Aero-engine Components Research Articles

Homogenization treatment-induced reduction of brittle intermetallic compounds in Ti2AlNb/Ti60 brazed joints: Effect on microstructure and mechanical properties

The microstructure evolution and mechanical properties of Ti2AlNb/Ti-36.5Zr–10Ni–15Cu-0.5Co-0.5Nb/Ti60 brazed joints before and after homogenization treatment were comprehensively evaluated. Initially, the joints brazed at 900 °C for 15 min exhibited a continuous coarse strip network of Ti2Cu, Ti2Ni, Zr2Cu and Zr2Ni brittle intermetallic compounds in the central region of the braze seam, which detrimentally affecting the mechanical properties. Subsequent homogenization at 600 °C for 1 h, the amount of continuous brittle intermetallic compounds significantly decreased through sufficient atomic diffusion. Concurrently, the inadequately transformed eutectic β-Ti within the intermetallic compounds underwent an active eutectoid reaction. This structural transformation established a coherent interface with elastic distortion between the α-Ti and intermetallic compounds, effectively minimizing the lattice mismatch. Thus, a widmanstatten structure (i.e., acicular intermetallic compounds within a eutectoid microstructure matrix) with improved mechanical properties was formed. Consequently, a mean shear strength of 350.3 MPa was attained after treatment at 600 °C for 1 h, marking a substantial 206 % increase compared to that before homogenization treatment. This finding underlines the pivotal role of homogenization treatment in optimizing the performance of aeroengine components.

Hot Corrosion Damage Modeling in Aeroengines Based on Performance and Flight Mission

Hot corrosion is a form of chemical damage that causes surface degradation, sound material loss, and reduced component life. A lifing analysis in aeroengines without considering hot corrosion can lead to unexpected damage findings and increased scrap rates due to blade thickness loss beyond repair. This paper presents a novel methodology to predict hot corrosion damage based on aeroengine performance and flight mission analysis while taking into account environmental exposure, fuel quality, and material factors. The participating mechanisms, from salt and sulfur ingestion to deposition and hot corrosion attack, are discussed to explain the phenomenon in aeroengine components. In the investigated engine type, the first stage of the low-pressure turbine is the most affected. The application of the new methodology provides insights into the damage progression during the flight, the most affected components and the importance of capturing variations in the fuel quality, environmental exposure at the flight region, and the thrust derate policy. For a representative 1500 n mile mission, the variations in environmental exposure, fuel quality, and derate policy within typical limits can result in up to +350% damage. The outputs of the new framework can inform the decision making for maintenance, repair, and overhaul contract costing and scheduling.

Simulation of fatigue crack initiation life under compressive loading based on material meso-characteristics

The fatigue problem under compressive loading is common in the key components of aeroengine. Most of the fatigue life prediction methods studied in the past are macroscopic phenomenological methods based on a large amount of experimental data. Due to the difficulty in carrying out the fatigue test under compressive loading, the S-N curve or Manson-Coffin formula commonly used on the fatigue life prediction of material under compressive stress were obtained through fatigue tests under conditions where the average stress was greater than or equal to zero. Although the average stress can be corrected by some formulas, whether the life estimation formulas fitted by those fatigue test data are applicable to the fatigue life analysis with average stress less than zero has not been solved in engineering. And there were few studies on the influence of compressive stress on fatigue life at home and abroad. On the basis of Tanaka-Mura fatigue crack initiation theory, a mesoscopic model suitable for fatigue characteristic analysis under compressive loading was established in this article. The model verified the feasibility in predicting the fatigue initiation life using test data of 7075-T651 aluminum alloy under the compressive fatigue load.

An Engine-Level Safety Assessment Approach of Sustainable Aviation Fuel Based on a Multi-Fidelity Aerodynamic Model

Safety is essential for sustainable aviation fuels (SAFs). However, evaluating SAFs’ impacts on aero-engine safety is challenging because it involves multiple space scales and the strongly coupled relationships of aero-engine components. Aiming at addressing this problem, a model-based approach is proposed to establish the relationship between the fuel-level physical properties and engine-level safety parameters. Firstly, a unified modeling criterion is proposed to consider the interrelations of aero-engine components. Under this criterion, aero-engine secondary air system (SAS) components are included in SAF safety assessment, since they have non-neglectable influences on aero-engine safety. Secondly, this paper proposes a surrogate-based iteration strategy to embed the combustor’s high-dimensional computational fluid dynamics (CFD) model into the aero-engine flow network model. Then, the proposed model-based safety assessment approach is applied to a Fischer–Tropsch hydro-processed synthesized paraffinic kerosine (F-T SPK) safety assessment case. The effects of fuel flow and blending ratio are considered. The results indicate the necessity to evaluate SAFs’ safety at the aero-engine level and consider the influences of SAS components. The proposed model-based approach may provide a preliminary screening before SAFs’ certification tests. This convenience may be beneficial for reducing the cost and accelerating SAFs’ application.

Anomalous enhancing effects of electric pulse treatment on strength and ductility of TC17 linear friction welding joints

Mechanical properties of TC17 titanium alloy undergo a significant reduction after linear friction welding (LFW), of which the strength and ductility are hard to be improved simultaneously by traditional aging heat treatment (AHT), seriously limiting the application of LFW in the manufacturing of TC17 titanium alloy blisks. To this end, the present work proposes to use electric pulse treatment (EPT) to enhance the strength and ductility of TC17 LFW joints simultaneously by improving its microstructure. The results show that, in comparison to the uneven distribution of α phases in the welding zone (WZ), heat-affected zone (HAZ), and base metal (BM) zone after AHT, EPT can selectively homogenize the α phase distribution of WZ and HAZ without impacting the BM. The selective effect of EPT is reflected as the synergistic influence of the local Joule heating effect and the electron wind effect, which promotes the diffusion of β phase stabilizing element Mo and leads to a competitive precipitation of β phase and α phase in the α phase transition temperature range. The ratio of α phase to β phase in the WZ and HAZ finally approaches an equilibrium point which is similar to that of BM, leading to a uniform distribution of α phase and realizing the synergy of strength-ductility of LFW joint: the maximum strength increase observed is 12.9 %, accompanied by a corresponding elongation increase of 122 % (by AHT & EPT), and the maximum plasticity improvement is 185 %, accompanied by a corresponding strength increase of 4.3 % (by EPT for 1 h). This study provides essential insights for improving the strength and ductility of LFW TC17 titanium alloy blisks and enhancing the applications of LFW in aeroengine components.

Failure behavior of SiC/SiC with BSAS-based EBC in gas combustion environment

To promote the application of SiC/SiC in hot-end components in aero-engines, the thermal shock behavior of SiC/SiC and SiC/SiC-EBC were investigated in gas combustion environment between 450 °C and 1200 °C. The tensile strength change, microstructure evolution and thermal cycling performance of samples after 200, 400, 1000 and 2000 cycles were studied respectively to reveal the damage and failure mechanism. Cracking damage induced by thermal stress was the main reason for performance degradation of SiC/SiC after initial thermal cycles, then oxidation/corrosion damage dominated with increasing cycles. The oxidation/corrosion damage to SiC/SiC was effectively inhibited by EBC with almost no reduction in strength. However, the poor crack resistance of SiC/SiC was the main factor affecting its thermal shock performance. Improving the crack resistance or self-healing ability of SiC/SiC and combining it with EBC is considered to be an effective strategy to increase its service lifetime.

Improving fatigue properties of normal direction ultrasonic vibration assisted face grinding Inconel 718 by regulating machined surface integrity

Fatigue properties are crucial for critical aero-engine components in extreme service environments, which are significantly affected by surface integrity (SI) indexes (especially surface topography, residual stress σ res, and microhardness) after machining processes. Normal-direction ultrasonic vibration-assisted face grinding (ND-UVAFG) has advantages in improving the machinability of Inconel 718, but there is a competitive relationship between higher compressive σ res and higher surface roughness R a in affecting fatigue strength. The lack of a quantitative relationship between multiple SI indexes and fatigue strength makes the indeterminacy of a regulatory strategy for improving fatigue properties. In this work, a model of fatigue strength (σ f)sur considering multiple SI indexes was developed. Then, high-cycle fatigue tests were carried out on Inconel 718 samples with different SI characteristics, and the influence of ND-UVAFG process parameters on SI was analyzed. Based on SI indexes data, the (σ f)sur distribution in the grinding surface layer for ND-UVAFG Inconel 718 samples was determined using the developed model, and then the fatigue crack initiation (FCI) sites were further predicted. The predicted FCI sites corresponded well with the experimental results, thereby verifying this model. A strategy for improving the fatigue life was proposed in this work, which was to transfer the fatigue source from the machined surface to the bulk material by controlling the SI indexes. Finally, a critical condition of SI indexes that FCI sites appeared on the surface or in bulk material was given by fitting the predicted results. According to the critical condition, an SI field where FCI sites appeared in the bulk material could be obtained. In this field, the fatigue life of Inconel 718 samples could be improved by approximately 140%.

Impact of interfacial texture on failure behaviour of 8YSZ thermal barrier coatings under thermal cyclic loading

Thermal barrier coatings (TBCs) provide thermal protection for high-temperature components of aero-engines and steam turbines, but are susceptible to failure under thermal cyclic loads. The effective design of interfacial texture is one of the approaches to enhance the lifetime of ceramic TBCs. In this study, four types of interfacial textures were designed on the surface of substrates, and 8YSZ coatings of equal thickness were prepared using atmospheric plasma spraying, and thermal cycling tests were carried out. Meanwhile, the thermal cycle lifetime, microstructure evolution, residual stress, and energy release rate criterion were also studied. The optimal morphology and parameters of the texture are determined. It is found that the lifetime of the concave trapezoidal texture is the longest, with a deflection angle of 67°, diameter of 700 μm, depth of 500 μm, and spacing of 5 mm. This study can provide theoretical guidance for the interfacial texture design of ceramic TBCs.

Effect of constraint removal on single-crystal blade dimensions during investment casting

Single-crystal turbine blades, manufactured via investment casting, are key components of high-performance aero engines and gas turbines. The final dimensions of the blades depend on the cumulative effect of heat shrinkage at different stages, including the superalloy, mold shell, and ceramic core, etc. To clarify the effect of these stages on the blade dimensions, we first use the heat transfer coefficient calculated from inverse heat-conduction analysis to simulate the directional solidification. By extracting and mapping the final physical field from the previous stage to the next stage, four successive constraint removal processes are performed at different time points. Subsequently, we establish a time-varying model for the entire process of directional solidification and constraint removal. The errors of the extraction/mapping are consistently less than 0.002 mm in all three directions, thus confirming the feasibility of data transfer. Secondly, we analyze the dimensional deviation patterns of the casting after sequentially removing the mold shell, process rib, grain selector, and ceramic core constraints, with emphasis on the dimensional accuracy of the airfoil while disregarding the dovetail generated from machining. The final average deviations of Section 1, Section 2, and Section 3 of the airfoil are 0.1896 mm, 0.1723 mm, and 0.1909 mm, and the maximum deviations occur consistently at the trailing edge near the tip. Finally, three section deviations are validated via measurements performed using a vacuum furnace and industrial computed tomography. The results show that the proposed method improves the dimensional accuracy of hollow turbine blades.

Effect of laser shock peening without coating on fretting wear behavior of GH4169 superalloy at high-temperature

Improving the high-temperature wear resistance of aero-engine components is critical. Laser shock peening without coating (LSPwC) seems to have more potential because of lower morphology changes and simpler processing compared with conventional laser shock peening (LSP) techniques. The influence of LSPwC process and test temperature on the fretting wear behavior of GH4169 superalloy for 2 × 105 cycles was investigated. The results indicate that a molten layer and a hardened layer are formed sequentially below the periodic dimple structure under the combination of shock wave and thermal effect. The formation of hardened layer enhances the difficulty of abrasive ploughing and the increase in test temperature promotes the generation of a glaze layer. Therefore, both LSPwC process and high temperature cause a reduction in wear volume. However, the main wear mechanism is similar before and after LSPwC. Besides, the protective mechanism of glaze layer under the gross slip regime may be ascribed to the frictional sintering process.

Study on aluminium oxide doping modification of indium oxide and thermoelectric properties

Traditional thermocouples face challenges in meeting the temperature measurement requirements of crucial aero-engine components, such as low interference, high temperature resistance, quick response, and in-situ measurement. The temperature measurement requirements mentioned above can be satisfied by thin film thermocouples, thanks to advancements in thin-film preparation technology. In this study, the indium oxide (In2O3) was calculated using first principle calculation both before and after doping. It was found that doping Aluminium oxide (Al2O3) can increase the Seebeck coefficient of In2O3. Solid-phase sintering was used to create aluminum oxide-doped In2O3 nanopowders, and thin film thermocouples were fabricated using screen printing. The thermoelectric characteristics of thin film thermocouples were studied to analyze the effects of different annealing procedures and doping concentrations. The optimal annealing process was determined to be at 1300 °C for 1 h, and the ideal doping ratio was found to be Al0.1In2O3.15. The indium oxide film doped with aluminum oxide exhibits a Seebeck coefficient that is 45.9% higher than that of the undoped indium oxide film. The thin film thermocouples underwent tests for impact and vibration resistance, laser pulse response, surface and cross-section morphology, X-ray diffraction, and thermoelectric output. It has been confirmed that doping indium oxide with aluminum oxide can increase a material's Seebeck coefficient. Additionally, a thin-film temperature sensor that has been prepared can be used in challenging conditions to meet the temperature measurement requirements of an aero engine.

Dynamic modeling and vibration analysis of bolted flange joint disk-drum structures: Theory and experiment

Bolted flange joint disk-drum structures (BFJDSs) are key components in aero-engines, in which fatigue damage often occurs due to structural vibration. However, no study has so far given an effective theoretical dynamic model of BFJDSs. In this study, an effective dynamic model of BFJDSs is proposed. The present model considers the flange effect, non-uniform contact pressure of bolted joint, and bolt mass. The theoretical expression for describing the stiffness variation of bolted joints is presented. During the modeling process, the energy functions of disk, drum, and flange are obtained according to the Kirchhoff plate, Sanders’ shell, and Euler-Bernoulli beam theories, respectively. The non-uniform artificial spring technique is adopted to simulate the pressure distribution of bolted joint. This method is capable of taking into account the contact pressure, which is closer to real contact situation. The displacement functions of disk and drum are expanded by using the Chebyshev orthogonal polynomials, and the motion equations are obtained by employing the Lagrange equation. Then, the experimental studies are performed on BFJDSs to illustrate the effectiveness of the theoretical model. Finally, the frequency veering and coupling vibration phenomena are revealed. The comparison studies indicate that the established theoretical model can well predict the vibration characteristics of BFJDSs under both bolt looseness and no-looseness conditions. The maximum error can be reduced from 46.81% to 3.61% by using the present dynamic model.

Computational Study on Compressor-Combustor Interactions for Improved Matching of a Micro Gas Turbine

ABSTRACT Understanding complex aerodynamics/combustion interactions across major components of aeroengines, including the compressor, combustor and turbine is demanded for the improved performance of advanced propulsion systems. This study employs high-fidelity CFD simulations to investigate the coupled compressor-combustor interactions for the improved performance of a micro gas turbine. In the prototype configuration, the deficient compressor/combustor matching causes insufficient air/fuel mixing and delayed combustion inside the combustor chamber. By reducing the flow area of the compressor and redistributing the air holes on the combustor chamber, a stable united recirculation structure is formed to promote air/fuel mixing and efficient combustion, achieving a over 50% decrease in OTDF and RTDF at the combustor exit and a total 20% decrease of pressure loss in the entire engine through-flow. Compared with the single-component combustor simulation, the coupled compressor-combustor results exhibit stronger turbulence mixing, increased temperature uniformity inside the combustor chamber. Compared with the single-component compressor simulation, the compressor in the multi-component simulation shows the reduced flow angle and Mach number at the compressor outlet, leading to increased stall critical angles and delayed flow separation at the compressor diffuser and vanes. This demonstrates that the added combustor endues the compressor with an expanded operating range and a lower mass-flow rate instability boundary. This work implies that 3D multi-component CFD simulation is necessary to understand the aerodynamic and combustion interactions and improve the multi-component matching in gas turbine engines.

Experimental and numerical study on projectile impact damage of EBC-SiC/SiC

The ingested particles or debris will cause high-speed & low-energy impact to the hot end components of aeroengine. Few literature focus on the projectile impact of SiC/SiC with Environmental Barrier Coating(EBC). The impact of 2 mm projectile was specially studied for Si/Yb2Si2O7/Yb2SiO5-SiC/SiC. Gas gun system was used to conduct impact testing at the speed of 3 m/s∼353m/s. The optical microscope, Scanning Electronic Microscope(SEM), micro CT were used to observe the impact and tension after impact(TAI) damage. For TAI, Digital Image Correlation (DIC) was used to conduct the full-field strain detection around the impact point. Meanwhile, the Smoothed Particle Hydrodynamics (SPH) methodology was programmed in C++ and developed to simulate SiC/SiC impact problem by introducing modified Hashin-Rotem criteria, extending interface method to 3D and adding GPU accelerated algorithm. The results indicate that threshold velocities of EBC huge peeling and SiC/SiC penetration are different, TAI residual strength also has extinct threshold at the velocity of 246 m/s with 37.1% declining extent, and the impact pit depth of SPH simulation is in good agreement with the experiment.

Investigation of neighboring grain effects on load shedding in titanium alloys under cold dwell fatigue

Load shedding under cold dwell fatigue in titanium alloys has been a threat to the safety of the aero-engine since it is responsible for crack formation. The strain incompatibility between the soft-hard grain pair arises from the significant creep in the soft grain during the dwell period, which is fundamentally related to the load shedding. Such deformational heterogeneity is not only affected by the soft-hard grain itself but also by its neighboring grains, but the effect of the latter is rarely reported. In this paper, the effect of neighboring grains, including location, crystal orientation, texture, etc., on load shedding is quantitatively investigated using a three-dimensional rate-dependent crystal plasticity model by controlling the microstructural morphology of the grains. In particular, the mechanism that the dwell fatigue cracks tend to nucleate from the sample's interior rather than the free surface is revealed. Restriction of rigid body rotation from neighboring grains can further enhance load shedding. The crystal orientation of individual and group neighboring grains sitting at different locations with respect to the soft-hard grain pair is evaluated. The grain directly attached to the hard grain has the strongest influence. The statistics of load shedding are analyzed by conducting hundreds of simulations with different textures. Stronger textures are found to be able to enhance load shedding and thus deteriorate the service lifetime. The change of locations showing the highest stress during the dwell period arising from plastic deformation in neighboring grain may lead to the phenomenon of reverse load shedding. Understanding the neighboring grain effects from both the mechanistic basis and the statistical point of view is important in integrity evaluations and service life assessments of titanium components in aero-engines.

Enhanced heat transfer in 3D printed ball-end grinding tool with blade-shaped structure

Complex curved components in aero-engines often require demand small-sized compliance ball-end (BE) grinding tools to achieve smooth machining in interference-prone-area. Nevertheless, heat accumulation during the grinding of difficult-to-process materials challenges the weak stability of flexible materials, thereby limited tool life and grinding performance. In this study, a novel rotary-enhanced heat transfer structure (REHT) based on aero-engine blades was introduced into the BE. The blade-shaped structure tools were fabricated using multi jet fusion (MJF). Numerical simulations were conducted to investigate the effects of blade type, rotating direction, rotating speed, and forced cold air velocity on the heat transfer mechanism of BE. The grinding performance of these tools, along with conventional structures, was comparatively investigated through robot-assisted grinding of titanium plates. The results indicated that the REHT improved the heat transfer capability of the tool by introducing high momentum fluids into the tool cavity and increasing the flow rate of the grinding interface. The REHT structure reduced the grinding temperature by over 19.84%, extending tool life by more than 40%. Furthermore, compliant grinding tools with REHT exhibited higher cumulative material removal alongside lower and more consistent surface roughness (Ra).

A novel thermal history sensor for thermal barrier coatings based on europium (III) ion self-reduction in barium aluminate

Thermometry of thermal barrier coatings (TBCs) is essential for designing effective thermal management systems and prolonging the service life of components in aeroengines and gas turbines. However, conducting on-line thermometry in these systems is challenging because it requires physical or optical access to components during operation. A promising alternative is off-line thermometry, in which phosphorescent thermal-history sensors are used to record exposure temperatures after operation. In this study, we synthesized europium (III)-ion-doped barium aluminate (BaAl2O4:Eu3+) via a solid-state reaction and confirmed its ability to function as a thermal history sensor. Subsequently, we developed a new thermal-sensing strategy based on the self-reduction of Eu3+ to Eu2+ in a BaAl2O4:Eu3+ phosphor exposed to high temperatures (900–1300 °C) in ambient air. The Eu3+ concentration decreased as the heat exposure temperature increased due to self-reduction of Eu3+ to Eu2+, which irreversibly changed the phosphorescence of BaAl2O4:Eu3+, i.e., its spectra, the ratio of the phosphorescence intensities of Eu3+ and Eu2+, and the lifetime of Eu3+. As the heat exposure temperature increased, the ratio of the phosphorescence intensities of Eu3+ and Eu2+ monotonically decreased due to self-reduction, whereas the phosphorescence lifetime of Eu3+ initially gradually increased until the heat exposure temperature reached 1000 °C because of the degeneration of concentration quenching, and then decreased as the temperature exceeded 1000 °C because of the degeneration of lattice distortions. This demonstrates that heat exposure temperatures exceeding 1000 °C can be quantitatively evaluated through off-line analysis of Eu3+ and Eu2+ phosphorescence intensity ratios and Eu3+ lifetime at room temperature. Thus, BaAl2O4:Eu3+, a self-reducing phosphor, is a potential sensor for off-line thermometry of TBCs.

High-temperature dry sliding wear behavior of additively manufactured austenitic stainless steel (316L)

Austenitic steels are commonly used in aero-engine components owing to their higher thermal resistance to wear and corrosion. This study investigates the high-temperature dry sliding wear behavior of direct metal laser-sintered (DMLS) austenitic stainless steel (SS 316L) up to 400°C. A dry sliding reciprocating wear test conducted on a high-frequency tribometer with a ball-on-flat setup. High-hardness chrome steel (900 HV0.5) and alumina (1165 HV0.5) balls of 6 mm diameter were used as the steel and ceramic counterbodies against 316L specimen (320 ± 5 HV0.5). When tested against the chrome steel, the initial average coefficient of friction (CoF) decreased to a lower value (0.45) during the initial running-in period. However, with rising temperatures, the CoF also increased, stabilizing at 0.6 under steady-state conditions. A similar CoF pattern was observed in experiments with the alumina counterbody. Wear resistance declined at 100°C, followed by an increase at 200°C due to the development of an oxide glaze on the surface. Subsequently, wear resistance decreased as the temperature continued to rise. Up to 200°C , the wear mechanism was characterized by a combination of abrasive and oxide debris wear, transitioning to adhesive wear with plastic deformation at higher temperatures (400°C). Diffraction (XRD) analysis revealed the formation of oxide layers and martensite phase at higher temperatures, contributing to the observed improved wear resistance around 200°C . However, these oxide layers were progressively removed from the surface at 400°C, leading to an increased wear. The measured wear rates are two times lower, up to 200°C, compared to the reported values for conventional 316L due to the fine microstructure and higher hardness of DMLS SS 316L. This results in potentially rendering it suitable for applications where high-temperature (up to 200°C) wear-resistant SS is needed. This study provides further insights into the importance of post-treatment of DMLS parts for temperature beyond 400°C.

A phase-field fatigue fracture model considering the thickness effect

Modern aero-engine turbine structures often have variable-thickness geometries (e.g., turbine structures with complex cooling systems) that pose a significant challenge to crack analysis. The phase-field fracture method is a numerical method that has recently received considerable attention. However, phase-field fatigue fracture models that consider the effect of thickness are still lacking. To address this problem, this paper proposes a modification of the fracture toughness in the phase-field fatigue model by introducing a relationship function between material fracture toughness and thickness (Gc-B curve). We use reference experimental data to fit the Gc-B curve of GH4169 and predict the fracture toughness of specimens with other thicknesses. The predictions of the proposed model agree well with the experimental results reported in the literature. This achievement contributes significantly to the structural safety design and life prediction of aero-engine components.

Design aspects of a CMC coating-like system for hot surfaces of aero engine components

Ceramic Matrix Composite (CMC) is an emerging material system that can be a game changer in the aerospace industry, both civil and military. CMCs components are, in fact, lighter and less prone to fatigue failure in a high temperature environment. However, at high temperatures, the diffusion of oxygen and water vapour inside the CMC can have detrimental effects. Therefore, the presence of protective coating is necessary to extend the life of CMC components. In the present work, a three-layers coating, consisting of a silicon bond (BND), adhesively bonded to the CMC, an Environment Barrier Coating (EBC) and a softer layer 3 (LAY3), is investigated for a CMC component. An aero-engine high pressure turbine seal segment was considered. Two design aspects are covered: (i) creep law is determined and calibrated in environment Abaqus from the experimental data of each coating layer available in the open literature, to provide a suitable instrument for the creep relaxation analyses of hot components; (ii) thickness sensitivity study of each layer of the coating is conducted to minimise the interface stresses of coating with substrate in order to mitigate cracking and removal/spalling phenomena when exposed to temperature gradients and to increase their service life. These two different aspects are combined together to predict the coating stress field as a function of service time.