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EP-4303416-B1 - TURBO EXPANDERS FOR TURBINE ENGINES HAVING HYDROGEN FUEL SYSTEMS

EP4303416B1EP 4303416 B1EP4303416 B1EP 4303416B1EP-4303416-B1

Inventors

  • HOLLEY, BRIAN M.
  • STAUBACH, JOSEPH B.
  • MULDOON, MARC J.
  • LENTS, CHARLES E.

Dates

Publication Date
20260506
Application Date
20230707

Claims (8)

  1. An aircraft propulsion system (400; 500), comprising: aircraft systems (404; 504) comprising at least one hydrogen tank (432) and an aircraft-systems heat exchanger (458); and engine systems (402; 502) comprising at least a main engine core, a high pressure pump (416), a hydrogen-air heat exchanger (424), and a turbo expander (420; 520; 602; 700; 800; 900), wherein the main engine core comprises a compressor section (24), a combustor section (26) having a burner (410), and a turbine section (28), wherein: hydrogen is supplied from the at least one hydrogen tank (432) through a hydrogen flow path (444), passing through the aircraft-systems heat exchanger (458), the high pressure pump (416), the hydrogen-air heat exchanger (424), and the turbo expander (420 ... 900), prior to being injected into the burner (410) for combustion; the turbo expander (420; ...; 900) comprises a rotor (704; 804) separated into a first expander portion (712; 812; 902) and a second expander portion (714; 814; 904) arranged about an output shaft (610;722; 806; 908); the aircraft propulsion system further comprises a generator (422; 522; 606; 906) configured to generate electrical power; and the output shaft (610;722; 806; 908) is operably connected to the generator (422; 522; 606; 906), characterised in that : the aircraft propulsion system (400; 500) further comprises a secondary tank (628) configured to contain a secondary fluid that is supplied to the turbo expander (420; ...; 900) and configured to provide at least one of seal buffering and starting flow for the turbo expander (420; ...; 900); and the secondary tank (628) contains high pressure gaseous hydrogen.
  2. The aircraft propulsion system (400; 500) of claim 1 wherein the first expander portion (712; 812; 902) of the rotor (704; 804) includes at least five blade rows (710; 810; 910) and the second expander portion (714; 814; 904) of the rotor (704; 804) includes at least seven blade rows (710; 810; 912).
  3. The aircraft propulsion system (400; 500) of any preceding claim, further comprising a gear box (604) operably coupled between the output shaft (610) and the generator (606).
  4. The aircraft propulsion system (400; 500) of any preceding claim, wherein the generator (422; 522; 606; 906) is configured to generate at least 1 MW of electrical power.
  5. The aircraft propulsion system (400; 500) of any preceding claim, wherein electrical power from the generator (422; ...; 906) is supplied to a wing anti-ice system (436; 536).
  6. The aircraft propulsion system (400; 500) of claim 1, wherein the turbo expander is a radial inflow turbo expander (900).
  7. The aircraft propulsion system (400; 500) of any of claims 1 to 5, wherein the turbo expander is an axial inflow turbo expander (700; 800).
  8. The aircraft propulsion system (400; 500) of any preceding claim, wherein a flow through the first expander portion (712) is in a first direction and a flow through the second expander portion (714) is in a second direction opposite the flow through the first expander portion (712).

Description

TECHNICAL FIELD The present invention relates generally to turbine engines and aircraft engines, and more specifically to turbo expanders for use when employing hydrogen fuel systems and related systems with turbine and aircraft engines. BACKGROUND Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section drives the compressor section to rotate. In some configurations, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine. Typically, liquid fuel is employed for combustion onboard an aircraft, in the gas turbine engine. The liquid fuel has conventionally been a hydrocarbon-based fuel. Alternative fuels have been considered, but suffer from various challenges for implementation, particularly on aircraft. Hydrogen-based and/or methane-based fuels are viable effective alternatives which may not generate the same combustion byproducts as conventional hydrocarbon-based fuels. The use of hydrogen and/or methane, as a gas turbine fuel source, may require very high efficiency propulsion, in order to keep the volume of the fuel low enough to feasibly carry on an aircraft. That is, because of the added weight associated with such liquid/compressed/supercritical fuels, such as related to vessels/containers and the amount (volume) of fuel required, improved efficiencies associated with operation of the gas turbine engine may be necessary. EP 3623603 A1 relates to a gas turbine engine using non-traditional cooled liquid fuel to fuel the engine, cool electronics, and drive a turbo-generator. BRIEF SUMMARY According to a first aspect of the invention, aircraft propulsion systems are provided as claimed in claim 1. Some embodiments of the invention are as claimed in the dependent claims. The foregoing features and elements may be executed or utilized in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, that the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting. BRIEF DESCRIPTION OF THE DRAWINGS The subject matter is particularly pointed out and distinctly claimed at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: FIG. 1 is a schematic cross-sectional illustration of a gas turbine engine architecture that may employ various embodiments disclosed herein;FIG. 2 is a schematic illustration of a turbine engine system in accordance with an embodiment of the present disclosure that employs a non-hydrocarbon fuel source;FIG. 3 is a schematic diagram of an aircraft propulsion system that is not in accordance with the claims of this application;FIG. 4 is a schematic diagram of an aircraft propulsion system in accordance with an embodiment of the present disclosure;FIG. 5 is a schematic diagram of an aircraft propulsion system in accordance with an embodiment of the present disclosure;FIG. 6 is a schematic diagram of a generator system in accordance with an embodiment of the present disclosure;FIG. 7 is a schematic diagram of a turbo expander of a generator system in accordance with an embodiment of the present disclosure;FIG. 8 is a schematic diagram of a turbo expander of a generator system in accordance with an embodiment of the present disclosure; andFIG. 9 is a schematic diagram of a turbo expander of a generator system in accordance with an embodiment of the present disclosure. DETAILED DESCRIPTION FIG. 1 schematically illustrates a gas turbine engine 20. As illustratively shown, the gas turbine engine 20 is configured as a two-spool turbofan that has a fan section 22, a compressor section 24, a combustor section 26, and a turbine section 28. The illustrative gas turbine engine 20 is merely for example and discussion purposes, and those of skill in the art will appreciate that alternative configurations of gas turbine engines may employ embodiments of the present disclosure. The fan section 22 includes a fan 42 that is configured to drive air along a bypass flow path B in a bypass duct defined in a fan case 23. The fan 42 is also configured to drive air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbo