Valve for diverting fluid flows in turbomachines

A valve for selectively changing the direction of flow of working fluid through a variable cycle engine which comprises a first and second compressor 14,16 spaced along a flow duct 26. The duct 26 having air intakes 42 leading to the second compressor 16 and outlets leading to nozzles 44. The valve comprising a sleeve 30 axially movable along the duct. The sleeve 30 having openings 31 in it in which are located doors 32. Links 36 are connected to each of the doors 32 so that as the sleeve 30 is moved axially the doors 32 are pulled open to open the air inlets 42 and the outlets and simultaneously obturate the flow duct 26. In a second position of the sleeve 30 the doors 32 and the sleeve 30 close off the air inlets and outlets and open the duct to allow the first compressor to supercharge the second compressor.

This invention relates to variable cycle gas turbine engines, and in 
particular to valves for selectively changing the direction of flow of the 
working fluid through the engine. 
The present invention is particularly concerned with variable cycle engines 
of the type generally disclosed in U.S. Pat. Nos. 3,913,321 or 4,038,818 
(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, and the engine is capable of operating in two 
distinct modes. These modes are namely a "series flow.revreaction. 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 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 aircraft requiring vertical take-off and leanding 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 AV8B) 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 prior 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. It is very difficult to 
achieve transition from the series flow mode to the parallel flow mode 
without inducing variations in flow conditions at the inlet to the second 
compressor. These variations may be localised and vary circumferentially 
around the fluid flow annulus of the second compressor and have a 
deleterious effect on the performance of the engine. In general some of 
these problems are due to the speed of the operation of the diverter valve 
means 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 lightweight, relatively inexpensive and provides a relatively 
uncluttered flow path during both modes of operation.

Referring to FIGS. 1 and 2, there is shown schematically a generally round 
nacelle 10 which houses a gas turbine engine 12 comprising, in flow 
series, a first axial flow compressor 14 and a core engine. The core 
engine comprises a second axial flow compressor 16, combustion means 18, 
turbine means 20 connected to the compressor means 14 and 16 to drive the 
compressor means, a jet pipe 22 and afterburning equipment 24. 
The first and second axial flow compressors are spaced axially from each 
other by a duct 26 which houses a diverted valve 28. The diverter valve 28 
consists of a hollow tubular sleeve 20 which is free to move axially along 
the duct 26. A plurality of doors 32 are circumferentially spaced around 
the engine axis inside the duct 26. One of the ends of each door is 
pivotally mounted on the sleeve 30 at hinge 34. A plurality of telescopic 
links 36 are connected at one of their ends to a tube 37 which covers the 
shaft 130 that connects the front compressor 14 to the turbine means 20. 
The other end of each link 36 is connected to one of the doors at a point 
which is adjacent to the centre of pressure of the gases acting on each 
door, when the doors are in the position shown in FIG. 2. The doors 32 are 
shaped and positioned such that in a first position of the sleeve 30 they 
lie along the length of the duct 26 and obturate the openings 31 in the 
sleeve 30. In this position the sleeve 30, together with the doors 32, 
provide a flow path from the first axial flow compressor 14 to the second 
axial flow compressor 16, as shown in FIG. 1. This position of the sleeve 
30 enables the engine to operate in a "series flow" mode of operation. 
FIG. 2 shows the engine in the "parallel flow" mode. In this mode the 
sleeve 30 is translated axially along the duct from the position shown in 
FIG. 1. The telescopic links 36 are caused to pivot about their pivotal 
attachment to the tube 37 and cause the blocker doors 32 to be swung 
inwards until each door co-operates with its adjacent door and with the 
tube 37 to obturate the duct 26. The doors 32 thereby uncover the openings 
31 in the sleeve 30 and thereby open auxilliary air intakes 42. 
Translation of the sleeve 30 also uncovers the outlet to the front nozzles 
44. This provides a flow path from the first axial flow compressor 14 to 
the front nozzles 44 and also provides a flow path whereby ambient air may 
flow into the auxilliary air intake 42 and into the inlet of the second 
compressor 16. It should be noted the nozzles 44 shown in the schematic 
views of FIGS. 1 and 2 are out of position relative to the positions of 
the air intakes 42. However, by reference to FIG. 3, the relative 
positions of the axially upstream nozzles 44 to the air intakes 42 shows 
that the same are angularly spaced therefrom. 
Referring to FIG. 3 there is shown a vertical cross-sectional view of an 
engine similar to that of FIG. 2 taken along the staggered line III--III. 
The left hand side of FIG. 3 shows the doors 32 in the "parallel flow" 
position of FIG. 2; the right hand side of FIG. 3 shows the engine in the 
"series flow" mode of FIG. 1. The aircraft fuselage 46 is effectively a 
double skin construction. The air intakes 42 are provided with shutters 48 
which blow inwards when the sleeve 30 is moved to unblock them. The air 
intakes 42 open into a plenum chamber defined by a wall 49 which conducts 
the air from the intakes 42 into the second compressor 16 downstream of 
the blocker doors 32. The wall 49 houses tangential rollers 50 which are 
arranged along the length of the duct 26 to support the sleeve 30 as it is 
translated axially. 
The front nozzles 44, which in FIGS. 1 and 2 are shown schematically as 
projecting into the free air stream, are in fact housed within the double 
skinned fuselage 46 as shown in FIG. 3 so as to reduce drag and provide a 
streamlined air frame for fast forward flight. In a preferred embodiment, 
the nozzles 44 take the form of a plurality of vanes 52 pivotally mounted 
in a ring 54 supported in a bearing 56. The nozzles are capable of 
rotation in the bearing by means of an actuator 57, and all the vanes 52 
of each nozzle are linked together by a rod 58 which is moved by an 
actuator 60 (shown schematically). All the vanes of each nozzle are moved 
in unison for the purpose of varying the outlet area of the nozzles and by 
rotating the nozzles in their bearings 56 the thrust produced by the 
nozzles can be vectored thereby selectively to produce upward thrust (for 
VTOL) or forward thrust for forward flight. During the series mode of 
operation the vanes 52 close off the nozzles 44 to provide a streamlined 
fuselage. 
Upstream of the nozzles 44 there is provided a thrust augmentor in the form 
of a plurality of reheat gutters 62. In this way additional fuel can be 
burnt upstream of the nozzles to increase the thrust produced by the 
nozzles and the nozzle outlet area increased by opening the vanes 52 to 
cope with the increased flow. 
Referring to FIG. 4 the engine is almost identical to that shown in FIGS. 1 
and 2 except that the links 36 for actuating the doors 32 are provided 
outside of the sleeve 30. Furthermore the links 36, in this modification 
are not carried by the tube 37 that covers the shaft 130. Instead the 
doors 32 are opened and closed by means of telescopic links 36 housed in a 
recess which allows the links to be extended in the series mode to pull 
the door outwards, and compressed in the parallel mode to push the doors 
inwards. In other respects the construction of the engine is identical to 
that of FIGS. 1 to 3.