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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.

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.

A comprehensive series of tests have been made on an experimental single‐stage turbine to determine the cooling characteristics and the overall stage performance of a set of air‐cooled turbine blades. These blades, which arc described fully in Part I of this paper had, internally, a multiplicity of passages of small diameter along which cool air was passed through the whole length of the blade. Analysis of the test data indicated that, when a quantity of cooling air amounting to 2 per cent, by weight, of the total gas‐flow through the turbine is fed to the row of rotor blades, an increase in gas temperature of about 270 dcg. C. (518 deg. F.) should be permissible above the maximum allowable value for a row of uncoolcd blades made from the same material. The degree of cooling achieved throughout each blade was far from uniform and large thermal stresses must result. It appears, however, that the consequences of this are not highly detrimental to the performance of the present type of blading, it being demonstrated that the main effect of the induced thermal stress isapparently to transfer the major tensile stresses to the cooler (and hence stronger) regions of the blade. The results obtained from the present investigations do not represent a limit to the potentialities of internal air‐cooling, but form merely a first exploratory step. At the same time the practical feasibility of air cooling is made apparent, and advances up to the present arc undoubtedly encouraging.

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.

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.

The authors aim to present a procedure for the parallel, steady and unsteady conjugate, Navier–Stokes/heat-conduction rotor-stator interaction analysis of multi-blade-row, film-cooled, turbine airfoil sections. A new grid generation procedure for multiple blade-row configurations, including walls, thermal barrier coatings, plenums, and cooling tubes, is discussed.

Steady, multi-blade-row interaction effects on the flow and wall thermal fields are predicted using a Reynolds’s-averaged Navier–Stokes (RANS) simulation in conjunction with an inter-blade-row mixing plane. Unsteady, aero-thermal interaction solutions are determined using time-accurate sliding grids between the stator and rotor with an unsteady RANS model. Non-reflecting boundary condition treatments are utilized in both steady and unsteady approaches at all inlet, exit and inter-blade-row boundaries. Parallelization techniques are also discussed.

The procedures developed in this research are compared against experimental data from the Air Force Research Laboratory’s turbine research facility.

The software presented in this paper is useful as both the design and analysis tool for fluid system and turbomachinery engineers.

This research presents a novel approach for the simultaneous solution of fluid flow and heat transfer in film-cooled rotating turbine sections. The software developed in this research is validated against experimental results for 2D flow, and the methods discussed are extendable to 3D.

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.

Maintaining the turbine blade’s temperature within the safety limit is challenging in high-pressure turbines. This paper aims to numerically present the conjugate heat transfer analysis of a novel approach to mini-channel embedded film-cooled flat plate.

Numerical simulations were performed at a steady state using SST k – ω turbulence model. Impingement and film cooling are classical approaches generally adopted for turbine blade analysis. The existing film cooling techniques were compared with the proposed design, where a mini-channel was constructed inside the solid plate. The impact of the blowing ratio (M), Biot number (Bi) and temperature ratio (TR) on overall cooling performance was also studied.

Overall cooling effectiveness was always shown to be higher for mini-channel embedded film-cooled plates. The effectiveness increases with increasing the blowing ratio from M = 0.3 to 0.7, then decreases with increasing blowing ratio (M = 1 and 1.4) due to lift-off conditions. The mini-channel embedded plate resulted in an approximately 21% increase in area-weighted average overall effectiveness at a blowing ratio of 0.7 and Bi = 1.605. The lower uniform temperature was also found for all blowing ratios at a low Biot number, where conduction heat transfer significantly impacts total cooling effectiveness.

To the best of the authors’ knowledge, this study presents a novel approach to improve the cooling performances of a film-cooled flat plate with better cooling uniformity by using embedded mini-channels. Despite the widespread application of microchannels and mini-channels in thermal and fluid flow analysis, the application of mini-channels for blade cooling is not explored in detail.

THE effect of higher turbine inlet temperature on the performance of an aircraft gas turbine engine can be quite dramatic. It can be used to increase the exhaust velocity of a dry turbojet to providea higher specific thrust; to increase the bypass ratio of a turbofan engine to improve its propulsive efficiency; to optimize the thermodynamic cycle at a higher pressure ratio to improve its specific fuel consumption; to reduce the amount of afterburner fuel flow in an augmented turbojet to improve its specific fuel consumption, or to increase the work output of a turboshaft engine. If the thrust or power of the engine is held constant, a size, cost and/or weight reduction can result. If the size of the engine is held constant growth capability can be provided.