Aircraft Engine Blades Research Articles

Surface modeling method for aircraft engine blades by using speckle patterns based on the virtual stereo vision system

A blade is one of the most important components of an aircraft engine. Due to its high manufacturing costs, it is indispensable to come up with methods for repairing damaged blades. In order to obtain a surface model of the blades, this paper proposes a modeling method by using speckle patterns based on the virtual stereo vision system. Firstly, blades are sprayed evenly creating random speckle patterns and point clouds from blade surfaces can be calculated by using speckle patterns based on the virtual stereo vision system. Secondly, boundary points are obtained in the way of varied step lengths according to curvature and are fitted to get a blade surface envelope with a cubic B-spline curve. Finally, the surface model of blades is established with the envelope curves and the point clouds. Experimental results show that the surface model of aircraft engine blades is fair and accurate.

First Report on the Deformation Mechanism Mapping of First and Second Generation Ni-Based Single Crystal Super Alloys

Nickel based single crystal super alloys are widely used as aircraft engine blades and are still in the process of development to improve the high temperature capabilities. The typical microstructure of these alloys consists of gamma phase as the matrix strengthened by the gamma prime precipitates. In recent times, addition of refractory metals like Re are tried to improve the creep strength. The first generation of these alloys does not contain any such additions. The second generation of the alloys contains around 3–6% Re. These additions and modifications have been found to improve the temperature capability of the CMSX, though there are limitations in their addition. In this article the microstructural developments, the metallurgical and high temperature behavior of the first and second generation NiSX family subjected to stress annealing, are discussed from engineering point of view. Based on experimental findings reported in literature, deformation maps illustrating the different mechanisms over a wide range of temperature and stress levels for the first and second generation alloys are presented.

Idealized vs. Realistic Microstructures: An Atomistic Simulation Case Study on γ/γ' Microstructures.

Single-crystal Ni-base superalloys, consisting of a two-phase γ/ microstructure, retain high strengths at elevated temperatures and are key materials for high temperature applications, like, e.g., turbine blades of aircraft engines. The lattice misfit between the γ and phases results in internal stresses, which significantly influence the deformation and creep behavior of the material. Large-scale atomistic simulations that are often used to enhance our understanding of the deformation mechanisms in such materials must accurately account for such misfit stresses. In this work, we compare the internal stresses in both idealized and experimentally-informed, i.e., more realistic, γ/ microstructures. The idealized samples are generated by assuming, as is frequently done, a periodic arrangement of cube-shaped particles with planar γ/ interfaces. The experimentally-informed samples are generated from two different sources to produce three different samples—the scanning electron microscopy micrograph-informed quasi-2D atomistic sample and atom probe tomography-informed stoichiometric and non-stoichiometric atomistic samples. Additionally, we compare the stress state of an idealized embedded cube microstructure with finite element simulations incorporating 3D periodic boundary conditions. Subsequently, we study the influence of the resulting stress state on the evolution of dislocation loops in the different samples. The results show that the stresses in the atomistic and finite element simulations are almost identical. Furthermore, quasi-2D boundary conditions lead to a significantly different stress state and, consequently, different evolution of the dislocation loop, when compared to samples with fully 3D boundary conditions.

Modeling of shrinkage during investment casting of thin-walled hollow turbine blades

The shrinkage ratio is a key parameter in designing the investment casting die for hollow turbine blades of high-performance aircraft engines. To avoid extensive modifications to the die shape, we took a single-crystal hollow turbine blade as the typical part to determine the nonuniform shrinkage distribution while considering its structural characteristics during investment casting. By using the structural identification method, different geometrical structures were identified, and the displacement field was established and verified via numerical prediction and experimental measurement. The deformation characteristics, including the wall-thickness distribution and shrinkage ratio, can be established by deformation decoupling analysis. The optimized die profile designed on the basis of the calculation results is in good agreement with the investment casting die in actual use, which indicates that the proposed method is beneficial for improving the geometrical accuracy of hollow turbine blades.

Application of structure-energy failure theory for determining the probability of failure and durability of turbine and compressor blades of aircraft engines

The aim of the present work was to develop a methodology for assessing reliability on the basis of physical characteristics of an element. To this goal, we used the structure-energy power failure theory. The use of the theory makes it possible to determine the probability of failure using the information on the internal damage of the material. The gas turbine engine blades of the first, second and third turbine stages, new and reconstructed high-pressure compressor blades of stage 5 were the object of the analysis. We used an industrial computer tomograph XTH 450 LC to carry out non-destructive testing. We carried out destructive tests on stands of the electrodynamic type. The probabilities of failure calculated by using the structure-energy failure theory correspond to the values of reliability indices obtained by the destructive experiments. We developed a computer program based on a statistical method for estimating the reliability of an aircraft engine. The program takes into account the reliability index of rotor blades calculated with the use of the structure-energy failure theory.

Defect-caused alteration of a composite unit modal parameters

Abstract This paper presents research on diagnostic parameters for monitoring conditions of a light aircraft engine blade made of composite materials (CM) in flight. This study covers the modal parameters investigation of the blades in the frequency range of 30–60 Hz. Using the Doppler Effect the vibration response between the used blades and a new blade was compared. The used blades demonstrated a fourfold increase in amplitude of resonant vibration in the first bending mode. The amplitude growth of the first bending mode resonant vibration can be used as a diagnostic sign of emerging defects. The ease of amplitude and frequency analysis of a blade resonant vibration in flight conditions will allow researchers to develop an express-control device for technical condition analysis of CM blades while in air.

The assessment of geometric accuracy of aircraft engine blades with the use of an optical coordinate scanner

Purpose The purpose of this paper is to present coordinate measuring system possibilities in the meaning of the geometric accuracy assessment of hot zone elements in aircraft engines. The aim of the paper is to prove that this method, which uses blue light and is most sufficient and cost-saving method, can to be used in the production line for serial manufacturing of elements, for which a high level of accuracy is required. Design/methodology/approach The analysis of the geometric accuracy of the blades was performed using non-contact optical coordinate scanner ATOS Triple Scan II Blue Light, manufactured by GOM Company, at the Department of Mechanical Engineering, Rzeszow University of Technology. Geometric analysis was conducted for blades manufactured from different waxes (A7Fr/60 and RealWax VisiJet CPX200), thus comparing injection technique and rapid prototyping (RP) method, and for casting made of Inconel 713C nickel-based superalloy. Findings The analysis of the criteria for the method of blades’ measuring selection showed that the chosen system successfully met all criteria for the verification of blades’ geometry at the selected stages of the process. ATOS II optical scanner with blue light technology allows measurement almost regardless of daylight or artificial (white) light. This allows the application of the measurement system in the production cycle, thus eliminating the need to create special conditions for measurements. Practical implications Requirements related to the accuracy of measured values, diversity and allowable measurement time are linked with the methods of production. Modern manufacturing methods based on computer-aided design systems/manufacturing/engineering systems require a non-contact optical measurement method based on the computer-aided-based coordinate measuring technique. In case of the non-contact optical scanning method based on the ATOS GOM measuring system, time and measurement costs depend on the methodology of measurement and the possibility of its automation. This is why the presented paper has a practical impact on possibilities for the automation of geometric accuracy measurements of obtained elements in the series production line. Social implications The use of ATOS Triple Scan II Blue Light by GOM Company allows the reduction of cost and time of production because of the possibility of the introduction of this system in an automated production line. Additionally, the measurement of hot section blades of aircraft engines by using the blue light method is much more accurate and has implication as it impacts safety of further used manufactured elements. Originality/value This paper presents the possibility of using the ATOS Triple Scan II Blue Light measuring system for geometric accuracy measurements in case of hot section blades of aircraft engines. This research is original because it describes three model geometric accuracy measurements, wax model obtained using the injection technique, wax model obtained using the I RP process and casting made of Inconel 713C nickel-based superalloy.

Degradation of nonmodified and rhodium modified aluminide coating deposited on CMSX 4 superalloy.

The Ni-base superalloy CMSX 4 used in the turbine blades of aircraft engines was coated with rhodium layer (0.5-μm thick). Next coated CMSX 4 superalloy was aluminized by the CVD method. The rhodium modified aluminide coating and nonmodified aluminide coating were oxidized at 1100°C at the air atmosphere. The rhodium modified aluminide coating showed about twice better oxidation resistance than the nonmodified one. The spallation equal 62% of the total area was observed on the surface of the nonmodified coating whereas only 36% spallation area was observed on the surface of the rhodium modified aluminide coating after the oxidation test. The oxide layer formed on the surface of the nonmodified coating was composed of nonprotective (Ni,Cr)Al2 O4 and (Ni,Cr)O phases. Aluminium in the coating reacts with oxygen, forming a protective α-Al2 O3 oxide on the surface of the rhodium modified aluminide coating. When the oxide cracks and spalls due to oxidation, additional aluminium from the coating diffuses to the surface to form the oxide. The presence of protective Al2 O3 oxide on the surface of the rhodium modified aluminide coating slows coating degradation. Therefore, rhodium modified aluminide coating has better oxidation resistance than the nonmodified one.

Experimental fatigue strength determination of damaged aircraft engine blades

Experimental fatigue strength determination of damaged aircraft engine blades

Differential excitation eddy current sensor testing for aircraft engine blades defect

Differential excitation eddy current sensor testing for aircraft engine blades defect

Numerical study on vibration stress of rotating fan blade under aerodynamic load at critical speed

The vibration stress caused by resonance is the main cause of failure in aircraft engine blades. The study investigated the vibration stress of a single-stage fan blade influenced by the interaction of aerodynamic and structural behavior at multiple critical speeds. A numerically based analysis through the interfacing of computational fluid dynamics and finite element modeling analysis, referred to as fluid–structure interaction was performed. A three-dimensional finite element model of the blade which was verified by modal test was established to carry out response analysis when blade resonates. A single flow channel model of the blade was constructed to simulate the aerodynamic pressure loads that result in vibration of the fan blades. The Campbell diagram was introduced to determine the critical conditions of the blade. The results show that resonance phenomenon appears for several times between 2000 and 3600 r/min, and further analyses of the blade mode shape as well as dynamic stress at these speeds indicate that vibration stress at critical speed is obviously higher than those at the adjacent speeds. But there is no positive correlation between the peak stress and critical speed, while it is determined by the comprehensive action of critical speed, vibration mode and synchronous excitation. The peak stress distribution is closely linked to the mode shape of the blade, and it always appears at the leading edge and changes along the blade height. It also illustrates that the leading edge of the blade is most prone to vibration fatigue.

The Study on Erosive and Abrasive Wear Resistance of Different Multilayer Coatings Obtained on Titanium Alloy Ti6Al4V

One of the areas in which increased durability is of particular significance is the use of machine elements in conditions of intensive exposure to erosive wear, for example, turbine blades of aircraft engines and power turbines in ventilation systems, and machine elements working in a dusty environment. As it is apparent from the analysis of the state of the art, high possibilities to increase the erosive and abrasive wear resistance of machine elements are created by appropriately designed coatings and layers, especially multilayer coatings composed on the base of materials with different plasticity properties. This is associated with particular characteristics of this type of multilayer coatings that allow the absorption of external energy reaching the surface without causing permanent damage to the coating, while significantly reducing the erosive wear rate.The article presents the results of material research and the results of erosion test of three different materials solutions of surface engineering, i.e. multilayer coatings (Cr / CrN) x8, (TiN / ZrN)multinano and a (Ti-Al) intermetallic / AlCrTiN hybrid layer, obtained on titanium alloy Ti6Al4V. The (Cr / CrN) x8 and (TiN / ZrN)multinano multilayer coatings were obtained by arc evaporation, and the (Ti-Al) intermetallic / AlCrTiN hybrid layer was prepared using a hybrid technology combining diffusion saturation in aluminium powders and arc evaporation in a single technological process. The selected material solutions of the surface layer were subjected to the analysis of the structure by scanning electron microscopy (Hitachi 3000). Measurements of hardness and Young's modulus were conducted using the nanohardness Tester CSM. The results of this study allowed the authors to compare the erosive wear resistance and wear resistance of selected material solutions, and this led to the conclusion that the (Ti-Al) intermetallic / AlCrTiN hybrid layer is the most effective solution.

RF Sputtering Thermal Barrier Coating for Enhance Corrosion Efficiency of Aero-Engines Components

Nickel base superalloys have been widely used for manufacturing key structural materials for aircraft engine blades. These alloys are developed to withstand the most deleterious degradation mechanisms as creep damage, fatigue, loss of thermal stability and environmental attacks as oxidation and hot corrosion during service. The damages can compromise the structural integrity of the aircraft and in the next generation of advanced engines they will require even higher gas temperature. Consequently, much work has been done in order to develop protective approaches that will increase component lifetimes and meet the development of advanced gas turbine engines. One of such is the used of thermal barrier coatings (TBCs). Thus the aim of the present investigation is to study the electrochemical behavior of YSZ coatings deposited by physical vapor deposition (PVD) on an aeronautical nickel based superalloy to improve the oxidation resistance. The electrochemical techniques conducted on Incoloy 800HT substrates, show that the corrosion velocity decreased with the YSZ coating system.

Conjectural bifurcation analysis of the contact-induced vibratory response of an aircraft engine blade

This paper deals with the numerical investigation of the unilateral contact-induced dynamics of a turbomachine blade rotating within a perfectly rigid yet distorted casing. This investigation is motivated by unelucidated vibratory behaviours observed experimentally. The simulations are based on an in-house time-marching strategy incorporating Lagrange multipliers for the unilateral contact treatment, as well as centrifugal stiffening and abradable coating removal. Significant extensions are proposed through the implementation of (1) aerodynamic loading on the blade and (2) post-processing techniques involving the empirical mode decomposition which provides fruitful insights on important transient phenomena. A thorough bifurcation analysis with and without aerodynamic loading highlights the existence of flip bifurcations with period-doubling and period-halving sequences over a broad angular speed range. Numerical simulations with external aerodynamic loading yield quasi-periodic and likely to be chaotic motions that could not be observed under vacuum. The proposed numerical investigations underline the key role of the aerodynamic loading in the blade dynamics and suggest that unexplained experimental vibratory behaviours are related to the vacuum conditions of the experiment.

Improvement in Casting Technology of the Blades of Aviation Gas-Turbine Engines Made of TNM-1B1 Alloy in Conditions of Induction Crucible Smelting

Properties of the TNM-B1 alloy based on γ-titanium aluminide are investigated in order to find reliable results of simulation of casting processes used to fabricate the blades of gas-turbine aircraft engines. Thermal properties of the TNM-B1 alloy are experimentally determined using the methods of differential scanning calorimetry (DSC), laser flash (LFA), and dilatometry (DIL). Various variants of gate-feeding systems for the production of cast blades using a smelting-pouring installation for vacuum induction smelting («Consarc», Scottland) in conditions of JSC Ufa Engine Production Association. The test casting is obtained and its defects are analyzed. Good correspondence of results of modeling to actual centrifugal pouring of blades is shown.

Fabrication and Evaluation of an Amino Modified Epoxy Coating for Anti-Scouring and Anti-Corrosion Application

Excellent performance is necessary for compressor blade when aircraft engine serves in harsh environment, where it will go through scouring from jet-stream and extraneous sand, shaking of blades and corrosion from corrosive medium, taking moisture, salt, SO2 for example. Due to the shortage of scouring resistant and anti-corrosion property, shedding of damaged coating is a fundamental reason for the breakdown of aircraft engine blades, which is also the technique bottleneck that hinders the engineering application of compressor blades. Aiming at such problems, an amino modified epoxy coating which performed excellent scouring resistent and anti-corrosion property was fabricated on blades in this paper. Physical properties, anti-corrosion performance, abrasion resistant profiles and scouring resistant test were comprehensively evaluated. Results showed that the coating displayed a high hardness (pendulum hardness 0.9), overlong pot life, well anti-corrosion performance and excellent scouring resistant property. Pot life extends from 8h which is that of common coatings to 36h, so that it is benefit to large scale coating in engineering application. The anti-corrosion performance is similar to traditional aluminizing/inorganic silicate protect coatings, and abrasion resistant property is superior to common protect coatings. Moreover, scouring resistance performance is 3~4 times that of traditional coatings. Hence, such protective coatings could ensure the long lasting of blades protection.

Improvement in the casting technology of blades for aviation gas-turbine engines made of TNM-B1 titanium aluminide alloy produced by induction crucible melting

Properties of the TNM-B1 γ-titanium aluminide alloy are investigated in order to find reliable results of a simulation of casting processes used to produce the blades of gas-turbine aircraft engines. Thermal properties of the TNM-B1 alloy are experimentally determined using the methods of differential scanning calorimetry (DSC), laser flash analysis (LFA), and dilatometry (DIL). Various variants of gate-feeding systems for the production of cast blades using a vacuum induction melting furnace (Consarc, Scotland) in conditions of the Ufa Engine Production Association (OAO UMPO) are considered. The test casting is obtained and its defects are analyzed. A good correspondence of results of simulation to the actual centrifugal filling of blades is shown.

Porosity of Solid and Cored Turbine Blades of Aircraft Engines

Paper presents results of quantitative evaluation of porosity conducted on big, thin walled airfoil turbine blades made from Inconel 713C alloy. To decrease mass, blades are design and manufacture like thin walled cored castings. Manufacturing of big thin walled casting airfoil blades is extremely difficult. During exploitation casting work undergo cycle fatigue. In that cause casting should be free of casting defects, including porosity. Conducted research focused on Inconel 713C superalloy pouring temperature effect on porosity level of cored casted turbine blade. Results were compared to porosity of solid casted turbine blade. It was found that porosity of cored blades is lower than solid blade porosity. In cored blades higher porosity is located on airfoil convex side. Airfoil concave side has lower porosity.

A study for stainless steel fan blade design with metal foam core

The pursuit for cheap, low-density and high-performance materials in the design of aircraft engine blades raises wide-ranging challenges to the materials and structural design engineers. Traditionally, these components have been fabricated using expensive materials such as lightweight titanium alloys and polymer composite materials composites. The present study investigates the use of a sandwich foam fan blade made of solid face sheets and a metal foam core. The face sheets and the metal foam core material were an aerospace grade precipitation-hardened 17-4 stainless steel with high strength and high toughness. The stiffness of the sandwich structure is increased by separating the two face sheets by a foam core. The resulting structure possesses a high stiffness while being lighter than a similar solid construction. Since the face sheets carry the applied bending loads, the sandwich architecture is a viable engineering concept. The material properties of 17-4 precipitation-hardened metal foam are briefly reviewed to describe the characteristics of the sandwich structure for a fan blade application. Vibration characteristics and design criteria on the 17-4 precipitation-hardened metal foam core sandwich blade design with different combinations of skin thickness and core volume are presented with a comparison to a solid titanium blade.

Зносостійкий сплав для захисту контактних поверхонь робочих лопаток авіаційних двигунів від окислення при високих температурах

Зносостійкий сплав для захисту контактних поверхонь робочих лопаток авіаційних двигунів від окислення при високих температурах