

during the test were exceptional as they were anticipated at the corners of the apertures and the

paint was removed for easy inspection. It is problematical whether under normal service conditions

the cracks would have been detected so early. Even if it were certain that cracks of this length could be found easily, the curves show that a

crack length of 0.25 in. corresponds to about 90 per cent of the total life when no remedial action

is taken. Coupled with this fact is the indication that the rate of propagation is probably greater in

service than on test, since the crack measured at a window in G-ALYU, with its more representative

loading 2, developed 2 to 6 times as fast as the cracks in G-ALYR. The delaying effect of the adjacent riveted frame provided a temporary barrier when the cracks

were about 4 in. long. Nevertheless the opinion was formed that special inspection procedures would have to be used to ensure reliable detection. In this connection inspection would be greatly eased if

the fuselage were partly pressurized to open up the cracks. Comparison of these findings with observations from tests on flat sheets and simple cylinders

shows at once that the flat-sheet tests were unrealistic because of the absence of radial pressure. There appears to be some measure of agreement between the window-corner cracks and those

induced in 12 ft diameter cylinders in that critical lengths and numbers of cycles to failure were of

the same order for a roughly comparable nominal stress cycle, but upon consideration of the differing

conditions between the window corners and in the simple unstiffened cylinders, such agreement

is perhaps fortuitous.

5.4. Interpretation of the Fatigue-Test Results. The meagre data make reliable analysis

impossible, but certain features need comment. First, the initial failure in G-ALYR occurred at approximately twice the number of cycles as that in G-ALYU. Second, a closer grouping is evident

of the failures in G-ALYU compared with G-ALYR (Fig. 37). Third, rate of crack growth in

G-ALYU was about four times that in G-ALYR. From this evidence it appears that fatigue performance is adversely affected to an appreciable

extent by other-than-pressure loads reaching the cabin, an effect already noted by Walker 1, z. This means that where accurate life estimates are required to be obtained from tests the general flying

and landing loads should be reproduced as faithfully as practical considerations allow. Furthermore,

since perfection in this respect is unlikely, an allowance must be made for inadequacies of

representation.

Y ~

6. Concluding Remarks. This Report contains material from which various conclusions may well be drawn, and especially if combined with later work. Three points, however, appear to be established.

(i) The simplifications in fatigue loading which are generally accepted to make a full-scale test practicable are likely to give a longer life than would be realised in service.

(ii) Tile attachment of reinforcing material inevitably introduces its own stress concentrations; the example of the countersunk rivet holes at the window corners illustrates this important principle.

(iii) Nearly all the fatigue life associated with a particular crack may have been expended by the time the crack first becomes noticeable.

No.

1

Author

P. B. Walker

P. B. Walker

P. B. Walker

H.M.S.O.. .

REFERENCES

Title, etc.

Pressure cabin fatigue. 5th International Conference, Los Angeles, 1955.

Pressure cabin fatigue. Aircraft Engineering. Vol. XXVIII. January, 1956.

Static strength tests of a Comet I pressure cabin. A.R.C. 18,359. December, 1955.

Civil Aircraft Accident. Report of the Court of Inquiry--C.A.P. 127. December, 1955.

6

APPENDIX

The Materials Used in the Structure

. D.T.D.687A.--Clad, high-tensile aluminium-alloy sheet.

(i) Chemical composition (nominal)

Copper 0 .4 per cent Zinc 5.3 per cent

Magnesium 2 .7 per cent

Manganese 0.5 per cent

Aluminium The remainder.

(ii) Heat treatment Quenched after 2 to 10 hours at 455 to 465 deg C.

Aged 4 to 30 hours at 120 to 140 deg C, or appropriate to suit requirements.

(iii) Strength properties (a) 0" 1 per cent proof stress: not less than 27 tons/sq in.

(b) Ultimate tensile stress: not less than 32 tons/sq in.

2. D.T.D.610

(i)

(ii)

.

Chemical composition Copper Not less than 3

I ron

Silicon

Magnesium

Manganese

Ti tanium Aluminium

Heat treatment

• 5, nor more than 4 .8 per cent Not more than 1.0 per cent

Not more than 1.5 per cent

Not more than i . 0 per cent

Not more than 1.2 per cent

Not more than 0.3 per cent The remainder.

Quenched from 500 to 510 deg C at 2 to 4 hours. Aged 5 days at room temperature (W condition).

(iii) Strength properties (a) 0.1 per cent proof stress: not less than 14 tons/sq in. (b) Ultimate tensile stress: not less than 24 tons/sq in.

(c) Elongation: not less than 12 per cent on sheets up to a in. thick.

D.T.D.546B.--Clad, high-tensile aluminium-alloy sheet.

(i) Chemical composition. Same as for D.T.D.610.

(ii) Heat treatment Quenched from 500 to 510 deg C at 2 to 4 hours.

Aged at 155 to 205 deg C for an appropriate time. (WP condition.)

(iii) Strength properties (a) 0.1 per cent proof stress: not less than 20 tons/sq in. (b) Ultimate tensile stress: not less than 26 tons/sq in.

(c) Elongation: not less than 8 per cent for sheets thicker than 12 s.w.g.

7

T A B L E 1

Highest Stresses Measured at the Edges of the Apertures for an Internal Pressure of 8.25 p.s.i.--G-AL YR

Aperture corner

Stress obtained by extrapolating aperture edge, p.s.i.

Highest 'strain-gauge' stress,, p.s.i. (see Figs. 14-17)

to

Stress obtained by extrapolating to aperture edge, p.s.i.

Highest 'strain-gauge' stress, p.s.i. (see Figs. 18-19)

Third starboard window

Top rear

47,700

40,800

Bottom rear

35,500

32,800

Rear aerial hatch

Port forward

33,000

32,850

Starboard rear

34,400

30,450

Forward port escape hatch

Top rear

40,000

33,250

Bottom rear

32,700

27,500

8

T A B L E 2

Chronological Occurrence of the Fatigue Cracks at the Aperture Corners

Aperture

Windows between

. spar frames (Centre section)

Other windows

I Escape hatch

Side of

fuselage

Corner of

aperture

Origin of fatigue crack

rivet hole

Crack length

I when first seen (inches)

Pressure cycles

3rd Port Bottom A 0" 20 5,248 forward

1st Port BottOm B 0- 50 6,542 rear

1st Starboard Top C 3 "409 6,90i forward

• 3rd Starboard Bottom D 0" 17" 6,901 forward

2nd Port Bottom C 0" 06 6,901 rear

6th Starboard Top E 0-14 ~ 6,901 rear (Rivet hole

oversize)

2nd Port Top C 0.10 6,959 forward

3rd Starboard Top C 1" 62 7,692 rear

2nd Port Top A 0" 04 8,564 rear

Forward "Port Bott6m B 0- 08 8,564 forward

4th Port Top A ,0- 51 8,941 rear

4th Port Bottom F 0" 70 9,225 forvcard

6th Port Bottom A 0" 06 9,350 forward

2nd Starboard Top A ,0- 7i ;10,016 forward

]Forward Port Bottom A 0- 31 10,016 rear

l%rward Starboard l Top '(2 '0-t0 1t,286 rear

O 0 0 0 a /

. , ~ 0 Oc

~ ' 0 0 ©

O O

- O F

°ou Rivet ,holes .at which the ,cracks .occur.red.

* Between rivet hole and edge of aperture. ]" Crack had spread to frame before it was discovered.

(83824) B

TABLE 3

Growths of the Fatigue Cracks

Location of the fatigue cracks

Window

Between Outside spars spars

3rd

1st

1st

6th

3rd

2nd

4th

2nd

Escape hatch

Forward

Forward

Corner

Bottom forward

Bottorh rear

Top forward

Top rear

Top r e a r .

Top forward

Bottom forward

Top r e a r -

Top forward

Bottom rear

Side

Port

Port

Starboard

Starboard

Starboard

Port

Port

Port

Starboard

Port

When first seen

Crack Pressure length

(in.) cycles

0" 20 5,248

0" 50 6,542

3" 40" 6,901

0" 14 6,901

1" 62 7,692

0.10 6,959

0"08 8,564

0" 16 8,941

0"75 10,016

0"31 10,016

i

When the crack had reached the adjacent frame

Crack length

(in.)

4"40

4"70

3 "40

3"15

3"35

Additional cycles

fromwhen first seen

700

359

0, had reached adjacent

frame when first seen

168

1,501

Did crack spread into

a frame rivet hole?

No

No

Yes

No, f r ame not

riveted

Yes

Yes

Growth in inches per cycle

0.009

0.021

Stopped at rivet

hole

Stopped at rivet

hole

Stopped at rivet

hole

Delaying effect

of frame in cycles

90

60

>1,820

240

890

i

Crack spread catastrophically before reaching the cir- cumferential frame--was about 2"75 in. long when failure occurred.

3' 40 669 No, frame No discontinuous kink not observable in crack

riveted growth curve

Crack grew to a length of 3' 50 in., but was not allowed to reach the adjacent frame, as the aperture was reinforced.

3 "35 1,189 No 0.1 appro-%

Length (in.)

6.25~

6.75~

6.41~

About 12

feet

5.85~

15"0

About 15

feet

7 '10

3 '50

Final details

Additional Growth cycles in inches

from when per cycle first seen

794 0-10

417' 0.06

2,040 0" 05

2,040

873 0'03

2,449 Instant- aneous failure

2,682 .Instant- aneous failure

725 Instant- aneous failure

229

Total pressure

cycles

6;042

6,959

8,941

8,941

8,564

9,350-

11,246

9,666

10,245

Remarks and action taken

Catastrophic failure imminent. Aperture repaired.

Catastrophic failure imminent. Aperture repaired.

Catastrophic failure imminent. Aperture repaired.

Catastrophic failure requiring major repair involving three star- board apertures.

Catastrophic failure imminent. Aperture repaired.

Catastrophic failure, only stopped by the patch around neighbouring window. Major repair.

Catastrophic failure from front spar frame (18) to between frames 8 and 9. Major repair.

Catastrophic failure in 9,666th cycle running from 7.10 in. to about 12 feet. Major repair.

Likely that crack would have grown further•

When this crack had reached the circumferential frame at 11,205 cycles, a reinforcing strap was riveted close to the frame to prevent further crack progress but at 11,246 cycles the crack at the forward corner caused a catastrophic failure.

* See Figure 25. Approximate critical length.

10

T A B L E 4

Critical Crack Lengths

Aperture

Port escape hatch (Bottom for- ward corner)

Port escape hatch (Bottom rear corner )

First window, port side

First window, starboard side

Second window, port side

Second window, starboard side

Third window, port side

Third window, starboard side

Fourth window, port side

Sixth window, port side e (G-ALYU)

Location

Forward section

Forward section

Centre section

Centre section

Critical crack length between the aperture

and the adjacent frame (in.)

Critical crack length attained during final

development (in.)

2 .75--at this length the crack caused a catastrophic failure

2.80

2.45

Crack had spread to the adjacent frame before any measurements were taken

No result here because of the occurrence ofthe above failure

6"75

6"41

Centre section

Centre section

Centre section

No observable critical stage

No observable critical stage

3.10

No final length as this crack spread catastrophi- cally from the frame

This aperture was repaired before crack had spread to the adjacent frame

Centre section

Aft section

Aft section

No intermediate critical stage occurred

No intermediate critical stage occurred

No intermediate critical stage occurred

6"25

5'85

7"10

5"60

A previous fatigue test in which wing loads were applied as well as pressure loads.

11

(88824) C

T A B L E 5

Measured Stresses at the Edges of Apertures in Three Fuselages, for an bzternal Pressure of 8" 25 p.s.i.

Port forward Starboard forward Aerial hatch * Third window escape hatch escape hatch

Fuselage Strain Extra- Strain Extra- Strain Extra- gauge polated gauge polated gauge polated

Strain Extra- gauge polated

28,000 32,200

32,850 34,400

30,400 34,000

G-ALYU 23,000 t

33,250

34,000

26,500~

39,600

38,800

34,500

No reading

31,600

No reading

No reading

36,300

35,100

40,800

38,300

38,000

G-ALYR

G=ANAV

47,700

43,500

* The strain gauges attached at this aperture were cemented to the 22 s.w.g, reinforcing plate riveted over the skin.

t" These stresses were measured after the escape hatch had been repaired following a catastrophic failure and may differ, therefore, from those that would have been measured in the original structure.

12

C.3

C~EW ~OOR

ATTACHMENT TO WIN~ AT F'RAME6 18 AND P-6

AERIAl HATCHES /

WINDOW8 ~.3.4.5 AND "7

I0 II 14, I',5 I 7 IS 18 25 86 27 ~ ~9 31 32 33 34- 35 36 3"7 3S

FREIGHT HATCHES

FIC. 1. The fuselage of Comet G-ALYR.

,. ........, o,, !

=uC

O3

20SW.Ct. O.T.D. 610. IL

FRAME RESTS ON THE STRINGER FLANGES, ELSEWHERE THERE IS CLEARANCE. B' ..~; ----.;WL

S~KIN

~O s.w.cq. ID.T.O. 687A.

~'X'R ED U X 3-OIINT.

Fro. 2. Typical frame and stringer sections.

SPAR F R A M E 18 /

OF 5KIN JOINT.

13 ~ 5 ~ C ] ' _ _ \ ~ "~¢:... OF ,SKIN ...TOINT.

THIS SKIN PANEL ~05.W.G. D.T.D. 546.

SPAR F R A M E ~6 /

C~5 ~ 6

FIG. 3. Skin panel containing windows and escape hatches.

14

ge (No.:,,)

ledio~ I ment

bolted at r fkmge$

hat' r

Fzc. 4. Attachment of the skin to the frames.

Skin removed

Escape hatch frame Window frame

FIQ. 5. Relative sizes of window and escape-hatch apertures.

15

(83824) D"

. . . .

Circumferential fram~ 16 S.W.G . - U-section partial frames

FIG. 6. Internal partial frames at forward escape hatches.

Failure at port forward escape hatch at 11,246 Cycles.

LOCATION OF SECTIONS

I II

~SKIN / ~ DIA RIVET C'SUNK -90 °

. . . . . . . . . . . . . . . . . . . . / . . . . . . APERTURE FRAME i6 s.w.¢- ~.tD. olo) ~EBUXED TO THE ~ I N

STRIP REDUXE m TO THE FRAME

m

.L FIG. 7. S E C T I O N THROUGH A WINDOW F R A M E .

/~" 'OIA. ~wT, C'SUNK so ° / S K I N A ~. APERTURE FRAME

] /.~ SEAI:IN~ 5TRIP REDUXED

r

FIG.& SECTION THROUGH AN ESCAPE HATCH FRAME.

Fits. 7 and 8. Sections through window and escape-hatch frames.

17

OO

EDGE OF REDUX. ~TOINT

- - / - - - O - - - O - - 0 0 o - O . "

,, o 0 6

/ Q i I I

DT"

4 B.A. BOLT AT 5TR. FLANGE. \

ul

<

h_ J

b- Z

p ADDITIONAL RIVETED ~L~ ATTACHMENT AT

WINDOWS P-~3 AND | 4 ONLY.

EDGE OF

O

RE DUX JOINT.

"" -..

0 \ \

0 \ \

o

0

\

ol I I

~-0"

W I N D O W C O R N E R E S C A P E H A T C H

RIVETS

4 B.A. BOLT AT STRINGER FLANFaE.

la_

J

z

< THIS ADDITIONAL ATTACHMENT AT THE FORWARD ESCAPE HATCHES ONLY.

CORNER

J_ ii ARE B DIA. CSK, HD.

FIG. 9. External structural details at w indow and escape-hatch-aper ture corners.

Fuselage skin Circumferent ial f rame

R e d u x / joint between aperture frame and skin

frame

" " °-: . Section X -X

h

FIc. 10. Local structure at a window corner, viewed from outside with skin removed.

-- / /

f /

/ /

Q

O . f -S Q ' .

: i I . . ~

R4zdux I . . . .

joint between aper tu re

k - - ~ n n e r row o f ~ r ivet holes

Outei- row of r ivet holes

0

2

3

-4

-5

6 Inches

5~ct ion X - )

FIG. 11. Local structure at an escape-hatch corner, viewed from outside with skin removed.

t I o Z - 6 O -..>.

_i

n, 4 D if) If) M nt O_

l.d

.J ILl LO

h

8

PRESSURE RELEASE

0 0 ZO 40 60 80

T I M E IN 5EC01ND5 ~-

LEAKAGE OF FU51ELA~E

I00

FIG. 12. The pressure cycle.

21

18 S.W.G. P late 16 S.W.G. Plate

t O ix)

£,.!;

, ~ , . , & - o , - ° o ."

I : ~ * . ~ . . . . . . . - ~ - - °

" Or

-

,~ x / . . - ',.at

' , . r id

FIG. 13. Typical window repair.

¢,

A F T

0 ~z

INNER EDqE OF WINDOW FRANE

a,

O - - - 0 0 0

I © so, oDD 47,700 EXTRAPOLA

I 0 4O)OO0 IqHEST ' STIll

Z ~ ~' 3o, oao ~ i

z_

l0 NX~-- - - INNER E r OF

WINI~OW FRAME

O I 2. RADIAL DISTANCE FRQP1

GAU~E'3TRE~S

EDgE- INCHE..S

FIG. 14. Stresses in the skin at the top corner of the third starboard window for an internal pressure of 8-25 p.s.i.

oo

w N o o w A M E

4o, ooo I i - - . . . ~ o o - O r e - -

EXTRAPOLATED EDGE STRESS. I I AFT

__ HIGH~'ST "STRAIN "~ 3 0 , O O O ~ - - G A U G E " 5TRESS.

O2 ]

z \ '~

-- \ ' t uJ IO~OOO ~- \ \ [

~ \ I . / I N N E R EDGE OF M" I WINDOW FRAME

\ \ 1 ] O RADIAL DISTANCE FROM

EDGE - I N C H E S .

FIG. 15. Stresses in the skin at the bottom corner of the third starboard window for an internal pressure of 8.25 p.s.i.

b ~

Ag'T >

O

4OtOOO - -

8o, ooo

i

.z_

~0~000 z

I0,000

0 ~ ~ /.INNER EDt~E OF ESCAPE HATCH FRAME

O

EXTRAPOLATED EDgE .STRESS ~ 0

co ,I io - - - H I q H E S T ..~TRAIN ~AUr~E' STREE6 0

"--_____.

j l N N E f EO~EOF ESCAPE HATCH FRAME

O I a RADIAL DkSTANCE FROM

EDgE- INCHES

O

%

FIC. 16. Stresses in the skin at the top corner of the forward port escape hatch for an internal pressure of

8.25 p.s.i.

40,000

~oo EXTRAPOL!TED \ \ Eo~E ~TRESS

30;ooo ~H,OHES.'~'T.A,. \ \~0A°OE' S,~E,E. 0

20,000 ~ . ;

10,OOO ~ . ~ ] ~ ' ~ j . i N- - i N -~N EN E R E Dr4E 0 F ]tu" ('~ I ~2~~ E HATCH FRAME. / q" " J i

i

C) I :~

0 RAO,AL DISTAINCE. FROM / / EDrqE - INCHES. / ~ o 0 ~ I I

oo/

0 0 O

0 0 ~ " ~ . - - - - - ~ INNER EIDGE OF

ESCAPE HATCH FRAME.

AFT

FIG. 17. Stresses in the skin at the bottom corner of the forward port escape hatch for an internal pressure of

8.25 p.s.i.

b~

J~ S, ?'50

&o ooO / © © , O © O

" ~ 0 O 4- ~,~oo

/0 O / \ , /oO/ \ /oO \

O O Io

INNER EDgE OF AERIAL HATCH FRAME

© © © 0 . 0 0

4O~OOO

FIG. 18.

~ / ~XTRAPO-, EDGE ~TRESS

Y / A 30) 000

20)000 ~ ~ %

io, ooo

INNE~ E£)~E OF AERtAL -- HATCH FRAME

O RAOIAL DISTANCE. FROM EO(~E - INCHES

Stresses in the skin at the port forward corner of the rear aerial hatch.

°°°°°~1 I 34,400 EXTRAPOLATED EDGE STRESS.

3O ooo a: '

,

Z

~ ~O~DOO

IO~OOO

\ \ ,-j

o e G RADIAL DISTANCE FROM ] O /

EDGE-INCHES. / C ) /

"'~'- ~ F-", ~ AERIAL HATCH - - ~ _ + ~ - © . ~ / \ FRAME.

° °-L°~ 0 / " 0 O ~ "~ ,?

FIG. 19. Stresses in the skin at the starboard rear corner of the rear aerial hatch.

E.H.

FRONT SPAR FRAME 18 ~- - ' -

I 8,941

I

j CENTRE

10~016 \

SECTION REAR 5PAR FRAME ~6

10~245 8~,564 7:'i9 ~

/ 6,901

8,941 8~94-1 i OI

8,941

STARBOARD SIDE

bO

[ 8,564

E.H.

11~'46

10~016

6~959

6~54 ?

NOTES :-

6,959 \

9,350 8164

6~901 5~74-8

6~04E ./ J 9;~?_ 5

9,666 1941'

(I) NUMBERS AT CORNERS ARE THE PRESSURE CYCLES WHEN THE FATIGUE CRACKS WERE FIRST OB3ERVED

(2) NUMBER5 INSIDE ARE THE PRESSURE CYCLES WHEN THE APERTURE WAS REPAIRED OR REINFORCED AND THEREFORE ELIMINATED FROM THE TEST.

9~666

WINDOW

/J 97666 9~350

FIC. 20. Distribution of fatigue cracks at window and escape-hatch-aperture corners.

ESCAPE HATCH

97666

PORT SIDE

v

b.) ---I

[ FORWARD

• ° / - 7

!o ° "o\ / / © 0 ~ "

SKIN FATI~LJE CRACK A / ' /

O

4 3

D ISTANCE FROPI

FIG. 21.

8,~00 PRESSURE CYCLES - -

/< 6,000

5,400 I

! , g, aO0

~- o

CENTRE- LINE OF RIVET ' A t IN INCNES Crack growth at third window, port side.

(Bottom forward corner)

>O @, \ / 0 0 O~

~ ' ~ ~ ~ . o' .~--______ - - ~ ~ ~

O

~ 6 , E I

i o PRESSURE CYCLES

, , , f

- - 6 , 4 0 0 3

DISTANCE FROM C E N T R E - L I N E OF R IVET ~J~ - - INCHE5

FIG. 22. Crack growth at first window, port side. (Bottom rear corner)

f

bO CO

I

O ~) O~ 0 ~ ( : :::o ; % ,~

/ A T 8,478 CYCLES "\

AILED AT 9;350CYCLE3.

+ kJ / ;-'ORWARD

9,500

CRACK OEVELOPED I "]--~ CATA 5T rO FH I C ALLY

AT 9~350 CYCLE&

97000

, CRACK REAC"EO d I RIVET HOLE , ~"

BISO 0 0

7~500

i °° G~500

I DISTANCE FROM CENTRE - LINE OF"

RIVET 'C * - iNCHES.

FIG. 23. Crack growth at second window, port side. (Top forward

corner)

B

O A

FINAL LENCTTH OF CRACK : 3"50 (WINDOW REPAIREO AT THI.S ~TA~E)

- - I O ~ ;

- - I o ~ o o o

DISTANCE FROM CENTRE- LINE OF RIVET 'A ~ INCHES

FIG. 24. Crack growth at second window, starboard side. (Top for-

ward corner)

e~

~O w

I O

O

O

ba

O 5

FORWARD

1 o©

1 - ~ © ' . <

I T ~ o c R ! C A L LENGTH = G'41" D

/ol

" ' 8,500

~J 8~000

7,500

__-2°1° o l ~ LEd tEl"i~H OF CRACK

/

I T O]$TANCE FROM CENTRE - LINE OF RIVET ' C ' -

°l o

<

o

o

<

WHEN FIRST 5EEPt =3"40".

3 4 iNCHES.

FIo. 25. Crack growth at first window, starboard side. (Top forward corner)

FATIG~UE CRACK IN 5KIN

//____<> CI~ITICAL LENC-I'I'H = S '85"

AQ O

O

©

©

F I G . 2 6 .

\

\

® I

I

\\

"4-.

600 PRESSURE CYCLES

8, L O 0

0 0 - -

O0

O0

7,6o0 ,q- 3 ~. I

DISTANCE FROM CENTRE-LINE OF RIVET C IN INICHES

Crack growth at third window, starboard side. (Top rear corner)

O

O

O

O

% FORWARD

©

4 3 ~:

~RITICAL LENGTH ~.75"APPIOX.

CRACK i 0 0 " ~

AT 1l,~46 CYCLES. I "-"'~""k i, o

1 )7i00 - -

9,000 ~--

8 ~ 5 0 0 - -

DISTANCE FROM CENTRE-LINE OF RIVET 'B' -- INCHES.

FIG. 27. Crack growth at forward escape hatch, port side. (Bottom forward corner)

AFT / O

Oc • o , O

x d x 0 \_ , ~D--n

Q B -- __ ~ ~ CATASTROPHIC FAILURE I "~----~-.-- TOOK PLACE" ALONE / ~

/ / ~ THIS PATH / \

t O~ DISTANCE FROM #E.NTRE- LIN~" OF RIVET ~A ~ - INCHES

FIG. 28. Crack growth at forward escape hatch, port side. (Bottom rear corner)

9,700

A A F T i

0 c

AL LENGTH = 7"10"

SKIN FATI6UE CRACK.

9,6Q0

r J

u

9,SOO ~" o / M

9,400

I

9'TO0

9~200

FIe. 29.

. j l f J J CYCLE 9,66

I A CATASTROPHIC FAILI OCCURRED AT THE 9~G66~. PRESSURE CYCLE.

I

I I

4 5 6

DISTANCE FROM CENTRE- LINE OF RIVET 'C'. - INCHES.

Crack growth at fourth window, port side. (Top rear corner)

() o

<

()

SKIN FATIGUE CRACK

CF, TICAL LENGTH = 5"6" 7 - -

Forward

• ]:

/ ':. ! • f,4,-xl

iJ

~. ~i~,~. ~? • % , , ,

.~.~,?~ 2 .~ :~

Fzo. 31. Third port window corner, 5,951 cycles.

Crock running between rivet holes

\ -- Forward

FIG. 32. Third port window corner, 6,042 cycles.

32

- - Forword

• ..;~" . . ¢ j

FIc. 33. First port window corner, 6,901 cycles.

Forword w

• ' • o • N

o E L i t . ¸

I ' "

B ° • •

o

FIG. 34. First port window corner, 6,959 cycles.

33

ForwQrd =

Fxc. 35.

Crack stopped at rivet hole

e, C

Crack stopped at rivet hole in frame, 6,901 cycles. First starboard window.

Forword

u

• * -4" 4 u , ..

D

i , m

~Q LL., -

FIG. 36. Cracks stopped at riveted frames, 8,564 cycles. Third starboard window.

34

.q

SERVICE HI.STORY

I ~ - ALYU -- IP_31 FLIGHTS

r ~ - ALYR -- -/4"7 FLICxHTS

[EACH POINT REPRESENT5 THE FIRST FATIGUE CRACK DETECTED AT A PARTICULAI~ APERTU/~E ]

ESCAPE HATCHES +

WINDOWS IN AFT SECTION

WINDOWS IN CENTRE SECTION + t"

0

"1-

I I I I I~ 0 0 0 2~000 8)000 4",000.

+ + +

+ -t- i - •

-t- -he + •

w

I I I S) 0 0 0 , 6~ 0 0 0 "7, 0 0 0 8 ~ 0 0 0

@ @

KEY J + = i ~ - ALYU

L e = CT- A L Y R

I I ,9,000 I0,000

PRESSUt~E CYCLE8

I I1~000

FIG. 37. Distribution of fatigue cracks at apertures.

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