Gas turbine engine using cryogenic fuel

Abstract

A gas turbine arrangement is disclosed, comprising a gas turbine engine and a cryogenic fuel pre-processing system. Also disclosed is a method of operating a gas turbine engine. In axial flow series, the gas turbine engine comprises: a compressor; a combustor; and a turbine configured to drive the compressor. The cryogenic fuel pre-processing system is configured to heat and / or vaporise cryogenic fuel to provide processed fuel for delivery to the combustor. The cryogenic fuel pre-processing system comprises: a gas outlet configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along a bleed air gas path; a heat exchanger configured to reject heat from the bleed turbine gas to the cryogenic fuel to provide cooled bleed turbine gas and heated and / or vaporised processed fuel; and a processed fuel inlet configured to deliver the processed fuel to the combustor.

Description

[0001] 008891418

[0002] 1

[0003] GAS TURBINE ENGINE USING CRYOGENIC FUEL

[0004] Field of the Invention

[0005] The present invention relates to gas turbine engines and methods for their operation. The invention has particular relevance to gas turbine engines which use cryogenic fuel, particularly, although not necessarily exclusively, liquid hydrogen fuel, for at least a part of their operation.

[0006] Background

[0007] The use of sustainable aviation fuels is one possible route to decarbonise aviation. Sustainable aviation fuels are alternative fuels made from non-petroleum feedstocks. Some available sustainable aviation fuels, such as hydrogen and ammonia, benefit from or require processing before they can be injected into the combustion chamber of a gas turbine. These processing steps generally require a heat input.

[0008] Currently, there are two sources of heat proposed to perform fuel processing. A first option is to combust a small fraction of the fuel to create heat which can be used to process the rest of the fuel. A system like this for vaporising cryogenic hydrogen fuel is described in US 2022 / 0099020 A1 ‘Hydrogen Fuel Vaporiser’. One problem with such arrangement is that it is thermodynamically inefficient; whenever a fuel is burned there is the potential to extract some work by using a thermodynamic cycle; in such arrangements, that potential to do work is being lost.

[0009] A second option is to use the waste exhaust heat from the engine. This is more thermodynamically efficient than the option described above, as it utilises waste exhaust rather than burning extra fuel; however, there are still some disadvantages with this approach. In particular, when the exhaust of an engine is used to process fuel, there must be a compromise between the efficacy of fuel processing and the production of efficient thrust in the engine: Firstly, the more efficient a heat engine is, the colder its exhaust is, which reduces the efficacy of many of the fuel processing processes. Secondly, the heat capacity of the fuel is typically much smaller that the heat capacity of the engine exhaust, as the air / fuel ratio of the engine is fixed by the stoichiometry of the engine. This means any heat exchange between the exhaust and the fuel not well matched, so only a small fraction of the waste exhaust heat can be utilised for fuel processing. This problem becomes worse as the engine is made more efficient as the engine consumes less fuel, which further reduces the heat capacity of the fuel compared to the heat capacity of the exhaust. These two effects mean that an engine optimised to provide efficient thrust will always be suboptimal for fuel processing using exhaust heat, and an engine optimised for fuel processing using exhaust heat will always be suboptimal for providing efficient thrust.

[0010] A third option available for cryogenic fuels is the use of an advanced thermodynamic cycle which exploits the low temperature of the fuel. One possible way to extract this work with an advanced thermodynamic cycle is to run an open cycle with the fuel used as a working fluid. Systems employing this arrangement are described in US 2016 / 0123226A1 ‘Gas Turbine using a Cryogenic Fuel and Extracting Work008891418

[0011] 2

[0012] Therefrom’, and EP4239170A1 ‘Combined Cycles’. In such arrangements, the heat input for these open fuel cycles is provided by the exhaust of the engine. As discussed above, using the exhaust of the engine means that there is a compromise between the operation of the engine and the operation of the open fuel cycle. The more efficient the engine is, the colder its exhaust, which reduces the performance of the open fuel cycle. The heat exchange between the exhaust air and the fuel in the prior art is also suboptimal, as the heat capacity of the air stream is not well matched to the fuel stream. These two effects mean that an engine optimised to provide efficient thrust will always be suboptimal for providing exhaust heat to an open fuel cycle, and an engine optimised for providing exhaust heat to an open fuel cycle will always be suboptimal for providing efficient thrust.

[0013] The present invention has been devised in light of the above considerations.

[0014] Summary of the Invention

[0015] The present inventors have realised that it may be possible to make use of one or more bleed air streams from the turbine of a gas engine in order to provide suitable processing of a cryogenic fuel prior to its injection into the combustion chamber of a gas turbine.

[0016] Accordingly, in a first aspect, the present invention provides a gas turbine arrangement comprising a gas turbine engine and a cryogenic fuel pre-processing system, the gas turbine engine comprising, in axial flow series:

[0017] a compressor;

[0018] a combustor; and

[0019] a turbine configured to drive the compressor;

[0020] wherein the cryogenic fuel pre-processing system is configured to heat and / or vaporise cryogenic fuel to provide processed fuel for delivery to the combustor and comprises:

[0021] a gas outlet configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along a bleed air gas path;

[0022] a heat exchanger configured to reject heat from the bleed turbine gas to the cryogenic fuel to provide cooled bleed turbine gas and heated and / or vaporised processed fuel; and

[0023] a processed fuel inlet configured to deliver the processed fuel to the combustor.

[0024] In a second aspect, the present invention provides a method of operating a gas turbine engine comprising, in axial flow series:

[0025] a compressor;

[0026] a combustor; and

[0027] a turbine configured to drive the compressor;

[0028] the method including providing a supply of cryogenic fuel, and processing the cryogenic fuel to provide processed fuel by passing the cryogenic fuel through a cryogenic fuel pre-processing system comprising:

[0029] a gas outlet configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along a bleed air gas path;008891418

[0030] 3

[0031] a heat exchanger configured to reject heat from the bleed turbine gas to the cryogenic fuel to provide cooled bleed turbine gas and heated and / or vaporised processed fuel; and

[0032] a processed fuel inlet configured to deliver the processed fuel to the combustor; and thereafter directing said processed fuel into the combustor for combustion therein.

[0033] The present inventors have realised that there are a number of potential advantages which arise from providing arrangements in which bleed air from the turbine is used to provide the heat required for cryogenic fuel processing.

[0034] Firstly, it may be possible to separately optimise the pressure and mass flow rate of the bleed air for fuel processing with minimal changes to the core engine architecture and overall pressure ratio, which are optimised to produce efficient thrust. Accordingly, the gas turbine arrangement may be more efficient than alternative known gas turbine arrangements that employ different methods of fuel processing.

[0035] Secondly, such arrangements may allow for more compact gas turbine arrangements as compared with those that employ different methods of fuel processing - in particular, the location of the heat exchanger can be selected in order to provide more compact arrangements than known systems, as will be discussed in greater detail below. Additionally, as the bleed air from the turbine has a higher density than the air in the exhaust of the engine, a smaller cross-sectional area for the heat exchanger can be used for the same velocity and mass flow rate of flow vs known arrangements that employ gas streams from the engine exhaust for fuel processing. This allows use of a more compact heat exchanger design, reducing heat exchanger size and weight.

[0036] Finally, using a turbine bleed stream for fuel processing allows this bleed stream to be at, or close to, the optimal temperature and mass flow rate for fuel processing, with minimal impact on the ability of the engine to provide efficient thrust. This allows the fuel processing to be performed more efficiently than is possible when using e.g. exhaust gas for fuel processing, which increases the efficiency of the overall propulsion system. The reasons why the bleed stream can be made more optimal than the exhaust gasbased processing arrangements are twofold: firstly, the temperature of the bleed stream can be made much hotter than the exhaust, which generally increases the efficacy of fuel processing. Secondly, the air / fuel ratio is no longer fixed by the stoichiometry of the engine as only a proportion of the air is bled off the turbine. This means that the heat capacity of the air which is bled from the turbine can be set by the designer to a desirable value.

[0037] In the documents US 2016 / 0123226 A1 and EP4239170A1 identified above, the relevant disclosure is limited to the use of heat from the exhaust to pre-process the fuel. The use of bleed air from the turbine is advantageous over this for the reasons explained above and in more detail below. Furthermore, as will be understood, the use of bleed air from the turbine is significantly different to taking bleed air from the compressor (whether low pressure or high pressure stages), and the present invention provides advantageous effects compared with an approach using bleed air from the compressor.

[0038] The cryogenic fuel pre-processing system may alternatively be referred to as a cryogenic fuel preprocessing and delivery / injection system, in that the purpose of the system is both to pre-process the fuel008891418

[0039] 4

[0040] to provide processed fuel which is at a suitable thermodynamic condition for injection into the combustor, but also to inject / deliver the processed (heated and / or vaporised) fuel into the combustor via the processed fuel inlet.

[0041] Further optional features of the invention will now be set out.

[0042] The compressor and turbine of the gas turbine engine may be generally conventional. In some embodiments, the compressor may comprise, in axial flow series, a low-pressure compressor and a high-pressure compressor. An intercase split may be present between the low-pressure compressor and a high-pressure compressor. In some embodiments, the turbine may comprise, in axial flow series, a high-pressure turbine and a low-pressure turbine. The high-pressure compressor may be driven by the high-pressure turbine via a first shaft. The low-pressure compressor may be driven by the low-pressure turbine via a second shaft. The low-pressure compressor may be in fluid connection with a fan, the fan being located axially upstream of the low-pressure compressor, in an axial flow direction. The low-pressure turbine may be in fluid connection with an exhaust, the exhaust being located axially downstream of the low-pressure turbine, in an axial flow direction.

[0043] The term “cryogenic fuel” is used herein to refer to fuels at very low temperatures, e.g. temperatures of below 123 K (around -150 °C). The cryogenic fuel may be a liquid. The cryogenic fuel may be selected from liquid hydrogen, liquid natural gas (LNG) or liquid methane. Preferably the cryogenic fuel comprises liquid hydrogen. Where the cryogenic fuel comprises liquid hydrogen, it may have a temperature of -253 °C or lower.

[0044] The term “processed fuel” is used herein to refer to cryogenic fuel that has been processed e.g. by heating and / or vaporising the cryogenic fuel. The processed fuel will typically be at a non-cryogenic temperature, e.g. a temperature of greater than 123K (i.e. above around -150 °C), although it is contemplated that in some arrangements, the processed fuel may still be cryogenic, but heated relative to the temperature of the cryogenic fuel input into the fuel pre-processing system. For example, in some arrangements, the cryogenic fuel input into the cryogenic fuel pre-processing system may have a temperature of around 20K (around -253 °C), and the processed fuel may have a temperature of around 80K or more (around -193 °C or more). As discussed in further detail below, in some arrangements, the processed fuel may have a temperature significantly greater than 123K, e.g. may have a temperature of 250K or more.

[0045] The cryogenic fuel may be stored in a cryogenic fuel storage unit, prior to being input into the cryogenic fuel pre-processing system. Where the cryogenic fuel is stored in a cryogenic fuel storage unit, the cryogenic fuel storage unit may be arranged to be in fluid connection with the cryogenic fuel preprocessing system. Conveniently, the gas turbine arrangement may comprise a main fuel conduit for delivery of cryogenic fuel to the cryogenic fuel pre-processing system. The main fuel conduct may fluidly connect the cryogenic fuel storage unit to the cryogenic fuel pre-processing system. The gas turbine arrangement may comprise a fuel pump configured to pump cryogenic fuel from a cryogenic fuel storage unit to the cryogenic fuel pre-processing system, e.g. via the main fuel conduit.008891418

[0046] 5

[0047] Where the gas turbine arrangement comprises a main fuel conduit for delivery of cryogenic fuel to the cryogenic fuel pre-processing system, the heat exchanger may be disposed at an intersection of the main fuel conduit and the bleed air gas path. This may keep the length and weight of required ducting to a minimum, thereby allow for a more compact arrangement.

[0048] The heat exchanger may comprise two or more heat exchange units. For example, the heat exchanger may comprise two or more heat exchange units configured in a series or parallel arrangement. Where such an arrangement is used, preferably the heat exchanger comprises two or more heat exchange units arranged in parallel, wherein each heat exchange unit is configured to receive a predetermined fraction of the total turbine bleed air, to heat up a predetermined fraction of the total fuel flow. For example, where the heat exchanger comprises two heat exchange unit arranged in parallel, each of the two heat exchange unit may be configured to receive approximately half of the total amount of turbine blood air and half of the total amount of cryogenic fuel flowing through the heat exchanger assembly. Such arrangements may allow the heat exchanger to more easily fit around other components in the gas turbine, and thereby allow for a more compact arrangement.

[0049] As noted above, the present invention involves pre-processing the cryogenic fuel prior to injection of the fuel into the combustor, by heating and / or vaporising the cryogenic fuel to provide processed fuel for delivery to the combustor. Preferably, the cryogenic fuel is heated to a temperature at which it is a gaseous fuel. In other words, where the cryogenic fuel is a liquid cryogenic fuel, preferably the preprocessing of the fuel comprises vaporising the liquid cryogenic fuel to produce a gaseous fuel. It will be appreciated that the precise amount of heat exchange required in order to vaporise the liquid cryogenic fuel to produce a gaseous fuel will depending on a number of factors including the temperature of the cryogenic fuel prior to processing, and the boiling point of the cryogenic fuel. Where the cryogenic fuel is liquid hydrogen, the step of heating and / or vaporising the cryogenic fuel may include heating the cryogenic fuel to a temperature of greater than -253 °C, e.g. to a temperature of at least 80 K. A minimum temperature of at least 80K for the processed fuel may be desirable, as this is just greater than the boiling point of air: if hydrogen were to be injected into the combustion chamber at a temperature lower than 80K, there may be a risk that air would liquefy on the hydrogel fuel, preventing efficient combustion. In preferred arrangements, the hydrogen may be heated to temperatures of 100K or more, 150K or more, 200K or more, or most preferably 250K or more. It is hypothesised that a temperature of >250K may provide most effective for good combustion of the hydrogen.

[0050] In some arrangements, the bleed turbine gas may be extracted from the turbine of the gas turbine engine at a pressure in a range of from 1.4 times the exit pressure of the turbine to 0.5 times the inlet pressure of the turbine, preferably in a range of from 1.6 times the exit pressure of the turbine to 0.2 times the inlet pressure of the turbine, more preferably at a pressure around 3 times the exit pressure of the turbine. In these ranges, the lower bound is defined as a fixed factor larger than the exist pressure of the turbine, and the upper bound is defined as a fixed factor times the inlet pressure of the turbine (which is typically much higher than the exit pressure of the turbine. For example, in an arrangement where the exit pressure of the turbine is 0.3 bar, and the inlet pressure of the turbine is 15 bar, then the bleed turbine008891418

[0051] 6

[0052] gas may be extracted from the turbine of the gas turbine engine at a pressure in a range of from 0.42 bar to 7.5 bar, preferably in a range of from 0.48 bar to 3 bar, more preferably around 0.9 bar. It has been found that extracting the bleed turbine gas at a pressure within these ranges may provide greatest system efficiency.

[0053] In some arrangements, the bleed turbine gas may be extracted from the turbine of the gas turbine engine ata temperature in a range of from 1.1 times the exit temperature of the turbine to 0.84 times the inlet temperature of the turbine, preferably in a range of from 1.14 times the exit temperature of the turbine to 0.66 times the inlet temperature of the turbine, more preferably at a temperature of around 1.3 times the exit temperature of the turbine. It has been found that extracting the bleed turbine gas as a temperature within these ranges may provide greatest system efficiency. If the bleed gas has a higher temperature, it may allow for improved efficacy of fuel processing.

[0054] As noted above, the cryogenic fuel pre-processing system of the gas turbine arrangement comprises at least one gas outlet configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along a bleed air gas path. This gas outlet may conveniently be referred to as a first gas outlet. The bleed air gas path may conveniently be referred to as a first bleed air path. In some arrangements, the cryogenic fuel pre-processing system comprises one or more further gas outlets configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along one or more further bleed air gas paths. For example, the cryogenic fuel pre-processing system may comprise at least a second gas outlet configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along a second bleed air gas path, in addition to the first outlet. Where the cryogenic fuel pre-processing system comprises one or more further gas outlets, the one or more further bleed air gas paths may be entirely distinct from the first bleed air path. Alternatively, the one or more further bleed air gas paths may share at least a portion of a common path with the first bleed air path. The one or more further bleed air gas paths may exhaust into any suitable location: they may exhaust into the same location as the first bleed air gas path, or into a different location. Further discussion of possible exhaust locations for the bleed air gas flow path(s) is set out below.

[0055] Providing multiple bleed gas flows may be particularly advantageous for certain gas turbine arrangements. For example, some fuel processes, such as the fuel open cycle described in EP4239170A1 ‘Combined Cycles’ require the fuel to be heated multiple times. Accordingly, provision of two or more independent bleed streams could be used to implement arrangements as disclosed therein. A further advantage of using two or more separate bleed streams is that the temperature and mass flow rate of each bleed stream can be optimised for its intended purpose, increasing the overall performance of the engine.

[0056] The one or more further gas outlets (where present) may be configured to extract bleed turbine gas from the turbine of the gas turbine engine at a location which is aligned with the location of the first gas outlet in the axial flow direction (i.e. where the second or further bleed gas flow is extracted from the turbine at a pressure substantially equal to the pressure at which the first bleed gas flow is extracted from the008891418

[0057] 7

[0058] turbine). Alternatively, the one or more further gas outlet may be located upstream or downstream of the first gas outlet in the axial flow direction (i.e. where the second or further bleed gas flow is extracted from the turbine at a pressure higher or lower than the pressure at which the first bleed gas flow is extracted from the turbine).

[0059] Where the gas turbine arrangement comprises multiple bleed gas flows, the bleed gas flows may be directed to the same heat exchanger. In such arrangements, the heat exchanger may be configured to reject heat from the bleed turbine gas of one or multiple (e.g. all) of the respective bleed air flows to provide cooled bleed turbine gas and heated and / or vaporised processed fuel. Alternatively, one or more auxiliary heat exchangers may be provided, and the multiple bleed gas flows may be respectively directed to the first heat exchanger and the one or more auxiliary heat exchangers. In some arrangements, an auxiliary heat exchanger is provided for each respective bleed gas flow. In such arrangements, each respective heat exchanger may be configured to reject heat from its respective bleed gas flow to provide cooled bleed turbine gas and heated and / or vaporised processed fuel.

[0060] In some arrangements, the bleed fraction (the mass flow of the bleed air relative to the total mass flow through the engine), is selected to be in a range of from 1 to 20 %. For example the bleed fraction may be 2 % or more, 3 % or more, 4 % or more, 5 % or more, 6 % or more, 7 % or more, 8 % or more, 9 % or more, or 10 % or more. The bleed fraction may be 19% or less, 18% or less, 17% or less, 16% or less, 15 % or less, 14% or less, or 13 % or less. Preferably, the bleed fraction may be in a range of from 7 to 13%, e.g. about 7%, about 8 %, about 9%, about 10 %, about 11 %, about 12 % or about 13 %. Where the arrangement comprises multiple bleed gas flows, this bleed fraction may be calculated as the total mass flow of bleed air across all of the bleed gas flows. It has been found that providing a bleed fraction in this range may offer suitable or improved engine performance and efficiency.

[0061] Subsequently to being passed through the heat exchanger, the cooled bleed turbine gas may be exhausted to one or more of:

[0062] the turbine;

[0063] the compressor;

[0064] the environment;

[0065] one or more further turbines of the gas turbine arrangement;

[0066] one or more auxiliary systems.

[0067] Preferably, the cooled bleed turbine gas is initially exhausted to one or more of the turbine, the compressor, one or more further turbines of the gas turbine arrangement, or one or more auxiliary systems: in other words, preferable the cooled bleed turbine gas is not initially exhausted to the environment. In this way, some further utility can be derived from the cooled bleed turbine gas flow. In some arrangements, the bleed air gas path may be configured to exhaust the cooled bleed turbine gas into the compressor. Preferably, the cooled bleed turbine gas is exhausted into the compressor at a location at which the compressor pressure is close to or substantially equal to the pressure at which the bleed gas is extracted from the turbine. This can help to ensure there is no significant pressure differential across the bleed gas flow path. When the cooled bleed turbine gas is mixed into the008891418

[0068] 8

[0069] compressor, if it is a lower temperature than the compressor air then it will intercool the compressor air as it mixed with it. Intercooling the compressor air will increase the power output of the engine and may increase the efficiency of the engine. Intercooling the compressor also lowers the temperature of the air and the components in the compressor, which may increase the lifetime of the components, particularly in the hottest section at the rear of the compressor.

[0070] The cooled bleed turbine gas may be exhausted into the compressor at an intercase split between low pressure (LP) and high pressure (HP) stages of the compressor. Such arrangements may be easiest to implement in view of overall engine architecture (i.e. may allow for less complex arrangements vs arrangements in which the cooled bleed turbine gas is exhausted into the compressor at a location other than at the intercase split). Furthermore, the pressure within the compressor at the location of the intercase split may be similar or substantially equal to the optimal pressure of extraction of the bleed gas flow from the turbine. Accordingly, exhausting the cooled bleed turbine gas into the compressor at an intercase split between low pressure (LP) and high pressure (HP) stages of the compressor can offer a good compromise between optimal system performance and simplicity of engine architecture.

[0071] In some arrangements, the bleed air gas path may be configured to exhaust the cooled bleed turbine gas into the turbine. The cooled bleed turbine gas may be exhausted into the turbine downstream of the gas outlet configured to extract bleed turbine gas. Exhausting the cooled bleed turbine gas into the turbine downstream of the gas outlet which extracts the bleed turbine gas may allow further work extraction from the cooled bleed turbine gas, improving overall efficiency of the system.

[0072] In some arrangements, the bleed air gas path may be configured to exhaust the cooled bleed turbine gas into one or more further turbines of the gas turbine arrangement. This may allow further work extraction from the cooled bleed turbine gas, improving overall efficiency of the system.

[0073] In some arrangements, the bleed air gas path may be configured to exhaust the cooled bleed turbine gas into one or more auxiliary systems. The one or more auxiliary system may include e.g. a cabin pressurisation system, and / or an air conditioning system. Using the cooled turbine bleed air for cabin pressurisation and / or air conditioning has an additional advantage over conventional arrangements: in conventional aircraft, bleed air for the cabin is typically taken from the compressor. This has a very low humidity, which may create discomfort for passengers. In comparison, cooled bleed turbine gas may have a higher humidity, in particular where the gas turbine arrangement is a hydrogen-fuelled gas turbine arrangement, as the cooled bleed turbine gas will comprise the water created from the combustion of hydrogen. Accordingly, where the cabin pressurisation system and / or an air conditioning systems utilise cooled bleed turbine gas, the higher humidity of this air may provide a more comfortable environment for passengers.

[0074] The invention includes the combination of the aspects and preferred features described except where such a combination is clearly impermissible or expressly avoided.008891418

[0075] 9

[0076] Summary of the Figures

[0077] Embodiments and experiments illustrating the principles of the invention will now be discussed with reference to the accompanying figures in which:

[0078] Figure 1 is a schematic diagram of a conventional (prior art) liquid hydrogen fuelled gas turbine engine.

[0079] Figure 2 is a schematic diagram of a first gas turbine arrangement which utilises bleed air fuel processing according to the present invention.

[0080] Figure 3 is a temperature-entropy (T-s) diagram of a gas turbine as shown in Fig. 2

[0081] Figure 4 is a schematic diagram of a second gas turbine arrangement which utilises bleed air fuel processing according to the present invention. The gas turbine arrangement employs multiple bleed gas flows.

[0082] Figure 5 is a temperature-entropy (T-s) diagram of a gas turbine as shown in Fig. 4.

[0083] Figure 6 is a graph showing the relative change in cycle efficiency for different engine configurations in the cruise flight condition based on simulation, including results of simulations based on: (I) a comparative engine configuration which burns a fraction of the cryogenic fuel to create heat which is then used to process the rest of the fuel as disclosed in US 2022 / 0099020 A1 ‘Hydrogen Fuel Vaporiser’ (ii) a comparative engine configuration which utilises exhaust gas fuel processing and (ill) an engine configuration which utilises turbine bleed air fuel processing according to the present invention.

[0084] Detailed Description of the Invention

[0085] Aspects and embodiments of the present invention will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art. All documents mentioned in this text are incorporated herein by reference.

[0086] Fig. 1 is a schematic diagram of a conventional (prior art) liquid hydrogen fuelled gas turbine engine 1. The gas turbine engine comprises, in axial flow series, a fan 3, a compressor 5 including a low pressure compressor 7 and high pressure compressor 9, a combustor 11 , a turbine 13 including a high pressure turbine 15 and low pressure turbine 17, an exhaust 19, and a fuel pre-processing system 21 which is configured to convert the liquid hydrogen to gaseous hydrogen before it enters the combustion chamber. The liquid hydrogen travels to the combustor 11 along a main fuel conduit 20 before entering the combustor via an inlet 22.

[0087] Fig. 2 is a schematic diagram of a first gas turbine arrangement 100 which utilises bleed air fuel processing according to the present invention. The arrangement is generally similar in a number of ways to that shown in Fig. 1 : the gas turbine engine comprises, in axial flow series, a fan 103, a compressor 105 including a low pressure compressor 107 and high pressure compressor 109, a combustor 111 , a turbine 113 including a high pressure turbine 115 and low pressure turbine 117, an exhaust 119, and a fuel pre-processing system 121. The fuel pre-processing system 121 is configured to heat and / or008891418

[0088] 10

[0089] vaporise cryogenic fuel to provide processed fuel for delivery to the combustor 111. The cryogenic fuel is stored in a cryogenic fuel storage unit (not shown) which is conveniently a cryogenic storage tank which stores the cryogenic fuel (e.g. hydrogen) in a liquid state - for example, at a temperature of -253 °C or less. The cryogenic fuel storage unit (not shown) is in fluid connection with the cryogenic fuel preprocessing system via the main fuel conduit 120. The main fuel conduit 120 can be considered to be part of the fuel pre-processing system 121 , with the fuel pre-processing system 121 constituting both a cryogenic fuel pre-processing and delivery / injection system, which is operable both to pre-process the fuel to provide processed fuel suitable for injection into the combustor 111 , but also to inject / deliver the processed (heated and / or vaporised) fuel into the combustor via a processed fuel inlet 129.

[0090] The fuel pre-processing system 121 comprises a first gas outlet 123 configured to extract bleed turbine gas from the turbine 113 of the gas turbine engine and direct the bleed turbine gas along a first bleed air gas path 125. In this diagram, the bleed air is shown as being extracted from the low-pressure turbine 117, however it will be appreciated that the bleed air could be extracted from the turbine at any suitable point. The fuel pre-processing system 121 comprises a first heat exchanger 127 configured to reject heat from the bleed turbine gas to the cryogenic fuel to provide cooled bleed turbine gas and heated and / or vaporised processed fuel. The heat exchanger 127 is located at the intersection of the main fuel conduit 120 and the bleed air gas path 125. The main fuel conduit 120 continues past the heat exchanger to the processed fuel inlet 129 configured to deliver the heated and / or vaporised processed fuel to the combustor.

[0091] In the arrangement shown here, the bleed air gas path 125 is arranged to exhaust the cooled bleed turbine gas into the compressor 105, in particular, at the intercase split between low pressure (LP) and high pressure (HP) stages of the compressor 107, 109. However, it will be appreciated that alternative arrangements for exhaust of the cooled bleed turbine gas are possible.

[0092] Fig. 3 is a temperature-entropy (T-s) diagram of a gas turbine as shown in Fig. 2, including the bleed gas flow. Bleed air is taken from the turbine at temperature Tc and used to supply heat for fuel processing, as described above in relation to Fig. 2, which cools the bleed air down to temperature Ta. The cooled bleed air is then supplied to the compressor, where the cooled bleed air mixes with the compressor air to temperature Tb. The temperature of the bleed air is determined by which stage of the turbine the air is bled from. The temperature and mass flow rate of the bleed air are optimised for fuel processing, while the pressure ratio of the gas turbine is optimised for the efficient production of thrust. Using a turbine bleed stream for fuel processing allows this bleed stream to be the optimal temperature and mass flow rate for fuel processing, with minimal impact on the ability of the engine to provide efficient thrust. This allows the fuel processing to be performed more efficiently than possible when using the exhaust for fuel processing, which increases the efficiency of the overall propulsion system.

[0093] Fig. 4 is a schematic diagram of a second gas turbine arrangement 200 which utilises bleed air fuel processing from multiple bleed gas flows according to the present invention. Fig. 5 is a temperatureentropy (T-s) diagram of a gas turbine as shown in Fig. 4, including the bleed gas flows. The first bleed stream is taken at temperature Te and used to supply heat for fuel processing, which cools the first bleed008891418

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[0095] air stream down to temperature Ta. The cooled first bleed air stream is then supplied to the compressor, where the cooled bleed air mixes with the compressor air to temperature Tb. The second bleed stream is taken at temperature Tf and used to supply heat for fuel processing, which cools the second bleed air stream down to temperature Tc. The cooled second bleed air stream is then supplied to the compressor, where the cooled bleed air mixes with the compressor air to temperature Td.

[0096] The arrangement of Fig. 4 is generally similar to the arrangement shown in Fig. 2. Features described above in relation to the arrangement shown in Fig. 2 will therefore not be described again here. The only difference between the arrangement shown in Fig. 4 and the arrangement shown in Fig. 2 is the presence of an additional bleed gas flow path. In this arrangement, the fuel pre-processing system 221 comprises a first gas outlet 223a configured to extract bleed turbine gas from the turbine of the gas turbine engine at a first location, and a second gas outlet 223b configured to extract bleed turbine gas from the turbine of the gas turbine engine at a second location. Here, the second location is upstream of the first location in an axial flow direction. The first gas outlet 223a is in fluid connection with a first bleed air gas path 225a. The second gas outlet 223b is in fluid connection with a second bleed air gas path 225b.

[0097] The fuel pre-processing system 221 comprises a first heat exchanger 227 configured to reject heat from the bleed turbine gas flowing along each of the first and second bleed gas flow paths 225a, b to the cryogenic fuel flowing along the main fuel conduit 220 to provide cooled bleed turbine gas and heated and / or vaporised processed fuel. The heat exchanger 227 is located at the intersection of the main fuel conduit 220 and each of the first and second bleed air gas paths 225a, b. The main fuel conduit 220 continues past the heat exchanger 227 to the processed fuel inlet configured to deliver the heated and / or vaporised processed fuel to the combustor.

[0098] In the arrangement shown here, the first bleed air gas path 225a is arranged to exhaust the cooled bleed turbine gas into the compressor at a first exhaust location, in particular, at the intercase split between low pressure (LP) and high pressure (HP) stages of the compressor. The second bleed air gas path 225b is arranged to exhaust the cooled bleed turbine gas into the compressor at a second exhaust location which in this arrangement is in the high pressure compressor, downstream of the first exhaust location in an axial flow direction. The advantage of using two or more separate bleed streams as shown in this arrangement is that the temperature and mass flow rate of each bleed stream can be optimised for its intended purpose, increasing the overall performance of the engine.

[0099]

[0100] To demonstrate the benefit of using bleed air for fuel processing, the following engine configurations were simulated for a large, high bypass ratio civil jet engine that uses cryogenic liquid hydrogen as a fuel. The following engine configurations were compared:

[0101] An engine with no fuel preprocessing system to serve as a baseline. In this engine, the cryogenic fuel is supplied directly to the combustion chamber. Note that it is not possible to combust liquid hydrogen in practice and this only serves as a theorical baseline case against which the other systems can be compared.008891418

[0102] 12

[0103] An engine that uses a fuel preprocessing system like the one described in US 2022 / 0099020 A1 ‘Hydrogen Fuel Vaporiser’.

[0104] An engine that uses the exhaust gas to vaporise and preheat the hydrogen before it enters the combustion chamber.

[0105] An engine that uses bleed air from the turbine to vaporise and preheat the hydrogen before it enters the combustion chamber, as shown in Fig. 2.

[0106] The results of the simulation are shown in Fig. 6. The engine that uses a fuel preprocessing system like the one described in US 2022 / 0099020 A1 ‘Hydrogen Fuel Vaporiser’ reduces the overall efficiency of the propulsion system by 0.5 % compared to the theoretical baseline case. Compared to the ‘Hydrogen Fuel Vaporiser’, preheating the hydrogen fuel using the engine exhaust increases the overall efficiency by 3.6 %. However, preheating the hydrogen fuel with turbine bleed air, according to the present invention, is even better that using the exhaust gas, delivering a 4.2 % increase in overall efficiency compared to the ‘Hydrogen Fuel Vaporiser’. This demonstrates that arrangements according to the present invention demonstrate significant improvements in system efficiency in comparison to known arrangements.

[0107] Effect of turbine bleed stream on Engine Temperatures and Flow Rates

[0108] The use of a turbine bleed stream for fuel processing affects the temperatures and flow rates in the engine. Based on modelling, it has been found that the main changes on temperatures and flow rates within a system that utilizes a turbine bleed stream for fuel processing are as follows:

[0109] When the cooled bleed stream is mixed into the compressor, the inlet temperature to the high pressure compressor is reduced as a result of the intercooling effect provided by this flow arrangement. This reduces the amount of power consumed by the high pressure compressor, which has a positive impact on overall efficiency.

[0110] The exit temperature of the high pressure compressor is also reduced: this is a direct consequence of the inlet temperature being reduced.

[0111] The mass flow rate through the fan and the low pressure compressor is reduced. This is because the efficiency of the core has increased, so to meet the same engine thrust requirement, a lower mass flow rate of air is required.

[0112] Where the bleed stream is mixed into the intercase split, the mass flow rate through the high pressure compressor, combustor, high pressure turbine and low pressure turbine (before the bleed is taken) is increased. This is because the mixing of the bleed stream into the intercase split between the low pressure and high pressure compressor increases the mass flow rate through the downstream components.

[0113] The mass flow rate through the low pressure compressor (after the bleed is taken) and the core exhaust nozzle is reduced. This is because the efficiency of the core has increased, so to meet the same engine thrust requirement, a lower mass flow rate of air is required.008891418

[0114] 13

[0115] of Turbine Bleed Location

[0116] One significant advantage of using bleed air from the turbine is that both the amount of air taken as bleed and the location that the bleed air is taken from can be separately optimized.

[0117] Using the simulation described above, the optimal location to take the bleed stream from was investigated. The largest performance benefits were achieved when the bleed stream was taken from the location in the turbine where the pressure is between 2.3 and 4.2 times exit pressure of the turbine, with the optimal location to take the bleed stream from being the location in the turbine where the pressure that is 3.0 times larger than the exit pressure of the turbine.

[0118] The pressure at which the bleed stream is removed from the turbine also determines the pressure at which the bleed stream is mixed back into the compressor. As discussed above, preferably, the cooled bleed turbine gas is exhausted into the compressor at a location at which the compressor pressure is close to or substantially equal to the pressure at which the bleed gas is extracted from the turbine. This can help to ensure there is no significant pressure differential across the bleed gas flow path.

[0119] From an engine architecture point of view, the easiest location to mix the bleed stream into the compressor is at the intercase split, which is the split between the low pressure (LP) and high pressure (HP) compressor. The pressure of the intercase split in the simulated engine is 4.0 times the exit pressure of the turbine, which corresponds to taking the bleed stream at a location within the turbine which is within the range where the largest performance benefits are seen. As a result, mixing the turbine bleed flow into the intercase split offers a good compromise between optimal performance and simplicity of engine architecture.

[0120] ***

[0121] 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.

[0122] 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.

[0123] 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.008891418

[0124] 14

[0125] Any section headings used herein are for organizational purposes only and are not to be construed as limiting the subject matter described.

[0126] Throughout this specification, including the claims which follow, unless the context requires otherwise, the word “comprise” and “include”, and variations such as “comprises”, “comprising”, and “including” will be understood to imply the inclusion of a stated integer or step or group of integers or steps but not the exclusion of any other integer or step or group of integers or steps.

[0127] 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%.

Claims

1. 00889141815Claims:

1. A gas turbine arrangement comprising a gas turbine engine and a cryogenic fuel pre-processing system, the gas turbine engine comprising, in axial flow series:a compressor;a combustor; anda turbine configured to drive the compressor;wherein the cryogenic fuel pre-processing system is configured to heat and / or vaporise cryogenic fuel to provide processed fuel for delivery to the combustor and comprises:a gas outlet configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along a bleed air gas path;a heat exchanger configured to reject heat from the bleed turbine gas to the cryogenic fuel to provide cooled bleed turbine gas and heated and / or vaporised processed fuel; anda processed fuel inlet configured to deliver the processed fuel to the combustor.

2. The gas turbine arrangement according to claim 1 wherein the cryogenic fuel pre-processing system comprises one or more further gas outlets configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along one or more further bleed air gas paths.

3. The gas turbine arrangement according to claim 1 or claim 2 wherein the bleed fraction (the mass flow of the bleed air relative to the total mass flow through the engine), is selected to be in a range of from 1 to 20 %.

4. The gas turbine arrangement according to any one of the preceding claims wherein the bleed turbine gas is extracted from the turbine of the gas turbine engine at a pressure in a range between 1.4 times the exit pressure of the turbine and 0.5 times the inlet pressure of the turbine.

5. The gas turbine arrangement according to any one of the preceding claims wherein the bleed turbine gas is extracted from the turbine of the gas turbine engine at a temperature in a range between 1.1 times the exit temperature of the turbine and 0.84 times the inlet temperature of the turbine.

6. The gas turbine arrangement according to any one of the preceding claims wherein the bleed air gas path is configured to exhaust the cooled bleed turbine gas into one or more of:the turbine;the compressor;the environment;one or more further turbines of the gas turbine arrangement;one or more auxiliary systems.

7. The gas turbine arrangement according to claim 6 wherein the bleed air gas path is configured to exhaust the cooled bleed turbine gas into the compressor, and wherein the cooled bleed turbine gas is00889141816exhausted into the compressor at an intercase split between low pressure (LP) and high pressure (HP) stages of the compressor.

8. The gas turbine arrangement according to claim 6 wherein the bleed air gas path is configured to exhaust the cooled bleed turbine gas into the turbine, and wherein the cooled bleed turbine gas is exhausted into the turbine downstream of the gas outlet configured to extract bleed turbine gas, in an axial flow direction.

9. The gas turbine arrangement according to claim 6 wherein the bleed air gas path is configured to exhaust the cooled bleed turbine gas into one or more auxiliary systems, the one or more auxiliary systems being selected from a cabin pressurisation system, and / or an air conditioning system.

10. The gas turbine arrangement according to any one of the preceding claims wherein the gas turbine arrangement comprises a main fuel conduit for delivery of cryogenic fuel to the cryogenic fuel preprocessing system from a cryogenic fuel storage unit.

11. The gas turbine arrangement according to claim 10 wherein the gas turbine arrangement comprises a fuel pump configured to pump cryogenic fuel from the cryogenic fuel storage unit through the main fuel conduit to the cryogenic fuel pre-processing system.

12. The gas turbine arrangement according to claim 10 or claim 11 wherein the heat exchanger is disposed at an intersection of the main fuel conduit and the bleed air gas path.

13. The gas turbine arrangement according to any one of the preceding claims wherein the heat exchanger comprises two or more heat exchange units.

14. The gas turbine arrangement according to claim 13 wherein the heat exchange units are arranged in parallel, and are each configured to receive a predetermined fraction of the total turbine bleed air, and a predetermined fraction of the cryogenic fuel.

15. The gas turbine arrangement according to any one of the preceding claims wherein the cryogenic fuel is liquid hydrogen.

16. A method of operating a gas turbine engine comprising, in axial flow series:a compressor;a combustor; anda turbine configured to drive the compressor;the method including providing a supply of cryogenic fuel and processing the cryogenic fuel to provide processed fuel by passing the cryogenic fuel through a cryogenic fuel pre-processing system comprising:a gas outlet configured to extract bleed turbine gas from the turbine of the gas turbine engine and direct the bleed turbine gas along a bleed air gas path;00889141817a heat exchanger configured to reject heat from the bleed turbine gas to the cryogenic fuel to provide cooled bleed turbine gas and heated and / or vaporised processed fuel; anda processed fuel inlet configured to deliver the processed fuel to the combustor; and thereafter directing said processed fuel into the combustor for combustion therein.

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