US2536062A

US2536062A - System of blade cooling and power supply for gas turbines

Info

Publication number: US2536062A
Application number: US68106A
Authority: US
Inventors: Kane Saul Allan
Original assignee: Individual
Current assignee: Individual
Priority date: 1948-12-30
Filing date: 1948-12-30
Publication date: 1951-01-02
Anticipated expiration: 1968-01-02

Description

Jan. 2, 1951 s. A. KANE SYSTEM oF BLADE COOLING AND POWER SUPPLY FOR GAS TURBINES Filed Dec. V30, 1948 ma "ffm Fung .....m

MNG H Mmmm MMM CYCLE A www Mmm@ Patented Jan. 2, 1951 SYSTEM F BLADE COLING AND POWER SUPPLY FOR GAS TURBINES Saul Allan Kane, Washington, D. C. Application December 30, 1948, Serial No. 68,106

(Cl. fio-49) (Granted under the act of March 3; 1883, as

8 Claims.

This invention relates generally to gas turbine systems and more particularly to use of a part of the system for turbine blade cooling.

In gas turbine systems it is of importance from the viewpoint of thermodynamic efficiency that the temperature of the expanding gas at `the rotor blades be as high as possible. In prior art systems th's end has been sought by making the turbine rotor bladesof heat resistant material and by passing a cooling iluid through or over the rotor or rotor blades so that the average blade temperature may be reduced over that of the heated gas alone while maintaining a high operating temperature `for the Working ilu'd. f

A defect in the turbine system of this prior type lies in the marked loss of energy required to pump the coolant through the system.

It is proposed to reduce this loss of energy by utilizing the coolant not only for rotor blade cooling but also as yan active power agent. It is proposed also to `increase the cooling effect of the coolant by direct contact means.

Generally stated, the invention comprises a turbine system wherein the cooling air is subjected to additional compression over air suppl'ed the combustion unitand this highly compressed air is cooled in a heat exchanger and delivered to the turbine at a high pressure and a moderate temperature where it is expanded at a velocity equivalent to that of the heated but moderately pressurized gases when expanded through the nozzle sector. 'Ihe coolant is supercooled and utilized both as a cool'ng medium and for contributing power to the turbine rotor.

The objects of the invention include provision of means for stepping up the normal compression of cooling gases; the provison of means for cooling said super-compressed coolant prior to application to the turbine nozzles; the provision of means for obtaining the same velocity for the coolant at the rotor blades as, that obta=ned from the heated gases following the expansion thereof. I

Other objects and many of the attendant advantages of this invention will be lreadily appreciated as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawing wherein- Figure 1 is a schematic drawing showing the turbine system;

Fig. 1A is a fragmentary View, with parts broken away, illustrating structure of gas turbine and nozzle chest combination.

amended April 30, 1928; '370 O. G. 157) I Figure 2 is an enthalpy-entropy diagram of the turbine system showing the main heating cycle;

Fig. 3 `is a similar type diagram showing the cooling cycle; and

Fig. 4 is a view showing the nozzle chests for the heating and cooling cycles, the same being a section taken on line 4 4 of Fig. 1.

The gas turbine which is indicated 'by the numeral I0, is mechanically connected by shaft II to compressors I2 and I3 which may be respectively of the axial and radial type. Supports I4 are also indicated on which the unit is rotatably mounted. Motor I5 is provided for starting the unit.

A combustor I6 having a fuel line Il controlled by valve I8 supplies fuel to the combustor. Air is supplied the combustor from compressor I2 by means of pipe sections I9 and 2U, a heat exchanger 2I being interposed between these pipe sections. From the combustor, heated gases are led by pipe 22 into the main nozzle chest 23 of the turbine. The branch `pipe system including pipe sections I9, 2U and 22, heat exchanger 2| and combustor I6 will be referred to hereinafter as branch A.

A fraction of the compressed air output of compressor I2 is led to the compressor It where the pressure is increased. Accompanying this pressure increase is a rise in temperature which is removed in heat exchanger 2| to which connection is made by pipe section 25. This relatively cool and highly compressed air is led by pipe 25 through heat'l exchanger coil 26 and pipe 2l to the nozzle chest 28 of the turbine, where it is expanded and directed against the rotor blades supplementing the gas supply from branch A. The branch including compressor I3. heat l exchanger 2l and

pipe sections

25 and 21 will be Areferred to herenafter as branch B.

It will now appear that power is supplied to the turbine through the two branches A and B. the heatedl gas at lower compression through branch A and the cooled gas at higher compression through branch B. The two branch gas streams are expanded through nozzles to approximately the same velocity and since the work converted by an impulse turbine is a function of mass and velocity a single optimum combination of design characteristics will be satisfied by both hot and cold gas streams.

The thermodynamics involved in the .system operation may best be understood by reference to the enthalpy-entropy diagrams of Figures 2 and 3. It is understood that the end sought is 3 l the utilization at the turbine of branch A gases at a temperature as high as may be safely accommodated by the rotor materials in order to obtain a high operating efliciency. Obviously a gas temperature destructive to the rotor material may not be continuously applied. but by employing with the heated gas of branch A the cooling air of branch B a weighted averagedtemperature at the rotor blades is obtained which can be below the critical value for the rotor blade material although the branch A gases themselves l might be higher.

It is pointed out that the branch gases are applied to the rotor at spaced intervals of the nozzle sector so that the mixture of the heated gas with the cooling gas does not occur until they have left the nozzles proper. (Bee Fig. 4.)

Thus to combine the two temperature branches is a useful operation. In addition, however, two important conditions affecting the gas streams are introduced. In the one condition a heat exchange is obtained between the two branches through the heat exchanger 2|. By this means a proportion of the heat developed in the branch B air by the added compressor action is removed so that, on expansion at the turbine nozzle, the cooling effect is increased. lIn addition, heat is absorbed by the branch A air thus increasing the internal energy of this gas stream. The heat exchanger therefore serves to increase the enectiveness of both heating and cooling.

A second desirable condition arises from the utilization of the cooling air as a power agent for driving the turbine rotor. This results from stepping up the pressure of the air and applying it by nozzles directly to the rotating blades in parallel with the branch A gas nozzles. Preferably the cooling air velocity at the nozzles should be the same as that secured at the nozzles of branch A so that proper introduction of the working fluid into the blade entrance channels may be accomplished without undue shock or spill over. This velocity equivalence is secured by proper selection of compression values in branch B and nozzle design.

Referring to Figure 2 there is shown a heating cycle with enthalpy-entropy coordinates and designating letters corresponding with letters on Figure 1. Following the first compression ab with resultant temperature rise, heat is absorbed at constant pressure due to the heat exchanger, as indicated by bc, and the combustor increases gas temperature to peak d along cd, a temperature normally in excess of rotor material resistance. Adiabatic gas expansion does work along de. Point m on curve bd represents usual maximum temperature which will result in a. temperature after expansion below the highest safe temperature at the rotor, as employed in prior systems.

Figure 3 indicates the cooling cycle, with heat v addition ab and bf in branch B due to compressors l2 and Il. Expansion through the heat exchanger reduces air temperature along fg' so that on expansion through the turbine nozzle a drop of temperature gh is obtained.

Stated in terms of equation the net work theoretically developed by the system per pound, for example, would be and the efliciency is W/A ,(d-c). The average temperature leaving the turbine nozzles is A(Te) +B(Th) 4 Y whichsinceAandBarefractionslessthanlmlty results in an average T substantially less than Te, the peak temperature value.

Thus it appears that heated gases at supercriticalpeak temperatures and superoooled coolant may be applied to the turbine rotor at equal velocities to develop an increased overall efflciency for the system.

Obviously, many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically describedg The invention describedvherein may be manufactured and used by or for tl, Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

What is claimed is:

l. A gas turbine unit comprising a first air compressor, a combustion unit for receiving air from said first compressor, a fuel supply for said combustion unit, a turbine having a set of velocity nozzles and a rotor provided with blades. a shaft forming a common driving element between the turbine rotor and first compressor, a second compressor mounted on and rotatable with said rotor shaft, a heat exchanger, a rst conduit connecting the first rotor through the heat exchanger and the combustion unit to one sector oi' the turbine nozzles, and a second conduit connecting said first compressor through the second compressor and heat exchanger to another sector of said turbine nozzles whereby the turbine rotor is subjected to the rotative effect of both gas and air streams and to the cooling eiect of said air stream.

2. A gas turbine unit comprising a turbine having a rotor provided with edge blades, a set of stationary nozzles and an enclosing casing; a combustion unit; conduit connections between the combustion unit and one sector of said nozzles; a source of fuel for said combustion unit; means for supplying air under reduced compression to said combustion unit; means for supplying air under increased compression directly to another sector of said nozzles; and a heat exchanger operatively connected to both compression means for removing heat from said increased compression means, whereby the cooling eil'ect thereof on said turbine rotor is increased.

3. A gas turbine unit comprising a turbine having a rotor provided with edge blades; a set of l stationary nozzles and an enclosing casing; a

combustion unit; a source of fuel for said combustion unit; first and second compressors; shaft means connecting said compressors and turbine for simultaneous operation; a heat exchanger including a casing and a heat transfer medium therein; ilrst conduit connections from the rst compressor to one nozzle sector of the turbine nozzles and through said heat exchanger and combustion unit; second conduit connections from the first compressor to another nozzle sector of the turbine nozzles and through the second compressor and the heat exchanger, said compressors acting cumulatively on air passing through said second conduit connections to increase the pressure over the air in said first conduit connections.

4. A gas turbine unit comprising a turbine having a rotor provided with edge blades, a set of stationary nozzles and an enclosing casing; a combustion unit; conduit connections between the combustion unit and one sector of said nozzles; a source of fuel for said combustion unit; means for supplying air under reduced compression to said combustion unit; means for supplying air under increased compression directly'to another sector of said nozzles; and a heat exchanger operatively connected to both compression means for removing heat from said increased compression means, whereby the cooling effect thereof on said turbine rotor is increased, said heat exchanger including a casing, a heat transferring medium. and two heat transfer elements each connected in series with one of said conduit connections.

5. In a gas turbine system, a gas turbine including a rotor with radial peripheral blades and adjacent circumferential stationary nozzles, means for supplying heated gases to a sector of said stationary nozzles, and added means for supplying coolant under pressure to another sector of said stationary nozzles, said added means including compressor apparatus and a cooling device whereby theeective drop of temperature of the coolant at the rotor is increased.

6. In a gas turbine system, a gas 'turbine including a rotor, radial peripheral blades and adjacent circumferential stationary nozzles. a combustor for supplying heated gases to a sector of said stationary nozzles, a ilrst compressor for supplying compressed airto said combustor, and means including a second .compressor for supplying coolant under pressure to another sector `of said stationary nozzles.

'7. In a gas turbine system, a gas turbine including a rotor having radial peripheral blades.

and adjacent circumferential stationary nozzles; a combustor; a rst compressor for supplying compressed air to said combustor; a conduit connection between the combustor and a sector ot said stationary nozzles for supplying to said sector compressed heated gases; and means including a second compressor for supplying a coolant, at a pressure in excess of said compressed air, to another sector of said stationary nozzles..

8. In a. gas turbine system, a gas turbine including a rotor having radial peripheral blades. and adjacent circumferential stationary nozzles; a combustor; a first compressor for supplying compressed air to said combustor; a conduit connection between the combustor and a sector of said stationary nozzles for supplying to said sector compressed heated gases; and means including a second compressor for supplying a coolant, at a pressure in excess of said compressed air, to another sector of said stationary nozzles, said pressure and nozzle capacities being such as to secure equal velocities at the nozzles of both noz zle sectors.

SAUL ALLAN KANE.

REFERENCES CITED The following references are of record in the file of'this patent:

UNITED STATES PATENTS Johnson Mar. 22, 1949