Patent Application
20120282101
1. Technical Field
This invention relates generally to gas turbine engines and particularly to a gas turbine engine rotor construction.
2. Background Information
Gas turbine engines, such as those which power aircraft and industrial equipment, employ a compressor to compress air which is drawn into the engine and a turbine to capture energy associated with the combustion of a fuel-air mixture which is exhausted from the engine's combustor. The compressor and turbine employ rotors which typically comprise a multiplicity of airfoil blades mounted on, or formed integrally into the rims of a plurality of disks. The compressor disks and blades are rotationally driven by rotation of the engine's turbine. It is a well-known prior art practice to arrange the disks in a longitudinally axial stack in compressive interengagement with one another which is maintained by a tie shaft which runs through aligned central bores in the disks. It is a common practice to arrange the disks so that they abut one another in the aforementioned axial stack along side edges of the disk rims. The disk rims are exposed to working fluid flowing through the engine and therefore are exposed to extreme heating from such working fluid. For example, in a gas turbine engine high pressure compressor, the rims of the disks are exposed to highly compressed air at a highly elevated temperature. The exposure of disk rims to such elevated temperatures, combined with repeated acceleration and deceleration of the disks resulting from the normal operation of the gas turbine engine at varying speeds and thrust levels may cause the disk rims to experience low cycle fatigue, creep and possibly cracking or other structural damage as a result thereof This risk of structural damage is compounded by discontinuities inherent in the mounting of the blades on the rims. Such discontinuities may take the form of axial slots provided in the rims to accommodate the roots of the blades or, in the case of integrally bladed rotors wherein the blades are formed integrally with the disks, the integral attachment of the blades to the disks. Such discontinuities result in high mechanical stress concentrations at the locations thereof in the disks, which intensify the risks of structural damage to the disk rims resulting from the low cycle fatigue and creep collectively referred to as thermal mechanical fatigue, experienced by the disks as noted hereinabove. Moreover, the high compressive forces along the edges of the disk rims due to the mutual abutment thereof in the aforementioned preloaded compressive retention of the disks in an axial stack further exacerbates the risk of structural damage to the disk rims due to the aforementioned low cycle fatigue and creep.
Therefore, it will be appreciated that minimization of the risk of disk damage due to thermal-mechanical fatigue, and stress concentrations resulting from discontinuities in the disk rim is highly desirable.
In accordance with the present invention, a gas turbine engine rotor comprising a plurality of blade supporting disks adapted for longitudinal compressive interengagement with one another includes at least one disk comprising a medial web and an annular rim disposed at a radially outer portion of the web, the rim including longitudinally extending annular shoulders and further comprising an annular spacer extending longitudinally from the disk proximal to the juncture of the web and rim, and being spaced radially inwardly from one of the shoulders for abutment at a free edge of the spacer with an adjacent disk for transmission of compressive preloading force from the one disk to the adjacent disk, the spacer and the one shoulder defining an annular slot in which a base of a segmented annular blade cluster is received. The spacer allows the compressive preloading of the disks to be transmitted therebetween radially inwardly of the disk rim so as to not exacerbate thermal mechanical rim fatigue. The blade cluster thermally shields the rim from at least a portion of the destructive heating thereof by working fluid flowing through the engine.
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As set forth hereinabove, the catenary shape of spacer 75 causes the spacer to act as a compression spring for preservation of the compressive preload of each disk against an adjacent disk for effective torque transmission therebetween. Since disk compressive preloading forces are transmitted through the spacers, the disk rims which experience severe thermal loading from the heat of the working fluid are not subjected to the compressive preloading forces which would otherwise exacerbate the thermal mechanical fatigue discussed hereinabove which the disk rims experience from the high temperature working fluid flowing therearound. The blade clusters themselves provide some insulative properties, thereby protecting the disk rims from heat carried by the working fluid flowing past the rotor. The segmented nature of the annular blade cluster bases reduces hoop stress therein from levels thereof which would be inherent in full, annular blade clusters. The definition of slots 90 and 92 by the rim shoulders and spacers eliminate the need for the formation of slots directly in the disk rims to accommodate individual blade roots. As set forth hereinabove, stress concentrations associated with such individual slots would otherwise exacerbate the thermal-mechanical fatigue associated with low cycle rim fatigue and creep. Furthermore, since individual blade slots are not necessary with the present invention, the disk rim portions may be efficiently and economically coated with any appropriate thermal barrier coating such as zirconium oxide or the like. Further disk stress reduction is achieved by the retention of the blade clusters by the rim shoulders which are more compliant than that portion of the disk rim which is in radial alignment with the disk web.
While a specific embodiment of the present invention has been shown and described herein, it will be understood that various modification of this embodiment may suggest themselves to those skilled in the art. For example, while the gas turbine engine rotor of the present invention has been described within the context of a high pressure compressor rotor, it will be appreciated that invention hereof may be equally well-suited for turbine rotors as well. Also, while specific geometries of portions of the disks and blade clusters have been illustrated and described, it will be appreciated that various modifications to these geometries may be employed without departure from the present invention. Similarly, while a specific number of compressor disks have been shown and described, it will be appreciated that the rotor structure of the present invention may be employed in rotors with any number of blade supporting disks. Accordingly, it will be understood that these and various other modifications of the preferred embodiment of the present invention as illustrated and described herein may be implemented without departing from the present invention and is intended by the appended claims to cover these and any other such modifications which fall within the true spirit and scope of the invention herein.
