1. Field of the Invention
This invention relates to turbine machines and particularly to rotor disks of gas turbine engines.
2. Description of the Prior Art
The gas turbine engine is typical of turbine machines in which the concepts described herein may be advantageously employed. In a gas turbine engine pressurized air and fuel are burned in a combustion chamber to add thermal energy to the medium gases flowing therethrough. The effluent from the chamber comprises high temperature gases which are flowed downstream in an annular flow path through the turbine section of the engine. The high temperature gases of the effluent induce radial thermal gradients across the components of the engine. The gradients exist both inwardly and outwardly of the flow path and are responsive to temperature and flow rate variations of the medium in the flow path. The thermal gradients across each individual turbine component are dependent upon both the position of the component relative to the flow path and of the influence which turbine cooling systems exert upon that component.
The gradients in a turbine rotor disk develop radially across the disk from the rim, which is located nearly adjacent the flow path, to the disk bore, which is located remotely from the flow path. In an uncooled disk the rim responds more rapidly to deviations in the flow path thermal characteristics than does the bore. Resultantly, the thermal profile between the rim and the bore is altered with each significant variation in a thermal characteristic. Each alteration in thermal profile causes a corresponding change in the disk thermal stress. The time period bracketing each stress peak, such as is illustrated in FIG. 7, is known as a "stress cycle". In each aircraft engine the combination of all stress cycles experienced by a component from aircraft takeoff through landing is termed "flight cycle".
Fatigue failure as a result of stress cycling adversely limits the life of components such as rotor disks. The minimum surface life expected is, according to industry standards expressed in terms of "low cycle fatigue life". The low cycle fatigue life for a rotor disk of an aircraft turbine engine is the minimum number of aircraft flight cycles to which the disk may be exposed without experiencing fatigue failure. As discussed above, the fatigue life of a disk is a function of all the individual stress cycles within a flight cycle. The low cycle fatigue life of a component is, then, calcuable as follows. EQU 1/N=1/N.sub.1 +1/N.sub.2 +1/N.sub.3 +. . .
where
N is the low cycle fatigue life expressed in flight cycles; PA1 N.sub.1 is the low cycle fatigue life based upon the first stress cycle; PA1 N.sub.2 is the low cycle fatigue life based upon the second stress cycles; and PA1 N.sub.3 is the low cycle fatigue life based upon the third stress cycle.
The number of stress cycles in each flight cycle varies according to the mission of the aircraft, but for a gas turbine engine in one typical commerical flight cycle the number of significant stress cycles is three. The first stress cycle occurs on takeoff; the second stress cycle occurs during climb to cruise; and the third stress cycle occurs after engine deceleration on descent. The effect of each stress cycle upon the overall low cycle fatigue life is largely dependent upon the magnitude of the included stress excursion or, in other words, the difference between the maximum and minimum stresses within each stress cycle. Additional factors such as the duration of individually imposed stresses also effect the fatigue life of the disk. The magnitude of the stress excursion in the first stress cycle, however, generally predominates.
In most modern gas turbine engines secondary flow systems are operative in the turbine region to cool various components including rotor blades. The cooling systems inherently cause variations in the thermal gradients imposed upon a disk by the medium gases of the flow path. Blade cooling systems such as that shown in U.S. Pat. No. 3,742,706 to Klompas entitled "Dual Flow Cooling Turbine Arrangement for Gas Turbine Engines " have directed cooling air through a rotor disk to interior cavities of the blades mounted peripherally upon the disk. Coolant passages in rotor disks such as that shown in Klompas have not, however, been contoured for optimum cooling of the rotor disk.
Substantial efforts are currently underway to provide rotor disk cooling which optimizes stress patterns during operation of the engine by reducing the number of stress cycles in each flight cycle and reducing the magnitude of the remaining stress excursions.