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THE demand of the aircraft designer has been, and presumably always will be, for his engines to operate better in three basic respects. To give more thrust, to have less weight, and to require less fuel.

ALTHOUGH numerous papers and lectures presented to the Royal Aeronautical Society have mentioned the upward trend in turbine inlet gas temperatures, there has been no review of the status of blade cooling technology since 1956, when Ainley's classic paper ‘The High Temperature Turbo‐jet’ was published. Accordingly it is the aim of this paper to present such a review. Before doing so it is worth while to compare the engine situation today with what it was in 1956. At that time, much of the available experience in the U.K. on air cooled turbines was based on experimental units, designed for the express purpose of measuring blade temperatures under controlled conditions of cooling airflow and high gas temperature. These research turbines had also yielded some useful preliminary data on the aerodynamic effects of cooling air discharge and on thermal stress and creep problems. Some engine experience had been attained, mainly (in the U.K.) with engines such as the Avon, Conway and Tyne. Whereas many of the research turbine and cascade blades had fairly complex patterns of relatively small cooling passages, the blades which had been submitted to engine running usually had a few comparatively large passages. Rotating blades were made exclusively by forging and extrusion processes from wrought nickel‐base alloys. Some nozzle guide vanes were cast.

Fundamentally, it is advantageous to operate an aeroengine's thermodynamic cycle at as high a turbine entry temperature as practical for the current metallurgical limits of the turbine blades in order to achieve peak cycle efficiency and thus lower specific fuel consumption. However, achieving the highest possible turbine entry temperature requires accurate knowledge of the turbine blade temperatures for control purposes to prolong component life as frequent excursions beyond the design limits of the blades can severely reduce their service life. The optical pyrometry technique represents the best method for providing this crucial temperature data needed for blade condition‐based monitoring. This paper presents the general operating principles, system aspects and design considerations for the application of the optical pyrometer instrument for inflight service use on gas turbine aeroengines.

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.

After briefly outlining the main features of the variable‐pitch propeller, this paper proceeds to describe the development of the piston‐engined hydraulically operated propeller as a brake, both in the air and on the ground. Examples are given of the magnitude of the braking effort of a propeller when windmilling under controlled conditions and when in reverse pitch under power. The advent of the gas turbine, originally intended as a means of jet propulsion, opened up a new field of application for the variable‐pitch propeller and this application with its attendant problems and their solution is discussed. Three types of gas‐turbine power plant, together with the appropriate propeller arrangements are reviewed. These arc: (I) the direct‐connected turbine; (2) the compound‐compressor turbine; and (3) the free‐propeller turbine.

The purpose of this paper is to present coordinate measuring system possibilities in the meaning of the geometric accuracy assessment of hot zone elements in aircraft engines. The aim of the paper is to prove that this method, which uses blue light and is most sufficient and cost-saving method, can to be used in the production line for serial manufacturing of elements, for which a high level of accuracy is required.

The analysis of the geometric accuracy of the blades was performed using non-contact optical coordinate scanner ATOS Triple Scan II Blue Light, manufactured by GOM Company, at the Department of Mechanical Engineering, Rzeszów University of Technology. Geometric analysis was conducted for blades manufactured from different waxes (A7Fr/60 and RealWax VisiJet CPX200), thus comparing injection technique and rapid prototyping (RP) method, and for casting made of Inconel 713C nickel-based superalloy.

The analysis of the criteria for the method of blades’ measuring selection showed that the chosen system successfully met all criteria for the verification of blades’ geometry at the selected stages of the process. ATOS II optical scanner with blue light technology allows measurement almost regardless of daylight or artificial (white) light. This allows the application of the measurement system in the production cycle, thus eliminating the need to create special conditions for measurements.

Requirements related to the accuracy of measured values, diversity and allowable measurement time are linked with the methods of production. Modern manufacturing methods based on computer-aided design systems/manufacturing/engineering systems require a non-contact optical measurement method based on the computer-aided-based coordinate measuring technique. In case of the non-contact optical scanning method based on the ATOS GOM measuring system, time and measurement costs depend on the methodology of measurement and the possibility of its automation. This is why the presented paper has a practical impact on possibilities for the automation of geometric accuracy measurements of obtained elements in the series production line.

The use of ATOS Triple Scan II Blue Light by GOM Company allows the reduction of cost and time of production because of the possibility of the introduction of this system in an automated production line. Additionally, the measurement of hot section blades of aircraft engines by using the blue light method is much more accurate and has implication as it impacts safety of further used manufactured elements.

This paper presents the possibility of using the ATOS Triple Scan II Blue Light measuring system for geometric accuracy measurements in case of hot section blades of aircraft engines. This research is original because it describes three model geometric accuracy measurements, wax model obtained using the injection technique, wax model obtained using the I RP process and casting made of Inconel 713C nickel-based superalloy.

The purpose of this paper is to introduce the multidisciplinary design optimization method using approximation model for the aircraft engine fan blade based on the airworthiness compliance such as stress, vibration, and bird impact.

Firstly, the airworthiness analysis of the typical fan blade was carried out based on the numerical simulation. Secondly, the design of experiment (DOE) was utilized to construct the approximation model of the fan blade. Finally, the airworthiness optimization of fan blade was carried out based on Kriging approximation model.

The numerical simulation result shows that the analysis method can show the airworthiness compliance in the design stage. And the optimization result shows that structure, bird impact and vibration characteristics improve obviously, satisfying the constraints conditions of optimization.

The multidisciplinary design optimization method of fan blade based on the airworthiness and approximation model is presented and achieved.

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.

TWENTY‐FIVE years ago, the gas turbine was successfully applied to aircraft propulsion for the first time and it was not long after this event that powerful engines of reasonable efficiency appeared. In strong contrast with the early success of large engines, the evolution of equally efficient small engines has proved both difficult and protracted. That few of the major manufacturers have persisted with their development may be surprising but is also indicative of the technical problems involved. Thus it reflects great credit upon Turboméca that they have produced the first turbines to offer a serious challenge to piston engines for small aircraft installations. In this connection, the Astazou occupies a significant position as the first small propeller turbine to achieve any real measure of success, and it has been developed progressively to maintain its early lead over competitors. Consequently this engine was the obvious choice to power the Turbo‐Skyvan.

The purpose of this paper is to present the advantages of computer-aided design/rapid prototyping (CAD/RP) usage in designing and manufacturing of the core models used for precise casting with direct and single solidification of aircraft engine turbine blade cores.

The process of modelling three-dimensional CAD geometry of research blade in relation to the model of the core was presented with different wax types used in the RP technique.

The geometry of the blade model has been designed in a way which allows making a silicon mould on the basis of a base prototype in the process of rapid tooling (RP/RT). Filing by different wax types was investigated in mean of the impact on filling accuracy of the mould cavity.

The resulting models were used to make ceramic moulds and carry further work on the development of casting technology in the process of directional solidification and single crystal solidification of core blades of aircraft engines.