4. Discussion
The results of studies on the regularities governing the propagation of
fatigue cracks in the structural components of RRJ-type aircraft wing
indicate that cracks propagated for a long period under variable
amplitude loading conditions. This fact allows to organize effective
monitoring for timely in-service crack detection.
At the same time, the question of how to reproduce the duration of
fatigue crack growth in the RRJ-95 aircraft (No. 89051) bracket remains
unresolved, since both MBM blocks (see crack growth in section No. 1)
and fatigue striations (other sections) were identified simultaneously.
The analysed regularity of the fatigue striation formation with
different spacing under variable amplitude loading conditions did not
allow to demonstrate convincingly the block of variable loads
responsible for the flight of the aircraft. Thus, the number of flights
in the indicated bracket was estimated based on the data of the aircraft
wing bench tests using the blocks that reproduced the schematized flight
loading cycle.
The program of bench tests consisted of sequence of loading blocks. Each
block was equivalent to the set of variable loads acting on the aircraft
wing for the full flight-cycle and simulated the entire sequence of the
aircraft flight regimes and operation on the ground – taxiing, warming,
etc. It was taken into account that the loading of the aircraft wing
resulted in the realization of a biaxial stress state. In this regard,
in-phase loading was carried out along two axes (Fig. 11). The
preliminary results of the strain gauge measurement for stress-state
estimation in stiffened panels of the RRJ-95 aircraft wing had shown
that for all the stages of flight the in-phase biaxial loading takes
place. Therefore, the bench-test program consisted of only the in-phase
loading.
The resulting crack growth duration was about 50% of the wing operating
time on the device at the moment of test termination. As shown by the
results of the fractographic analysis, the origins of the cracks are
located near the hole edges at a distance of about 1 mm similarly in the
bracket of the RRJ-95 aircraft (No. 95075) wing airframe and in the wing
of the aircraft (Fig. 12). This fact allows to suppose that the stress
concentration is close in the compared cases.
The tensile strength of both aluminium alloys is as close as the
chemical composition and microstructure of the deformable heat-treated
alloy.
All of the above-mentioned facts allow to assume that at the durability
near the low-cycle fatigue regime, the ratio between the crack growth
period and durability under comparable conditions is also similar, i.e.
about 50%.
The validity of such an analysis is consistent with the results of the
RRJ-95 aircraft (No. 4862) wing previous tests, in which fragments of
the lower panel were fractured during the operating time of 28180
laboratory flights. The development of the crack was accompanied by the
MBM formation (Fig. 13a), and an assessment of the crack growth duration
in the blocks of test flights showed that it was approximately 16000
blocks (Fig. 13b). The obtained assessment indicates the fact that the
crack growth duration is about 16000/28180=57%.
Therefore, it can be assumed that for the fractured bracket of the
RRJ-95 aircraft (No. 89051) the duration of crack growth at the
operating time of 2345 flights is about 2345×0.5 = 1172 flights.
As a result of the study, it was found that the longest period of crack
growth was in section No. 3 for about 97000 unit loading cycles. This
duration is slightly less than that corresponding to the full period of
crack growth because the first section was No. 1, in which the crack
began to propagate. However, it could not reach the maximum length,
because due to the operating other sections, the load was redistributed
from section No. 1 to the remaining sections of the bracket lug.
It is worth noting that with the performed assessment of the block
duration per flight, the error will not exceed 15%.
So, if we use the duration of 1172 flights and the number of striations
equal to 97000, it appears that for one flight the bracket was loaded
approximately 97000/1172 = 83 times. The obtained result of estimating
the duration and the number of unit loading cycles per flight does not
contradict the fractographic data (see, for example, Fig. 5b). On the
fracture surfaces of the bracket in different sections, the extended
areas with fatigue striations were revealed, which do not have clearly
defined block boundaries due to the fact that the number of single
cycles in the blocks reached many numbers. Moreover, the fatigue
striation spacing varies irregularly due to the random loading of the
bracket.
A comparison of the operating time according to the schematized flight
cycle and the bracket in-service duration indicates that the loading on
the device does not fully reproduce the structural component loading
intensity for the flight.
Nevertheless, the duration of the crack propagation in the structural
components of the RRJ-type aircraft wing is significant, which allows
cracks to be reliably detected with an organized inspection frequency
exceeding 500 flights. The start of monitoring can be arranged after the
operating time of 1000 flights. By reducing the stress concentration on
the holes in which cracks occurred, it is possible to significantly
increase the duration up to the moment of crack initiation.