The research progress of computed tomography (CT) technology in defect detection for aero-engine composite components was systematically reviewed. It addressed the bottleneck problem of accurately detecting and evaluating the cross-scale and multi-form defects generated in composites during their application in aero-engines, focusing discussions from the perspectives of the process mechanisms of defect formation, the technical principles of CT detection, and intelligent recognition algorithms. It provided a detailed analysis of typical defects in key components made of resin-based, metal matrix, ceramic matrix, and carbon-carbon composites. By reviewing the application scenarios of CT technology and starting from the typical defect characteristics and structural morphology differences revealed by CT detection, the effectiveness and limitations of different technologies, such as micro-focus CT and synchrotron radiation CT, were compared in identifying various types of defects. The results indicated that intelligent recognition technology based on deep learning is an effective approach for achieving accurate characterization of cross-scale defects, while multi-modal fusion detection is an important development direction for solving the challenge of micro-defect detection in thick-walled components. This review systematically investigated the compatibility of different matrix composites in CT detection, providing a theoretical basis and technical support for the full-lifecycle intelligent detection and reliability assessment of composite components in aero-engines.

In response to the design requirements of high thrust-to-weight ratio aero-engines, a single-web turbine disk of an aero-engine was selected as the benchmark. The temperature field was numerically calculated using ANSYS Fluent and the stress field was calculated using ABAQUS. The distribution characteristics of circumferential stress, radial stress and Von-Mises equivalent stress of single-web and twin-web turbine disks were compared and analyzed. The influence law of the web inclination angle (θ) variation (range: 8°≤θ≤ 12°) on the stress level of the twin-web turbine disk was mainly studied. Results showed that compared with the reference single-web turbine disk, the structural weight reduction effect of the twin-web turbine disk was obvious. The weights of the twin-web turbine disks with θ=8°, 9°, 10°, 11° and θ=12° were reduced by 22.78%, 23.23%, 23.42%, 23.2% and 23.16%, respectively. The structural characteristics of the twin-web transmission path made the load and stress distribution of the twin-web turbine disk more uniform. The stress level of the twin-web turbine disk with θ = 12° was the best within the study range, and its maximum circumferential stress, maximum Von-Mises equivalent stress, maximum average circumferential stress at the key points and maximum Von-Mises equivalent stress at the key points were respectively reduced by 15.66%, 16.03%, 15.65% and 17.09% compared with the twin-web turbine disk with θ= 8°. In the future, the web inclination angle θ should be regarded as a key factor in the design of the structure of the twin-web turbine disk.

Bolted connections are widely used in aerospace equipment assembly due to their structural simplicity and high reliability. The dynamic characteristics of bolted connections significantly influence the dynamic behavior of assembled structures. While detailed finite element models can analyze vibration characteristics, their computational cost is often prohibitively high, making them unsuitable for analyzing complex bolted assemblies. To address this issue, this study proposes an equivalent bolted connection unit model, which simplifies the bolted connection into a stepped double-ring thin-layer unit to simulate the stress distribution at the bolted interface. The model consists of three parts: the upper joint, the double-ring thin-layer unit, and the lower joint. The size of the thin-layer unit model is determined based on the bolt stress distribution, and key parameters of the ring-shaped thin-layer unit—including thickness, elastic modulus, Poisson’s ratio, and density—are derived using Hertzian contact theory. The accuracy of the proposed model in simulating dynamic characteristics is validated through comparison with vibration experimental data from a typical aerospace tooling structure. The results demonstrate that, compared to the traditional virtual material method, the proposed model significantly improves simulation efficiency while maintaining vibration frequency simulation errors within 10%. This indicates that the model is effective for dynamic simulation analysis of bolted connection structures.

To enhance the fatigue life assessment accuracy for critical regions of aero-engine turbine disks and support high-reliability, long-life disk design, the challenge of accurate low-cycle fatigue (LCF) life prediction for disk eccentric bore structures was addressed. An innovative three-parameter critical distance (TCD) model was proposed to overcome the limitations of traditional TCD models, which suffered from restricted fitting accuracy due to the complex relationship between stress gradients and fatigue life. The core innovations lied in: introducing normalized maximum stress gradient and high-stress zone dimensions to characterize stress gradient effects, establishing the three-parameter critical distance expression; revealing a strong correlation between stress gradients and critical distance, and validating through notch specimen LCF tests demonstrating differential evolution of critical distance with life under contrasting high and low stress gradients; establishing a comprehensive life assessment framework integrating elastoplastic finite element simulation with the Smith-Watson-Topper fatigue life model, enabling direct critical distance calculation and accurate life prediction. Validation via LCF testing on simulated disk eccentric bore specimens demonstrated that the proposed model significantly reduced life prediction error, bringing results from outside the triple error dispersion band of traditional methods to within a 1.5-fold dispersion band. A substantial improvement in prediction accuracy for life-critical regions with stress concentrations was achieved, directly supporting the aero-engine turbine disk strength design, safety margin optimization, and life extension decision-making, and underscoring its significant practical engineering value.

To evaluate the feasibility of beam seal fittings as connectors for aviation hydrogen pipelines, an analysis of the self-sealing performance of beam seal fittings was conducted. A multi-scale finite element model of the beam seal fitting, incorporating the surface topography characteristics of rough sealing surfaces, was established. Based on this model, simulations were performed to calculate the real contact area and contact pressure in the sealing region under various working conditions. The law of the self-sealing performance of beam seal fittings varying with preload force, sealing medium temperature, and medium pressure was investigated. The results indicated that the self-sealing performance of beam seal fittings was primarily attributed to the sealing performance of the first seal. An increase in axial preload force and sealing medium pressure enhanced the self-sealing effect and improved the sealing performance of the first seal. In contrast, the temperature of the fluid medium had no significant impact on self-sealing performance, as the sealing performance of both the first and second seals remained largely unchanged within the investigated temperature range. The findings demonstrated the stability of the sealing performance of beam seal fittings across a wide temperature range and their adaptability to high-pressure environments.

By establishing a mathematical model of the bird strike process of aero engine fan blades, an analytical method to quantify the bird strike damage of fan blades was proposed. Based on the kinetic energy of the bird strike process and the key structural parameters of the blade’s resistance to bird strike, the equivalent stress of bird strike was defined to reflect the damage level of the fan blades and their resistance to bird strike capability. And a systematic analysis of bird strike damage to the actual fan blades of an engine under different working conditions was conducted, clarifying the most critical working conditions for bird strike on fan blades under different engine states. The research shows that the fan rotate speed, aircraft flight speed, as well as the blade leading edge angle and thickness all jointly influence the bird strike damage to fan blades. The most critical impact position under different working conditions may vary. Changes in the mass of the bird do not affect the most critical impact position. The equivalent stress of bird strike can quickly analyze the resistance of fan blades to bird strike and determine the most critical working conditions, providing an efficient analytical tool for the evaluation during the design optimization phase and the selection of test assessment schemes.

Multi-physics numerical simulations of gas turbine rotor blades are computationally intensive, making it difficult to be directly applied for online condition monitoring and life prediction. In order to address this challenge, a reduced-order modeling method for multi-physics-based low-cycle fatigue life prediction of gas turbine rotor blades was developed, aiming to improve both computational efficiency and prediction accuracy. Based on the multi-physics results of fluid-thermal-solid coupled numerical simulations of turbine blades under representative operating conditions, a reduced-order model was constructed using proper orthogonal decomposition combined with data-driven regression techniques, enabling rapid and accurate prediction of temperature, stress, and strain fields. On this basis, the Manson-Coffin and Smith-Watson-Topper methods were employed for efficient evaluation of the blade’s low-cycle fatigue life. Results showed that the average relative error of the constructed reduced-order model was 0.11% for the temperature field, 1.01% for the stress field, and 0.75% for the strain field. The prediction speed was only 0.005 s for the temperature field, 0.03 s for the stress field, and 0.31 s for the strain field. The average relative error of the low-cycle fatigue life prediction was less than 3.5%, providing important theoretical and methodological support for online condition monitoring and life assessment of gas turbine rotor blades.

The ceramic matrix composites (CMCs) and metal connection structures in aero-engine combustion chambers and turbine guide vanes were systematically examined, covering structural design, numerical simulation, and experimental validation. Current studies put focus on four thermal mismatch mitigation strategies: transition region coordination, clearance design, interlocking fit, and thermal stress-free fastener. Multi-scale modeling and progressive damage analysis were widely employed in simulations, alongside a stepwise experimental verification framework. Structural configurations varied with specific application scenarios and operational conditions. Future development should combine multiple design strategies and shift from component-level connection design to system-level thermo-mechanical-cooling multifunctional coordinated design, so as to ensure long-term structural reliability under extreme conditions.

Thermal contact resistance at the interfaces of turbine shafts and blades is critical for the thermal analysis and durability design of aero-engines. Accordingly, an experimental apparatus based on the steady-state heat flux method was developed to measure thermal contact resistance in IN718 superalloy contacts. The measurement uncertainty of the setup was first evaluated using a single-point thermal contact resistance method, demonstrating that the relative errors for dimensionless single-point thermal contact resistance were within 18%. Subsequently, experimental investigations were conducted to examine the effects of two levels of surface roughness, interface pressure (0.05~10 MPa), and average interface temperature (20~160 °C) on the thermal contact resistance of IN718 specimens. The results indicated that thermal contact resistance decreased with the increasing pressure and temperature, but increased with higher surface roughness. A comparison of the experimental data with existing classical semi-empirical models revealed significant prediction inaccuracies, particularly within specific pressure and roughness ranges. Consequently, considering both the Gaussian and non-Gaussian distribution characteristics of the surface profile, a dimensionless empirical correlation was proposed. The maximum error between the predicted values from this correlation and the experimental data was less than ±20%.

To meet the regulatory requirements for the operational performance of feedback-free self-excited sweeping nozzles across various application scenarios, this paper proposes a parametric design method for such nozzles. A two-dimensional numerical simulation approach was employed to elucidate the effects of structural parameters on the internal flow mechanisms and macroscopic operating characteristics of the nozzle. The study revealed that the formation of self-excited oscillatory flow within the nozzle critically depended on the optimization of three key parameters: the impingement angle, the inlet-to-outlet ratio, and the distance between the impingement point and the outlet. Frequency control exhibited a strong correlation with the impingement angle, impingement point position, and the chamber length-to-width ratio. The frequency characteristics were primarily governed by vortices formed at the geometric dome; reducing the chamber length-to-width ratio, adjusting the impingement angle, and modifying the distance between the impingement point and the nozzle exit enhanced the sweeping frequency. Furthermore, the inlet-to-outlet ratio significantly impacted the sweeping angle and flow characteristics. A decrease in this ratio led to an increase in the sweeping angle and amplified gas-liquid phase velocity gradients. However, it also resulted in increased flow loss.

To address the thermal protection requirements of combustor liners under high-temperature and elevated heat loads, the single-tube combustor model incorporating tangential effusion cooling technology was developed. A conjugate heat transfer analysis was employed to numerically investigate the influence of varying inlet temperature, inlet pressure, and fuel-air ratio on liner wall temperature and cooling performance. The results show that the higher inlet temperatures reduced cooling film coverage effectiveness, weakened heat transfer intensity within effusion holes and on the outer liner wall, decreased average integrated cooling efficiency, and increased both wall temperature and thermal gradients. Under the condition of this paper, variations in inlet pressure exhibited minimal influence on near-wall flow and heat transfer, with negligible differences in effusion cooling effectiveness. Under extreme conditions that the proportion of cooling air is only about 19%, fuel-air ratio of 0.042, and combustor temperature rise approaching 1300 K), the liner maintained a high average integrated cooling efficiency of 70%—90%, demonstrating the significant potential of optimized tangential effusion cooling designs for next-generation ultra-high-temperature combustors.

For the cooling design for the complex and compact structure of the swirl-cup combustor head, three cooling design schemes were proposed. Numerical simulations were conducted to calculate the flow field, temperature field, and wall temperature distribution for different schemes. The results showed that, without cooling, corner vortices near the wall in the head led to the formation of high-temperature stagnation vortices by drawing in burning gases, causing a wide range of high temperatures on the surface of the flame tube head with a maximum wall temperature of 1178 K. The head cooling designs involved introducing cooling air between the inner and outer fairings to form a protective air film and disperse the near-wall corner vortices, effectively reducing the head wall temperature to varying degrees. The use of inclined multi-hole in the thermal shield was limited by geometrical dimensions, resulting in uneven development and poor adhesion of the air film, and leading to ineffective cooling. By designing a guide ring structure, it is possible to control the diversion of cooling in a compact space and to adjust the flow path design to increase the cooling airflow rate and velocity, forming a uniform air film adhering to the wall at the thermal shield and the swirler housing, with a wall temperature below 870 K. Under high temperature and high pressure conditions, the optimized scheme was validated through combustor component tests. The results demonstrated that the combustor head wall temperature did not exceed the material's long-term permissible operating temperature, proving that the head cooling design could effectively control the wall temperature.

To investigate the ignition, extinction and flame crossover characteristics of a newly designed dual-vortex-controlled radial staged combustor, an experimental study was conducted to examine the ignition and flame propagation mechanisms in the concave cavity and mainstream zone under different operating modes, as well as stable combustion patterns and the ignition/extinction characteristics of the combustor under various operating conditions. The results indicated that the combustor could be successfully ignited when fuel was independently supplied to either the outer-ring concave cavity or the mainstream zone, flame propagation and stable combustion were also achievable between the outer-ring concave cavity and the mainstream zone. Under both flame propagation modes, the stabilized combustion flame exhibited consistent pattern, and distributed within the concave cavity, its downstream region and the downstream of the swirl cup. The outer-ring concave cavity achieved a minimum fuel-air ratio of 0.0094 for ignition and 0.0034 for extinction under room temperature and atmospheric pressure conditions, demonstrating significantly superior ignition/extinction performance compared with the mainstream zone.

To investigate the influence of main stage swirl number on NOx emissions in a centrally-staged lean premixed prevaporized combustor across wide operating conditions, three-dimensional numerical simulations were used to analyze the turbulent combustion and NO formation characteristics during the reference emissions Landing and Take-Off (LTO) cycle. And the simulation methodology was validated against experimental NOx emission data. Furthermore, the influence of main stage swirl number on NOx emissions under LTO cycle conditions was simulated and analyzed. The results revealed that the influence of main swirl number on NOx emissions varied under different operating conditions. At idle conditions, an increase in main swirl number enhanced fuel-air mixing in the pilot combustion zone, leading to higher combustion temperature and elevated NOx emissions. Under high-load conditions, as the main swirl number increased, additional airflow was entrained into the re-circulation zone, further reducing the fuel-to-air ratio in the lean primary combustion zone and thereby decreasing NOx emissions. With increasing main swirl number over the range of 0.5 to 0.9, total NOx emissions during the reference emissions landing and takeoff cycle decreased.

An experimental system was built to study the effects of RP-3 and RP-5 fuels at different fuel temperatures on the flow rate characteristics, cone angle, Sauter mean diameter (SMD) and droplet size distribution. The results showed that under the same fuel pressure, as the fuel temperature decreased from 20 ℃ to −30 ℃, the density, kinematic viscosity, and surface tension of RP-3 and RP-5 fuels increased, and the flow rate also increased. The flow rate of RP-5 fuel was 2% to 4% higher than that of RP-3; the cone angle decreased by 14% for RP-3 fuel and 23% for RP-5 fuel; as SMD increased, the droplet size distribution index first decreased and then increased, indicating that the decrease in fuel temperature led to an increase in the average droplet size, but the unevenness of droplet size decreased and the droplet size distribution became more concentrated.

To investigate the mixing characteristics of fuel and supersonic inflow, a numerical simulation study was conducted based on the annular-slot-orifice injection structure in the combustion chamber of an open-barrel-type rotating detonation ramjet engine. The study explored the effects of fuel injection angle, injection method, and axial position of the injection ring on the mixing characteristics of fuel and supersonic inflow. The results showed that increasing the fuel injection angle improves the uniformity of fuel mixing, reduces total pressure loss, and leads to more intense collisions between transverse jets. When the fuel is injected in a non-colliding manner, the mixing uniformity increases, and the collision area between the supersonic inflow and the fuel enlarges. Moving the injection ring upstream enhances the uniformity of fuel mixing, and the mixing distance between the transverse jet and the supersonic inflow becomes longer. Analysis of the flow field structure revealed that the supersonic inflow and ethylene transverse jets undergo cylindrical wake flow within the expansion section, forming a fan-shaped continuous expansion region extending downstream. Upon entering the combustion chamber, the primary mixing flow forms a series of fragmented small vortex structures, thereby promoting the mixing effect of ethylene and air.

Rotating seal ring are used in aero engines to block low-temperature bleed air in the bearing cavity from high-temperature bleed air in the turbine. In order to obtain the influence of rotation speed and inlet and outlet pressure ratio on the sealing effect, an experimental system with high pressure ratio and high speed was built. Through the measurement of multi-section parameters, combined with the comparative experiments of sealing clearance and contact surface lubrication, the evolution law of leakage characteristics was analyzed. The results show that the leakage is reduced by 36.4% when the rotation speed is increased from 0 r/min to 10000 r/min (pressure ratio 2.5), and the leakage increases by 91.8% when the pressure ratio is increased from 1.1 to 2.5 (10000 r/min). The assembly clearance affects less than 5% within the 0.5 mm tolerance, and the addition of lubricating coatings reduces leakage by 15%—60%. Numerical simulations show that the average contact stress increases by up to 1.24 MPa under the experimental conditions, which improves the sealing effect, and the increase of friction coefficient reduces the average contact stress by more than 1.35 MPa, which is not conducive to sealing. The quantitative relationship between rotational speed and seal performance and the efficiency mechanism of lubrication are revealed, which provides a basis for the design of rotary seals for aero engines.

A single-slotted Fowler flap high-lift device for a hybrid wing-body (HWB) configured amphibious aircraft was designed, incorporating blowing/suction zones on the upper surface of the main wing. It systematically validated the lift-enhancement effectiveness of the co-flow jet (CFJ) technology on this novel aerodynamic configuration and elucidated the mechanisms behind its aerodynamic performance improvements. Numerical simulation results demonstrated that: when the flap deflection angle reached 40°, the maximum lift coefficient of the high-lift configuration attained 2.7; after implementing CFJ, the flow velocity over the upper wing surface increased, accompanied by a significant rise in negative pressure. The initial flow separation on the wing surface was effectively controlled, with the maximum lift coefficient increasing to 3.2, stall angle of attack extended by 4°, drag coefficient reduced by over 20%, and available range of pitching moment expanded. Additionally, CFJ enhanced the surface flow velocity on the flaps and suppressed the potential local separation under high angle-of-attack conditions, increasing the leading-edge negative pressure peak by up to 30% and significantly improving the flap lift-enhancement efficiency.

The complex flow field around aircraft engines during carrier landing and boltering was analyzed by establishing a full-scale numerical model of both carrier and aircraft. The overset grid and moving pressure measurement techniques combined with delayed detached-eddy simulation (DDES) were developed to resolve the complex vortex systems generated by the advancing carrier hull. The effects of wind speed, wind direction angle, and angle of attack on the total pressure distortion at the engine inlet were analyzed, and the dynamic evolution characteristics of the total pressure distortion were revealed. Key findings indicated that under headwind conditions, counter-rotating vortex pairs formed on both sides of the carrier hull, while crosswinds caused significantly asymmetric flow separation, with turbulence intensity enhanced on the leeward side. With the increase of wind angle, the total pressure distortion index exhibited decreased mean values but elevated standard deviations. Specifically at 90° yaw angle (relative to 0°), the mean value dropped from 0.0649 to 0.0498 (23.3% reduction), while the standard deviation surged from 1.04×10⁻³ to 4.97×10⁻³ (377.9% increase). Particularly during touchdown, the distortion transients became especially prominent due to ground effects and pitch adjustment.

To address the significant contradiction between the optimal blade shapes of tiltrotor aircraft in hover and cruise modes, a mission scenario-driven aerodynamic optimization design method for tiltrotor blades was developed. This approach comprehensively considered the total energy consumption across different mission scenarios to ensure overall improvements in flight performance and efficiency. By incorporating typical mission phases including vertical takeoff, cruise, and hover and their respective durations, the method transformed the multi-point multi-objective optimization problem into a single-objective optimization problem with the total energy consumption of the entire mission scenario as the objective. The shape differences of the optimized blades under cruise and hover-focused mission scenarios were analyzed. Results indicated that the developed method achieved high aerodynamic performance of tiltrotor blades across multiple operating conditions, with optimization time reduced by 69.67% compared with multi-objective optimization.

Aero-engine usually has a large number of freely-bent and interlaced pipes. To avoid faults caused by friction or resonance between pipes during operation, the pipe spacing must be strictly controlled during assembly. However, the existing detection methods rely on manual operation and have problems such as high labor intensity, low efficiency, and unstable measurement accuracy. Using a laser scanner to collect point cloud data of aero-engine pipes and calculate the pipe spacing can significantly improve the degree of automation and efficiency of pipe spacing measurement. However, the point cloud data of aero-engine pipes is huge and difficult to segment, which brings challenges to the subsequent pipe spacing measurement. Therefore, a hierarchical multi-feature-based pipe segmentation algorithm is proposed by taking curvature, normal vector angle, and normal vector cross product as the segmentation judgment conditions. The calculation results show that this algorithm can accurately segment the pipe point cloud through a layer-by-layer refinement method, improving the over-segmentation and under-segmentation problems of existing algorithms, with an accuracy rate of 94.65%, which lays a foundation for subsequent pipe spacing calculation.

The intercooled recuperated turbofan engine has more health parameters than the ordinary three-shaft turbofan engine, which leads to a larger calculation amount of fault diagnosis. To enhance the speed of gas path fault diagnosis of intercooled recuperated turbofan engine under multi-component degradation, a fast gas path fault diagnosis method for intercooled recuperated turbofan engine was developed. Based on the traditional nested iterative diagnosis architecture, a non-nested iterative architecture was designed, which considered the degradation of 6 rotating components and 2 heat exchanger components. By dividing the components, the engine component matching and fault diagnosis could be completed in the same Newton-Raphson iterative algorithm. The sensor selection optimization was carried out to improve the ability of measuring parameters to identify the degradation of different components. The results of simulation test on all conditions of engine life cycle showed that the maximum diagnostic error of this method was 0.0036%, the average error was 0.0008%, and the average calculation time was less than 0.063 seconds. Compared with the traditional diagnostic architecture, this non-nested architecture can significantly reduce the number of sub-component calls and fault diagnosis time while ensuring the diagnostic accuracy. It provides the method and theoretical support for overcoming the conflict between high-precision fault diagnosis and rapid diagnosis of multi-component faults.

To study the design approaches for enhancing lubrication effect of aero-engine lubrication system and achieving precise, on-demand lubrication at multiple locations, an axial under-race lubrication structure with oil dam was presented. Numerical study was conducted on the internal oil-gas two-phase flow process and oil flow-rate distribution characteristics. Based on verifying the accuracy of the numerical methods, the influences of oil flow-rate, rotational speed, oil temperature, and the circumferential angle of oil dam on the oil flow-rate and distribution ratio of the lubricated spline and bearing were analyzed, and the mechanism of the factors was also revealed. The results indicated that the oil formed a continuous film with marked phase-separated flow characteristics under rotational centrifugation, and was delivered to the bearing and spline after being segregated by the oil dam. The relative deviation between the lubrication flow-rate of spline and bearing and the circumferential angle ratio of oil dam under different conditions remained within 6%, showing good consistency. The larger circumferential angle ratio of oil dam indicated the more oil trapped in the oil collection chamber during the initial state, resulting in reduced outlet flow-rate. The outlet flow-rate varied approximately linearly with the circumferential angle ratio of oil dam during the steady state, with a maximum deviation of 4.72%. Relevant studies validated the effectiveness of controlling flow-rate distribution by adjusting the circumferential angle of oil dam under varying conditions, achieving precise lubrication requirements for both bearing and spline as needed.

To address the failure caused by multi-point contact between steel balls and raceways in three-point contact ball bearings, this study took a specific model of three-point contact ball bearing as the research object, and a dynamic model was adopted to analyze the production mechanisms of multi-point contact under variations in structural and operating parameters. And then change rule of the number of contact points between steel balls and the non-main loading inner-half ring (NMLIHR) was obtained. The results reveal that under the basic working conditions, multi-point contact occurs when the inner ring groove curvature coefficient drops below 0.5058 or the outer ring groove curvature coefficient rises above 0.5134, or the gasket thickness exceeds 0.60 mm. Under non-basic working conditions, the multi-point contact phenomenon can disappear when the load ratio is more than 0.35. Further, higher load ratios are required to prevent multi-point contact under conditions of elevated rotational speeds and large radial loads. In addition, the quantitative maps of critical conditions under multi-parameter influences were developed, which could determine the critical threshold values of inner and outer groove curvature coefficients corresponding to different gasket thicknesses, and the critical threshold values of load ratio and rotational speed corresponding to different radial loads. The research findings provide the guidance for avoiding multi-point contact and wear failure in three-point contact ball bearings from both design and service perspectives.