Combustion system

Abstract

A combustion system comprising: a combustion chamber extending in an axial direction between an inlet and an outlet, the combustion chamber configured to receive an airflow through the inlet and to discharge the airflow through the outlet; a fuel injection port configured to inject fuel into the airflow to form an air-fuel mixture; an ignition system for igniting the air-fuel mixture in the combustion chamber, the ignition system comprising an array of electrical plasma initiation points disposed downstream of the fuel injection port, and distributed radially and circumferentially around the combustion chamber, wherein each electrical plasma initiation point comprises a pair of electrodes configured to apply a voltage across an electrode gap between the pair of electrodes to produce plasma within the air-fuel mixture passing between the electrodes, thereby igniting the air-fuel mixture.

Claims

1. A combustion system comprising: a combustion chamber extending in an axial direction between an inlet and an outlet, the combustion chamber configured to receive an airflow through the inlet and to discharge the airflow through the outlet; a fuel injection port configured to inject fuel into the airflow to form an air-fuel mixture; an ignition system for igniting the air-fuel mixture in the combustion chamber, the ignition system comprising an array of electrical plasma initiation points disposed downstream of the fuel injection port, and distributed radially and circumferentially around the combustion chamber; a vane structure in the combustion chamber, the vane structure configured to guide the airflow from the inlet, wherein the array of electrical plasma initiation points are disposed on the vane structure and distributed radially and circumferentially around the vane structure, wherein each electrical plasma initiation point comprises a pair of electrodes configured to apply a voltage across an electrode gap between the pair of electrodes to produce plasma within the air-fuel mixture passing between the electrodes, thereby igniting the air-fuel mixture; and wherein the vane structure comprises a circumferentially distributed plurality of vanes, wherein each electrical plasma initiation point comprises a first electrode on a first vane and a second electrode on a circumferentially adjacent second vane.

2. The combustion system according to claim 1, wherein the electrodes of each electrical plasma initiation point are embedded in, and lie flush with, the respective vanes on which they are disposed.

3. The combustion system according to claim 1, wherein the vane structure comprises a plurality of integrated fuel injection ports, each fuel injection port being disposed upstream of the plurality of electrical plasma initiation points on the vane structure.

4. The combustion system according to claim 1, wherein the electrical plasma initiation points are distributed in rings around the vane structure, with each ring disposed at a different radial extent.

5. The combustion system according to claim 1, comprising a controller configured to control the electrodes to discontinuously energise.

6. An afterburner for a gas turbine engine, duct burner, ramjet, or scramjet comprising a core duct, a bypass duct, and the combustion system according to claim 1.

7. The afterburner according to claim 6, wherein a controller controls the electrodes disposed within the bypass duct to be continuously energised.

8. The afterburner according to claim 6, wherein a controller controls the electrodes disposed within the core duct to be discontinuously energised.

Description

DESCRIPTION OF THE DRAWINGS

(1) Embodiments will now be described, by way of example only, with reference to the accompanying Figures, in which:

(2) FIG. 1 is a sectional side view of a gas turbine engine;

(3) FIG. 2 schematically shows an axial cross-sectional view of the gas turbine engine with an afterburner having a combustion system;

(4) FIG. 3 schematically shows an oblique view of a vane structure of the combustion system; and

(5) FIG. 4 schematically shows a close-up cross-sectional view of a part of the vane structure.

DETAILED DESCRIPTION

(6) FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

(7) In use, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 where further compression takes place. The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low-pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high-pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

(8) Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example, or to a duct burner, ramjet, or scramjet. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

(9) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

(10) FIG. 2 shows the gas turbine engine 10 with an afterburner 100 at the exhaust of the gas turbine engine 10. In this example, the afterburner 100 comprises a core duct 102 and a bypass duct 104 which receive airflow respectively from the core 11 and the bypass duct 22 of the gas turbine engine 10. In some examples, the afterburner may not separate a core duct and a bypass duct, such as in a turbojet which has only a single main gas path.

(11) The afterburner 100 comprises a combustion system 120 which is configured to combust fuel to provide additional thrust in the afterburner 100. The combustion system 120 in this disclosure is sufficient to anchor a flame in the afterburner, such that no flameholders are required to slow the speed of the airflow. Therefore, it can be seen that, in this example, the afterburner 100 does not comprise any flameholders.

(12) The combustion system 120 comprises a combustion chamber 122 extending in an axial direction between an inlet 124 and an outlet 126. The combustion chamber 122 is configured to receive an airflow through the inlet 124 (i.e., the core airflow A and bypass airflow B from the gas turbine engine 10). The combustion chamber 122 is configured to discharge the airflow through the outlet 126.

(13) The combustion system comprises a fuel injection port 128 configured to inject fuel into the airflow in the combustion chamber 122 to form an air-fuel mixture, and an ignition system 130 (best seen in FIGS. 3 and 4) for igniting the air-fuel mixture in the combustion chamber 122. The ignition system 130 is disposed downstream of the fuel injection port 128.

(14) In this example, the fuel injection port 128 and the ignition system 130 are integrated into a single component in the form of a vane structure 150, which in this example is an exit guide vane. This minimises the disruption to the airflow in the afterburner. In other examples, the fuel injection port and the ignition system may be on separately mounted components.

(15) In this example, the vane structure 150 is mounted at the inlet of the combustion chamber 122 such that it is configured to guide the airflow from the inlet 124. In other examples, the vane structure may be mounted in any suitable location to guide the airflow from the inlet 124.

(16) In this example, the vane structure 150 spans across the core duct 102 and the bypass duct 104 of the afterburner 100. In other examples, the vane structure may span only the core duct or only the bypass duct, and/or only part of either duct.

(17) FIG. 3 shows the vane structure 150 independently of the combustion system 120. The vane structure 150 in this example comprises a plurality of circumferentially distributed vanes 152, which have an aerodynamic profile to minimise disruption to the airflow.

(18) The ignition system 130 on the vane structure 150 comprises an array of electrical plasma initiation points 132. In this example, the electrical plasma initiation points 132 are disposed on the vane structure 150 and distributed radially and circumferentially around the vane structure 150. The electrical plasma initiation points 132 use alternating current (AC) or direct current (DC) to energise a space, and thereby to create plasma within the space from the air-fuel mixture. This merely requires a simple pair of electrodes and is therefore very space efficient. Further, this enables many electrical plasma initiation points 132 to be distributed around the combustion chamber, both radially and circumferentially, to enable plasma to be generated at many different points from the air-fuel mixture in the airflow which is guided by the vane structure 150. Integration of the electrical plasma initiation points 132 on the vane structure further improves the design of the engine, since any number of electrical plasma initiation points 132 can be disposed on the vane structure 150 and powered through electrical components sheltered within the vane structure 150, without impeding the air flow through the vane structure 150.

(19) In an example in which there is no vane structure, or the ignition system is not integrated with a vane structure, the ignition system may simply comprise an array of electrical plasma initiation points which are distributed radially and circumferentially around the combustion chamber 122.

(20) Each of the electrical plasma initiation points 132 can generated plasma which enables anchoring of a flame. The radial and circumferential distribution of a plurality of electrical plasma initiation points 132 enables simple combustion staging, to optimise sub-system performance, by fuelling and energising sub-groups of fuel injector ports 128 and electrical plasma initiation points 132. This allows efficient combustion at low reheat settings, and also permits a low thrust step when reheat is initiated.

(21) As shown in FIG. 4, each of the electrical plasma initiation points 132 comprises a pair of electrodes 132a, 132b which are configured to apply a voltage across an electrode gap 134 to produce plasma within the air-fuel mixture passing between the electrodes 132a, 132b, thereby igniting the air-fuel mixture. The electrode gap may be between 1-20 mm wide.

(22) In this example, the electrodes 132a, 132b are disposed on the vane structure 150 within both the core duct 102 and the bypass duct 104 of the afterburner 100. In other examples, they may only be disposed in the core duct 102 or only in the bypass duct 104.

(23) When the ignition system 130 is integrated in the vane structure 150 as shown, the vane structure 150 can then be used to protect the electrodes 132a, 132b and the services to the electrodes 132a, 132b, whilst minimally impeding the airflow through the combustion chamber 122.

(24) In this example, each electrical plasma initiation point 132 comprises a pair of electrodes 132a, 132b on a single vane 152 separated along the axial direction 180 (which, in this example, is parallel to the principal rotation axis 9 of the gas turbine engine 10) such that an upstream electrode 132a from each electrical plasma initiation point 132 is upstream of a downstream electrode 132b from the respective electrical plasma initiation point 132. The upstream electrode 132a may be configured to be a negatively charged electrode, while the downstream electrode 132b may be configured to be a positively charged electrode, which improves function, as positive ions would move in counter-flow, thereby increasing time of exposure.

(25) In this example, each of the electrical plasma initiation points 132 are distributed in rings around the vane structure 150, with each ring disposed at a different radial extent. In other examples, the radial extent of the electrical plasma initiation points may form any suitable pattern on the vane structure.

(26) In some examples, each electrical plasma initiation point 132 may comprise a pair of electrodes on a single vane 152, and separated along a radial direction 170, perpendicular to the axial direction 180. Therefore, one electrode from each electrical plasma initiation point may be disposed radially outwardly of the other electrode from the respective electrical plasma initiation point.

(27) In other examples, each electrical plasma initiation point 132 may comprise a first electrode on a first vane 152 and a second electrode on a circumferentially adjacent second vane 152, such that the electrode gap spans a space between each vane 152.

(28) In this example, each of the electrodes 132a, 132b of each electrical plasma initiation point 132 is embedded in, and lies flush with, the respective vane 152 on which they are disposed. In other examples, the electrodes may extend out from the surface of the vane. For example, when the pair of electrodes of an electrical plasma initiation point are on adjacent vanes, the electrodes may extend out of the respective vane surface towards one another to reduce the electrode gap between them.

(29) In this example, the vane structure 150 also comprises a plurality of integrated fuel injection ports 128, where each fuel injection port 128 is disposed upstream of the plurality of electrical plasma initiation points 132 on the vane structure 150. In some examples, only the fuel injection ports may be on the vane structure or only the electrical plasma initiation ports may be on the vane structure, or the fuel injection ports and the electrical plasma initiation ports may be separately mounted from the vane structure.

(30) In this example, the combustion system 120 comprises a controller which is configured to control the electrodes 132a, 132b within the core duct 102 of the afterburner 100 to discontinuously energise to reduce the electrical power required for flame anchoring, as continuous ignition is not required, which reduces electrode erosion. In other examples, the controller may control the electrodes within the bypass duct 104 of the afterburner 100 to continuously energise. The pulse frequency may be any suitable frequency which may be dependent on the velocity of airflow and the scale of the combustion system 120.

(31) In examples where a core duct and bypass duct are not separated in an afterburner, the combustion system may comprise a controller which is configured to control all of the electrodes to discontinuously energise, or all of the electrodes to continuously energise or otherwise control the electrodes based on the gas stream conditions. Discontinuous energisation may be with a pulse frequency of at least 1 kHz, or any suitable pulse frequency dependent on the velocity of airflow and the scale of the combustion system 120.

(32) Although the disclosure relates generally to a gas turbine engine, it will be appreciated that the combustion system and afterburner can be equally applied to any suitable engine, such as a duct burner, a ramjet, or a scramjet. Further, although the disclosure relates generally to the fuel injection ports and the electrical plasma initiation points being integrated with a vane structure, this is not essential.

(33) It will be understood that the disclosure is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.