Failure of a Helicopter Main Rotor Blade

Lawrence Kashar, Kashar Technical Services, Inc.


From: Handbook of Case Histories in Failure Analysis, Vol 1, K.A. Esakul, Ed., ASM International, 1992

Abstract: A Marine Corps helicopter crash was investigated. Efforts were directed to the failure of one of the main rotor blades that had apparently separated in the air. The apparent failure of a blade integrity monitor (BIM) system was also considered. The rotor blade comprised a long, hollow 6061-T651 aluminum alloy extrusion and 26 fiberglass “pockets” that provided the trailing-edge airfoil shape. Visual examination of the fracture surface of the aluminum extrusion indicated fatigue crack growth followed by ductile overload separation. Examination of the fatigue fracture region revealed several pits that appeared to have acted as fracture origin sites. Time to failure was determined using fracture mechanics. It was concluded that failure was caused by a fatigue crack that grew to critical length without detection. The crack originated at pits that resulted from the use of an improperly designed heating element used to cure fiberglass repairs.

Keywords: Extrusions; Fatigue failure; Military planes

Material: 6061-T651 (6xxx series, wrought aluminum-magnesium silicon), UNS A96061

Failure type: Fatigue fracture


Background

About an hour after takeoff, a Marine Corps helicopter crashed, killing all on board. In the ensuing investigation, a large portion of one of the main rotor blades was found about a mile from the crash site, indicating that it separated in the air.

Applications

Because this helicopter was designed for use by the U.S. military, its survivability under hostile fire was a design concern. The main rotor blade was designed to withstand hits, and a system was included that would indicate if the blade had been damaged. This system, referred to as the blade integrity monitor (BIM), took advantage of the hollow 6061-T651 aluminum extrusion used as the spar of each blade. The internal cavity of the extrusion was filled with nitrogen at a slight positive pressure, and a visible pressure indicator was placed on each blade near the hub. If the structural portion of the blade was punctured by hostile fire, the nitrogen would escape, and the BIM pressure indicator would show that the spar no longer contained pressurized nitrogen and should be replaced. Unfortunately, the pressure indicators could only be inspected when the helicopter was shut down on the ground. However, design calculations indicated that, after being punctured by gunfire, the blade could be operated for 9 to 16 h before failure.
Because of the BIM system and the calculated operating time between loss of nitrogen pressure and failure, investigation of the time to failure in this accident (after about 1 h of flight) was of great interest. Because the BIM system indicators were supposed to be checked prior to every flight, one aspect of the investigation focused on the BIM system itself, to ensure that it was not defective and had not given a false indication of blade reliability. No evidence was found of any defect or malfunction. The mechanical properties and composition of the extrusion were verified to be within the specified limits.

Visual Examination of General Physical Features

Figure 1 shows a schematic of the rotor blade, indicating the approximate fracture location. The rotor blade is primarily an assembly of two types of components: a long, hollow spar made from a 6061-T651 aluminum alloy extrusion, which provides the necessary structural strength, and 26 fiberglass “pockets” that provide the trailing-edge airfoil shape. Figure 2 presents a cross section through the rotor blade, showing the shape of the extrusion and the epoxy attachment to part of a pocket.
Figure 1

Fig. 1  Plan diagram of the main rotor blade, showing the aluminum extrusion spar, the 26 fiberglass trailing-edge pockets, the hub and endcap fittings, and the approximate location of the separation.

Figure 2

Fig. 2  Transverse diagram of the main rotor blade close to the point of separation, showing the extent of the flat fracture (arrows), the set-back for fairing-in the fiberglass pockets, and the location of strain gages in a prior test program.

Most of the fracture surface around the extrusion was oriented about 45° to the extrusion walls, a characteristic of a ductile overload separation. However, the portion of the fracture surface including the bottom-aft corner of the spar was flat and 90° to the extrusion walls, a characteristic of an area of slow crack growth. This area is denoted by the arrows in Fig. 2; also shown in Fig. 2 are locations at which strain gage data were obtained.

Testing Procedure and Results

Surface examination

Macrofractography. The appearance of the flat fracture region showed a concentric ring pattern between the arrows in Fig. 2, with its center adjacent to the set-back allowing the “fairing-in” of the fiberglass pocket (Fig. 3). The general appearance of this area was characteristic of an origin area for fatigue crack growth. Along the bottom edge of the fracture just forward of the set-back were several pits that appeared to have acted as fracture origin sites (Fig. 3).
Figure 3

Fig. 3  Concentric ring appearance, typical of fatigue crack growth, emanating from the pits along the bottom edge of the main rotor blade adjacent to the set-back for the fiberglass pockets.

Metallography

Similar pits were also observed in the general area of fracture initiation, but not on the fracture plane (Fig. 4). A metallographic section parallel to the length of the rotor blade spar was prepared through the pit shown in Fig. 4 and extending through the fracture surface (Fig. 5). The appearance of the pit, especially in cross section, was not that of a corrosion pit. In the cross section, the material forming the pit contained voids and had a different appearance than the base alloy; this material appeared to have been melted and resolidified.
Figure 4

Fig. 4  Pit located about 0.25 mm (0.0l in.) from the fracture plane near the origin on the bottom surface of the main rotor blade.

Figure 5

Fig. 5  Longitudinal section through the pit shown in Fig. 4 and the fracture surface (horizontal at top), showing that the material forming the pit contained voids and had a microstructure different from the base material.

Two hypotheses were possible for the intense, local heating necessary to melt the material forming the pits without affecting the surrounding material—a lightning strike or electrical arcs. The latter was deemed possible: a search of the maintenance records had revealed that the fiberglass pocket at this location had recently been replaced. Replacement of a pocket requires curing of the epoxy bond with a contact electrical heater.
Microstructural Analysis. Several areas in the resolidified material in the cross section exhibited unusual microstructures. One of these areas was analyzed using an energy-dispersive X-ray (EDX) detector on a scanning electron microscope (SEM). The results of this analysis are shown in Fig. 6; the major element in the spectrum was iron. The area analyzed is indicated by the arrow in Fig. 7, which is a slightly higher-magnification view of the same cross section. An X-ray map was obtained of this same area to show the distribution of iron (Fig. 8). This map indicated that all of the unusual-appearing microstructural areas in the resolidified material were areas of high iron concentration.
Figure 6

Fig. 6  EDX spectrum obtained from an anomalous-appearing area in the material forming the pit (arrow in Fig. 7).

Figure 7

Fig. 7  Higher-magnification view of the cross section shown in Fig. 5. Arrow indicates the location at which the spectrum shown in Fig. 6 was obtained.

Figure 8

Fig. 8  X-ray map showing the distribution of iron in the area shown in Fig. 7. Note that the material forming the surface of the pit contained a considerable amount of iron, far in excess of that found in the base material.

The presence of these iron contamination concentrations was sufficient evidence to conclude that the source of the pits was an electrical arc from an iron object, not from lightning. Investigation of the heating elements used to cure the epoxy when replacing the pockets showed that their exterior was iron and that it was quite possible for the outer surface to become “live.” These pits acted as stress concentrations that were the origin sites for the fracture.

Stress analysis

SEM examination of the fracture surface confirmed that the flat portion of the blade fracture was generated by fatigue crack growth. The load cycles in the main rotor blade of any helicopter can be generated from a variety of sources: the differences in relative air speed encountered during a revolution, the perturbations in air flow when the blade passes over the fuselage, the natural frequencies of the blade, and so forth. Because the main rotor blades are rotating wings giving lift to the aircraft, they are subjected to bending stresses during use, with the lower side normally in tension.
It was decided that the flight time between the loss of nitrogen pressure to the final separation of the blade could be estimated by counting the number of fatigue striations on the fracture surface between the point at which the fatigue crack penetrated the wall of the spar (allowing the nitrogen to escape) and the transition from fatigue crack growth to the final overload failure.
The striation counts were performed at two laboratories by examination of the fracture using both SEM and transmission electron microscope (TEM) replica techniques. The general appearances of the fatigue striations are shown in Fig. 9 and 10. The results of the striation counts are given in Table 1. Although a statistician might dwell on the variation in striation counts obtained, these results were considered to be in good agreement, especially considering that they were achieved by a sampling technique (rather than by actually counting every striation) on three different electron microscopes by three different investigators.
Figure 9

Fig. 9  SEM micrograph of the fracture surface of the crack at a location close to the point at which the crack penetrated the wall of the hollow extrusion, showing uniformly spaced fatigue striations. 3700×.

Figure 10

Fig. 10  SEM micrograph of the fracture surface of the crack at a location close to the point at which the transition from fatigue to overload occurred, showing nonuniform fatigue striation spacing (a largest striation followed by two or three small ones). 3700×.

Table 1   Striation counts and equivalent flight time

Method
Striation counts
Flight time(a), min
1 cycle/rev
4 cycles/rev
TEM/replica
40,200
217
54
SEM Lab 1
50,200
271
68
SEM Lab 2
48,000
259
65
(a) Calculated on the basis of 185 blade revolutions per minute in level cruise flight and either one or four load cycles per blade revolution
As shown in Table 1, although the three different striation counts were in good agreement, an enormous disparity in the flight time resulted from differing theories regarding the number of load cycles experienced by the blade per revolution—about 1 h if there are four stress cycles per revolution and about 4 h if there is only one stress cycle per revolution. The results of strain gage tests of the blade in operation were obtained; Fig. 11 shows the results from the strain gage location that was near the fatigue crack area. These results indicated that there were four stress cycles per revolution of the blade; however, they varied greatly in magnitude. Some of these stress cycles were sufficiently small that they may not have caused fatigue crack growth, particularly in the initial stages of crack growth.
Figure 11

Fig. 11  Output of strain gages attached to an operating main rotor blade at the locations shown in Fig. 2, indicating that in each revolution of the blade four tensile stress cycles of different magnitude were encountered.

To determine how many stress cycles per revolution would cause crack growth as a function of the crack length and to determine how long it would take for failure to occur, a computer simulation of the fatigue crack growth process was used. Fortunately, fatigue crack growth data were available (Ref 1) for 6061-T651 alloy (Fig. 12). The cyclic stress data and the handbook crack growth data were input into a computer program that was similar to the USAF “CRACKS” program, based on Foreman's equation (Ref 2) for crack growth (below), using numerical integration and including the Willembourg/USAF crack retardation model.
Figure 12

Fig. 12  Cyclic crack growth rate as a function of the change in stress intensity for 6061-T651 aluminum alloy. Source: Ref 1.

Foreman's equation for crack growth is expressed as:
Unknown image
where da/dn is the crack growth per cycle of stress, ΔK is the change in stress intensity associated with that cycle of stress, R is the stress ratio, K Ic is the plane-strain fracture toughness (28.6 MPa $\sqrt{m}$, or 26 ksi $\sqrt{\mathrm{in.}}$), and C and m are constants that can be determined from the da/dn versus ΔK data for the material.
The results of the computer simulation showed that, given the properties of the 6061-T651 aluminum (Fig. 12) and the stress cycles as shown in Fig. 11, it would require slightly less than 1 h of flight operation for the fatigue crack to grow from the point at which it broke through the wall to the final overload failure. Additionally, the computer simulation showed that at short crack lengths only the highest stress cycle was active in causing crack growth, whereas as the crack lengthened, the other stress cycles started to play a role in crack growth. Ultimately, near the end of the fatigue crack growth, all four stress cycles were causing crack growth.
These computer results were confirmed by careful examination of the fatigue crack surface. Striation structure and spacings were found to vary as the crack grew. As shown in Fig. 9, the striation spacing near the point of breakthrough (relatively short crack length) was very uniform, indicating that only one magnitude of stress was causing crack growth. As shown in Fig. 10, the striation spacing near the transition to the final overload failure was very different; the striations were grouped—one large striation followed by two or three smaller ones, indicating that several stress cycles of different magnitude were now causing crack growth. These observations reflected the computer results with surprising accuracy.

Conclusion and Recommendations

Most probable cause

Failure was caused by a fatigue crack that grew to critical length without detection. An improperly designed heating element that was prone to shorting to its case caused the electrical arc pitting, which acted as the stress-concentrating origin for the fatigue crack. Calculations and striation counting showed that the main rotor blade failed after slightly less than 1 h of use, before the BIM system would have been activated.
The discrepancy between the calculations performed during this failure analysis and those done during the design of the blades is most probably related to the hypotheses used as to the cause of the loss of nitrogen pressure from the blade. In this analysis, the leak was caused by the growth of a fatigue crack through the wall; in the design analysis, the leak was caused by a hole generated by hostile fire. In the design analysis, a fatigue crack had to initiate and grow to final failure after the leak, whereas in the case being studied the fatigue crack had already been initiated and only the time for crack growth was critical. Obviously, in this accident, the time to final failure would be considerably less than expected by design analysis.

Remedial action

To improve the reliability and safety of these helicopters, a modification is now available in which a loss of nitrogen pressure in any of the main rotor blades is immediately indicated on the aircraft control panel. This allows the crew to abort the flight as soon as the problem occurs.

References

  1. Damage Tolerant Design Handbook, Air Force Materials Laboratory, Wright-Patterson Air Force Base, Jan 1975.
  2. R.G. Foreman, V.E. Kearney, and R.M. Engle, Numerical Analysis of Crack Propagation in Cyclic Loaded Structures, J. Basic Eng., Vol. 89, 1967, p 459–464.

Related Information

Fatigue Failures, Failure Analysis and Prevention, Vol 11, ASM Handbook, ASM International, 2002, p 700–727