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This paper aims to introduce an aircraft engine inspection robot (AEIR) which can go in the internal of the aircraft engine without collision and detect damage for engine blades.

To obtain the position and pose information of the blades inside the engine, a novel tactile sensor based on electrical impedance tomography (EIT) is developed, which could provide location and direction information when it contacts with an unknown object. In addition, to navigate the continuum robot, a control method is proposed to control the continuum robot, which can control the continuum robot to move along the pre-planned path and reduce the deviation from the planned path.

Experiment results show that the average error of contact location measurement of the tactile sensor is 0.8 mm. The average error relative to the size (diameter of 18 mm) of the sensor is 4.4%. The continuum robot can successfully reach the target position through a gap of 30 mm and realize the spatial positioning of blades. The validity of the AEIR for engine internal blade detection is verified.

The aero-engine inspection robot developed in this paper can replace human to detect engine blades and complete different detection tasks with different kinds of sensors.

THE use of investment castings for aerospace applications dates back to the Second World War when components made by this process were utilized in the construction of superchargers for piston aero engines and nozzle components for turbojet engines. The technique has developed enormously since that time so that today investment castings are used widely for aircraft and missile applications—particularly on the power plant side for blading.

The purpose of this paper is to evaluate the effect of film cooling holes on the vibration characteristics of a turbine blade, and provide the design basis for the blade, which may reduce computing costs.

Modal analysis of the blades with and without film cooling holes is performed to evaluate the effect of film cooling holes on its natural frequency. Harmonic analysis of the blade is performed to calculate the stress concentration factors of film cooling holes for different modes.

The frequency differences between two blades with and without film cooling holes are insignificant, while the differences of the vibration stress cannot be neglected. For the first three modes of the blades, the stress concentration factor is sensitive to the hole’s shape and position on the blade. With the help of the stress concentration factor defined in this work, the concentration of stresses induced by different film cooling holes can be accurately described when evaluating HCF life of the turbine blade.

The effect of film cooling holes on a turbine blade's natural frequencies was confirmed to be insignificant and the stress concentration factors around the holes are calculated. Therefore, the simplified model of the blade without film cooling holes can be used to evaluate the natural frequencies and vibration stress, which saves a lot of time and cost.

This paper aims to provide SIF calculation method for engineering application.

In this paper, the stress intensity factors (SIFs) calculation method is applied to the anisotropic Ni-based single crystal film cooling holes (FCHs) structure.

Based on contour integral, the anisotropic SIFs analysis finite element method (FEM) in Ni-based single crystal is proposed. The applicability and mesh independence of the method is assessed by comparing the calculated SIFs using mode of plate with an edge crack. Anisotropic SIFs can be calculated with excellent accuracy using the finite element contour integral approach. Then, the effect of crystal orientation and FCHs interference on the anisotropic SIFs is clarified. The SIFs of FCH edge crack in the [011] orientated Ni-based single crystal increases faster than the other two orientations. And the SIF of horizontal interference FCHs edge crack is also larger than that of the inclined interference one.

The SIFs of the FCH edge crack in the turbine air-cooled blade are innovatively computed using the sub-model method. Both the Mode I and II SIFs of FCHs edge crack in blade increase with crack growing.

The purpose of this paper is to develop a novel process of integral ceramic molds for investment casting of hollow turbine blades.

At first, a resin pattern of a hollow turbine blade prototype is fabricated by stereolithography (SL). And then aqueous gelcasting process is utilized to fill the resin pattern with ceramic slurry of low viscosity and low shrinkage, through in situ polymerization of the slurry a ceramic mold is formed. At last, the ceramic mold for investment casting of hollow turbine blade is obtained by vacuum drying, pyrolyzing and sintering.

An integral ceramic mold is successfully fabricated by combining SL and gelcasting process, cores and shell are connected with each other and thus high relative position accuracy is guaranteed. Properties of integral ceramic mold at room temperature and high temperature satisfy the requirements of directional casting for complex‐shaped thin‐walled blades.

Because the integral ceramic mold is a close body, it is very difficult to directly measure its inner dimensions and the relative position accuracy of cores and shell, and the further research is needed.

This method enhanced the versatility of using SL prototype in the fabrication of integral ceramic mold for investment castings. Although this paper took a hollow turbine blade as an example, this method is also capable of fabricating integral ceramic molds for other complex investment castings.

A new prototyping technology named QuickCast, developed by USA‐based 3D Systems, has been introduced to the European market. Since its USA introduction 12 months ago, the method has been adopted by investment‐casting foundries and end‐user industries such as aerospace.

The application of thermal barrier coatings (TBCs) allows aero-engine blades to operate at higher temperatures with higher efficiency. The preparation of the TBCs increases the surface roughness of the blade, which impacts the thermal cycle life and thermal insulation performance of the coating. To reduce the surface roughness of blades, particularly the blades with small size and complex curvature, this paper aims to propose a method for industrial robot polishing trajectory planning based on on-site measuring point cloud.

The authors propose an integrated robotic polishing trajectory planning method using point cloud processing technical. At first, the acquired point cloud is preprocessed, which includes filtering and plane segmentation algorithm, to extract the blade body point cloud. Then, the point cloud slicing algorithm and the intersection method are used to create a preliminary contact point set. Finally, the Douglas–Peucker algorithm and pose frame estimation are applied to extract the tool-tip positions and optimize the tool contact posture, respectively. The resultant trajectory is evaluated by simulation and experiment implementation.

The target points of trajectory are not evenly distributed on the blade surface but rather fluctuate with surface curvature. The simulated linear and orientation speeds of the robot end could be relatively steady over 98% of the total time within 20% reduction of the rest time. After polishing experiments, the coating roughness on the blade surface is reduced dramatically from Ra 7–8 µm to below Ra 1.0 µm. The removal of the TBCs is less than 100 mg, which is significantly less than the weight of the prepared coatings. The blade surface becomes smoothed to a mirror-like state.

The research on robotic polishing of aero-engine turbine blade TBCs is worthwhile. The real-time trajectory planning based on measuring point cloud can address the problem that there is no standard computer-aided drawing model and the geometry and size of the workpiece to be processed differ. The extraction and optimization of tool contact points based on point cloud features can enhance the smoothness of the robot movement, stability of the polishing speed and performance of the blade surface after polishing.

This paper aims to improve the wear resistance of titanium alloy using a high-hardness boride layer, which was fabricated on Ti6Al4V by a high-temperature boronizing process.

The boride layers on Ti6Al4V were obtained at 1000°C for 5–15 h. Scanning electron microscopy, energy dispersive analysis and X-ray diffractometer were used to characterize the properties of the boride layer. The tribological performance of the boride layer at room and elevated temperatures was investigated.

The X-ray diffraction analysis showed that the boride layers were a dual-phase structure of TiB and TiB₂. When the boronizing time increased from 5 h to 15 h, the microhardness increased from 1192 HV_0.5 to 1619.8 HV_0.5. At 25°C and elevated temperatures, the friction coefficients of the boride layers were higher than that of Ti6Al4V. The wear track areas of T-5 at 200°C and 400°C were 2.5 × 10^–3 and 1.1 × 10^–3 mm², respectively, which were 6.1% and 2.6% of that of Ti6Al4V, indicating boride layer exhibited a significant wear resistance. The wear mechanisms of the boride layer transformed from slight peeling to oxidative wear and abrasive wear as the temperature was raised.

The findings provide an effective strategy for improving the wear resistance of Ti6Al4V and have important implications for the application of titanium alloy in a high-temperature field.

In the search for on or off‐line methods for condition monitoring of wear sensitive components, Thin Layer Activation (TLA) has much to offer. It is possible to measure very small surface losses directly, and by the use of double layer, or direct, or indirect sentinel layer modifications, wear of material to any depth can be measured. The technique can be applied to most materials and extended to virtually all materials and components by the use of treated plugs or inserts. Material loss can often be measured under operating conditions without dismantling equipment. The technique is reliable and of known useful life, ie there can be no electrical or mechanical failures of the implanted layer. The total radioactivity is very low and no modification of material surface properties is likely. The use of TLA in condition monitoring will speed up the identification of incipient faults.

The purpose of this paper is to develop a method to predict the multi-failure risk of aero engine in service and to evaluate the effectiveness of different corrective actions.

The classification of failure risk level, the determination of hazard ratio and the calculation of risk factor and the risk per flight have been proposed. The multi-failure risk assessment process of aero engine has been established to predict the occurrence of failure event and assess the failure risk level. According to the history aero engine failure data, the multi-failure risk, i.e., overheat, blade wounding, pump failure, blade crack, pipe crack and combustor crack, has been predicted considering with and without corrective action. Two corrective actions, i.e., reduce the maintenance interval and redesign the failure components, were adopted to analyze the decreasing of risk level.

The multi-failure risk of aero engine with or without corrective action can be determined using the present method. The risk level of combustor crack decreases from high-risk level of 1.18×1e−9 without corrective action to acceptable risk level of 0.954×1e−9 by decreasing the maintenance interval from 1,000 to 800 h, or to 0.912×1e−9 using the redesign combustor.

It should be noted that probability of detection during maintenance actions has not been considered in the present analysis, which would affect the failure risk level of aero engine in service.

The method in the present analysis can be adapted to other types of failure modes which may cause significant safety or environment hazards, and used to determine the maintenance interval or choose appropriate corrective action to reduce the multi-failure risk level of aero engine.

The maintenance interval or appropriate corrective action can be determined using the present method to reduce the multi-failure risk level of aero engine.