GB2064008A

GB2064008A - A Cooling System for a Gas Turbine Engine

Info

Publication number: GB2064008A
Application number: GB8034559A
Authority: GB
Original assignee: FN Herstal SA; Herstal SA; Alsthom Atlantique SA
Current assignee: FN Herstal SA; Herstal SA; Alstom SA
Priority date: 1979-11-07
Filing date: 1980-10-27
Publication date: 1981-06-10
Also published as: IT8068698A0; FR2469565A1; FR2469565B1; DE3041175A1; BE885794A; IT1129374B

Abstract

A cooling system for a gas turbine engine which turbine engine includes a compressor unit (1, 2), a combustion chamber (3) and a turbine (4), the cooling system comprising discontinuous cooling by passing at least some of the blades of said turbine alternately through a hot flux from the combustion chamber (3) and a cold flux that does not pass through the combustion chamber with the hot and cold fluxes being separated by partitions at the upstream and downstream ends of the path of the cold flux through the turbine, in conjunction with continuous cooling by circulating air (A) along ducts located inside at least some of the blades of said turbine. <IMAGE>

Description

SPECIFICATION A Cooling System for a Gas Turbine Engine The present invention relates to a cooling system for a cyclically cooled gas turbine engine which includes a compressor, a combustion chamber and a turbine.

Background of the Invention French patent application No. 2,087,405 (corresponding to British patent No. 1,347,514 and U.S. patent No. 3,731,486) describes a gas turbine engine which includes a centrifugal compressor provided with an outlet diffuser, a combustion chamber and a centripetal turbine whose inlet is provided with volute and a distributor, the rotors of the compressor and of the turbine being integral with each other and forming a continuous shaft, the turbine rotor being discontinuously cooled by alternately passing through a hot flux which comes from the combustion chamber and a cold flux which comes from the compressor and is conveyed directly from the outlet of the moving blade assembly of the compressor to the inlet of the moving blade assembly of the turbine without passing through the diffuser nor through the volute nor through the distributor, the hot and cold fluxes being separated by upstream and downstream partitions in the turbine.

This disposition provides excellent cooling of the turbine and therefore makes it possible to increase the temperature of the gases at the turbine inlet. However, although during testing it turned out that cooling was better than expected, power and fuel consumption were not.

The present invention aims to improve the system so as to increase its power and reduce its fuel consumption.

Summary of the Invention The present invention provides a cooling system for a gas turbine engine which turbine engine includes a compressor unit, a combustion chamber and a turbine, the cooling system comprising discontinuous cooling by passing at least some of the blades of said turbine alternately through a hot flux from the combustion chamber and a cold flux that does not pass through the combustion chamber, with the hot and cold fluxes being separated by partitions at the upstream and downstream ends of the path of the cold flux through the turbine, in conjunction with continuous cooling by circulating air along ducts located inside at least some of the blades of said turbine.

According to one particular embodiment of the invention, said cold flux is conveyed in a closed circuit by a separate compressor and is cooled by cooling means.

According to another embodiment, the heated cold flux which leaves the turbine passes through a heat exchanger as a first fluid and then through a cooling system before returning to the compressor unit, and after passing through a part of the compressor unit, the flux passes again through said heat exchanger as a second fluid before entering the combustion chamber.

Preferably, the turbine is an axial-flow turbine and the compressor unit includes a two-part axialflow compressor, the second part of which also conveys air sucked in from outside, the cold flux for cooling the turbine being taken from air sucked into said second part.

According to another variant, the turbine is an axial-flow turbine and the compressor unit includes an axial-flow compressor supplied with air from the outside and a centrifugal compressor supplied with the heated cold flux which leaves the turbine.

The system in accordance with the invention combines two known cooling systems: the alternate hot and cold flux system, and the system of circulating air through ducts inside the blades.

This gives excellent results, since, with the alternate hot and cold flux system alone, the turbine inlet temperature can be raised from 900O to 11 OOC, i.e. 2000C more than in a non-cooled turbine; and with the system which circulates air through ducts inside the blades on its own, the turbine inlet temperature can be raised from 900O to 1 380C, i.e. 2380C more than a non-cooled turbine; while with the apparatus in accordance with the invention which combines the two systems, the inlet temperature can be raised to 1 4050C; this is a gain of 5050C relative to a noncooled turbine plant.

Brief Description of the Drawings Three embodiments of the invention are described by way of example with reference to the accompanying drawings in which: Figure 1 schematically illustrates a turbine engine in accordance with the invention in which the compressor unit comprises both an axial-flow compressor for feeding the combustion chamber and a centrifugal compressor for circulating a cold flux; Figure 2 illustrates another gas turbine which is a variant of that shown in Figure 1; and Figure 3 schematically illustrates a gas turbine engine for an aeronautical application in which the compressor unit comprises a single, two-part, axial-flow compressor.

Detailed Description With reference to Figure 1, a turbine engine comprises a power turbine 5 driven by a gas generator which includes an axial-flow turbine 4, a combustion chamber 3 and an axial-flow compressor 2 and a centrifugal compressor 1. The inlet end of the axial flow turbine 4 is divided into two sectors such that the inlet turbine wheel receives hot gases from the combustion chamber 3 over about 70% of its circumference and cold gases from the centrifugal compressor 1 over the remaining 30% of its circumference. The cold gases do not pass through the combustion chamber 3. Thus, the axial flow turbine 4 is discontinuously cooled by alternately passing through a hot flux and a cold flux.In accordance with the invention, conventional cooling is superposed on this cyclic cooling; said conventional cooling consists in circulating air through ducts formed inside the blades of the turbine. Such circulation is symbolically represented by arrows A. The cold flux undergoes a small pressure drop in the axial-flow turbine 4 and is separated from the hot flux at a cold flux outlet 6 from the turbine unit 4. The heat transferred to the cold flux by the turbine unit 4 is then extracted in a heat exchanger 7 and the cold flux then returns to the suction end of the centrifugal compressor 1 via conduit 8.

Naturally, the person skilled in the art knows how to manufacture partitions so as to separate the fluxes at the inlet of the turbine 4 and at its cold flux outlet 6.

In the case illustrated by Figure 2, the heat transferred to the cold flux by the turbine is reinserted into the cycle. In this example, the exhaust from the axial-flow compressor 2 is separated into two zones. A first exhaust sector of about 70% directs compressed air towards the combustion chamber 3 via a conduit 9 and a second exhaust sector of about 30% directs compressed air towards an inlet sector of about 30% to the axial-flow turbine 4 via a conduit 10.

From the separation point 6 of the two fluxes, after a small pressure drop in the turbine and before being returned to the inlet of the centrifugal compressor 1, said cold flux is sent via a conduit 11 through a heat exchanger 12 where it gives up its heat to the cold compressed air from the centrifugal compressor 1. A cooling apparatus 13 is disposed upstream from the inlet to the compressor. After leaving the compressor the cold compressed air passes through the heat exchanger 12 where it is heated, and is then sent via a conduit 14 to the combustion chamber 3. As in Figure 1, the arrows A symbolize a circulation of air in ducts in the turbine blades.

Thus, in this example, some of the heat transferred by the axial flow turbine 4 to the cold flux is recuperated. The heat is recuperated by a heat exchanger 12 but since the efficiency of the exchanger is not equai to unity, some of the heat coined by the cold flux on passing through the turbine must be lost in the refrigeration apparatus 13. However, the power absorbed by the compressor 1 is reduced and the more it is reduced, the greater the useful power.

Figure 3 illustrates a gas turbine engine for aeronautical applications in which the compressor unit is constituted by a single two-part, axial-flow compressor 2. The cold flux leaving the turbine 4 at its cold flux outlet 6 is reinserted by the conduit 11 into the second part of the axial compressor 2.

In this example, the hot gases at the outlet of the turbine unit 4 are used in a nozzle 1 5.

Claims

1. A cooling system for a gas turbine engine which turbine engine includes a compressor unit, a combustion chamber and a turbine, the cooling system comprising discontinuous cooling by passing at least some of the blades of said turbine alternately through a hot flux from the combustion chamber and a cold flux that does not pass through the combustion chamber, with the hot and cold fluxes being separated by partitions at the upstream and downstream ends of the path of the cold flux through the turbine, in conjunction with continuous cooling by circulating air along ducts located inside at least some of the blades of said turbine.

2. A cooling system according to claim 1, wherein said cold flux is conveyed in a closed circuit by a separate compressor and is cooled by cooling means.

3. A cooling system according to claim 1 or 2, wherein the heated cold flux which leaves the turbine passes through a heat exchanger as a first fluid and then through a cooling system before returning to the compressor unit, and wherein after passing through a part of the compressor unit, the flux passes again through said heat exchanger as a second fluid before entering the combustion chamber.

4. A cooling system according to claim 1, wherein the turbine is an axial-flow turbine and wherein the compressor unit includes two-part axial-flow compressor, the second part of which also conveys air sucked in from outside, the cold flux for cooling the turbine being taken from air sucked into said second part.

5. A cooling system according to any one of claims 1 to 3, wherein the turbine is an axial-flow turbine and vvherein the compressor unit includes an axial-flow compressor supplied with air from the outside and a centrifugal compressor supplied with the heated cold flux which leaves the turbine.

6. A cooling system for a gas turbine engine substantially as herein described with reference to and as illustrated in the accompanying drawings.