Wydawnictwo SIGMA-NOT wydaje czasopisma fachowe informujące swoich czytelników o najnowszych osiągnięciach naukowych i nowoczesnych rozwiązaniach technicznych w Polsce i na świecie, popularyzuje problemy techniczne oraz poszerza wiedzę i kulturę techniczną.

Abstract Despite of nearly 100 years of turbine engine development and design, blade vibrations remain a great engineering challenge. The rotating turbine blades’ vibrations lead to cyclic oscillations, which result in alternating stress and strain in harsh environments of high temperature and pressure. In modern aeroengines, high hot flow velocities might generate erosion and corrosion pitting on the metal surfaces, that leverage remarkably mean stresses. The combination of both mean and alternating stresses can lead to unexpected engine failures, especially under resonance conditions. Then, alternating stress amplitudes can exceed the safety endurance limit, what accelerates the high cyclic fatigue leading quickly to catastrophic failure of the blade. Concerning the existing state-of-the-art and new market demands, this paper revises forced vibrations with respect to excitation mechanisms related to three design levels: (i) a component like the blade design, (ii) turbine stage design consisting of vanes and blades and (iii) a system design of a combustor and turbine. This work reviews the best practices for preventing the crotating turbine and compressor blades from High Cyclic Fatigue in the design process. Finally, an engine commissioning is briefly weighed up all the pros and cons to the experimental validations and needed measuring equipment.

Abstract Low Cycle Fatigue (LCF) is one of most common mechanisms behind turbine blade failures. The reason is high stress concentration in notch areas, like fir-tree root groves, which can cause cyclic stress beyond the safe threshold. The stress levels strictly depend on the manufacturing accuracy of the fir-tree lock (for both fitted together: blade root and disk groove). The probabilistic study aimed at determination of stress was performed using Finite Element Method (FEM) simulation on a population of 1000 turbine models (disk + blades +friction dampers), where fir-tree lock dimensions were sampled according to the normal distribution, within limits specified in the documentation. The studies were performed for different manufacturing quality levels: 3-Sigma, 6-Sigma and 3-Sigma with tolerance ranges reduced twice. Based on the results, the probabilistic distributions, probabilities and expected ranges of values could be determined for: material plastification, stress, strain, LCF lifetime, etc. The study has shown how each tooth of the root is loaded and how wide a stress range should be expected in each groove. That gives information on how the definition of tolerances should be modified to make the construction more optimal, more robust, with lower likelihood of damage, taking into account the cost-quality balance. It also shows how the Six Sigma philosophy can improve the safety of the construction, its repeatability and predictability. Additionally, the presented numerical study is a few orders of magnitude more cost- and time-effective than experiment.

The strength of mechanical joints is the weakness of laminates, therefore, mechanical joints require a special attention to be focused during both a designing and a manufacturing process. The aim of the paper is to present sensitivity analysis of a mechanical joint in a composite structure on selected parameters. Parameters of this analysis were classified into four groups: geometrical, stiffness, material failure and initial stress. Global stiffness and strength of laminate, two methods of the nut washer and the bolt head modelling, residual stiffness, initial delamination and other parameters were taken into account. A double-shear bolted joint with four fasteners is the object of analysis. The inner part of the specimen was made of a quasi-isotropic CFRP laminate. The outer elements were made of 2024T3 aluminum alloy sheet, typically used in aircraft structures. Nonlinear finite element analysis of a metal-composite joint was performed using Newton-Raphson method with MSC. Marc code. The numerical and the experimental results were compared.

Nowadays, large aircrafts such as Boeing 787 and Airbus A350 are built in fifty percent of composites (e.g. CFRP, GFRP or Glare). Composites, in general, are brittle and more notch sensitive material than metal alloys. Therefore, mechanical joints of composite parts require a special attention during both the designing and manufacturing process. In the paper, a review of modelling methods for a pin loaded laminate is presented. The simulation of a specimen made of quasi-isotropic CFRP laminate was carried out. Nonlinear analysis was performed using Newton-Raphson method with MSC. Marc. The specimen load capacity was estimated using the maximum stress and Hashin failure criteria in quasi-linear as well as in progressive failure analyses. Failure indices and bearing curves for various contact descriptions (node to segment analytical and discrete, segment to segment and using simplified kinematic ties) and two models of a laminate structure (single layer model and eight orthotropic plies) were compared. The assumed criterion and properties of a failure progressive analysis strongly influence the results. The gradual stiffness reduction analysis with displacement loading in a single layer model results in the upper estimation of the specimen load capacity.

High specific mechanical properties of composites are the reason for their use in various fields, e.g., the aerospace industry. Mechanical joints are still used in the aerospace industry to assembly large aircraft structures. The properties of laminate around the hole can be, however, weakened, compared to their nominal values as a result of a drilling process or cyclic loading. This paper aims at the classification and analysis of imperfections affecting mechanically fastened joints in a laminate structure. A method of modeling the hole vicinity, a gradient material model, as well as the numerical and experimental estimation of laminate deterioration in this area, were proposed and analyzed. Comparative analysis of numerical and experimental results based on displacements of the testing machine grip and the extensometer length confirmed the aforementioned results as consistent in linear ranges. Therefore, joint characteristics obtained based upon measurement of the grip displacement and the ratio of stiffness in linear ranges are sufficient to determine the parameters of a gradient material model. Some imperfections resulting from, e.g., asymmetry, were included in the gradient material model; thus, the obtained weakening of laminate properties in the hole vicinity can be overestimated. Therefore, further analyses of the gradient material model for laminate structures are necessary.

The object of analysis is the casing of F 124 turbofan engine. Analysis of load conditions indicates that a dimensioning load case is crash landing, while an axial component of inertial load is about 30g. The casing is a thin-walled ribbed shell made of titanium alloy. The aim of the paper is mass analysis of the casing subjected to axial compression. Comparative analysis of selected materials, i.a., a sandwich structure (metal-fibre laminate) made of titanium alloy faces and flax fibre core, is performed using a finite element method. In the case of a sandwich structure, a shift of titanium layers from a neutral surface of the shell is significant since it causes a decrease in titanium overall thickness compared to a reference plain model (without decreasing bending stiffness). The best material mass efficiency were obtained for a sandwich structure and it is over 40% higher compared to a plain isotropic structure, whereas, in the case of the original ribbed structure the efficiency is about 26% higher.

Materials of a high specific strength and stiffness are used in the aerospace industry to obtain the lowest possible aircraft mass. The object of analysis is the casing of the F124 turbofan engine. The axially compressed cylindrical part of this casing is considered. The aim of the paper is to analyse possible benefits of replacing the original ribbed metal casing with a sandwich structure. The sandwich structure (metal-fibre laminate) of titanium alloy faces and a flax fibre laminate core is proposed. Semi-analytical optimisation of a sandwich structure was performed including a polynomial approximation of the critical load with correction obtained based on numerical analysis. The best mass efficiency was obtained for a core to faces thickness ratio equal to about 4.

Mechanical joints (e.g. bolted) used for decades are proved to be reliable. They can be assembled and applied in very rough conditions since they are less sensitive to environmental effects than other types of joints (e.g. adhesive). Therefore, they are still employed in aircraft design. High specific stiffness and strength of composite materials (especially CFRP) cause a continuous increase in their usage in aircraft structures. In general, composites are brittle materials and more notch sensitive than metal alloys. Hole making is a necessary stage in manufacturing of a mechanical joint. Holes vicinities are the areas of high stress concentrations and determine load capability of the whole structure. Therefore, mechanical joints of composite parts require a special focus during both a designing and a manufacturing process. The aim of the paper is analysis of potential local material weakness/deterioration caused by a drilling process and its influence on the global response of a mechanical joint. The specimen in the form of a double-shear joint was analyzed. The weakened areas were identified on the basis of NDT ultrasonic analysis. A simple gradient material model was proposed to describe the hole vicinity. Numerical simulations were performed and compared to experimental results.

Abstract The motivation of the article is fatigue and fretting issue of the compressor rotor blades and disks. These phenomena can be caused by high contact pressures leading to fretting occurring on contact faces in the lock (blade-disk connection, attachment of the blade to the disk). Additionally, geometrical notches and high cyclic loading can initiate cracks and lead to engine failures. The paper presents finite element static and modal analyses of the axial compressor 3rd rotor stage (disk and blades) of the K-15 turbine engine. The analyses were performed for the original trapezoidal/dovetail lock geometry and its two modifications (new lock concepts) to optimize the stress state of the disk-blade assembly. The cyclic symmetry formulation was used to reduce modelling and computational effort.

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The paper deals with the analysis of residual stress fields in the riveted joint and the estimation of the internal stress magnitude releasing by partial and complete removing of the rivet material. Stress relieving causes deformations around the rivet hole, which can be measured and compared to the deformation state before removal. Numerical FE simulations of the upsetting process are carried out to determine the residual stress and strain fields. The contact with friction is defined between the mating parts of the joint. Non-destructive testing methods are used in combination with numerical calculations.

Various materials of high specific stiffness and strength are used in aircraft structures, therefore, methods of joining them are continually improved. Adhesive joining is the most popular method used in composite structures, however, mechanical joining is often necessary due to constructional reasons. Stress concentration around the hole (notch) and point load (fastener action) transfer into the structure are the main disadvantages of mechanical joints, especially, in composite structures. The bearing failure is the only global progressive mechanism of mechanically fastened laminates and therefore it is the acceptable one. Different attempts to improve bearing performance of laminates have been considered by many authors. One of the most interesting solutions is bonding the titanium foils (thin inserts) into the composite structure. The aim of the paper is analysis of a metal insert influence on the stress state around the hole. Specimens made of carbon fibre reinforced plastic and locally modified laminate (part of prepreg layer around the hole is replaced with a circular titanium foil) have been taken into account. The specimens were manufactured in a co-curing process using autoclave technology. Nonlinear analyses were carried out with Marc code. In a locally modified laminate, the main mechanism of load transfer is interlaminar shear stress. The area of load transfer is augmented by the area of metal-fibre adhesive joining. A significant part of the applied load is carried by titanium foils, then it is gradually transferred by means of shear stresses due to adhesion between titanium foils and laminate layers. Bearing performance depends on the insert material stiffness and its elasto-plastic behaviour. Bearing stiffness reduction is strongly influenced by parameters of the titanium-laminate adhesive interface.

Riveting is a commonly used (especially in aircraft structures) method of joining metal and composite components. The methods of forming solid shank rivets can be classified in two types: static and dynamic. The static method is the most efficient one. Regrettably, its application is limited. A popular upsetting tool used in an aircraft is a pneumatic riveter. The rivet driving requires a few hammer strokes. The total stress in a riveted joint depends on the residual and applied stress. Residual post-riveting stress fields are widely accepted to have a beneficial influence on the fatigue life of aircraft structures. The analysis is carried out for a solid mushroom rivet (made of PA25 alloy) joining two sheets (made of 2024T3 alloy). Nonlinear dynamic simulations of the upsetting process are carried out. Simulation of the riveting process is significantly influenced by a material model. The numerical calculations are performed for three different cases of upsetting described by the formed rivet head diameters 1.4d, 1.5d and 1.6d, respectively. The rivet head diameter and, consequently, the residual stress state depend on hammer stroke energy. It has a significant influence on a plastic region around the rivet hole, whereas the influence of a number of strokes can be neglected. The strain rate in both local and global (average) formulation is analysed in the paper. For one hammer stroke, the global strain rate of the rivet shank is about 1.0 thousand per second. The local strain rate is about two times greater than the global one, so a strain rate factor has an effect on the residual stress state. For a few hammer strokes, the strain rate is lower than for one stroke; however, it increases a little in each stroke. The hole deformation can be treated as a function of the internal energy of the sheet. The lower total energy of the part the greater influence of the strain rate on the internal energy is observed.

This paper presents numerical analyses of uniaxial tension tests on rods made of a material with elasticplastic characteristics with mapping the temperature fields generated in the fields of plastic strains. The use of thermo-mechanical coupling algorithms implemented in CAE programs that enable us to enter the parameters of materials dependent on temperature and used to visualize the damage of the samples is presented. The numerical models were validated with experimental research.