Pratt & Whitney Canada Patente

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An aircraft propulsion system (20) is provided that includes a turbine engine (24) and an intercooler, IC, (42). The turbine engine (24) includes a source of non-hydrocarbon fuel, and a low pressure compressor, LPC, (34A) having an LPC air inlet and outlet, and a high pressure compressor, HPC, (34B) having an HPC air inlet and outlet. The intercooler (42) has an IC air inlet (48), an IC air outlet (50), a U-shaped air passage (52), and crossflow tubes (58). The IC air inlet (48) is in fluid communication with the LPC air outlet. The IC air outlet (50) is in fluid communication with the HPC air inlet. The U-shaped air passage (52) is in fluid communication with the IC air inlet (48) and outlet (50). The crossflow tubes (58) extend through the U-shaped air passage (52). The crossflow tubes (58) are in fluid communication with the source of non-hydrocarbon fuel, and are configured to contain a flow of the non-hydrocarbon fuel therethrough.

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An aircraft engine diagnostic system (100) includes at least one sensor (102, 104, 106) and a diagnostic controller (110). The at least one sensor (102, 104, 106) measures a parameter of a gas turbine engine (20). The diagnostic controller (110) is in signal communication with the at least one sensor (102, 104, 106). The diagnostic controller (110) determines an actual thrust output of the gas turbine engine (20) and an execution of at least one control system limiter configured to limit control of the gas turbine engine (20) based on the measured parameter. The diagnostic controller (110) further determines an existence of an engine fault based on the actual thrust output of the gas turbine engine (20) and an execution of the at least one control system limiter.

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An engine system (10) is provided that includes an open propulsor rotor (12), an engine core assembly (14), a bypass flowpath (48) and a condenser (118). The engine core assembly (14) is configured to drive rotation of the open propulsor rotor (12). The engine core assembly (14) includes a core flowpath (30), a compressor section (26), a combustor section (27) and a turbine section (28). The core flowpath (30) extends through the compressor section (26), the combustor section (27) and the turbine section (28) from an inlet (38) into the core flowpath (30) to an exhaust (40) from the core flowpath (30). The bypass flowpath (48) bypasses the engine core assembly (14). The condenser (118) is configured to exchange heat energy between combustion products flowing within the core flowpath (30) downstream of the combustor section (27) and bypass air flowing within the bypass flowpath (48).

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A fuel nozzle (70) for a gas turbine engine (20) includes a first fuel nozzle body (72) and a second fuel nozzle body (74). The first fuel nozzle body (72) and the second fuel nozzle body (74) form a fuel passage (78). The fuel passage (78) includes a fuel accumulation chamber (80), an upstream fuel passage (112), a downstream fuel passage (114), and a plurality of tangential feed holes (116). The upstream fuel passage (112) extends along an upstream fuel passage centerline (118) to a terminal end (120) of the upstream fuel passage (112). The terminal end (120) is spaced from the accumulation chamber (80). The downstream fuel passage (114) extends from a passage inlet (126) along a downstream fuel passage centerline (124). The passage inlet (126) is disposed at the accumulation chamber (80). The plurality of tangential feed holes (116) are formed by the first fuel nozzle body (72). Each tangential feed hole (116) of the plurality of tangential feed holes (116) extends between and to a hole inlet (128) and a hole outlet (130). The hole inlet (128) is disposed at the upstream fuel passage (112) upstream of the terminal end (120). The hole outlet (130) is disposed at the accumulation chamber (80).

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An engine system (10) is provided that includes an engine core assembly (14) and a fuel system (16). The engine core assembly (14) includes a core flowpath (32), a core compressor section (28), a core combustor section (29) and a core turbine section (30). The core flowpath (32) extends through the core compressor section (28), the core combustor section (29) and the core turbine section (30) from an inlet (38) into the core flowpath (32) to an exhaust (40) from the core flowpath (32). The core compressor section (28) includes a core compressor rotor (44). The combustor section (29) includes a combustor (34). The fuel system (16) includes a fuel flowpath (62), a fuel turbine section (66) and a fuel injector (68). The fuel flowpath (62) extends through the fuel turbine section (66) to the fuel injector (68). The fuel turbine section (66) includes a fuel turbine rotor (76) coupled to and rotatable with the core compressor rotor (44). The fuel injector (68) is configured to direct fuel received from the fuel flowpath (62) into the combustor (34).

A rotary engine (10) has: housings (11,18) secured to one another, the housings (11,18) including a first side housing (11), a second side housing (11), and a rotor housing (18) disposed between the first side housing (11) and the second side housing (11); and a rotor (22) rotationally received within a rotor cavity (20) defined by the first side housing (11), the second side housing (11), and the rotor housing (18); wherein the first side housing (11), the rotor housing (18), and the second side housing (11) are cooled in parallel via respective coolant passages (11C,18C) including: a first side housing coolant passage (11C) extending through the first side housing (11); a second side housing coolant passage (11C) extending through the second side housing (11); and a rotor housing coolant passage (18C) extending through the rotor housing (18), and wherein the coolant passages (11C,18C) are free of inter-passage connection between the housings (11,18).

A dual bypass turbofan gas turbine engine (12) includes counter-rotating first (20) and second (22) fans through which ambient air is received and propelled in an aft direction; a first bypass duct (24) receiving first bypass air including a first portion of the propelled air; a second bypass duct (26) receiving second bypass air including a second portion of the propelled air; a core gas path (28) receiving core air including a third portion of the propelled air; a compressor (40,44) pressurizing the core air; an intercooler (42) facilitating heat removal from the pressurized core air; a combustor (46) in which the pressurized core air is mixed with fuel and ignited to generate a stream of combustion gas; a turbine (48,50) extracting energy from the combustion gas; and a differential gear train (54) apportioning an input torque received from the turbine (48,50) between the first fan (20) and the second fan (22).

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