GB2242172A

GB2242172A - Gas turbine engine.

Info

Publication number: GB2242172A
Application number: GB8328254A
Authority: GB
Inventors: Derek Aubrey Roberts
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1982-03-23
Filing date: 1983-03-18
Publication date: 1991-09-25
Anticipated expiration: 2003-03-18
Also published as: FR2665485A1; GB2242172B; DE3341126C2; DE3341126A1

Description

GAS TURBINE ENGINE This invention relates to variable cycle gas turbine engines and in particular to variable cycle gas turbine engines having vectorable outlets to give an aircraft fitted with such an engine a capability for vertical flight or short take-off and landing (V/STOL).

Variable cycle engines are generally disclosed in U.S. Patents 3,913,321 or 4038818 (assigned to Rolls-Royce Limited). In general, these engines comprise a first axial flow compressor and a core engine comprising, in flow series, a second compressor combustion equipment, and turbines to drive the first and second compressors. The engines are capable of operating in two distinct modes. These modes are namely a "series flow" mode and a "parallel flow" mode, In the series flow mode the first and second compressors are connected in flow series and the whole output flow of the first compressor supercharges the second compressor.In the "parallel flow" mode the output flow of the first compressor is prevented from supercharging the second compressor and is discharged to ambient air through either a by-pass duct or through fixed or vectorable discharge nozzles, and simultaneously an auxilliary air r intake is opened up to allow air to enter the second compressor.

Variable cycle engines of the type described above offer many advantages, particularly for air- craft requiring vertical take-off and landing and also supersonic forward flight capabilities. The engine performance can be optimised for vertical take-off and landing and subsonic flight during the parallel flow mode and optimised for forward supersonic flight during the series flow mode. In this way, for vertical flight, the well proven advantages of engines such as the Rolls-Royce Limited Pegasus engine (used to power the British Aerospace AV8A Harrier or the British Aerospace/McDonnell Douglas AV & ) can be exploited whilst enabling efficient use of the engine in the series flow mode for supersonic flight.

To enable the output flow from the first compressor to be redirected selectively for series or parallel modes of operation, it is usual to provide a diverter valve downstream of the first compressor but upstream of the second compressor.

Examples of such diverter valves are described in the above mentioned patents. The problems associated with these known diverter valves reside in their complexity, weight, cost and disruptive effect on the thermodynamic cycle of the engine during transition from the series flow mode to the parallel flow mode. In general, the disruption to flow is due to the sio speed of the operation of the diverter valve flaps and doors and the fact that many movable parts such as doors and flaps are positioned in the airflow path and have to be operated in unison.

An object of the present invention is to provide a variable cycle engine of the type described above with a diverter valve means which is simple to operate, is litwelght, relatively inexpensive and provides a relatively uncluttered flow path during both odes of operation.

The invention, as claimed, provides a substantially unrestricted flow path from the first compressor to the nozzles or to the inlet of the second compressor and utilizes the vectorable nozzles as a diverter valve.

This eliminates the provision of additional heavy and costly diverter valves.

Figures 1 and 2 show, respectively, a part sectioned plan view and side elevation of an engine constructed in accordance with the present invention, showing the engine in one configuration.

Figures 3 and 4 show the engine of Figures 1 and 2 in a second configuration.

Figure 5 illustrates schematically a known mechanism for varying the angles of a number of rows of stator vanes of one of the compressors of the engine of Figures 1 to 4.

Figure 6 illustrates schematically the mechanism for varying the outlet area of the front nozzles of the engine of Figures 1 to 4.

Figure 7 illustrates a modified air intake for the engine of Figure 1 which enables the aircraft to be flown in forward flight without supercharging the core engine.

Referring firstly to Figures 1 to 4 generally, a gas turbine engine 2 for powering an aircraft to give vertical or short take-off and landing capability, comprises a first compressor 4, a second compressor 6, a combustor 8 and a turbine 10 for driving the compressors. Airflow enters the first compressor 4 from atmosphere and exhausts through two vectorable nozzles 12,14. Airflow enters the second compressor 6 through two intakes 16,18 arranged (as will be described in greater detail hereafter) adjacent the vectorable nozzles 12,14 and exhausts via the combustor 8, and turbine 10 through two vectorable nozzles 20,22.

Referring now particularly to Figures 1 and 2, when the engine is required to produce vertical thrust on the aircraft, i.e. a thrust normal to the axis of the engine, the nozzles 12,14 and 20,22 are vectored downwards, as shown. In this configuration the airflow into the first compressor 4 is from atmosphere at the front of the engine and te discharge from the first compressor exhausts to atmosphere to produce vertical thrust via the nozzles 12,14. The airflow into the second compressor 6 is from atmosphere via the intakes 16,18 and the discharge from the second compressor exhausts to atmosphere via the combustion 8 and turbine 10 to produce vertical thrust via the nozzles 20,22.In this way the compressors 4 and 6 work in parallel producing a large mass flow at low. speed, suitable for take-off.

Referring now particularly to Figures 3 and 4, when the engine is required to produce forward trust on the aircraft, i.e. thrust parallel to the axis of the engine, the nozzles 12,14 are vectored rearwards, as shown. The intakes 16,18 are arranged relative to the nozzles 12,14 so that in this configuration each of the nozzles 12,14 registers with a respective one of the intakes 16,18. In this configuratIon the intake to the first compressor 4 is fro atmostrere at the cront on the engine and the discharge from t ne first compressor exhausts via the nozzles ~2, Into the intakes 16,18 to constitute the intake to the second compressor 6. The discharge from the second compressor 6 exhausts via the chamber 8 and turbine 10 to produce forward thrust via the nozzles 20,22. In this way the compressors 4 and 6 work in series, producing a smaller mass flow at high speed than when working in parallel, as in Figures 1 and 2, making this configuration suitable for high-speed forward flight.

In order to enable good transfer of flow from the nozzles 12,11; to the intakes 16,18 when the compressors are working in series, a resilient seal 24,26 is provided at the lip of each intake, The seals 24,26 are inflatable and are inflated, as shown in Figures 2 and 4 to seal between each intake and its respective nozzle when the nozzle is in registry with the intake.

It is necessary to reduce the exit area of the nozzles 12,1l;hen they are moved out of register with the intakes 16,18 to increase the pressure drop across the, and optirnise their function in this configuration.

Each nozzle 12,14 is provided with a set of variable exit guide vanes 28,30 (see Figure 6). The setting of the exit guide vanes 28,30 is controlled by a cam 32(a), cam track 32(b), bell crank 33, levers 35 and push rods 34. The nozzle exit guide vanes 28,30 restrict the nozzle exit area as soon as the nozzles move out of register with the intakes 16,18.

During transition from the rearwards position to the downwards position, the vanes 28,30 may be angled to turn the flow discharging from the nozzle rearwards to give the discharge a substantial rearward component.

It will be appreciated that the space formed between the first compressor 4 and the second compressor may be used to site bulky ancillary engine components such as oil tank 36.

It will be appreciated that the core engine, that is to say the compressor 6, combustor 8 and turbines 10, have to be designed to operate with widely different flow conditions.

In one mode of operation (the "serIes flow" mode) the inlet air to the compressor 6 is supercharged by the compressor 4. In the "parallel flow" mode the air is not supercharged. To maximise thrust it is necessary to provide a higher pressure ratio compressor and avoid the compressor stallin.

Referring to Figure 5, the compressor 6 may be provided with variable stator vanes 38, each of which is pivotally mounted at its outer end on a spigot 40.

Levers 42 are provided to connect each spigot 40 to a unison ring 44. There is provided one Fso ri for each row of stator vanes. A bear. 46 is provided which is mounted on a double pivot 48 of the type described in our British Patent No. 1,511,723. The beam 46 has links 50 mounted on universal connections 52 and the links 50 are connected to the unison rings 44 by universal connections 54.A motor 56, which drives a lead screw 56(a) which engages a nut 5 on the otherwise free end of the beam 46, is provided for moving the beam 46 about the axes 60,64 and, thereby, rotate the unison rings 44. The vanes 38 are connected to the unison rins by levers 66 and universal connectors 68. Rotation of the rings 44 causes simultaneous rotation of all the stator vanes 38 in each row.

This mechanism is fully described in our British Patent No. 1,511,723.

Additionally or alternatively, the turbines may be provided with variable position guide vanes 70 to control the flow through the turbines.

Additionally or alternatively, the flow through the core engine may be controlled by means of variable area nozzles. There are many well known ways of varying the area of the nozzles. One may employ an axisymmetric array of petals (not shown) or a two dimensional nozzle which employs one or more movable plates 72 which can be moved to alter the area of the nozzle by means of a motor 73 which drives a screw jack 74 by way of flexible drive shafts 76 and gearbox 77.

Additionally or alternatively, the inlet guide vanes 78 of the compressor 4 may be of variable geometry, The intakes 16 may be repositioned as shown in Figure 7 so that the nozzles 12,14 have to be rotated to a slightly upwards position when it is required for them to register with the intakes 16,18. Subsonic cruise can be obtained, without superchargin the core engine, by rotating the nozzles 12,li to a position where they discharge rearwards. If necessary, the exit guide vanes in the nozzles 12,lit may be angled to turn the effLux rearwards to enhance the forward thrust. In all positions of the nozzles 12,14 up until they register with the intakes 16,15, the exit areas are reduced.

In yet a further embodiment, (not shown) the air intakes 16 may be designed to allow for a "by-pass" mode of operation in which, when the nozzles 12,14 are directed rearward (but not fully mating with the intakes 16), part of the efflux of the nozzles 12,14 is discharged into the core engine and part is discharged rearwardly but not through the core engine. In this case, an obturator flap (not shown) would be provided at the lip of each intake. The obturator flap would operate to close down each air intake 16 to an area equal to the area of that part of each nozzle 12,14 directly communicating with the interior of the air intake.

Claims

1. A gas turbine engine comprising: a first compressor; a second compressor; a combustor; turbine means for driving the first and second compressors; first intake means for a flow to enter the first compressor; first outlet means for discharging the flow from the first compressor; second intake means for a flow to enter the second compressor, and second outlet means for discharging the flow from the second compressor via the combustor and the turbine means, wherein the first outlet means is vectorable between a first position in which the flow from the first compressor discharges to ambient and a second position in which the first outlet means registers with the second intake means and the flow from the first compressor discharges into the second compressor.

2. A gas turbine engine according to Claim 1 wherein resilient seal means is provided to seal between the first outlet means and the second intake means when the first outlet means is in the second position.

3. A gas turbine engine according to Claim 2 wherein the resilient seal means is inflatable.

4. A gas turbine engine according to Claim 1 in which the optimum flow of working medium through the engine is that when the first outlet means is in the second position and the first and second compressors are each running at a predetermined speed, wherein flow control means are provided for controlling the flow of the working medium through the first compressor, the second compressor and the turbine means to prevent the first and second compressors stall during operation at or below said predetermined speeds.

5. A gas turbine engine according to Claim 4 wherein the control means comprises stator vanes, the angle of attack of which is variable to vary the flow through the engine.

6. A gas turbine engine according to Claim 4 wherein the control means comprises means for varying the exit area of the first outlet means so as to produce a smaller exit area when the first outlet means is in any position other than the second position.

7. A gas turbine engine according to Claim 4 wherein the control means comprises means for varying the exit area of the second outlet means.

8. A gas turbine engine according to Claim 1 wherein the first outlet means is movable to a third position, intermediate the first and second positions, where the first outlet means discharges partly into the second air intake means and also rearwards alongside the second air intake means, and obturator means are provided to restrict the area of the second air intake means to an area equal to that part of the first outlet means which discharges into the second air intake means.

9. A gas turbine engine according to Claim 1 wherein the first outlet means is movable to a third position, intermediate the first and second positions, where the first outlet means discharges rearwards alongside the second air intake means but does not discharge into the second air intake means.

10. A gas turbine engine according to Claim 1 wherein the first outlet means is provided with a plurality of flow deflector vanes extending across the flow path through the first outlet means, and the vanes are movable to deflect an efflux of gas from the first outlet means rearwards during transition of the first outlet means from the second position to the first position,