FUEL SYSTEM

Abstract

Disclosed is a fuel system for a gas turbine engine. The system comprises a fuel pump for fluid communication with a fuel reservoir; a driving turbine for driving the fuel pump; and a source of compressed air flow to drive the driving turbine. The source of compressed air may be the engine core, a dedicated fuel system compressor or the compressor of a cabin blower system.

Claims

1. A gas turbine engine comprising a fuel system, the gas turbine comprising a core compressor configured to provide core air flow to a combustor and a core turbine; the fuel system comprising: a fuel pump for fluid communication with a fuel reservoir; a driving turbine for driving the fuel pump; and a fuel system compressor operatively coupled to the core compressor and configured to provide compressed air flow to drive the driving turbine.

2. A system according to claim 1, wherein the fuel pump comprises a turbo pump.

3. A system according to claim 1, further comprising a cryogenic fuel reservoir.

4. A system according to claim 1, wherein the system comprises a fuel system transmission system, the fuel system transmission system operatively coupled to a compressor stage within the engine core.

5. A system according to claim 1, wherein the fuel system comprises a cabin blower bleed line from a cabin blower compressor of a cabin blower system.

6. A system according to claim 1, wherein the fuel system comprises a core bleed line from a compressor stage of the core of the gas turbine engine.

7. A system according to claim 1, further comprising a fuel evaporator downstream of the fuel pump to evaporate the fuel for introduction into the gas turbine engine.

8. A system according to claim 1, further comprising an exhaust downstream of the driving turbine and upstream of the fuel pump for venting any upstream fuel leakage.

9. A system according to claim 1, further comprising a throttle upstream or downstream of the driving turbine.

10. An aircraft comprising a fuel system according to claim 1.

11. An aircraft according to claim 10, further comprising a cabin blower system comprising a cabin blower compressor, the fuel system comprising a cabin blower feed line for channelling compressed air from the cabin blower compressor to the driving turbine.

12. A gas turbine engine comprising a fuel system, the gas turbine comprising a core compressor configured to provide core air flow to a combustor and a core turbine; the fuel system comprising: a fuel pump for fluid communication with a fuel reservoir; a driving turbine for driving the fuel pump; and a bleed air line from the core compressor configured to provide compressed air flow to drive the driving turbine.

13. An aircraft comprising a fuel system according to claim 12.

14. An aircraft according to claim 13, further comprising a cabin blower system comprising a cabin blower compressor, the fuel system comprising a cabin blower feed line for channelling compressed air from the cabin blower compressor to the driving turbine.

Description

DESCRIPTION OF THE DRAWINGS

[0050] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0051] FIG. 1 is a sectional side view of a gas turbine engine; and

[0052] FIG. 2 is a schematic representation fuel system according to an embodiment of the present disclosure.

DETAILED DESCRIPTION

[0053] With reference to FIG. 1, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate-pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

[0054] The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate-pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate-pressure compressor 14 compresses the air flow directed into it before delivering that air to the high-pressure compressor 15 where further compression takes place.

[0055] 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 com busted. The resultant hot combustion products then expand through, and thereby drive the high-, intermediate- and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high—17, intermediate—18 and low—19 pressure turbines drive respectively the high-pressure compressor 15, intermediate-pressure compressor 14 and fan 13, each by suitable interconnecting shaft.

[0056] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.

[0057] FIG. 2 shows a fuel system 100 for a gas turbine engine 10. The system comprises a fuel pump 101 for fluid communication with a cryogenic fuel reservoir (not shown) via a cryogenic pipework 102. A driving turbine 103 for driving the fuel pump 101 is provided upstream of the fuel line 102. The fuel pump 101 is a turbo-pump.

[0058] The driving turbine 103 is powered by compressed air flow within an airflow channel 104 generated by a fuel system compressor 105 and a fuel system transmission system 106. The fuel system transmission system 106 is operatively coupled to a compressor stage within the engine core 120. The fuel system transmission system 106 comprises a summing epicyclic gearbox 107 operatively coupled to the intermediate- and/or high-pressure compressor 14, 15 of the gas turbine engine core 120. The fuel system transmission system 106 comprises a first accessory gearbox 130 operatively coupled between one of the intermediate- or high-pressure compressor 14, 15 of the gas turbine engine core 120 and the summing epicyclic gearbox 107 to provide a first mechanical input into the summing epicyclic gearbox 107.

[0059] The fuel system transmission system 106 comprises a first electrical machine 108 connected to the first accessory gear box 130. The first electrical machine 108 is configured to convert mechanical power from the first accessory gear box 130 into electrical power. The first electrical machine 108 is configured to provide this electrical power to a second electrical machine 109 which is configured to convert the electrical power (from the first electrical machine 108) to mechanical power to provide a second mechanical input into the epicyclic gear box 107.

[0060] The fuel system also comprises a fuel system power management system 110 to control the transfer of electrical power between the first and second electrical engines 108, 109.

[0061] The summing epicyclic gearbox 107 has an output that is a function of the difference between the speeds of the first and second inputs. The second input is a continuously variable positive or negative input (as a result of the control by the power management system 110) which can be used to increase or decrease the compressed air output of the fuel system compressor 105 as desired and as required by operating conditions.

[0062] In addition, a first core bleed line (not shown) channels compressed air from the intermediate pressure compressor 14 in the gas turbine engine into the compressed air flow channel 104. A second core bleed line (not shown) channels compressed air from the high-pressure compressor 15 in the gas turbine engine into the compressed air flow channel 104.

[0063] In other embodiments (not shown), the source of compressed air may additionally or alternatively be provided from a cabin blower system of an aircraft i.e.

[0064] the fuel system may comprise a cabin blower bleed line from a cabin blower compressor of the cabin blower system.

[0065] The cabin blower compressor may be provided with air from the core compressors (as shown in FIG. 2) or from the engine fan bypass duct.

[0066] By using compressed air flow from the fuel system compressor 105 and from the first and second core bleed lines to drive a driving turbine 103 which, in turn drives the fuel pump 101, the fuel pump 101 can be operated at rotational speeds high enough to pump liquid hydrogen within the gas turbine engine without the need for traditional drives/gears and without the need for burning fuel to drive the driving turbine. Any upstream leakage of hydrogen to the driving turbine 103 can be exhausted from the driving turbine 103 along with the compressed air via an exhaust 111.

[0067] A throttle (not shown) is provided in the compressed air flow channel 104 to vary the volume/flow rate of the compressed air flow and thus control the driving power of the driving turbine 103 (and thus the output of the fuel pump 101). Such a throttle can negate the need for a fuel metering system.

[0068] The summing epicyclic gearbox 107 can also be used to reverse the fuel system compressor 105 from a blower mode to a start-up mode in which air from the reversed compressor 105 can still be used to drive the driving turbine 103 and allow operation of the fuel pump during start-up. In addition, isolation of the second electrical machine 109 allows the reversed compressor 105 to drive first accessory gearbox and, in turn, the operatively coupled engine core compressor 14, 15 to assist in engine start up.

[0069] The features disclosed in the foregoing description, or in the following claims, or in the accompanying drawings, expressed in their specific forms or in terms of a means for performing the disclosed function, or a method or process for obtaining the disclosed results, as appropriate, may, separately, or in any combination of such features, be utilised for realising the invention in diverse forms thereof.

[0070] While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.

[0071] For the avoidance of any doubt, any theoretical explanations provided herein are provided for the purposes of improving the understanding of a reader. The inventors do not wish to be bound by any of these theoretical explanations.

[0072] Any section headings used herein are for organizational purposes only and are not to be construed as limiting the subject matter described.

[0073] It must be noted that, as used in the specification and the appended claims, the singular forms “a,” “an,” and “the” include plural referents unless the context clearly dictates otherwise. Ranges may be expressed herein as from “about” one particular value, and/or to “about” another particular value. When such a range is expressed, another embodiment includes from the one particular value and/or to the other particular value. Similarly, when values are expressed as approximations, by the use of the antecedent “about,” it will be understood that the particular value forms another embodiment. The term “about” in relation to a numerical value is optional and means for example +/−10%.