On the Fatigue and Fracture of Bladed and Integrally Bladed Rotors of Aircraft Engine Compressors

The fatigue and fracture for bladed and integrally bladed rotors (IBR) of aircraft engine compressors has been studied. For IBRs, a new distinct finite element technique was developed to model crack propagation under combined low cycle and high cycle fatigue loading. The crack trajectory, aspect ratio, and shape resulting from the method agreed very well with airfoils which fractured in service. The technique can be extended on other compressor disk applications. For bladed rotors limited by fretting fatigue, a unique fracture mechanics based methodology was developed for obtaining an evolved coefficient of friction (COF) resulting from fretting motion between the fan blade and hub. The predicted nucleation location, nucleation life, crack trajectory, shape and propagation life agreed well with the fractured components. The study confirms that the fretting-specific modified Smith-Watson-Topper (SWT) parameter more accurately predicts the nucleation location and life of the crack compared to the plain fatigue SWT parameter.

On the Fatigue and Fracture of Bladed and Integrally Bladed Rotors of Aircraft Engine Compressors

The fatigue and fracture for bladed and integrally bladed rotors (IBR) of aircraft engine compressors has been studied. For IBRs, a new distinct finite element technique was developed to model crack propagation under combined low cycle and high cycle fatigue loading. The crack trajectory, aspect ratio, and shape resulting from the method agreed very well with airfoils which fractured in service. The technique can be extended on other compressor disk applications. For bladed rotors limited by fretting fatigue, a unique fracture mechanics based methodology was developed for obtaining an evolved coefficient of friction (COF) resulting from fretting motion between the fan blade and hub. The predicted nucleation location, nucleation life, crack trajectory, shape and propagation life agreed well with the fractured components. The study confirms that the fretting-specific modified Smith-Watson-Topper (SWT) parameter more accurately predicts the nucleation location and life of the crack compared to the plain fatigue SWT parameter.

INVESTIGATING POSSIBILITIES OF CRACK INITIATION LIFE EXTENSION IN JET ENGINES COMPRESSOR DISKS

This paper deals with jet engines compressor disks which have dovetail joints with blades. A compressor disk with reduced fatigue resistance was taken as an example. Two simplified conceptual solutions of the dovetail joint with blades were devised. Based on the low cycle fatigue theory, the crack initiation life of their critical parts with newly-proposed transition rounding at the bottom of dovetail grooves was estimated. Two different flank angles in the dovetail grooves (60° in the critical part that belongs to the first dovetail joint conceptual solution and 55° in the critical part that belongs to the second dovetail joint conceptual solution), two different aviation steels selected for workmanship (13H11N2V2MF and AISI 304 steel) and two load histories (load history LH1 and load history LH2), were taken into account. By load history LH2 an overload of the critical parts was simulated. The results of crack initiation life estimation of the critical parts in the dovetail joint conceptual solutions show that there is a possibility for the crack initiation life extension of the observed compressor disk. In all analyzed variants, it has been shown that the critical part in the second dovetail joint conceptual solution has longer crack initiation life than the critical part in the first dovetail joint conceptual solution. For example, the critical part in the second dovetail joint conceptual solution made of AISI 304 steel, in the case of load history LH1 has 141.55% longer crack initiation life than the critical part in the first dovetail joint conceptual solution made of 13H11N2V2MF steel. In the case of load history LH2 (an overload case) that percent is greater and amounts to 173.15%.

Fracture resistance parameters for the compressor disk imitation model

This paper presents a calculation and experimental technique for determining stress intensity factors in an imitation model of a titanium alloy disk. We studied a low-pressure compressor disk of a gas turbine engine (GTE) D-36. During operations, there occur fatigue cracks initiated and developed in the slot fillet under the blade at the place of transition of the bottom to the lateral surface of the inter-groove projection, which lead to a separation of the disk’s part within its rim. The mixed-mode crack growth occured in the compressor disks. Based on the principles of imitation modeling, the geometry and loading condition of the imitation model of the compressor disk was developed. The fatigue test of the imitation model was carried out with a frequency of 5 Hz, at room temperature and with stress ratio Rc = 0.1, by means of a biaxial testing machine. The crack growth was monitored using an optical microscope. The criterion for failure was the condition for reaching a growing crack of the compensation hole. During the test, the positions and sizes of the crack fronts were fixed, which are the basis for the numerical calculation of the fracture resistance parameters. In the order of the numerical studies, six three-dimensional finite element models with different positions and sizes of the crack fronts are considered. The results of the numerical calculations based on the finite element method were used to determine the distributions of elastic and plastic stress intensity factors along each crack front. We demonstrated the advantages of the calculation and experimental methods for solving the problems of interpretation and prediction of the crack growth in the rotating disks of turbomachines using the methods of fracture mechanics.

Fatigue Reliability Analysis of a Compressor Disk Based on Probability Cumulative Damage Criterion

The reliability of aero engine has a direct impact on the flight safety of the whole plane. With the continuous improvement of performance requirements of aero engines, the related fatigue and reliability problems also appear. For the fatigue failure characteristics of the typical component (compressor disk) in an aero engine, the fatigue reliability of its multi-site damage structure in service is analyzed by using probability cumulative damage criterion in this paper. The probability distribution definitions of life, damage and damage threshold are discussed and the relationship among them is also introduced by the new proposed criterion. Meanwhile, a method to determine the probability distribution of cumulative damage threshold and probability life prediction is carried out, based on which a hierarchical index system of statistical analysis and reliability modeling principle on the system level is further constructed for compressor disk. At the end of the paper, a certain cruise of fighter plane is analyzed to verify the validity of the new model. Emphasizing the difference between the compressor disk and traditional component, the new reliability analysis model developed in this study is basically reasonable for most of the load histories for the compressor disk, other than the traditional one, especially for the changeable and complex cruise missions.