GB2263505A

GB2263505A - Shielded vectored thrust gas turbine engines

Info

Publication number: GB2263505A
Application number: GB8035103A
Authority: GB
Inventors: Roger Hurd
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1980-10-31
Filing date: 1980-10-31
Publication date: 1993-07-28
Anticipated expiration: 2000-10-31
Also published as: GB2263505B

Abstract

A gas turbine aero engine having vectorable hot and cold nozzles (23, 21) is provided with ducting (24) which directs the cold air ejected from the cold nozzles (23) when they are directed rearwardly over the hot nozzles (21). The ducting (24) is extended rearwardly of the hot nozzles (23) and serves to suppress infra red radiation from the hot gas plume by constituting a barrier and by causing the cold gases to mix with the hot gases. <IMAGE>

Description

TITLE INFRA RED REDUCTION DESCRIPTION This invention relates to the suppression of infra red radiation emanating from the plume of the hot exhaust nozzles of a gas turbine powered aircraft. The invention is particularly concerned with suppressing the infra red radiation from the plume of vectored nozzle engines such as Rolls-Royce Limited's Pegasus engine fitted to the Quarrier or AV8A or AVBB aircraft.

Vectored thrust aircraft propulsion systems pose problems with regard to infra-red radiation and to uppression because of the unmixed exhaust arrangement.

The rear or core nozzle emits a hot jet plume independent y of the cold front fan exhaust. Consequently, when directed rearwards there is no in-built mechanism for cooling the core jet to attenuate its infra-red emissions, AS in the case of conventional mixed exhaust military turbo-fans).

Tt is an object of the present invention to provide -as turbine engines of the type having vectorable hot and cold nozzles with means for reducing the infra-red radiation from the hot gases ejected from the hot nozzles.

According to the present invention there is provided a gas turbIne engine of the type having one or more vectorable cold nozzles supplied with air from a compressor of the engine and one or more vectorable hot nozzles supplied with the efflux of hot gases from a @rbine of the en;;- ne, the engine being provided with c'uc+ing for receiving air issuing from the cold nozzle nozzles when the cold nozzle or nozzles are directed r1 a predetermided direction and for conveying the ar .rer the hot nozzle or nozzles to cause the air from the cold nozzle or nozzles to inter mix with the o iisess discharged from the hot nozzle or nozzles.

A cornbustor means may be provided between a compressor and the cold nozzle or nozzles in which case an obturator means is provided to close off the ducting and redirect the gases issuing from the cold nozzles away from the hot nozzle or nozzles when the combustor means is operated and the cold nozzle or nozzles are directed in the predetermined direction.

The, or each, hot nozzle may be defined by a wall which is fluxed to provide interdigitated channels for conveying, alternately around the periphery of a discharge opening of the or each hot nozzle, cold air frorn the cold nozzles from outside of the nozzle and hot gases from within the nozzle, thereby to cause intermixing of the hot gases and the cold air downstream of the hot nozzle or nozzles.

A combustor means may be located in the ducting downstream of the hot nozzle or nozzles and the ducting provided with a variable area nozzle downstream of the hot nozzle or nozzles.

In a further aspect of the invention there is provided an aircraft Incorporating an engine constructed in accordance with the present invention.

The present invention will now be described by way of an example with reference to the accompanying drawing in which: Figure 1 illustrates schematically an aircraft incorporating an engine in accordance with the present invention.

Figure 2 illustrates a plan view of one engine for the aircraft of Figure 1.

Figure 3 illustrates a plan vie of an alternative engine ta b+ shove n figure 2.

Figure; 4 and 5 illustrate modifications to the ducting of the engines of Figures 2 and 3.

Referring to the drawings, Figure 1 shows an aircraft 10 having a gas turbine engine 11 of the bypass type. Compressed air from a compressor of the engine is supplied to two vectorable cold nozzles 21 (one on each side of the aircraft) and the hot efflux from a turbine the engine is supplied to two vectorable hot nozzles 23 located af of the cold nozzles (there being one hot nozzle each side of the aircraft). Further details of :ne engine 11 are shown in Figure 2.

Referring to Figures 1 and 2 the engine 11 comprises 4wo contra rotating spools defining, in flow series, a multi stage axial flow low pressure compressor 14, a gh pressure cointressor 15, a main combustor 16, a high pressure turbine 17, a low pressure turbine 18 an a jet @ipe 19.

Compressed air from the L.P. compressor 14 is fed to the H.R. compressor 15 and to a bifurcated by-pass @lenum chamber 20 which terminates at each side of the angine in a vectorable cold nozzle 21. The nozzles 21 are one in bearings 22 so as to be capable of rotating to and from a position directed slightly forward, through 2 vertically downward position, to a position where they point rearwards.

Similarly, the bIfurcated net pipe 19 terminates each side of the engine in a vectorable nozzle 23 which is movable in unison with the cold nozzles 21 in the same directions' as the cold nozzles 21. The nozzles 21 and 23 may be movable independently to trim the aircraft.

The engine 11 is provided at each side with ducting 24. The ducting 24 has an inlet opening positioned in the plane of the exit face of the cold nozzles 21 to receive cold air ejected through the nozzles 21 when they are directed rearwardly. The ducting 24 is shaped and positioned to convey this cold air rearwardly over the outside of the hot nozzles 23 to cause the cold air to mix with the hot efflux of the hot nozzles 23 downstream of the nozzle opening. The acting 24 is extended rearwardly of the exit plane of nozzles 23.

The nozzles 23 project into the ducting 24 and the ducting 24 is provided with an opening that allows the nozzles 23 to be rotated to a vertical position (for vertical thrust) or slightly forwards (to provide reverse thrust).

Referring to Figure 3 there is shown an engine similar to that shown in Figure 2 except that the at pipe is not bifurcated and there is a single vector ble hot nozzle 23 instead of two. Here again the engine is provided. with ducting positioned to receive cold air ajected frow the cold nozzles 21 and t direct the cold air over the outside of the hot nozzle 23 causing it to intermix with the hot efflux of the hot nozzle 23 down tream of the nozzle 23.

The ducting is provided with a single opening in its lower wall to enable the nozzle 23 to be swung through the vertical position to S slightly forward facing position.

With the engines of Figures 2 to 5, if desired, an additional combustor 25 may be provide in the ducting c'ownstream of the hot nozzle or nozzles 23 to provide a re-neat faci*y and in this case, a variable area 'j"zle 25 is provided at the outlet of the ducting.

It is to be understood that when re-heat is employed the benefit of the ducting shielding the infra red emission is lost.

If desired, additional combustors 26 (shown dotted) may be provided in the plenum chambers 20 at a location between the L.P. compressor 14 and the cold nozzles to crease the thrust from the cold nozzles. In this case : wou'd be necessary to provide an obturating device ;hich prevents the efflux of the nozzles 21 flowing over the nozzles 23 when they are pointed rearwards, and a variable area nozzle outlet for the cold nozzles 21.

Such an obturating device may be formed by flaps 27,37 ire the outermost side walls of the ducting as shown fn Figure 4. The flap 27 is moveable to close off the doting when the nozzles 21 are directed rearwards and plenum chamber burning is operative. The efflux of -he nozzles 2' is thus deflected away from the aircraft ~rl a rearward direction.

Referring now to Figure 4 the portion 28 of the diicting immediately adjacent the hot nozzles 23 may be oveable with the nozzle andhave flaps 29,30 in its lowermost wall immediately downstream of the nozzles 23.

@@e flaps 29,3 allow the efflux of gases from nozzles 23 to pass through the wall of the ducting during rotation the nozzles 23 from the horizontal to the vertical positions. In this way abrasion of the ducting 24 would id reduced as would undesirable thrust and loading effects.

To improve mixing of the hot and cold strearns, the @ot nozzles 23 are preferably defined by a wall 32 (shown in greater detail in the circle in Figure 2) @nich is fluted to provide interdigitated channels 33,3: lternately around the periphery of the discharge @pening of each hot nozzle 23. The cold air fror the aj'd nozzles 21 flows over the outside of the nozzles 23 long channels 33 whereas hot gas from within the nozzles 34 flow along channels 34 to intermix the hot and cold gases. The hot and cold nozzles 21,23 are scarfed to improve matingof the nozzles and the ducting 24.

Claims

1. A gas turbine engine of the type having one or more vectorable cold nozzles supplied with air from a compressor of the engine and one or more vectorable hot nozzles supplied with the efflux of hot gases from a turbine of the engine, the engine being provided with ducting for receiving air issuing from the cold nozzle or nozzles when the cold nozzle or nozzles are directed ln a predetermined direction and for conveying the air over the hot nozzle or nozzles to cause the air from the cod nozzle or nozzles to inter mix with the hot gases discharged from the hot nozzle or nozzles.

2. An engine according to claim 1 wherein a combustor weans is provided between a compressor and the cold nozzle or nozzles and an obturator means is provided to close off the ducting and redirect the gases issuing from the cold nozzles away from the hot nozzle or nozzles :tien the combustor means is operated and the cold nozzle r nozzles are directed in the predetermined direction.

An An engine according to claim 1 or claim 2 wherein the or each hot nozzle is defined by a wall which is fLuted to provide interdigitated channels for conveying, Iternately around the periphery of a discharge opening WT the or each hot nozzle, cold air from the cold nozzles rom outside of the nozzle and hot gases from witr1in the nozzle thereby to cause intermixing of the hot gases an7 the cold air downstream of the hot nozzle or nozzles.

An engine according to claim 1 or claim 2 wheren sombustor means are located in the ducting downstream of - 2 hot nozzle or nozzles and the ducting is provided with a variable area nozzle downstream of the hot nozzle or nozzles.

5. An engine substantially as herein described with reference to the accompanying drawings.

6. An aircraft incorporating an engine according to any one of claims 1 to 5.