Single-crystal superalloys serve as the pivotal materials for blades and vanes of aeroengines and industrial gas turbines, as they display superior capability at high-temperature, complex-stress, and oxidative environments with microstructural stability balanced with excellent process performance. Furthermore, these alloys exhibit complex compositions and are associated with high costs, while their design methodologies remain relatively underdeveloped, they exhibit complicate deformation mechanism and manufacture process. Hence, much effort is paid to the related research on single crystal superalloy, and many results are obtained based on the new methods and new techniques. This article provides a comprehensive review of the recent advancements in alloy development, design methods, deformation mechanisms and fabrication processes. Additionally, it offers a brief overview of future research directions for single crystal superalloys and potential alternative materials.

To meet the stringent requirements for high-performance structural materials in the aerospace sector, high-entropy alloys (HEAs) have emerged as critical candidate materials owing to their unique design philosophy and distinctive performance characteristics. This work focuses on three typical HEA systems, namely refractory HEAs, lightweight HEAs, and eutectic HEAs. It systematically analyzes the representative compositions, microstructural features, and performance advantages of the three alloy systems, as well as the major technical bottlenecks and corresponding breakthrough approaches in core aerospace application scenarios such as ultra-high temperature service and lightweight design. By analyzing the research and development progress in major countries and regions worldwide, this work identifies the developmental trend of HEAs shifting from theoretical exploration to demand-driven research. It also sorts out the common scientific and technical challenges concerning composition design, fabrication, and performance evaluation that urgently need to be addressed for the engineering application of HEAs in aerospace, which are summarized as follows: (1) design and screening of HEAs component systems that break through the performance limits of conventional materials; (2) integrated computation and high-throughput rapid optimization technologies for the vast compositional space of HEAs; (3) high-quality and high-efficiency fabrication and forming technologies for complex HEAs components; (4) determination of cost-effective composition ranges that realize the coordinated optimization of mechanical performance and production cost for HEAs; (5) comprehensive performance assessment and database construction targeting extreme service environments. Furthermore, this work puts forward a forward-looking prospect on the future technological development directions of aerospace HEAs, aiming to provide theoretical support and technical references for accelerating the transformation of HEAs from fundamental research to vital engineering applications.

Fabricating metal-matrix high-temperature friction-reducing and wear-resistant composite coatings on the surface of tribo-pairs represents an effective strategy to improve the service life, stability and reliability of mechanical components under high-temperature conditions, and has thus received extensive attention in the aerospace industry. This article provides a review of the advancements in metal-matrix high-temperature friction-reducing and wear-resistant composite coatings, with emphasis on their performance requirements, compositional and structural design, typical coating systems, friction and wear mechanisms, and optimization methods. The coefficients of friction and wear rates of typical coatings across a temperature range from room temperature to 1000 ℃ are summarized and analyzed, and the evolution of their wear mechanisms with increasing temperature is concluded: at low temperatures, abrasive grooving and adhesive transfer dominate; as temperature increases, the mechanism gradually shifts to oxidative delamination and the formation of a tribo-oxide layer. Currently, the design and optimization of the coatings mainly depend on empirical design and trial-and-error method through adjusting the compositions of hard phases and lubricating phases. In recent years, data-driven machine learning methods have emerged as a new direction for coating composition design and performance optimization.

The influence of low angle boundaries (LABs) on the tensile behavior of a third generation single crystal superalloy at temperatures of 760, 850 ℃, and 980 ℃ has been explored. The microstructure of the specimens after tensile rupture is examined using scanning electron microscopy (SEM) and transmission electron microscopy (TEM). The findings reveal that the third generation single crystal superalloy with LABs demonstrates favorable tensile properties. LABs with angles ranging from 0° to 8.9° have a negligible impact on the yield strength and ultimate tensile strength of the alloy at 760, 850 ℃, and 980 ℃. At 980 ℃, the tensile strength of the specimen with LABs of 11.4° decreases by 80 MPa compared to that of the specimen with LABs of 0°. As the angle of the LABs increases, the elongation of the alloy with LABs in the range of 0°-11.4° at 980 ℃ exhibits a downward trend. Additionally, the propensity for intergranular fracture in the alloy with LABs becomes more pronounced as either the angle of the LABs increases or the temperature rises.

Taking investment-cast MAR-M246 superalloy as the research material, a single aging treatment is adopted to replace the conventional solution plus double aging heat treatment, and the effects of single aging treatment on the microstructure, room-temperature/high-temperature tensile properties, and medium/high-temperature stress rupture properties of MAR-M246 superalloy are investigated. A high-resolution field-emission scanning electron microscope (FE-SEM) is used to characterize the microstructures and fracture morphologies of the as-cast and aged alloys, while tensile and stress rupture tests were carried out on alloys under different states. The results show that the microstructures of both as-cast and aged MAR-M246 superalloys consist of γ matrix, γ′ precipitates, blocky intragranular MC carbides, strip-like intergranular MC carbides, and γ+γ′ eutectic phases. After aging treatment, the volume fraction of γ′ precipitates increases from 42.68% in the as-cast alloy to 49.85% in the aged alloy, and the average size of γ′ precipitates rises from (301±9) nm to (321±12) nm with a more uniform size distribution; meanwhile, the cubicity of γ′ precipitates is markedly enhanced after aging. After aging, the room-temperature yield strength of MAR-M246 superalloy is slightly improved, and the elongation after fracture at 900 ℃ increases significantly. Under the condition of 760 ℃ and 724 MPa, the stress rupture life of MAR-M246 superalloy is remarkably extended from 18.15 h (as-cast) to 38.23 h (aged), and the stress rupture elongation rises from 3.14% (as-cast) to 5.4% (aged). However, at 980 ℃ and 225 MPa, the stress rupture life declines from 79.58 h (as-cast) to 45.48 h (aged). These phenomena are mainly attributed to the fact that the γ′ precipitates with higher cubicity generated during aging are prone to rapid coarsening and rafting.

K417G alloy finds extensive application in the fabrication of aero-engine turbine blades. To tackle damage, such as cracks that emerge after prolonged service, brazing is a commonly adopted repair method. In this research, two gap sizes, namely 0.05 mm and 0.2 mm, are deliberately designed to mimic cracks of varying widths in the K417G alloy. Vacuum brazing experiments are carried out using BCo-1 cobalt-based filler metal at 1160 ℃ with different holding times. The microstructure and interfacial evolution of the joints are thoroughly analyzed via scanning electron microscopy, and their high-temperature tensile properties are tested on universal testing machine. The results reveal that the BCo-1 filler metal can establish a high-quality metallurgical bond with the K417G alloy. No defects, including pores or unbonded area, are detected in the joint area. The joint phases primarily comprise γ + γ′ phases, Cr-Mo-rich boride phase, and silicide phases. As the holding time is extended, the morphology of the compound phases in the 0.05 mm narrow-gap joint undergoes significant changes, and their content exhibits a trend of initially increasing and then decreasing. In the 0.2 mm wide-gap joint, the superalloy powder particles and compound phases gradually coarsen. The holding time has a relatively minor impact on the high-temperature performance of the 0.05 mm narrow-gap joint, which demonstrates tensile strength of 298 MPa at 950 ℃. In contrast, an extended holding time leads to deterioration in the performance of the 0.2 mm wide-gap joint. The optimal properties, with tensile strength of 460 MPa at 950 ℃, are achieved under the parameter set of 1160 ℃ for 15 min holding time.

Residual compressive stress is a critical factor affecting the performance and service life of components. The active introduction and precise control of residual compressive stress have become one of the core technical approaches to improve component reliability and extend service life. Aiming at the difficulties in optimizing shot peening process parameters and insufficient prediction accuracy of residual stress fields for GH4169D superalloy, this work adopts a combined simulation and experimental method for investigation. A multi-shot peening simulation model considering the random distribution characteristics of shot particles is established to analyze the material strain rate as well as the correlation between residual stress and plastic strain under various process conditions, and the accuracy of the random shot peening residual stress model is verified. Shot peening strengthening experiments are carried out based on the response surface methodology. The influence laws of different shot peening intensities and coverage rates on the depth distribution of the residual stress-affected layer are compared, and parametric modeling is performed for the depth distribution of the residual stress-affected layer. The results reveal that the cosine attenuation function can realize parametric modeling of residual stress fields of GH4169D superalloy under diverse shot peening processes, with the fitted determination coefficient R² greater than 0.95. These findings provide an effective method for rapid optimization of shot peening strengthening processes and accurate modeling of residual stress fields of GH4169D superalloy.

Ti₂AlNb alloys possess promising application prospects for typical aero-engine components operating at 650-750 ℃, yet they suffer from oxidation embrittlement under the complex high-temperature service environment. In this work, as-rolled Ti₂AlNb alloy is selected as the research material, and as-cast alloy with the identical nominal composition is set as the control group. A series of experiments including 100 h static high-temperature oxidation at 650-800 ℃, room-temperature tensile tests on oxidized specimens, and high-temperature tensile tests at 600-900 ℃ are carried out. X-ray diffraction (XRD), scanning electron microscopy (SEM), transmission electron microscopy/high-angle annular dark-field scanning transmission electron microscopy (TEM/HAADF-STEM), and electron backscatter diffraction (EBSD) are employed to characterize oxidation products, interfacial structures, and elemental distribution, so as to clarify the oxidation damage and tensile failure mechanisms of the as-rolled microstructure. The results reveal that after 100 h oxidation at 800 ℃, the mass gain of the rolled alloy reaches 13.7 mg·cm⁻², which is 21.2% higher than that of the as-cast alloy (11.3 mg·cm⁻²). The oxidation rate constant increases from 0.063 for the as-cast alloy to 0.091 for the rolled alloy, representing an increment of 44.4%. A multi-layer oxide scale forms on the rolled alloy, consisting of an outer mixed oxide layer, a middle TiO₂ particle layer and an inner oxygen/nitrogen-enriched embrittlement zone. Microcracks preferentially nucleate at α₂/O phase boundaries and propagate along oxidation channels. After oxidation, at room temperature, the elongation of specimens with oxide layer removed decreases from 12.5% (unoxidized state) to 0.7%-5.0%, while specimens retaining intact oxide layer exhibit a further drop in elongation to 0.4%-0.6%. At high temperature, as the temperature rises from 600 ℃ to 900 ℃, the yield strength declines from 671 MPa to 148 MPa, the ultimate tensile strength decreases from 759 MPa to 169 MPa, and the elongation increases from 14.1% to 161.7%. The tensile failure of as-rolled Ti₂AlNb alloy is jointly governed by oxide layer cracking, oxygen/nitrogen-enriched embrittlement and high-temperature softening.

To address the challenge that a single surface strengthening technique is unable to simultaneously tackle the issues of crack initiation and propagation in the fretting fatigue failure of titanium alloys, a composite treatment involving initial laser shock peening followed by nitrogen ion implantation at 300 ℃ is employed to enhance the properties of TC6 titanium alloy. The residual stress distributions of four types of specimens, namely the untreated ones, those subjected to nitrogen ion implantation at 300 ℃, those treated with laser shock peening, and those receiving the composite treatment, are measured using an X-ray diffraction stress analyzer. The fretting fatigue life of the titanium alloy specimens is assessed on a self-designed surface-contact fretting fatigue testing rig. The fracture surfaces and wear scars at the crack initiation zones are characterized using scanning electron microscopy (SEM) and energy-dispersive X-ray spectroscopy (EDS). The results demonstrate that the composite treatment creates both a high-hardness nitrided layer and a deep gradient residual compressive stress field, with the residual compressive stress extending to a depth of approximately 1.4 mm. The average fretting fatigue life of the specimens subjected to the composite treatment reaches 2.98×10⁵ cycles, which is 161.4%, 108.4%, and 30.1% higher than that of the untreated specimens, those implanted with nitrogen ions at 300 ℃, and those treated with laser shock peening, respectively. Fracture analysis reveals that the composite strengthening transforms the damage mechanism in the fretting contact area from severe adhesive wear to predominantly abrasive wear, significantly postponing crack initiation. Meanwhile, the deep residual compressive stress effectively reduces the crack propagation rate.

This work investigates the alterations in the microstructure and mechanical properties of the Ti-48Al-2Cr-2Nb alloy (referred to as 4822 alloy) following simulated short-term superheated service at 800-950 ℃. The research focuses on the effects of thermal exposure temperature and holding time on the microstructure, room-temperature and high-temperature tensile properties, and high-temperature endurance properties of 4822 alloy. The results reveal that after exposure at 800-950 ℃, the α₂ phase within the α₂/γ lamellar colonies dissolves, forming coarse γ lamellae. As the thermal exposure time increases, the dissolution of the α₂ phase becomes more pronounced, the ends of the α₂/γ lamellar structures coarsen, and the lamellae grow into adjacent colonies. After treatment at 900 ℃, spherical B2 precipitates form at the boundaries between lamellar colonies and equiaxed γ grains. At 950 ℃, fine B2 precipitates appear within the coarse γ lamellae inside the colonies, and the number of B2 precipitates increases. Prolonging the exposure time leads to the precipitation of extremely fine γ lamellae within blocky α₂ phases, with their quantity increasing; equiaxed γ grains at colony boundaries gradually grow, and the number of equiaxed γ grains within the colonies also rises. After thermal exposure at 850 ℃/10 min, 850 ℃/120 min, and 900 ℃/10 min, both the room and high-temperature yield strengths of the alloy increase, while the room and high-temperature elongations decrease. However, after thermal exposure at 950 ℃/10 min, the room-temperature elongation of the alloy is (2.4±0.55)%, representing an increase of 5.73% compared to its original state. After thermal exposure at 850 ℃/10 min, 850 ℃/120 min, and 900 ℃/10 min, the stress rupture lives of the alloy under both 650 ℃/400 MPa and 700 ℃/350 MPa conditions are extended. However, after thermal exposure at 850 ℃/120 min, the stress rupture life under 650 ℃/400 MPa is (14.85±3.55) h, which decreased by 94.02% compared to its original state.

Silicon carbide ceramic matrix composites (SiC-CMC) are lightweight, high-strength and high-temperature resistant thermal structural materials, which have been widely applied in high-tech fields such as aerospace. Preceramic polymers are key materials for the fabrication of such composites via the process of precursor infiltration and pyrolysis. However, the high cost of the materials has prevented their widespread use. Developing new silicon carbide ceramic precursors and reducing their preparation costs are of great significance for the advancement and application of SiC-CMC. In this work, a new silicon carbide ceramic precursor polymer, PMSCS, was synthesized via the Wurtz coupling reaction using dichloromethylmethylsilane (Cl₂CHSiMeH₂) and chloromethylmethylsilane (ClCH₂SiMeH₂) as the monomer and end-capping agent, respectively. Owing to the participation of reactive Si―H bonds in the monomer during polymerization, PMSCS is a branched polymer with a main chain consists of C―C bonds along with Si―C bonds, which exhibits good solubility and storage stability, with a ceramic yield of 74% (mass fraction) upon pyrolysis at 1000  ℃ under inert atmosphere. Compared with the most widely used Yajima method polycarbosilane, PMSCS offers a higher ceramic yield, and its synthesis is simpler, eliminating the need for a high-temperature rearrangement step. Additionally, the synthesis of PMSCS does not require expensive lithium aluminum hydride as a reducing agent, which is essential for the preparation of hyperbranched liquid polycarbosilane. Therefore, PMSCS shows promising potential as a cost effective silicon carbide ceramic precursor.

With the rapid development of the aviation industry, advanced aero-engines have imposed increasingly stringent requirements on material properties. To meet the demand for enhancing the thermal resistance and load-bearing capacity of polyimide composites, this study investigates the interfacial compatibility between the surface characteristics of carbon fibers and the polyimide resin matrix. The surface characteristics of carbon fibers under three different treatment conditions are characterized by scanning electron microscopy, atomic force microscopy, X-ray photoelectron spectroscopy, and surface energy measurements, and the corresponding polyimide composites are fabricated. The interfacial compatibility between carbon fibers and the resin matrix is evaluated by testing the thermal resistance and interfacial strength of the composites. Results reveal that the surface activity of untreated carbon fibers is relatively low, leading to interfacial cracks in the as-prepared composites, with an interlaminar shear strength (ILSS) of merely 64.5 MPa. In contrast, electrochemically treated carbon fibers exhibit a high content of surface active functional groups, and the ILSS of the resultant composite reaches 111 MPa, with a weight loss rate of only 6.5% after thermal aging at 400 ℃ for 100 h. Composites prepared from electrochemically treated carbon fibers coated with epoxy sizing demonstrate superior interfacial properties; however, their glass transition temperature T_g decreases by approximately 40 ℃ compared to those reinforced with unsized carbon fibers. This indicates that epoxy sizing compromises the thermal resistance of polyimide composites. Therefore, optimizing the sizing agent for carbon fibers represents a viable strategy to improve the interfacial compatibility of polyimide composites.