Gas turbine engine

TURBINE BLADE COOLING

ABSTRACT

INTRODUCTION

It is well known from the thermodynamic analysis through literature survey that the performance of a gas turbine engine is strongly influenced by the temperature at the inlet to the turbine. Figure 1 illustrates the relation between the specific power output and turbine rotor inlet temperature. There is thus a growing tendency to use higher turbine inlet temperatures, implying increasing heat loads to the engine components. Engine manufacturers have recognised this for some time and have been continuously increasing turbine inlet temperature, especially during the last three decades. The blades are cooled by extracting air from the compressor stages.

Modern gas turbine engines are designed to operate at inlet temperatures of 1800-2000K, which are far beyond the allowable metal temperatures. Thus, to maintain acceptable life and safety standards, the structural elements needs to be protected against the severe thermal environment. This calls for the design of an efficient cooling system for these elements. Rotor blade of high pressure gas turbine is such a critical element and hence the blade metal temperature should not be allowed to exceed beyond a value at which the life or safety standards can’t be met. It is required to cool the blade in such a way that the amount of heat transferred from the externally flowing hot gas to the blade should be removed by an appropriate cooling design to limit the very high temperature.

STRESSES IN THE BLADE

Rotor blades of gas turbine are subjected to very high rotational speeds of the order of several thousand rpm and also are exposed to a variable thermal environment. Hence these blades are subjected to different types of stresses of different magnitudes and directions. As it is known, that the strength is a function of life and working temperature the net stress at any section of the blade should not exceed the maximum allowable value. The control on the blade metal temperature is the only way to sustain the stresses for the designed life of the blade for a specific operating condition and life requirement. Therefore to know about the cooling requirement, stresses should be predicted correctly on the blades at different sections.

There are mainly four types of stresses with that rotor blades are being subjected;

1. Centrifugal tensile stress
2. Gas bending stress and
3. Centrifugal bending stress
4. Thermal stress

1.1. Centrifugal tensile stress

Centrifugal stress in the rotor blade is due to the rotation of the blade. It is tensile in nature. This is the largest in magnitude but not necessarily the most important because it is almost a steady stress. When the rotational speed of the blade is specified, the allowable centrifugal tensile stress places a limit on the annulus area but does not affect the choice of blade chord. This stress is the basic cause of the blade failure due to the creep.

1.2. Centrifugal bending stress

If the blade design is such that the centroids of all the blade cross-sections at different radii, taken perpendicular to the radial direction, do not lie in the same radial plane, centrifugal stresses arising in the blade will try to bend the blade. This type of stress arising due to the different directions of the centrifugal stresses in different blade sections is called as centrifugal bending stress. It will produce compressive stress in one side of the blade whereas tensile stress in the opposite side. Any torsional stress arising from these centrifugal stresses is small enough to be neglected. Thus this stress is very sensitive to manufacturing errors.

1.3. Gas bending stress

The force arising from the change in angular momentum of the gas in the tangential direction, which produces the useful torque, also tries to bend the blade about the axis of rotation of the blades. The stress arising due to this bending force is called as gas bending stress. There may be change of momentum in the axial direction and in reaction turbines there will certainly be a pressure force in the axial direction. All these two will produce a bending moment in the blade about the tangential direction. The gas bending stress will be tensile in the leading and trailing edges and compressive in the back of the blade and with tapered twisted blades either the leading or trailing edge suffers with the maximum value of this stress. This is a fluctuating stress and its value becomes maximum when the rotor blade passes through the leading edge of the stator.

1.4. Gas bending stress

Turbine blade is subjected to three-dimensional temperature gradients, along the blade height, along the blade profile and along the thickness of the blade.

Due to these temperature gradients the blade fibres tend to deform unequally. This unequal deformation causes mainly two types of stresses to set up in the blade, compressive and tensile. As the blade considered is un-cooled therefore the contribution of the stress due to the temperature gradient along the thickness of the blade in net stress is not appreciable and can be neglected. Usually with the cooled blade this source of stress is main among all the sources of thermal stress.

Again the thermal stress due to the temperature gradient along the blade height would not come in picture because the blade is free to expand along the height. Only the stress due to temperature gradient along the chord of the blade will contribute in net blade stress but its magnitude would not be much because the temperature gradient along the chord is not so high.

BLADE MATERIAL AND STRENGTH

Gas turbine blades are exposed to a very severe thermal atmosphere. The temperature is so high that it is fairly much more than the melting points of the common high-strength materials. Besides high temperature the requirement of durability is also another factor, which makes common materials unsuitable for use. Only super alloys may be suitable for this purpose. But the current trend of continuously increasing the turbine entry temperature attracted the concentration of the designers not only towards the new materials with well-improved mechanical and thermal properties but also to restrict the temperature of the blade material by its proper cooling. So, the material should have sufficient strength to face the operating situations.

1.5. Strength of blade material

In ordinary temperature conditions the strength of the material under constant loads is estimated by tensile strength or yield strength. At high temperatures under action of constant loads in ordinary structural materials there appears the phenomenon of creep. It occurs as a result of prolonged exposure of materials to high stresses at high temperatures. This is particularly a acute problem on highly stressed rotating turbine blades and it occurs in the form of slowly and continuously developing plastic deformation. And excess of this plastic deformation causes the failure of the component. It is observed that at constant stress the higher the temperature the more quickly proceeds the process of creep i.e. the lesser the life of the component. It means that at a particular stress lesser will be the temperature higher will be the life of component. Therefore life of the component is a function of working temperature and stress. Hence to maintain the life of the component at a desire value it is required to lower the temperature of the component.

Gas turbines operate in conditions of high temperatures and therefore in highly stressed components like rotor blades there appears the phenomenon of creep. Therefore for these cases where creep is the main criterion behind component failure the ultimate tensile stress is defined as the stress at which the component fails at a certain working temperature after the expiry of a certain period of time. It means that the strength of the material subjected at high temperatures is a function of this temperature and its operational life.

PAST COOLING

The technology of turbine cooling was recognised by some almost from the inception of the first turbojet engine. Cooling studies were first performed in the 1940 and many investigations were carried on in the 1950s. Around 1960, turbine cooling was first used in a commercial aircraft engine. Since that time, there has been a very rapid rise in turbine inlet temperature that has placed an even greater emphasis on turbine cooling. A continuous improvement in high-temperature materials has also helped to increase the turbine inlet temperature. The cooling technique used during 1960’s was single internal passage convection cooling. The air used for cooling was injected through the root of the blade and to the internal aerofoil.