Turbine rotor cooling

Increased turbine inlet temperatures and/or extended life of turbine rotors in gas turbine engines is achieved by locating a plurality of passages 50 that extend axially in the disc 46 and turbine blades 32 of a so-called monorotor. The passages 50 have inlets 54 between compressor blades 18 on the rotor 20 and exit openings 52 in the free edges 38 of the turbine blades 32. Thus, air compressed by the compressor blades 18 is flowed through the passages 50 to cool the rotor 20.

FIELD OF THE INVENTION 
This invention relates to gas turbines, and more particularly, to the 
cooling of turbine rotors so as to increase rotor life and/or allow 
operation at higher turbine inlet temperatures. 
BACKGROUND OF THE INVENTION 
Generally speaking, and with all other things being equal, the efficiency 
of operation of a gas turbine increases as the turbine inlet temperature, 
that is, the temperature of the gas applied to the turbine blades, is 
increased. However, in practice some efficiency to be gained by high 
temperature operation is given up by using turbine inlet temperatures that 
are sufficiently low as to enable the turbine rotor to last a reasonable 
useful life without failure. In short, relatively higher turbine inlet 
temperatures, while increasing turbine efficiency decrease the useful life 
of the turbine while relatively lower turbine inlet temperatures provide 
for a long lived turbine, they result in decreased operating efficiencies. 
In order to increase turbine life or allow turbine operation at higher 
turbine inlet temperatures without sacrificing useful life, various 
schemes have been proposed for cooling the turbine rotor. A number of the 
turbine rotor cooling schemes proposed are unwieldy and expensive to 
implement and thus have not been altogether satisfactory. 
The present invention is directed to overcoming this difficulty. 
SUMMARY OF THE INVENTION 
It is the principal object of the invention to provide a new and improved 
gas turbine. More specifically, it is an object of the invention to 
provide a new and improved cooled rotor for use in gas turbines so that 
higher turbine inlet temperatures may be utilized without sacrificing 
useful life or, where similar turbine inlet temperatures are used, the 
turbine is provided with increased useful life. 
An exemplary embodiment of the invention achieves the foregoing object in a 
gas turbine including a combustion gas inlet, an annular diffuser 
downstream of the inlet, and a combustor for receiving combustion 
supporting gasses and burning fuel therewith to produce a high pressure 
turbine driving gas. An annular nozzle is located adjacent the diffuser 
and receives the turbine driving gas and directs the same generally 
radially inwardly. A rotor including rotary shaft means extends through 
and is centered within the annular diffuser and the annular nozzle. The 
rotor has a disc with compressor blades on one side which are operatively 
interposed between the inlet and the diffuser, and turbine blades on the 
other side. A plurality of passages extend at least from the disc one side 
through the disc and through the turbine blades. 
As a consequence of this construction, combustion supporting gas being 
compressed by the compressor blades flows through the passages in the disc 
and in the turbine blades. This compressed gas will be considerably cooler 
than the turbine driving gasses and will thus serve to cool the rotor to 
either extend its useful life or allow its operation at higher turbine 
inlet temperatures. 
In a preferred embodiment, there are a plurality of such passages in each 
of the turbine blades and the passages extend generally axially. 
In one embodiment of the invention, the passages have inlet ends in the 
disc one side and located between the compressor blades. 
The invention contemplates that the turbine blades have edges opposite the 
disc and that the passages have outlet ends in such edges. 
In a highly preferred embodiment, the disc, the compressor blades and the 
turbine blades define an integral rotor body, that is, a so-called 
monorotor. 
Other objects and advantages will become apparent from the following 
specification taken in connection with the accompanying drawings.

DESCRIPTION OF THE PREFERRED EMBODIMENT 
An exemplary embodiment of a gas turbine made according to the invention is 
illustrated in the drawings and with reference to FIG. 1 is seen to be in 
the form of a radial flow turbine. The same has an air inlet end generally 
designated 10 at which is located a compressor blade shroud 12. An annular 
diffuser 14 of conventional construction is supported by the shroud 12 and 
is operative to perform the usual diffusion operation on compressed gas 
received from radially outer ends 16 of curved compressor blades 18 
carried on a rotor, generally designated 20. The compressor blades 18 also 
have axial ends 22 adjacent the inlet 10. 
Compressed gas to support combustion from the diffuser 14 is directed to an 
annular combustor shown schematically at 24 which also receives fuel on a 
line 26. The fuel is combusted within the combustor 24 and then directed 
to an annular nozzle 28. The hot gasses of combustion exiting the nozzle 
28 are directed radially inwardly in a conventional fashion to impinge 
against radially outer ends 30 of a plurality of turbine blades 32 on the 
rotor 20. 
The rotor 20 may be fitted to a shaft 34 so as to be centered within the 
diffuser 14 and the nozzle 28. The shaft 34 is journalled by suitable 
bearings (not shown) so that hot gasses of combustion impinging upon the 
blades 32 from the nozzle 28 will cause the rotor 20 and thus the shaft 34 
to rotate. The rotary motion will be utilized to perform useful work in a 
known manner and will also drive the compressor blades 18 to compress the 
air received from the inlet 10 for use in the combustion process. Gas 
exiting the nozzle 28 is confined against the blades 32 by a conventional 
turbine blade shroud 36 which is in close adjacency to the free edges 38 
of the turbine blades 32. The turbine blades 32, like the compressor 
blades 18, are curved and terminate in axial ends 40. 
The turbine blades 32 have a second edge 42 which, according to the 
preferred embodiment is really no edge at all. Preferably, the rotor 20 is 
a so-called monorotor, meaning that the compressor blades 18, the turbine 
blades 32, the rotor hub 44, and the central disc 46 separating the blades 
18 and 32 are all cast as a unitary structure and are integral. As thus 
far described, the rotor 20 may be regarded as conventional. 
FIG. 3 illustrates the temperature distribution across various parts of a 
conventional turbine rotor configured as described above for a turbine 
inlet temperature of 2,860.degree. Rankine. Thus, FIG. 3 illustrates 
temperature distribution across an uncooled turbine rotor. 
In order to operate at the same turbine inlet temperature, but to reduce 
the temperature caused stresses on the rotor 20 to extend its life, the 
cooling means of the invention are utilized to cool the rotor 20. 
Returning to FIGS. 1 and 2, the cooling means of the invention will be 
described. They include a plurality of generally axially extending bores 
or passages 50 which extend through the disc 46 from the side thereof 
adjacent the compressor blades 18 past the opposite side of the disc 46 
and through the turbine blades 32 to emerge from the free edges 38 
thereof. Thus, the ends 52 of the bores 50 in the free edges 38 of the 
blades 32 are exit openings for a coolant gas while the ends 54 of the 
bores 50 located in the channels between the compressor blades 18 are 
inlet ends whereby gas from the inlet 10 and compressed by the blades 18 
may enter the passages 50 and flow through the same. 
Thus, the invention accomplishes cooling of the turbine blades 32 by 
bleeding air from the compressor side of the rotor through the passages 
50. In a preferred embodiment, perhaps about 5% of the air entering the 
inlet 10 is caused to flow through the passages 50. Thus, FIG. 4 
illustrates the temperature distribution across the rotor for a 5% flow of 
inlet gas through the passages 50 for cooling purposes. It will be noted 
that the same turbine inlet temperature, namely, 2,860.degree. Rankine, is 
employed and that substantial temperature reductions from the uncooled 
rotor whose temperature distribution is depicted in FIG. 3 are achieved. 
The formation of the passages 50 can be accomplished by the simple means of 
boring the holes parallel to the axis defined by the shaft 34. Thus, the 
manufacture of the cooled rotor is not complicated and the same enables 
the use of higher turbine inlet temperatures or, prolongs the life of the 
rotor. It is expected that turbine inlet temperature could be increased 
approximately 200.degree. Rankine without sacrificing useful life through 
utilization of the invention. Alternatively, it is believed that with no 
increase in turbine inlet temperature, use of the invention will 
approximately double the life of the rotor.