1. Introduction
The process of fatigue damage accumulation in different zones of
aircraft leads to the initiation of fatigue cracks due to various
reasons and their following propagation with different intensity
[1-3]. The process is realized under in-service conditions and at
different combinations of variable amplitude loading various [4-9].
It follows that the fundamental principle of in-service aircraft is the
knowledge about that the total process of damage accumulation should be
divided on two separate stages, i.e. crack nucleation and propagation
stages. It is worth noting that due to the cracks grow for a rather long
period of lifetime, the initiated cracks can be monitored by the
non-destructive test methods [10].
Therefore, the basic principle of aircraft operation is the Damage
tolerance approach [11-15]. Damage tolerance is a property of a
structure relating to its ability to sustain defects safely until a
repair can be effected. This approach is commonly used in aerospace
engineering, mechanical engineering, and civil engineering to manage the
extension of cracks in structure through the application of the
principles of fracture mechanics. In engineering, a structure is
considered to be damage tolerant if a maintenance program has been
implemented that will result in the detection and repair of accidental
damage, corrosion and fatigue cracking before such damage reduces the
residual strength of the structure below an acceptable limit. To
implement this principle of operation, the introduction of periodic
inspection of the crack propagation zones with reasonable frequency is
required. The introduction of monitoring is based on an understanding of
the intensity with which the crack growth process will be implemented.
Therefore, on the one hand, information on the development of in-service
cracks is necessary, and, on the other hand, full-scale fatigue testing
is required to correlate the expected and realized in-service loading
mode of load-bearing structural components.
The loads acting on the structural components of the aircraft wings and
airframe have a pronounced non-stationary character, and depending on
the flight conditions, wind gusts can occur that affect significantly
the material behaviour in the process of crack initiation and growth.
For example, on the one hand, during overloading the process of damage
accumulation is more intensive, but, on the other hand, following
deceleration of damage accumulation can arise [1-3]. Therefore, a
correlation of data on the realized in-service failure with data of
full-scale tests under variable amplitude loading conditions allows
introducing the reasonable estimation of inspection frequencies for
flight safety.
Development and improvement of quantitative fractography technique [1,
7, 16] allows to systematize fracture surface relief parameters
responsible for the different loading regimes of structures under
variable amplitude loading conditions for the flight. Therefore, it
makes possible to generalize the established regularities of fatigue
crack developed in the load-bearing structures of one of the most
in-service active RRJ-95 (Russian Regional Jet 95) aircraft.
During the operation of the RRJ-95 aircraft (No. 89051) the failure of
the bracket across four lug sections was found in the mount fitting of
the top rudder drive. The operating time from the new was 3785 hours or
2345 flights. To identify the factors influencing the process of crack
generation and to substantiate the inspection frequency for brackets in
operation, the estimation of crack growth duration was required.
At the same time, full-scale bench tests of the RRJ-95 aircraft (No.
95075) wing structure were carried out according to the schematized
program of the flight loading cycle in order to determine the crack
propagation duration in different zones of the structure.
The test program included cycles with variable amplitude and maximum
stress that simulate the loads during aircraft taxiing, take-off, climb,
cruising regime, descent, and landing. The schematized loading program
could not reproduce the entire spectrum of in-flight loads but reflected
the estimated damage accumulation per flight with an average duration of
1.5 hours.
The cracks revealed in the structural components had a considerable
length, indicating their long-term nature of propagation. However, by
using the results of quantitative fractography, it was necessary to
clarify the nature and duration of the crack growth, as well as to
specify the sequence of crack development over different sections of the
above-mentioned objects.
The results of studies are summarized below with an estimate of the
fatigue crack growth duration in accordance with the methodology for
systematizing fracture relief parameters reflecting the crack
propagation per flight or unit loading cycle.