Figure: Distribution of the fatigue cracks (from STA 2060 to STA 2120), from ASC-AOR-05-02-011, Figure 1.16-12

Eddie Sez:

There is a term in the safety business that isn't in a lot of literature because it illustrates a level of futility many of us are reluctant to face, and that term is "assumed safety." We assume a lot as pilots before releasing the brakes prior to takeoff. We assume the airplane was properly designed, maintained, and serviced. We assume those we share the skies with are also properly equipped. But sometimes those assumptions are wrong. Corporate aviation pilots can get into the nuts and bolts better than airline pilots, but we too are relying on a lot of other people to do their jobs properly. Sometimes we are just passengers in airplanes that are ticking time bombs. Case in point: China Airlines CI611.

This airplane had a tail strike that severely weakened part of the skin in the aft lower section. The airline did not repair the damage properly and failed to follow up with the necessary inspections over the years. I've left out much of the inspection history because we get enough of what went wrong through the repair. I've presented enough of the nuts and bolts to help us pilots understand what we should be looking for when inspecting damage repairs. A very thin crack or even a scrape can become a tear large enough to bring down the airplane. These imperfections are especially dangerous when around corners and holes. The repairs are painted over and sometimes there is nothing for you to see, as in this case. But you might see something that gives you a hint.

The Aviation Safety Council did an excellent job piecing this all together but I would have written the probable cause a little differently. I believe the cause was an airline management that encouraged short cuts to manufacturer repair procedures, mechanics who were either unaware they were not using correct repair techniques or agreed to these techniques despite their knowledge, and improper oversight from the Chinese authorities. This sort of maintenance malpractice is often deadly and in this case it was.

What follows are quotes from the relevant regulatory documents, listed below, as well as my comments in blue.


Accident Report


Narrative

Figure: CI611 radar track, radar sites, and debris field from ASC-AOR-05-02-011, figure 1.8-1.

[ASC-AOR-05-02-011, ¶1.1]

  • The takeoff and initial climb were normal. The flight contacted Taipei Approach at 1508:53, and at 1510:34, Taipei Approach instructed CI611 to fly direct to CHALI3. At 1512:12, CM-3 [the flight engineer] contacted China Airlines Operations with the time off-blocks, time airborne, and estimated time of arrival at Chek Lap Kok airport. At 1516:24, the Taipei Area Control Center controller instructed CI611 to continue its climb to flight level 350, and to maintain that altitude while flying from CHALI direct to KADLO 4 . The acknowledgment of this transmission, at 1516:31, was the last radio transmission received from the aircraft.

  • Radar contact with CI611 was lost by Taipei Area Control at 1528:03. An immediate search and rescue operation was initiated. At 1800, floating wreckage was sighted on the sea in the area 23 nautical miles northeast of Makung, Penghu Islands.

[ASC-AOR-05-02-011, ¶2.1] Based on the radar track data, the accident aircraft suffered an in-flight breakup as it approached its cruising altitude of 35,000 ft.


Analysis

Component Description

Figure: Description of the components of the lower lobe frame, from ASC-AOR-05-02-011, figure 1.6-1.

I flew the Boeing 747 for five years and was always impressed with its massive size. It is like seeing the Grand Canyon; photos do not do it justice. You get a false sense of security from the proportions. How can a ten inch crack bring down such a massive airplane? The key to understanding structures on any airplane is that every component works as part of the entire airplane. The skin itself plays a role in holding everything together.

[ASC-AOR-05-02-011, ¶1.6.1.2]

  • In the B747-200 fuselage, applied loads are supported by both the skin and by internal structure including frames, stringers, shear ties, and stringer clips.

  • The skin of the aircraft is constructed from sheets of aluminum alloy. The sheets are connected with lap joints and butt joints. Lap joints run longitudinally (along the length of the aircraft) and have one sheet overlapping the adjacent sheet. Butt joints run circumferentially (around the cross-section of the fuselage) and are constructed with a splice plate to which is attached both adjoining skin sheets. The butt joint is so named because the skin sheets butt up against one another but do not overlap.

  • Stringers are longitudinal stiffeners attached directly to the skin that run the length of the fuselage and are located around the periphery of the cross-section.

  • Individual fuselage frames are located approximately every 20 inches along the length of the fuselage and conform to the cross-section of the aircraft. The frames themselves can be considered as beams with an upper and lower chord separated by a stiffened web. However, because the entire frame is approximately circular in shape, the chords are referred to as the inner chord and fail-safe (outer) chord. The inner chord essentially defines the interior cross-section of the cabin while the fail-safe chord of the frame is adjacent to the stringers. The fail-safe chords are so-named because they serve to help carry cabin pressurization loads (hoop tension) should a longitudinal crack develop in the skin.

  • Shear ties connect fuselage frames to the fuselage skin and are located between stringers. Shear ties serve to transfer loads between the frame and skin and to transfer hoop tension loads from the skin to the frame fail-safe chord should a crack develop in the skin.

  • Stringer clips are located at frame/stringer intersections and serve to connect the frames to the stringers.

Tail Strike and Subsequent Repairs

[ASC-AOR-05-02-011, ¶1.6.2]

Examination of a Repair Doubler Between STA 2060 to 2180 and S-49L to S-49R

Figure: Exterior (up) and interior (down) of Item 640C1, from ASC-AOR-05-02-011, figure 1.16-1.

[ASC-AOR-05-02-011, ¶1.16.3]

  • Item 640C1 was a segment of Item 640 approximately from STA 2060 to 2180 and from S-49L to S-49R (Figure 1.16-1). A 23-inch wide, 125-inch long external repair doubler was attached to the skin by two rows of countersunk rivets around its periphery as well as by fasteners common to the stringer and shear tie locations. Universal head rivets were used at S-51R and S-49L while countersunk rivets were used at S-50L and S-51L.
  • A "faying surface" is the surface on a sheet of metal that faces another; once they are joined you cannot see the surface.

  • After disassembling the doubler from the skin and removal of the protective finishes, scratching damage was noticed on the faying surface of the skin. This damage consists of primarily longitudinal scratching distributed in an area of 120 inches by 20 inches. The most severe scratching typically occurred at the skin stiffening members such as skin stringers and body frame shear ties. Evidence of an attempt to blend out these skin scratches, in the form of rework sanding marks, was noted over much of the repair surface.
  • Scratching leaves stress points in the metal, even when the surface is sanded smooth.

  • Corrosion was noted at several shear tie locations on the skin inboard surface sometimes penetrating completely through the skin thickness.

  • One additional observation described in the BMT [Boeing Materials Technology Laboratory] report is the large percentage of the overdriven rivets on the repair doubler. Out of 402 rivets, 267 were found overdriven (66%), 15 were under driven (3.7%), and the rest 120 appeared to be normal (29.8%).
  • A rivet that is underdriven could allow movement of the joined surfaces and create undue stress at the rivet. A rivet that is overdriven can weaken the structure around the rivet.

[ASC-AOR-05-02-011, ¶1.17.2.3.1]

[ASC-AOR-05-02-011, ¶2.3.1.1]

Evidence of Rapid Depressurization

[ASC-AOR-05-02-011, ¶2.2.2]

[ASC-AOR-05-02-011, ¶2.2.6]

Summary

[ASC-AOR-05-02-011, ¶2.2.8]


Probable Cause

[ASC-AOR-05-02-011, ¶3.1]

  1. Based on the recordings of CVR and FDR, radar data, the dado panel open-close positions, the wreckage distribution, and the wreckage examinations, the in-flight breakup of CI611, as it approached its cruising altitude, was highly likely due to the structural failure in the aft lower lobe section of the fuselage.

  2. In February 7 1980, the accident aircraft suffered a tail strike occurrence in Hong Kong. The aircraft was ferried back to Taiwan on the same day unpressurized. and a temporary repair was conducted the day after. A permanent repair was conducted on May 23 through 26, 1980.

  3. The permanent repair of the tail strike was not accomplished in accordance with the Boeing SRM, in that the area of damaged skin in Section 46 was not removed (trimmed) and the repair doubler did not extend sufficiently beyond the entire damaged area to restore the structural strength.

  4. Evidence of fatigue damage was found in the lower aft fuselage centered about STA 2100, between stringers S-48L and S-49L, under the repair doubler near its edge and outside the outer row of securing rivets. Multiple Site Damage (MSD), including a 15.1-inch through thickness main fatigue crack and some small fatigue cracks were confirmed. The 15.1-inch crack and most of the MSD cracks initiated from the scratching damage associated with the 1980 tail strike incident.

  5. Residual strength analysis indicated that the main fatigue crack in combination with the Multiple Site Damage (MSD) were of sufficient magnitude and distribution to facilitate the local linking of the fatigue cracks so as to produce a continuous crack within a two-bay region (40 inches). Analysis further indicated that during the application of normal operational loads the residual strength of the fuselage would be compromised with a continuous crack of 58 inches or longer length. Although the ASC could not determine the length of cracking prior to the accident flight, the ASC believes that the extent of hoop-wise fretting marks found on the doubler, and the regularly spaced marks and deformed cladding found on the fracture surface suggest that a continuous crack of at least 71 inches in length, a crack length considered long enough to cause structural separation of the fuselage, was present before the in-flight breakup of the aircraft.

  6. Maintenance inspection of B-18255 did not detect the ineffective 1980 structural repair and the fatigue cracks that were developing under the repair doubler. However, the time that the fatigue cracks propagated through the skin thickness could not be determined.

See Also:

Abnormal Procedures & Techniques / Maintenance Malpractice


References

Aviation Occurrence Report Volume I, ASC-AOR-05-02-011, In-flight Breakup Over the Taiwan Strait Northeast of Makung, Penghu Island, China Airlines Flight CI611, Boeing 747-200, B-18255, May 25, 2002, Aviation Safety Council

Aviation Occurrence Report Volume II, ASC-AOR-05-02-001, In-flight Breakup Over the Taiwan Strait Northeast of Makung, Penghu Island, China Airlines Flight CI611, Boeing 747-200, B-18255, May 25, 2002, Aviation Safety Council