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The purpose of this paper is to present results of laboratory testing work on causes of a service failure/damage to an aircraft turbojet's gas‐turbine blade made of the EI 867‐WD alloy.

The tests comprised comparing the microstructure of a service‐damaged blade with microstructures of specimens drawn from a similar all‐new blade, both subjected to temperatures of different values for different annealing times.

Findings based on the comparison of experimentally gained results of microstructure examination of both the gas‐turbine blades were: the change in the microstructure of a damaged blade results from the growth and cuboidal‐to‐lamellar change of shape of the reinforcing phase γ′ (Ni₃Al); and the size and shape of this phase are comparable to those of the phase γ′ of a new blade subjected to annealing at temperature exceeding 1,223 K for 1 h. The results gained allowed for drawing the conclusion that the damaged turbine blade was operated in the exhaust‐gas temperature exceeding the maximum permissible value of 1,013 K for approximately 1 h in the course of an air mission.

The comparison‐oriented experimental testing work was carried out on a new blade manufactured in the way and from material identical to those of the damaged blade. The applied methodology enables us to gain qualitative results of investigating into the causes of a failure/damage to a gas‐turbine blade.

The presented methodology of identifying (origin‐finding of) a service‐induced damage to a gas‐turbine blade proves helpful in the case of an engine failure, when information on the operating conditions thereof is insufficient.

The paper is an original work by the authors. To the best of their knowledge, the issue has not been found in the literature, approached in this particular way. It has been based on research work on air accidents due to the service‐induced failures/damages to gas‐turbine blades in aircraft turbojet engines.

The purpose of this study a fast procedure for the structural analysis of gas turbine blades in aircraft engines. In this connection, investigations on the behavior of gas turbine blades concentrate on the analysis and evaluation of starting dynamics and fatigue strength. Besides, the influence of structural mistuning on the vibration characteristics of the single blade is analyzed and discussed.

A basic computation cycle is generated from a flight profile to describe the operating history of the gas turbine blade properly. Within an approximation approach for high-frequency vibrations, maximum vibration amplitudes are computed by superposition of stationary frequency responses by means of weighting functions. In addition, a two-way coupling approach determines the influence of structural mistuning on the vibration of a single blade. Fatigue strength of gas turbine blades is analyzed with a semi-analytical approach. The progressive damage analysis is based on MINER’s damage accumulation assuming a quasi-stable behavior of the structure.

The application to a gas turbine blade shows the computational capabilities of the approach presented. Structural characteristics are obtained by robust and stable computations using a detailed finite element model considering different load conditions. A high quality of results is realized while reducing the numerical costs significantly.

The method used for analyzing the starting dynamics is based on the assumption of a quasi-static state. For structures with a sufficiently high stiffness, such as the gas turbine blades in the present work, this procedure is justified. The fatigue damage approach relies on the existence of a quasi-stable cyclic stress condition, which in general occurs for isotropic materials, as is the case for gas turbine blades.

Owing to the use of efficient analysis methods, a fast evaluation of the gas turbine blade within a stochastic analysis is feasible.

The fast numerical methods and the use of the full finite element model enable performing a structural analysis of any blade structure with a high quality of results.

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.

Conventional investment casting of turbine blades is a time consuming and expensive process due to the complications in wax injection steps and the complex shape of airfoil surfaces. By using rapid investment casting, a substantial improvement in the gas turbine blade manufacturing process can be expected. However, this process needs to be able to compete with conventional investment casting from a dimensional accuracy view of point. The purpose of this paper is to investigate the manufacture of gas turbine blades via two indirect rapid tooling (RT) technologies, namely epoxy (EP) resin tooling and silicon rubber molding.

The second stage blade of a Ruston TA 1750 gas turbine (rated at 1.3 MW) was digitized by a coordinate measuring machine. The aluminum‐filled EP resin and silicon rubber molds were fabricated using StereoLithography master models. Several wax patterns were made by injection in the EP resin and silicone rubber molds. These wax patterns were utilized for ceramic shell fabrication and blade casting.

Dimensional inspection of cast blades showed that silicone rubber molding was not a suitable approach for production of blade wax patterns. The maximum deviation for the final cast blade made using the silicone rubber mold was +0.402 mm. The maximum deviation for the final cast blade made using the EP resin mold was lower at −0.282 mm. This showed that EP resin tooling could enable new cost‐effective solutions for small batch production of gas turbine blades.

The research results presented will give efficient industrial approach and scientific insight of the gas turbine blade manufacturing by use of rapid technologies.

There are some general research works related to utilization of rapid technologies for manufacturing of gas turbine blade. However, this paper presents a unique procedure of integrated reverse engineering and RT technologies for rapid investment casting of gas turbine blade through presenting comprehensive comparison between two techniques from dimensional accuracy view of point.

The purpose of this paper is to find the time dependent thermal creep stress relaxation of a turbine blade and to investigate the effect thermal radiation of the adjacent turbine blades on the temperature distribution of turbine blade and creep relaxation.

For this analysis, the creep flow behavior of Moly Ascoloy in operational temperature of gas turbine in full scale geometry is studied for various thermal radiation properties. The commercial software is used to pursue a coupled fields analysis for turbine blades in view of the structural force, materials kinematic hardening, and steady-state temperature field.

During steady-state operation, the thermal stress was found to be decreasing, whereas by considering the thermal radiation this rate was noticed to increase slightly. Also by increase of the distance between stator blades the thermal radiation effect is diminished. Finally, by decrease of the blade distance the failure probability and creep plastic deformation decrease.

This paper describes the effect of thermal radiation in thermal-structural analysis of the gas turbine stator blade made of the super-alloy M-152.

Blade failures in gas turbine engines often lead to loss of all downstream stages and can have a dramatic effect on the availability of the turbine engines. There are many components in a gas turbine engine, but its performance is highly profound to only a few. The majority of these are hotter end rotating components.

Three-dimensional finite element thermal and stress analyses of the blade were carried out for the steady-state full-load operation.

In the previous works the thermal radiation effects on creep behavior of the turbine blade have not performed.

Blade tip clearance has always been a concern for the gas turbine design and control. The numerical analysis of tip clearance is based on the turbine components displacement. The purpose of this paper is to investigate the thermal and mechanical effects on a real cooling blade rather than the simplified model.

The coupled fluid-solid method is used. The thermal analysis involves solid and fluid domains. The distributions of blade temperature, stress and displacement have been calculated numerically under real turbine operating conditions.

Temperature contour can provide a reference for stress analysis. The results show that temperature gradient is the main source of solid stress and radial displacement. Compared with thermal or mechanical effect, there is a great change of stress magnitude for the thermomechanical effect. Large stress gradients are found between the leading and trailing edge of turbine cooling blade. Also, the blade radial displacement is mainly attributed to the thermal load rather than the centrifugal force. The analysis of the practical three-dimensional model has achieved the more precise results.

It is significant for clearance design and life prediction.

The purpose of this paper is to found a life model for the single crystal (SC) turbine blade based on the rate‐dependent crystallographic plasticity theory.

This life model has taken into consideration the creep and fatigue damages by the linear accumulation theory. A SC blade was taken from an aero‐engine, which had worked for 1,000 hours, as the illustration to validate the life model.

The crystallographic life model has a good prediction to the life and damage of the SC turbine blade. In the mean time, the micro damage study of the miniature specimens showed that creep damage has more serious influence on the material performance in the blade body but it is fatigue damage in the blade rabbet.

The life model can reflect the crystalline slip and deformation and crystallographic orientation of nickel‐based SC superalloys.

UNDER steady load conditions, materials in gas turbines are subject to various forms of static and alternating stresses. Changes in the operating conditions such as starting, load variations and shut down cause additional thermal stresses which limit the permissible rate of these changes in service. In stationary plants these effects can be minimized by adjustment of the starting and shut down procedure or by protection of the sensitive parts with a cooling flow. In gas turbines for propulsion purposes load changes are governed by external conditions, are more frequent and take place at a higher rate. The consequent thermal stresses are then referred to as thermal shocks. Various methods for testing the resistance of materials to thermal shocks have already been suggested and applied. However, they differ very widely, and no quantitative, or even comparable figures are available as yet.

Turbine blades in gas turbine engines operate at elevated temperatures and in highly oxidising atmospheres that can be contaminated with fuel residues and sea water salts. These components, which are expensive to produce, are subjected to high stresses during operation but must be totally reliable during their design life. An economic way to maintain blade properties is to coat the base metal superalloy with a protective layer capable of resisting both high temperature oxidation and hot corrosion. Conventional aluminide coatings are widely used for this purpose but platinum aluminides offer improved corrosion resistance. A collaborative exercise involving Rolls‐Royce and Johnson Matthey has now resulted in the development of a platinum aluminide diffusion coating that offers some advantages over the commercial systems.

Numerical simulations were carried out for two cooling schemes, a circular hole and a louver cooling scheme, at the leading edge of a rotor blade in a complete turbine stage.

Two holes were positioned at the leading edge of a rotating blade, one on the pressure side and the other on the suction side. The methodology was validated with a circular hole case. Numerical results of cooling effectiveness for three blowing ratios at three rotational speeds were successfully obtained. Both blowing ratio and rotating speed of the rotor affect the cooling effectiveness level.

It was shown that for the circular hole, the blowing ratio is the dominant factor at low blowing ratios and the rotational speed is the dominant factor at high blow ratios when jet is prone to lift off in determining the cooling effectiveness level. For the louver scheme, a higher rotational speed leads to a higher level of cooling effectiveness since jet liftoff is avoided.

There are only a few studies of film cooling on a rotational turbine blade and very few studies of film cooling at the leading edge of a rotating turbine blade in the open literature. The present work presents a challenging CFD case. The analysis of film cooling at the leading edge of an airfoil was presented, which sheds light on the physics of film cooling and should prove helpful to the cooling designs of turbine blades.