GB2379483A

GB2379483A - Augmented gas turbine propulsion system

Info

Publication number: GB2379483A
Application number: GB0121763A
Authority: GB
Inventors: Michael John Provost; Peter Beardsley
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2001-09-08
Filing date: 2001-09-08
Publication date: 2003-03-12
Also published as: GB0121763D0

Abstract

The system 50 comprises a main engine 52 and either an auxiliary engine 54, axial flow fans (90, fig 4) or circumferential flow fans (100, fig 5) for injecting pressurised fluid into the main engine 52. The auxiliary engine 54 is connected to the main engine 52 by a duct 68 and the pressurised fluid from engine 54 enters either a bypass duct 34 via a manifold 70, or a core exhaust nozzle (fig 3) via a manifold (78, fig 3), and exhausts through a variable area nozzle (69, fig 3) or 74. The auxiliary engine 54 may be in fluid communication with two or more main engines 52, or two or more auxiliary engines 54 may be in fluid communication with one main engine 52. The axial flow fans or circumferential flow fans may be driven by electric , hydraulic or diesel motors, or by the main engine and are in flow communication with a bypass duct.

Description

A NOVEL PROPULSION SYSTEM FOR AN AIRCRAFT The present invention relates to an arrangement of a propulsion system for an aircraft incorporating a gas turbine engine and a method of operating the same.

For a gas turbine engine, in use and at a given aircraft flight speed, a specified net thrust can be achieved by either accelerating a small inlet air mass flow to a high exhaust velocity or by accelerating a large mass flow to a lower exhaust velocity. The former requires an engine with a small frontal area, which beneficially results in a low-drag engine installation but produces high jet noise, whereas the latter results in an engine which beneficially produces low jet noise but has a large frontal area, which results in a high-drag engine installation.

For some aircraft, which require low noise at some flight conditions for environmental acceptability and low drag at other flight conditions for efficient operation, a propulsion system is required to vary the exhaust gas velocity produced by the engine to values other than those that would be achieved with fixed engine geometry. There is much prior art involving so-called'variable cycle' engines that achieve this aim. However, these engines are generally extremely complex mechanically and have not yet achieved widespread acceptance.

W091/18199 discloses a gas turbine engine suitable for supersonic aircraft having supplementary air intakes which may be opened to admit ambient air into the bypass duct. The ambient air and bypass air mix with the core flow and pass though a common exhaust nozzle in an attempt to reduce the exhaust jet velocity and thus reduce its noise generating capacity. This arrangement is limited in that the amount of ambient air admitted to the bypass duct is restricted in accordance with the pressure differences between ambient and bypass gas flows.

It is therefore an object of the present invention to provide a propulsion system comprising a main gas turbine engine which provides the means to inject pressurised gases into the main engine to increase thrust output and/or reduce exhaust jet noise and/or reduce the diameter of the main engine for a given thrust requirement.

According to the present invention there is provided a propulsion system for an aircraft comprising a main engine, an exhaust nozzle assembly and a means for injecting pressurised fluid into said propulsion system wherein the means for injecting pressurised fluid is in fluid communication with the main engine so as to direct pressurised fluid into said propulsion system at a position upstream of the exhaust nozzle assembly.

According to a further aspect of the present invention, a method of operating such a propulsion system comprises the steps of: during aircraft take-off and climb, operating the main engine at a high output level and the means for injecting pressurised fluid at a high output level so that the nozzle flow of the main engine is augmented, whilst maintaining a variable area nozzle in its greatest area position; during aircraft cruise, operating the main engine at a reduced output level and ceasing operation of the means for injecting pressurised fluid, whilst maintaining a variable area nozzle in its smallest area position.

The present invention will now be described by way of example only with reference to the following figures in which: Figure 1 is a schematic section of a ducted fan gas turbine engine; Figure 2 is a schematic section of a first arrangement of a propulsion system for an aircraft in accordance with the present invention;

Figure 3 is a schematic section of a second arrangement of a propulsion system for an aircraft in accordance with the present invention; Figure 4 is a schematic section of a third arrangement of a propulsion system for an aircraft in accordance with the present invention; Figure 5 is a schematic section of a fourth arrangement of a propulsion system for an aircraft in accordance with the present invention; Figure 6 is a schematic isometric view of an aircraft incorporating the present invention.

With reference to figure 1 a ducted fan gas turbine engine 10 comprises, in axial flow series an air intake 12, a propulsive fan 14, a core engine 18 and a core exhaust nozzle assembly 20 all disposed about a central engine axis 22 and surrounded by an outer and generally annular bypass wall 17. The core engine 18 comprises, in axial flow series, a series of compressors 24, a combustor 26, and a series of turbines 28. The series of turbines 28 are drivingly connected to the compressors 24 and propulsive fan 14. The direction of airflow through the engine 10 in operation is shown by arrow A. Air is drawn in through the air intake 12 and is compressed and accelerated by the fan 14. The air from the fan 14 is split between a core engine flow and a bypass flow. The core engine flow passes through an annular array of stator vanes 30 and enters the core engine 18, flows through the core engine compressors 24 where it is further compressed, and into the combustor 26 where it is mixed with fuel which is supplied to, and burnt within the combustor 26. Combustion of the fuel mixed with the compressed air from the compressors 24 generates a high energy and velocity gas stream which exits the combustor 26 and flows downstream through the turbines 28. As the high energy gas stream flows through the turbines 28 it rotates turbine rotors extracting energy from the gas stream which is used to drive the fan 14 and

compressors 24 via engine shafts 32 which drivingly connect the turbine 28 rotors with the compressors 24 and fan 14. Having flowed through the turbines 28 the high energy gas stream from the combustor 26 still has a significant amount of energy and velocity and it is exhausted, as a core exhaust stream, through the core engine exhaust nozzle assembly 20 to provide propulsive thrust. The remainder of the air from, and accelerated by, the fan 14 flows within a bypass duct 34 around the core engine 18. This bypass air flow, which has been accelerated by the fan 14, flows to the exit nozzle assembly 16 where it is exhausted, as a bypass exhaust stream to provide further, and in fact the majority of, the useful propulsive thrust. The fan 14 comprises an annular array of fan blades 36 which are retained by a fan disc 38.

In order to meet engine exhaust noise requirements it is common practice to increase the fan 14 diameter to allow an increased flow area, which for a given thrust produced by the engine results in a reduced exhaust gas stream velocity and subsequently exhaust noise is less. This common practice is suitable where the size of the engine is not restrictive on aircraft performance. However, where the drag is critical to the performance of the aircraft, it is desirable to minimise the diameter of the engines. For a required engine thrust output level there is therefore a compromise between having a large diameter engine with subsequently high aerodynamic drag but lower exhaust noise and a smaller diameter engine with lower aerodynamic drag but higher exhaust noise. The requirement for quieter engines is principally during take-off and climb, when the engine is working at or near to its maximum output and near centres of population.

It is the object of the present invention to provide a gas turbine propulsive system for an aircraft which has both a reduced diameter and low noise for a relatively high thrust level.

With reference to figure 2 which shows a first embodiment of a propulsion system 50 in accordance with the present invention. The propulsion system 50 comprises a main gas turbine engine 52 and an auxiliary engine 54. The main engine 52 comprises a general configuration similar to the engine shown and described with reference to figure 1.

The auxiliary engine 54 comprises a rotational axis 60, about which an air intake 56, a compressor 58, a combustor 62 and a turbine 64 are disposed in axial flow series. A shaft 66 drivingly connects the turbine 64 to the compressor 60. The auxiliary engine 54 works in conventional manner as known in the art.

The propulsion system 50 is arranged so that the auxiliary engine 54 is connected via a duct 68 to a manifold 70 substantially surrounding and in fluid communication with the bypass duct 34. The outer bypass wall 17 comprises an annular array of apertures 72 and together with the manifold 70 distribute fluid exhausted, in use, from the auxiliary engine 54. The propulsion system 50 further comprises a variable area nozzle 74 disposed on to the downstream periphery of the nozzle assembly 16.

During aircraft take off and climb manoeuvres the main engine 52 is producing a high level of thrust, which is boosted by the gas flow produced by the auxiliary engine 54 injected into the main engine 52. During aircraft cruise, which is the majority of a normal flight cycle, the auxiliary engine 54 is switched off and the main engine 52 solely powers the aircraft. Optionally, valves (not shown) may be provided to close the apertures 72 during this phase of engine operation. For this embodiment the main engine 52 is specifically designed for efficiency at the cruise condition and the auxiliary engine 54 is specifically designed to boost thrust output or to assist in the

rec-.'c-cion of the noise produced by the main engine at a given thrust relative to a conventional gas turbine engine.

The noise is generated by the exhaust gases mixing with the ambient air. Exhaust noise increases relative to an increase in velocity of the exhaust gas stream and it is an aspect of the present invention, that will become apparent, to reduce the noise by reducing the velocity of the exhaust gas stream.

This embodiment is reliant on the operation of the variable area nozzle 74. As can be seen in Figure 2 the variable area nozzle 74 is in a first position indicated by the solid line marked 74. During cruise flight of the aircraft, the nozzle 74 is disposed in this first position where it converts the pressure of the gas flow, produced by the engine, into propulsive thrust. During take-off and climb where the auxiliary engine 54 is delivering its exhaust gas into the main engine 52, the nozzle shown by the dashed lines 74'is disposed in a second and open position, in order to accept the increased gas flow relative that produced by the main engine 52 working alone.

Referring now to Figure 3, the auxiliary engine 54 is arranged to be in fluid communication with the core exhaust nozzle 20. The propulsion system 50 is arranged so that the auxiliary engine 54 is connected via a duct 68 to a manifold 778 substantially surrounding and in fluid communication with the core nozzle 20. The core nozzle 20 defines an annular array of apertures 80 and together with the manifold 78 distribute fluid exhausted, in use, from the auxiliary engine 54 into the core nozzle 20. The propulsion system 50 further comprises a shortened by-pass duct 34 which terminates in a nozzle 74 which is fixed, but otherwise similar to that shown in Figure 2. However, a variable area nozzle 69 positioned at the downstream periphery of the nozzle assembly 16 which works in the same general manner as the variable exhaust nozzle 74 described with reference to Figure 2. Thus the nozzle 69 varies between the positions shown in solid and interrupted lines in Figure 3.

Furthermore the embodiments described with reference to Figures 2 and 3 may be combined so that one auxiliary engine 54 is in fluid communication with both the bypass duct 34 and core nozzle 20. It should be apparent that an auxiliary engine 54 may supply pressurised fluid to the bypass duct 34 and a further auxiliary engine 54 may supply pressurised fluid to the core nozzle 20. A single auxiliary engine 54 may also supply pressurised fluid to two or more main engines 52 and two or more auxiliary engines 54 may supply pressurised fluid to a single main engine 52.

Referring to Figure 4, a propulsion system 50 comprises a gas turbine engine 10 which is of a generally conventional configuration and operates in the same manner as that already described with reference to Figure 1 and so common reference numerals denote the same features. The propulsion system 50 further comprises a means for injecting a pressurised fluid 90 into the bypass duct 34 of the main engine 52. In this embodiment of the present invention the means for injecting a pressurised fluid 90 is an axial flow fan assembly 90 which comprises an axial fan 92, which rotates about a central axis 94 and is powered by a motor 96. The means for injecting pressurised fluid 90 is disposed to the bypass wall 17 so that gas flow B is in fluid communication with the bypass duct 34 and joins the bypass gas flow A. In this way the fluid flow through the main engine 10 is augmented. The motor 90 may be any one cf a group of motors comprising electric, diesel, hydraulic actuation means and the motor may be replaced by a mechanical drive. In this embodiment an electric motor is utilised. The electrical supply to the motor may be from any suitable source, however, a preferred supply is described and incorporated herein by reference to a recent UK Patent Application, GB0103216. 8, of the present applicant. The operation of this propulsion system is in

general accordance with the forgoing description and with reference to Figures 2 and 3.

Although two axial flow fan assemblies 90 are shown, any number of axial flow fan assemblies 90 may be used without departing from the scope of the present invention.

A further embodiment of the present invention is now described with reference to Figure 5, where the means for injecting pressurised fluid 100 is a circumferential fan 100. The circumferential fan 100 is disposed to the bypass wall 17 and when in operation the circumferential fan 100 is in fluid communication with the bypass duct 34, as shown by airflow C. The circumferential fan 100 comprises an annular array of blades 102 which are configured to draw in ambient air and pressurise it, thereby augmenting the bypass air flow A.

In this embodiment the circumferential fan assembly 100 is driven by an electrical motor 104, although any suitable motor may be used without departing from the scope of the present invention. The motor is mounted on the bypass wall 17 and is drivingly connected to the annular array of blades 102 via a drive arm 106 and an annular rack and pinion arrangement 108. The electrical supply to the motor may be from any suitable source, however, a preferred supply is described and incorporated herein by reference to a recent UK Patent Application, GB0103216. 8, of the present Applicant. The operation of this propulsion system is in general accordance with the forgoing description and with reference to Figures 2 and 3.

Referring to Figure 6, an aircraft 110 incorporates the propulsion system 50 as previously described with reference to Figure 2. For clarity, ducting has not been shown between the auxiliary engine 54 an the main engine 52. The auxiliary engine 54, which during take-off and climb provides the main engine 52 with a gas stream to augment the main engine's 52 gas stream, may also operate to provide power for other aircraft 110 systems. These

aircraft systems comprise: a wing thermal anti-icing device 112; an air driven pump 114; aircraft cabin air conditioning 116; electrical/electronic equipment cooling 118; panel instrument cooling 120 and; a hydraulic actuation system 122 for aircraft landing gear. All these aircraft systems are well known in the art. Conventional aircraft comprise an Auxiliary Power Unit (APU) 124 which provides pneumatic, electric and/or hydraulic power to some or all of the aircraft systems. It is an advantage of this embodiment of the present invention that functions of the APU 124 are replaced by the auxiliary engine 52.

Claims

1. A propulsion system 50 for an aircraft comprising a main engine 52, an exhaust nozzle assembly 16 and a means for injecting pressurised fluid 54,90, 100 into said propulsion system 50 wherein the means for injecting pressurised fluid 54,90, 100 is in fluid communication with the main engine 52 so as to direct pressurised fluid into said propulsion system at a position upstream of the exhaust nozzle assembly 16.

2. A propulsion system 50 for an aircraft as claimed in claim 1 wherein the means for injecting pressurised fluid 54 comprises an auxiliary engine 54.

3. A propulsion system 50 for an aircraft as claimed in any one of claims 1-2 wherein the main engine comprises a bypass duct 34, the bypass duct 34 defines an annular array of circumferentially spaced apart apertures 72 and a manifold 70, the manifold 70 substantially surrounds the bypass duct 34 and the apertures 72, the auxiliary engine 54 is in fluid communication with the fan bypass duct 34 via a duct 68 extending from the auxiliary engine 54 to the manifold 70.

4. A propulsion system 50 for an aircraft as claimed in any one of claims 1-2 wherein the main engine comprises an engine core nozzle 76, the core nozzle 76 defines an annular array of circumferentially spaced apart apertures 80 and a manifold 78, the manifold 78 substantially surrounds the core nozzle 76 and the apertures 80, the auxiliary engine 54 is in fluid communication with the core nozzle 76 via a duct 68 extending from the auxiliary engine 54 to the manifold 78.

5. A propulsion system 50 for an aircraft as claimed in any one of claims 3-4 wherein the auxiliary engine 54 is in fluid communication with both the manifold 78 and the manifold 70.

6. A propulsion system 50 for an aircraft as claimed in any one of claims 1-5 wherein the auxiliary engine 54 is in fluid communication with two or more main engines 52.

7. A propulsion system 50 for an aircraft as claimed in any one of claims 1-5 wherein two or more auxiliary engines 54 are in fluid communication with one main engine 52.

8. A propulsion system 50 for an aircraft as claimed in claim 1 wherein the means for injecting pressurised fluid 90 into the main engine 52 is one or more axial flow fan assemblies 90, each axial flow fan assembly 90 is in fluid communication with a bypass duct 34 of the main engine 52 and comprises an annular array of fan blades 92 radially extending from an axis 94 and about which a motor 96 rotates the blades 92 to provide, in use, a flow of pressurised fluid into the bypass duct 34.

9. A propulsion system 50 for an aircraft as claimed in claim 1 wherein the means for injecting pressurised fluid 100 into the main engine 52 is a circumferential fan assembly 100, the circumferential fan assembly 100 is in fluid communication with a bypass duct 34 of the main engine 52 and comprises a motor 104 drivingly connected to an annular array of blades 102 which are configured to rotate about a main axis 22 of the main engine 10 to provide, in use, a flow of pressurised fluid into the bypass duct 34.

10. A propulsion system 50 for an aircraft as claimed in either claim 8 or claim 9 wherein the motors 96,104 are electric and are supplied with electricity which is generated by the main engine.

11. A propulsion system 50 for an aircraft as claimed in any one of claims 8-9 wherein the means for injecting pressurised fluid 54 may be driven by any one of the group comprising an electric motor, a hydraulic motor, a diesel motor and a mechanical drive driven by said main engine.

12. A propulsion system 50 for an aircraft as claimed in claim 1 wherein a variable area nozzle 74 is disposed to the downstream periphery 71 of the fan bypass duct 34.

13. A propulsion system 50 for an aircraft as claimed in any one of claims wherein the means for injecting pressurised fluid 54 is also used to power an electrical, hydraulic, pneumatic or mechanical aircraft system.

14. A method of operating a propulsion system as claimed in any one preceding claim comprising the steps of: during aircraft take-off and climb, operating the main engine at a high output level and the means for injecting pressurised fluid at a high output level so that the nozzle flow of the main engine is augmented, whilst maintaining a variable area nozzle in its greatest area position; during aircraft cruise, operating the main engine at a reduced output level and ceasing operation of the means for injecting pressurised fluid, whilst maintaining a variable area nozzle in its smallest area position.

15. A propulsion system for an aircraft as hereinbefore described and with reference to figures 1 to 5.