GB2397347A

GB2397347A - Deriving electrical energy from waste heat in a gas turbine engine

Info

Publication number: GB2397347A
Application number: GB0301208A
Authority: GB
Inventors: James Richard Ieuan Lloyd
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2003-01-20
Filing date: 2003-01-20
Publication date: 2004-07-21
Also published as: GB0301208D0

Abstract

A gas turbine engine comprises a core engine and a core nozzle assembly 30 defined by a wall 42. In use hot gases are exhausted through the core nozzle assembly 30, creating a thermal gradient through the wall (42). The nozzle assembly 30 includes an array of thermionic converters 40, positioned such that they are subject to the thermal gradient thereby producing an electrical output.

Description

GAS TURBINE ENGINE ARRANGEMENT

The present invention relates to an arrangement of a gas turbine engine and in particular an arrangement for providing electrical power using waste heat recovery.

Gas turbine engines for industrial power generation, aircraft, marine or other applications produce significant amounts of heat from the combustion of fuel, a substantial proportion of which is exhausted and lost as waste. It has been a long felt want in the industry that at least some of this waste heat energy be recovered.

Conventional vehicular gas turbine engines extract power through a radial drive that is sized according to the load it needs to transmit for engine starting and power generation. The radial drive, drives the gearbox which drives a generator, or which is driven to start the engine.

The gearbox, drive shaft and generator all cause significant energy losses and reduce the efficiency of the engine.

Therefore it is an object of the present invention to provide an arrangement of a gas turbine engine capable of recovering a proportion of the waste energy by providing and reducing energy losses associated to the gearbox, drive shaft and generator.

Accordingly, a gas turbine engine comprises a core IS engine and a core nozzle assembly defined by a wall, in use the core engine produces hot exhaust gases that are exhausted through the core nozzle assembly thereby heating the nozzle assembly creating a thermal gradient through the wall, wherein the core nozzle assembly includes an array of thermionic converters, the thermionic converters positioned such that they are subject to the thermal gradient thereby producing an electrical output. This configuration reduces weight and cost of the engine and improves the aerodynamic drag thereby improving specific fuel consumption. À

The present invention will be more fully described by way of example with reference to the accompanying drawings in which: Figure 1 is a schematic section of a ducked fan gas turbine engine; Figure 2 is a perspective view on the rear of a core exhaust nozzle and 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 that 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 31 which drivingly connect 3s the turbine 28 rotors with the compressors 24 and fan 14.

Having flowed through the turbines 28 the high energy gas À 1 . . 3 À À stream from the combustor 26 still has a significant amount of energy and velocity and it is exhausted, as a core exhaust stream, through a core engine exhaust nozzle assembly 30 to provide propulsive thrust. The core engine exhaust nozzle assembly 30 is defined by a generally annular wall 42. 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 an exit bypass 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.

Conventional vehicular gas turbine engines extract power through a radial drive 32 that is sized according to the load it needs to transmit for engine starting and power generation. The radial drive 32, drives the gearbox 36 which drives a generator/starter 38, or which is driven to start the engine 10. The gearbox 36, radial drive 32 and generator 38 all cause significant energy losses and reduce the efficiency of the engine 10. The generator 38 is operable as an starter motor where it is driven by compressed air from an auxilliary power unit installed on the aircraft and as an electrical generator when it is driven through the radial drive 32 by the engine 10.

As For conventional engines 10 the outer nacelle wall 17 profile comprises a bulge 44 which is necessary over engine accessories (36, 38). The bulge 44 causes aerodynamic losses and as a result increases the specific fuel consumption of the installed engine 10. Further aerodynamic losses are incurred by a fairing 46 surrounding the radial drive 32 and spanning the bypass duct 34.

Figure 2 shows an array of thermionic converter devices 40 attached to the wall of the core exhaust nozzle 30. US5,994,638, US6,064,137 and US6,229,083 all disclose suitable thermionic converters which are capable of transforming a heat flux into electrical energy. The : e: À: thermionic converters 40 are attached over a significant portion of the core nozzle 30, however, for each application and required output the extent of the array 40 may be tailored accordingly.

The core nozzle 30 is a particularly suitable position for the thermionic converters 40 as there is a significant heat flux across the wall 42 of the core nozzle 30. Hot exhaust gases from the combustor 26, which pass through the turbines 28, then pass into and through the core nozzle 30 heating the radially internal side of the wall 42.

Relatively cooler air passes through the bypass duct 34, surrounding the core nozzle 30, thereby removing heat from the radially outer side of the wall 42. Thus there is a substantial heat gradient across the wall 42, which is required for the thermionic converters 40 to generate electrical power.

Electrical generation from the thermionic devices 40 enables the gearbox 36, radial drive 32 and generator/starter 38 to be substantially reduced in physical size. Where other means of engine start up are employed it is possible for the gearbox 36, radial drive 32 and generator 38 to be completely eliminated.

Where the air starter is removed a combined electric starter motor/generator is used thereby further reducing the nacelle bulge due to the accessories. This offers significant advantages for supersonic aircraft where a large nacelle profile can severly effect the drag of the aircraft.

The implementation of the thermionic converters 40 and subsequent electrical generation thereby enables a significant weight and cost saving for the engine. The aerodynamic efficiency is improved as the bulge 44 in the nacelle wall 17 is reduced or removed altogether.

Furthermore, the implementation of the present invention enables the failing 46 spanning the bypass duct 34 to be r # À # À À . # made substantially more slender which reduces aerodynamic losses.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims

: 8 À A C C À Claims: 1. A gas turbine engine comprising a core engine and

a core nozzle assembly defined by a wall, in use the core engine produces hot exhaust gases that are exhausted through the core nozzle assembly thereby heating the nozzle assembly creating a thermal gradient through the wall, wherein the core nozzle assembly includes an array of thermionic converters, the thermionic converters positioned such that they are subject to the thermal gradient thereby producing an electrical output.
2. A gas turbine engine substantially as described in this specification and with reference to and as shown in figures 1-2 of the accompanying drawings.