The turbine is a major component common to the gas turbine-propeller engine, and to the thermal jet engine. In the gas turbine-propeller engine the turbine must develop the shaft power for driving the air compressor, propeller, and the auxiliaries. In the thermal jet engine, however, it is required to furnish only sufficient power to drive the air compressor and the auxiliaries. It should be noted that, in general, gas turbine-propeller engines are designed to deliver auxiliary jet thrust from the exhaust gases in addition to the propeller thrust, the usual proportions being 80 per cent propeller thrust and 20 per cent auxiliary jet thrust.

The general characteristics of turbines are well understood, and a wealth of information concerning them has been gathered during the past decades. Particularly helpful to the development of turbines which operate with highly heated gases are the experiences gathered in the development of turbo-superchargers and also steam turbines for high-pressure and hightemperature applications. In many respects the turbine for gas turbine-propeller engines or turbo-jet engines is quite similar to the conventional steam turbine, the major difference being in the metallurgy, the means provided for cooling the bearings and highly stressed parts, and in the constructional features to safeguard against thermal distortion. The basic theory underlying their design and the evaluation of their operating characteristics is identical with that for steam turbines.

Basic Requirements

The basic requirements for the turbine are the same for either type of engine. Although the remarks which follow apply specifically to the turbine for a turbojet engine, it should be understood that they apply equally well to the turbines for turboprop engines. The principal requirements are: light weight; small frontal area; high efficiency; ability to operate for sustained periods at high temperature; and reliability and serviceability.

Light weight is secured by operating the turbine rotor with the highest permissible rim speed, using small-diameter rotors. Since the stresses in a given turbine disk increase approximately as the square of the rim speed¹, the maximum rim speed is limited by strength considerations, which are governed by the stress characteristics of the disk and blade materials at the operating temperature. Although low rim speeds are desirable from a stress standpoint, there is a lower limit to the rim speed imposed by the dictates of high efficiency which, in general, improves with rim speed. The lower limit of the rim speed in feet per second, which is dictated by flow conditions, must be larger than

 ,

where ∆T is the temperature drop through the turbine in degrees Fahrenheit. Since the turbine efficiency improves, in general, with increasing rim speed and permits using lower bucket temperatures for the same power output, the choice of rim speed is a compromise between allowable stress and turbine efficiency. The rim speeds of most turbojet turbines range from 820 to

1,000 fps.

The general design of the turbine passages is based primarily on considerations which are mainly fluid dynamical. The flow conditions must be so designed that, for the required thrust output and mass flow of gas, the acoustic velocity (unify Mach number)² is not reached at the outlet from the buckets (or in the ducting downstream leading the exhaust gases away from the turbine, or in the exhaust nozzle), for choking of the flow occurs if the Mach number in these flow passages attains the value unity. The critical Mach number is based on the axial velocity of the gas in the exit annulus from the turbine. The possibility of attaining unity Mach number in the outlet from the turbine buckets is a consideration to be investigated in Rateau stage turbines.

Disk and rim failures in turbo-jet turbines did occur in the early development stages of this propulsion engine. They have now been overcome by the application of such methods as improved gas seals, the incorporation of methods for cooling the disk, and improved metallurgy. One of the major factors has been a better understanding of the metallurgical problem. Research has shown that if the disk operates with high temperatures or steep temperature gradients it is likely to develop plastic deformation. If this occurs the stress distribution can no longer be based on conventional elastic theory, and when the disk cools off after operating it is subjected to large residual stresses. As a consequence of the residual stresses there is a change in the natural vibration frequencies, which are functions of the stress conditions. Furthermore, successive periods of plastic strain, cooling, and then heating again modify the stress-strain characteristics of the disk material and may lead to changes in its crystal structure. By applying the remedies mentioned above these difficulties can be avoided.

In turbo-jet engines employing a centrifugal compressor, the turbine imposes no problem in the securing of small frontal area. The frontal area of the turbine is much smaller than that of the compressor and combustion chamber assembly and has little influence upon the overall size in that type of application. Where the turbine drives an axial-flow compressor the frontal areas of the turbine and compressor become more nearly equal.

The turbine blades may be either solid or hollow, the type of construction being influenced by the material selected for their manufacture. The hollow blade offers the advantages of being adapted to cooling by flowing cold air through its interior and of reducing weight. The walls of the blade are usually tapered so that the outer extremity, where the stress vanishes, is quite thin. The greatest benefit derived from cooling is at the root of the blade where the stresses are high; the outer edge, because of its small stress, may be allowed to run hot³. In most designs the blades are twisted to maintain a favorable angle of attack for the fluid throughout its length. In the early development of the turbo-jet, blading failures did occur, but they are now a rarity. The difficulties were overcome by increased accuracy in the manufacture of the blades, avoidance of small radii at root junctions, better analysis of vibration problems, and improved metallurgy.

Since improving turbine efficiency and output are related to ability to operate with higher temperatures, developments aimed at raising the permissible operating temperature of the turbine are of great importance. One promising approach is the application of ceramic coatings on the turbine blades to take the impact of the hot gases. The problem here is to develop a ceramic coating of high melting point which will bond to the metal and will have a coefficient of expansion close enough to that of the metal to prevent the coating from cracking or flaking off. Another approach proposes to let cooling liquid flow through a passage in the root of the blades.

A general thermodynamic treatment can be applied to both impulse and reaction stages by considering an intermediate stage of a multistage reaction turbine; the intermediate stage typifies the general case of a turbine stage. In such a stage the stationary blade row is the counterpart of the nozzle of an impulse stage.

(Kates E. K.)