Aerospace Gas Turbine Engines Research Articles

Combustor shape optimization and NO emission characteristics for premixed NH3-CH4 turbulent swirling flame towards sustainable combustion

Ammonia is considered an attractive carbon-free fuel for energy generation in aerospace gas turbine engines. However, ammonia is often used with methane fuel blend due to its low combustion limits. The formation of nitrogen oxides (NOx) during the combustion of ammonia-methane fuel blend is a fundamental problem of interest in the scientific community. To date, there have been several studies of ammonia-methane fuel blend, but none of them have considered the effect of combustor shape on NO emission. In this work, large eddy simulations (LES) are carried out at stoichiometric condition to analyze the flame structure and NO emission characteristics of the ammonia-methane premixed turbulent swirling flame in box and cylindrical shaped combustors. The fuel blend used in this study is 70% ammonia and 30% methane by volume. The LES solver employs a dynamic thickened flame (DTF) combustion model along with Okafor's reduced mechanism, which consists of 43 species and 130 elementary reactions. The flame shape and stabilization are discussed for both combustors, along with the effects of flowfield on the flame. The characteristics of the swirl stabilized flame are illustrated using several time-averaged and instantaneous contours of parameters, including heat release rate, local equivalence ratio, axial velocity, and NO emission. The effect of mixing and reaction rate distribution in the premixed configuration is studied quantitatively, subsequently identifying several flame regions. It is identified from the LES results that the NO emission is 1.6 times higher in the cylindrical combustor compared to the box combustor. While a typical V-shape flame is observed for both combustors, a flame tail also co-exits for the cylindrical combustor. The heat release rate in the flame tail is about 33% higher than the main flame which reignites the unburnt gases for a secondary combustion and generates NO. Thus, a box combustion chamber design is beneficial to minimize the NO emission. The residence time is also analyzed in both combustor to understand the oxidation process and identify the key reaction pathways for NH3/CH4 premixed flames.

Perspectives on Environmental Barrier Coatings (EBCs) Manufactured via Air Plasma Spray (APS) on Ceramic Matrix Composites (CMCs): A Tutorial Paper

There are many sets of information in the literature (e.g., papers, books and websites) about the great achievements that are expected for aerospace gas turbine engines by the employment of ceramic matrix composites (CMCs) and thermally sprayed environmental barrier coatings (EBCs) in their hot zones (e.g., combustion chambers, vanes, shrouds, blades and afterburners). Among these achievements, it is typically highlighted (i) turbine weight reduction, (ii) reduced fuel consumption, (iii) higher operation temperatures, (iv) superior thrust-to-weight ratio and (v) lower emission of toxic gases to the atmosphere. Although these achievements are true, they are generally not well-explained to the reader on how together they come to be. In addition, according to “conventional wisdom”, some of these engineering feats are in fact opposing each other (e.g., higher operation temperatures versus lower emissions). The objective of this tutorial paper is to present the reader how these feats are achieved by the concomitant combination of imaginative engineering. It will explain the non-stop driving force for increasing combustion temperatures; show the basic concepts of CMCs, the paramount need of EBCs, and the complexity of creating EBC architectures via air plasma spray (APS). Finally, highlights on how EBCs/CMCs are tested at high temperature will be provided. The content of this paper shall be understood by anyone with basic knowledge in materials processing and surface engineering.

Direct laser metal deposition of Inconel 738

Inconel 738 is one of the widely used nickel-based superalloys in high-temperature applications, especially in land-based and aerospace gas turbine engines. This paper reports the feasibility of direct laser metal deposition (LMD) of Inconel 738. Cracks evolved during deposition at the substrate/deposit interface and within the deposit along high angle grain boundary for scanning speed of 6 and 12 mm/s due to the intense residual stress and incipient melting. Results showed liquation cracking due to low melting crack boundary γ′ and significant micro-segregation of Al and Ti along the crack boundaries. By maximizing the energy density and by reducing the scanning speed to 3 mm/s, crack-free single wall specimens were successfully manufactured. Microstructural evolution of primary, secondary, grain boundary γ′, MC carbides, and M2B borides in the as-deposited and heat-treat specimens are discussed. Mechanical properties and microstructural development were investigated using tensile testing and scanning electron microscopy. Energy dispersive spectroscopy confirmed significant micro-segregation on various elements along the interdendritic and grain boundaries. X-ray diffraction validated the presence of the observed carbides and boride in the as-deposited and heat-treated samples.

Inter-engine variation analysis for health monitoring of aerospace gas turbine engines

This study investigates the development of a new inter-enginevariation analysis method for the purpose of Equipment healthmonitoring, in which the similarity - in both system behaviourand external disturbances - across multiple (sister) engines isleveraged. The sister engine provides a baseline descriptionof the engine under observation, such that the challenge becomesthe differentiation between normal inter-engine variationand the anomalous behaviour, bypassing the need todescribe highly complex engine dynamics. The inter-engineresiduals are modelled directly with input data from both engines,using previous healthy data for training. The trainedmodel is used to compensate known differences between realengines. Anomalous data is detected by comparison of thesimulated output with the true residuals. The method is demonstratedon a real data set containing both nominal, healthy enginedata, and engine data containing anomalies.

English

The significant goal of this study is to develop a system to determine the margin of safety for ME3 super alloy in aerospace gas turbine engines. This margin of safety can be defined as critical temperature, force stress, exposure time, and cycles. In this research, by applying analytical solutions and using experimental data, equations are obtained to predict the life of a nickel-base super alloy in aerospace gas turbine engines. The experimental data that are applied to obtain the equations are achieved by gas turbine environment simulation experiments. This environment is contained with mechanical cycles at high temperature. The results that are achieved by these solutions are compared with the data from an experiment on a composite material. This comparison has proved that the results of this study seem logical. Ultimately, by employing of convex, concave, and linear equations, many results to predict the life of ME3 are obtained. Main reasons in ME3 failure at gas turbine engine environment is determined by this method. Based on the results of this research, the most important cause of ME3 failure is due to maximum mechanical force at high temperatures close to 763°C. Key words: Failure, nickel-base super alloy, gas turbine engines, mechanical fatigue, high temperature.

Gas turbine engine prognostics using Bayesian hierarchical models: A variational approach

Prognostics is an emerging requirement of modern health monitoring that aims to increase the fidelity of failure-time predictions by the appropriate use of sensory and reliability information. In the aerospace industry it is a key technology to reduce life-cycle costs, improve reliability and asset availability for a diverse fleet of gas turbine engines.In this work, a Bayesian hierarchical model is selected to utilise fleet data from multiple assets to perform probabilistic estimation of remaining useful life (RUL) for civil aerospace gas turbine engines. The hierarchical formulation allows Bayesian updates of an individual predictive model to be made, based upon data received asynchronously from a fleet of assets with different in-service lives and for the entry of new assets into the fleet.In this paper, variational inference is applied to the hierarchical formulation to overcome the computational and convergence concerns that are raised by the numerical sampling techniques needed for inference in the original formulation. The algorithm is tested on synthetic data, where the quality of approximation is shown to be satisfactory with respect to prediction performance, computational speed, and ease of use. A case study of in-service gas turbine engine data demonstrates the value of integrating fleet data for accurately predicting degradation trajectories of assets.

CMAS degradation of EB-PVD TBCs: The effect of basicity

Turbine entry temperatures of aerospace gas turbine engines will need to increase in future engine designs due to a constant desire to improve fuel efficiency and power whilst decreasing emissions which may be environmentally harmful. At present, thermal barrier coatings (TBCs) are employed together with a cooling film to protect engine hardware from the high temperatures seen within the turbine and combustion zones. Airborne particles such as dust, sand and volcanic ash, otherwise termed ‘CMAS’ (calcium–magnesium–alumino-silicates) are ingested into the engine, where at the elevated temperatures, this deposit can melt and induce degradation of the TBC.The columnar, high porosity microstructure of electron-beam physical vapour deposited (EB-PVD) TBCs, typically found on high pressure turbine blades, makes them particularly susceptible to CMAS molten deposit attack. This paper investigates the effect of modifying the chemical composition of molten CMAS deposits on the degradation mechanisms of EB-PVD TBCs. This paper identifies a parameter termed “basicity index” which can be used to evaluate the severity and damage caused by these molten deposits to TBCs. It is believed that the higher temperatures expected to be observed in next generation aero-engines coupled with sand/dust chemistry varying around the world, and then basicity-driven TBC degradation will become more prevalent.

Deposition and characterization of plasma sprayed Ni-5A1/ magnesia stabilized zirconia based functionally graded thermal barrier coating

Thermal barrier coatings (TBCs) are employed to protect hot section components in industrial and aerospace gas turbine engines. Conventional TBCs frequently fail due to high residual stresses and difference between coefficient of thermal expansion (CTE) of the substrate & coatings. Functionally graded thermal barrier coatings (FG-TBCs) with gradual variation in composition have been proposed to minimize the problem. In this work, a five layered functionally graded thermal barrier coating system was deposited by atmospheric plasma spray (APS) technique on Nimonic 90 substrates using Ni-5Al as bond coat (BC) and magnesia stabilized zirconia as top coat (TC). The coatings were characterized by SEM, EDS, XRD & optical profilometer. Microhardness and coefficient of thermal expansion of the five layers deposited as individual coatings were also measured. The deposited coating system was oxidized at 800°C. SEM analysis showed that five layers were successfully deposited by APS to produce a FG-TBC. The results also showed that roughness (Ra) of the individual layers decreased with an increase in TC content in the coatings. It was found that microhardness and CTE values gradually changed from bond coat to cermet layers to top coat. The oxidized coated sample revealed parabolic behavior and changes in the surface morphology and composition of coating.

High-temperature oxidation behavior of Mo–Si–B-based and Co–Re–Cr-based alloys

Improving the efficiency of aerospace gas turbine engines requires materials that can be used at increasingly higher temperatures in aggressive environments. This paper summarizes the current stage of alloy development of Mo–Si–B-based and Co–Re–Cr-based alloys regarding the high-temperature oxidation resistance. Since refractory metals, such as Mo and Re, suffer from catastrophic oxidation, the main task of research is to find alloying elements that improve the oxidation behavior of these alloys. For Mo–Si–B-based alloys, it was observed that an addition of Zr has a significant positive influence on the oxidation resistance by reducing the time necessary for the formation of a protective borosilicate layer. An addition of 0.2 at.% Y improves the viscous properties of the borosilicate increasing the protectiveness of the oxide scale. Macroalloying with Ti yields a strong positive effect on the oxidation behavior and, in addition, notably reduces the density of Mo–Si–B-based alloys. In Co–Re–Cr-based alloys, Cr is included to achieve favorable mechanical properties and to form a protective chromia layer during oxidation. As a consequence of the synergetic effect of Cr and Si, an addition of 2 at.% Si significantly improves the oxidation behavior of the alloy. Al addition further promotes the formation of the protective chromia layer at intermediate temperatures and exhibits the potential of the formation of a protective alumina scale suitable for applications at very high temperatures. The critical evaluation of the complex oxidation behavior of both metallic systems in a broad temperature range gives insight into the underlying fundamental mechanisms, reveals the potentials of particular alloying elements and, thus, guides future development of these material classes.

Properties and residual stress distribution of plasma sprayed magnesia stabilized zirconia thermal barrier coatings

Thermal barrier coatings (TBCs) are employed to protect hot section components in industrial and aerospace gas turbine engines. Conventional TBCs frequently fail due to high residual stresses generated within the coating systems owing to the difference between coefficient of thermal expansion (CTE) of the substrate & deposited layers. Functionally graded-thermal barrier coatings (FG-TBCs) with gradual variation in composition and microstructure have been proposed to minimize the problem. In the present study, three different TBC systems, one duplex and two FG-TBCs, having three and five layers, respectively, were prepared by atmospheric plasma spray (APS) process on Nimonic 90 substrates using Ni-5Al as bond coat (BC) and magnesia stabilized zirconia as top coat (TC) materials. Five layers comprising the three TBC systems were also deposited separately as individual coatings to study their characteristic properties. The coatings were characterized by a scanning electron microscope. Microhardness and CTE of the individual coatings were also measured. Finite element analysis (FEA) using ANSYS 14.5 was performed to estimate the thermal residual stresses generated within the as-sprayed coating systems. The microstructure, microhardness and CTE in the five-layer FG-TBC changed gradually as compared with duplex and three-layer FG-TBC. The lowest radial, axial and shear stresses were generated in the five-layer FG-TBC compared with those of the duplex and three-layer FG-TBC. Increasing the number of layers with a gradient in composition from BC to TC having similar thickness of coatings increased the bond strength of the FG-TBCs. The bond strength of five-layer FG-TBC was found to be almost 1.5 times as high as that of the duplex TBC.

On the Use of Wire-Coil Inserts to Augment Tube Heat Transfer

An experimental study was conducted to assess the impact of wire-coil inserts on the performance of compact fuel/air heat exchangers used in aerospace gas turbine engines. Experiments were conducted with JP-10 fuel using electrically heated tubes at relevant temperatures and flow rates for the application of interest. Three different coil pitches were considered and companion CFD calculations were performed to assess the flow fields qualitatively. Results indicate that the wire coils were very effective in augmenting heat transfer in this flow regime that straddles laminar/turbulent boundary layer characteristics. A Nusselt number correlation was developed for use in this application.

Response of titanium aluminide alloy to abrasive waterjet cutting: Geometrical accuracy and surface integrity issues versus process parameters

Due to their low thermal conductivity, high abrasion of cutting edges and a tendency to crack during machining, titanium aluminide (TiAl) alloys are notoriously difficult-to-cut materials when using conventional (e.g. turning, grinding, drilling, milling) cutting operations. Despite their low machinability TiAl alloys find niche application areas in the manufacture of complex shaped components for aerospace gas turbine engines. Abrasive waterjet (AWJ) machining is one of the most promising non-conventional machining processes for difficult-to-cut materials because of the reduced mechanical and thermal damage to workpiece surfaces produced using this technique. However, there is no information about the AWJ machining of TiAl alloys in the open literature. Based on a generic design of an aeroengine component, this paper investigates the response of a TiAl alloy to AWJ cutting process variables to enable generation of high-integrity surfaces. The effects of operating parameters (pump pressure, material removal rate, abrasive flow rate, and stand-off distance) on the output process quality measures have been investigated as follows: geometrical accuracy (kerf straightness); surface roughness homogeneity (measured along cutting direction on the cutting front surface to study the striation formation); workpiece surface integrity (grit embedment and possible recast micro-layers or material pull-out / micro-cracking after cutting). Advanced surface topography and metallurgical (e.g. SEM, EDX) analyses have been carried out to characterize the response of TiAl alloy to various AWJ conditions leading to the fulfilment of quality criteria of a group of targeted aeroengine components. It was found that the AWJ process has a very high capability (e.g. satisfactory geometrical accuracy, surface quality, minimum surface anomies) to cut TiAl alloy despite grit embedment. However, the authors have addressed this problem by plain waterjet cleaning know-how. In addition, a surprising phenomenon (bubble-like ‘TiO2-based spot’) caused by a strong exothermic reaction was found on the AWJ cut surfaces.

Prediction of maximum unbalance responses of a gear-coupled two-shaft rotor-bearing system

A general method was presented for obtaining the unbalance response orbit of a gear-coupled two-shaft rotor-bearing system, based on the finite element approach. Specifically, analytical solutions of the maximum and minimum radii of the orbit were proposed. The method was applied to the unbalance response analysis of a 600 kW turbo-chiller rotor-bearing system, having a bull-pinion speed increasing gear. Bumps in the unbalance responses were observed at the first torsional natural frequency because of the coupling between the lateral and torsional dynamics due to gear meshing. In addition, the analytical solutions were validated with results obtained by a full numerical approach. The proposed method can be generally applied to an analysis of the unbalance response orbits of dual-shaft rotor-bearing systems coupled by bearings as well, which are often found in aerospace gas turbine engines.

Effect of Copper on the Oxidation Behavior of Ti–48Al–2Cr–2Nb

Titanium-aluminide intermetallic alloys are candidate materials for some components in aerospace gas-turbine engines. Oxidation resistance and suitable joining techniques are two critical factors limiting the application of these materials. Transient-liquid-phase (TLP) bonding has been used successfully to join gamma TiAl-base alloys, but results in small amounts of copper being introduced into the alloy near the joint and this change in composition can affect oxidation resistance. The results of this investigation show that large copper additions (>2 at.%) to Ti–48Al–2Nb–2Cr result in the formation of copper-rich second phases, which can degrade oxidation resistance. On the other hand, smaller amounts of copper (1–2 at.%) improve the oxidation resistance of the alloy. Fortunately, the amount of copper in the alloy after completion of a successful TLP bond is relatively low (<1 at.%), so that the oxidation resistance of a TLP joint should be at least as good as that of the base alloy.

Property gradation for modification of response of rotating MMC discs

Aerospace and industrial gas turbine engine components such as discs, blades, and propeller shafts for which long fibre reinforced metal matrix composites are being made, operate under complex mechanical and thermal loading conditions. A major aim of functionally graded materials application is to optimise component response through appropriately tailored microstructures. This paper explores the influences of property gradation, centrifugal body force loading, and thermal loading on stresses in rotating discs. The discs were modelled as non-homogeneous orthotropic materials such as those obtained through non-uniform reinforcement of a metal matrix by long fibres. The results show how different temperature change distribution patterns and property gradation types correlate with hoop stresses developed in the disc.

The effect of surface modification on fretting fatigue in Ti Alloy turbine components

Severe fretting damage has been observed on the pressure surfaces of fan and compressor blade dovetails/disks in an aerospace gas turbine engine. A study has been carried out to evaluate the effect of an ion implantation technique in combination with the presently used surface treatments, such as shot peening and coating, on the fretting fatigue life of titanium alloy gas turbine engine components. The results from fretting fatigue tests, residual stress measurements, and nanoindentation tests were used to quantitatively evaluate the effect of various surface treatments on the fretting fatigue life of the fan blade and disk materials. Results from microstructural characterization and analyses of elemental and phase distributions within the implanted region are used to understand the effect of ion implantation on the surface properties of the alloys. Finally, an attempt has been made to evaluate the potential for improving the fretting fatigue life of the engine components using various surface modification techniques.