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.