GB2118248A

GB2118248A - Double flow path gas turbine engine

Info

Publication number: GB2118248A
Application number: GB08208522A
Authority: GB
Inventors: Clive Edward Olive
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1982-03-23
Filing date: 1982-03-23
Publication date: 1983-10-26

Abstract

In a gas turbine engine provided with two parallel gas flow paths, one of the flow paths (path I) is through the H.P. compressor 6, combustor 8 and H.P. turbine 10 and the second flow path (path II) by-passes the H.P. compressor 6 and H.P. turbine 10. For take-off, path I is opened and path II is closed by moving doors 18,20,22 to produce high thrust for take-off. For supersonic flight, path II is opened and path I is closed by moving the doors 18,20,22 (thus by-passing the H.P. spool 6). <IMAGE>

Description

SPECIFICATION Gas turbine engine This invention relates to gas turbine engines and in particular to gas turbine engines for powering aircraft which are required to operate over a wide range of speeds.

With existing gas turbine engines for powering aircraft, overall engine performance is compromised to give maximum thrust at takeoff, with the result that at high speeds engine performance is restricted by limitation of flow through the engine.

It is an object of the present invention to provide a gas turbine for powering an aircraft wherein the above disadvantage is overcome or at least alleviated.

According to the present invention a gas turbine engine for powering an aircraft comprises a first compressor; a second compressor; combustor means; a first turbine for driving the second compressor; a second turbine for driving the first compressor; a first flow path defined within the engine from the first compressor to the second turbine via in flow series the second compressor, the combustor means, and the first turbine; a second flow path defined within the engine from the first compressor to the second turbine via the combustor means, the second flow path bypassing the second compressor and the first turbine; and means for selectably varying gas flow through the engine between the first flow path and the second flow path.

It will be appreciated that in powering an aircraft such an engine can produce high thrust at take-off and also operate efficiently at high speeds, e.g. supersonicaliy.

Two gas turbine engines in accordance with the invention will now be desribed, by way of example only, with reference to the accompanying drawings, in which: Figure 1 shows a schematic, partly-sectioned elevation of the first engine; and Figure 2 shows a schematic, partly-sectioned elevation of the second engine.

Referring firstly to Fig. 1, a first gas turbine engine 2 for powering a supersonic aircraft comprises a first, low-pressure (L.P.) compressor 4, a second high-pressure (H.P.) compressor 6, a combustor 8, a first, H.P. turbine 10 which drives the H.P. compressor 6 and a second, L.P. turbine 12 which drives the L.P.

compressor 4.

Downstream of the L.P. compressor 4 an intermediate casing surrounding the H.P.

spool 6 divides gas flow through the engine into two paths designated I and II. Flow path I passes directly through the combustor 8 to the L.P. turbine 1 2. Flow path II passes through the H.P. compressor 6, exiting through a final centrifugal stage 1 4 thereof.

The flow path II then passes through the combustor 8 and enters the H.P. turbine 10 via a radial flow stage 1 6 thereof, passing through H.P. turbine 10 to the L.P. turbine.

Movable blocker doors 18,20 and 22 are provided to allow the gas flow through the engine to be selectably varied between the flow paths I and II.

Thus with the blocker doors 18,20 and 22 in the positions shown in full line in the drawing, flow path I is blocked and gas flow through the engine passes through the L.P.

compressor 4 and follows flow path li through the H.P. compressor 6, the combustor 8 and the H.P. turbine 10 before finally passing through the L.P. turbine 1 2. It will be appreciated that in this configuration of the engine high thrust is developed by the L.P. and H.P.

compressors working in series. This configuration is suitable for aircraft take-off.

With the blocker doors 18,20 and 22 in the positions shown in broken line in the drawing, flow path II is blocked off and gas flow through the engine passes through the L.P. compressor 4 and follows flow path I directly through the combustor 8 before finally passing through the L.P. turbine 1 2. It will be appreciated that in this configuration the engine, producing less thrust, can operate very efficiently at high speeds since gas flow through the engine by-passes the H.P. spool 6,10 which would limit flow at high speed and so restrict engine performance. This configuration is suitable for aircraft propulsion at supersonic speed.

It will be appreciated that the engine configuration may be varied between the two extreme configurations described by setting the blocker doors 18,20 and 22 to positions intermediate those shown in full-line and broken line to vary the proportions of flow through flow paths I and II as required.

If it is desired, when the engine gas flow is via flow path I, to block off completely any flow which may be induced from the centrigual stage 14 of the H.P. compressor 6 through the combustor 8 and into the radial flow state 1 6 of the H.P. turbine 10, this can be achieved by making the combustor 8 axially movable so as to block off the passages 24 and 26 between the combustor 8 and the H.P. spool 6,10.

It will be appreciated that the engine 2 ordinarily operates as a by-pass engine, some of the flow from the L.P. compressor passing to atmosphere through by-pass opening 28. If it is desired to reduce this by-pass flow when gas flow through the engine is via flow path I, this can be achieved by moving the blocker doors 20 to obturate, partially or totally, the by-pass opening 28.

It will be appreciated that in order to allow the performance of the engine to be optimised in its various configurations variable flaps 30 and 32 are provided at the inlet and outlet respectively of the engine to vary the inlet and outlet areas. Also, variable geometry compo nents, e.g. stator vanes or fan blades, may be employed to aid optimisation in various configurations: Referring now to Fig. 2, a second gas turbine engine for powering a supersonic aircraft is similar to the first engine described above, but instead of the engine flow paths I and II having a common portion in which a common combustor is situated, in the second engine the flow paths I and II have no common portion and separate portions of the combustor 8a and 8b are provided in each flow path. Having separate portions of the combustor in separate flow paths allows the flow paths to be straightened and to this end the radial stages 14 and 16 of the H.P.

compressor and H.P. turbine respectively may be replaced by axial flow components.

It will be appreciated that although in the above described engines blocker means in the form of blocker doors are used, alternative forms of blocker means, e.g. shutters or variable irises, may be used if desired.

Claims

1. A gas turbine engine for powering an aircraft, the engine comprising: a first compressor; a second compressor; combustor means; a first turbine for driving the second compressor; a second turbine for driving the first compressor; a first flow path defined within the engine from the first compressor to the second turbine via in flow series the second compressor, the combustor means, and the first turbine; a second flow path defined within the engine from the first compressor to the second turbine via the combustor means, the second flow path by-passing the second compressor and the first turbine; and means for selectably varying gas flow through the engine between the first flow path and the second flow path.

2. A gas turbine engine according to claim 1 wherein the first flow path and the second flow path have a common portion in which a common combustor means is situated.

3. A gas turbine engine according to claim 1 or 2 wherein the means for selectably varying gas flow comprises blocker door means situated in the first flow path and in the second flow path to block a selected one of the first and second flow paths.

4. A gas turbine engine substantially as hereinbefore described with reference to the accompanying drawing.