Pratt & Whitney Canada

A system comprising a sensor (204) connected to an aircraft system (102) configured to monitor an aircraft system characteristic and generate sensor data responsive thereto. A diagnostic data module (DDM) (202) is configured to receive the sensor data from the sensor (104) and transmit the received sensor data from the DDM (202) to a remote monitoring unit. The DDM (202) operates independently of a preexisting data monitoring system (106) of the aircraft system (102).

An apparatus may include a first terminal, a second terminal, a direct current (DC) motor (202a, 302a, 402a, 502a, 602a), a brake mechanism (204a, 304a, 404a, 504a, 604a), and a voltage blocking mechanism. The DC motor (202a, 302a, 402a, 502a, 602a) may be configured to operate when a voltage is applied across the first terminal and the second terminal. The brake mechanism (204a, 304a, 404a, 504a, 604a) may be configured to apply braking to the DC motor (202a, 302a, 402a, 502a, 602a) when no voltage is present across the first terminal and the second terminal, and remove the braking from the DC motor (202a, 302a, 402a, 502a, 602a) when the voltage is applied across the first terminal and the second terminal. The voltage blocking mechanism may be configured to prevent the DC motor (202a, 302a, 402a, 502a, 602a) from operating when the voltage is applied across the first terminal and the second terminal until the braking from the brake mechanism is removed from the DC motor (202a, 302a, 402a, 502a, 602a).

A system (10) and method (100) for manufacturing a component (14) with optimum geometry includes determining a target mass flow rate (102) through an aperture (48) of the component (14) and determining geometric variances (104) associated with the aperture (48). Optimum geometry is selected based on the geometric variances and the target mass flow rate. Input variables for a manufacturing tool are determined that correspond to the optimum geometry and the component (14) is manufactured based on the selected input variables.

An aircraft propulsion system (20) includes a turbocompressor unit (40), an engine exhaust assembly (32), and a tube assembly (118). The turbocompressor unit (40) includes a compressor (44), a turbine (42), a turbocompressor rotational assembly (46), and a static structure (48). The static structure (48) forms a first cavity (72) and a second cavity (84). The engine exhaust assembly (32) includes an annular scroll (38) disposed between the first cavity (72) and the second cavity (84). The annular scroll (38) includes a scroll body (86) forming a flow channel (90), and a plurality of hollow stator vanes (94). The plurality of hollow stator vanes (94) is disposed within the flow channel (90). The tube assembly (118) includes a tubular body (120). The tubular body (120) extends through one of the plurality of hollow stator vanes (94) between and to the first cavity (72) and the second cavity (84). The tubular body (120) forms an internal passage (112) extending from the first axial tube end (130) to the second axial tube end (132). The internal passage (112) connects the first cavity (72) in fluid communication with the second cavity (84).

An apparatus (20) for an aircraft powerplant includes an impeller rotor (30) and a seal runner (118). The impeller rotor (30) includes a plurality of fluid circuits (106). Each of the fluid circuits (106) includes a first outlet passage (110), a second outlet passage (112) and an inlet passage (108) fluidly coupled to the first outlet passage (110) and the second outlet passage (112) in parallel. The seal runner (118) includes a bore (132), an inner guide surface (144) and an outer land surface (156). The seal runner (118) extends circumferentially about the axis (44) and radially between the inner guide surface (144) and the outer land surface (156). The bore (132) is fluidly coupled to the first outlet passage (110) of each fluid circuit (106). The inner guide surface (144) forms a radial outer peripheral boundary of the bore (132). The inner guide surface (144) radially diverges away from the axis (44) as the inner guide surface (144) extends axially away from the impeller rotor (30) and to a distal end (136) of the seal runner (118).

A shaft assembly (42) for a gas turbine engine (20) including an outer shaft (46), an inner shaft (44), a lock nut (48), a locking shaft (50), and a washer (52). The outer shaft (46) includes a bore (56) extending between first and second ends. A portion (74, 78, 82) of the inner shaft (44) extends within the bore (56) of the outer shaft (46). The lock nut (48) is secured to the inner shaft (44) proximate a second axial end (74) of the inner shaft (44). The locking shaft (50) includes a body portion (100) extending between a coupling end (102) and a retaining end (104). The locking shaft (50) is disposed within the bore (56), with the coupling end (102) engageable with an end of the lock nut (48). A washer (52) is engageable with the retaining end (104) and the outer shaft second end. In the assembled state, the locking shaft (50) prevents rotation of the lock nut (48).

An impeller rotor (30) includes a plurality of fluid circuits (106) arranged circumferentially about an axis (44) within the impeller rotor (30). Each of the fluid circuits (106) includes a first outlet passage (110), a second outlet passage (112) and an inlet passage (108) fluidly coupled to the first outlet passage (110) and the second outlet passage (112) in parallel. A seal runner (118) includes a bore (132), a sidewall (126) and a rim (128). The bore (132) is fluidly coupled to the first outlet passage (110) of each of the fluid circuits (106). The rim (128) projects radially out from the sidewall (126) to an outer end (164) of the rim (128). The rim (128) is seated in a recess (204) in the impeller rotor (30). The rim (128) includes a flat disposed at the outer end (164) of the rim (128). The flat of the rim (128) is configured to engage a flat of the impeller rotor (30) to rotationally fix the seal runner (118) to the impeller rotor (30).

An aircraft propulsion system (20) includes a combustor section (28), an engine oil system (34), and a fuel system (36). The combustor section (28) includes a combustor (44) forming a first portion of a core flow path (68) through the aircraft propulsion system (20). The engine oil system (34) includes a fuel-oil heat exchanger (80). The fuel system assembly (36) includes a jet fuel system (86) and a hydrogen fuel system (88). The jet fuel system (86) is connected in fluid communication with the combustor (44) and configured to selectively direct a jet fuel to the combustor (44). The hydrogen fuel system (88) is connected in fluid communication with the combustor (44) and configured to selectively direct a hydrogen fuel to the combustor (44). The hydrogen fuel system (88) is further connected in fluid communication with the fuel-oil heat exchanger (80).

An aircraft engine (10), has: a turbine (15) rotatable about a central axis (A); a scroll case (20) having an inlet (22) and an outlet (23) connected to the turbine (15), and a conduit (21) extending around the central axis (A) from the inlet (22) to the outlet (23); a bearing housing (30) extending around the central axis (A) and including a support flange (41); and a turbine support case (40) secured to the bearing housing (30), the turbine support case (40) having spokes (44) distributed around the central axis (A) and extending along a direction having an axial component relative to the central axis (A), the spokes (44) extending through the scroll case (20) and radially supported by the bearing housing (30), one of the turbine support case (40) and the scroll case (20) defining lugs (26) circumferentially distributed around the central axis (A), the other of the turbine support case (40) and the scroll case (20) defining slots (48), the turbine support case (40) and the scroll case (20) circumferentially locked to one another via the lugs (26) engaging the slots (48).

An engine includes an engine oil system. The engine oil system includes a pump assembly (70) forming a portion of an oil flow path. The pump assembly (70) includes a drive shaft (72), a plurality of pumps (76), and a pump housing (74). The drive shaft (72) extends along a rotational axis (82). The plurality of pumps (76) are arranged axially along the rotational axis (82). Each of the plurality of pumps (76) includes a pump rotor (104) coupled with the drive shaft (72). The pump housing (74) extends along the rotational axis (82) between a first housing end (86) and a second housing end (88). The pump housing (74) circumscribes and houses the drive shaft (72) and the plurality of pumps (76) between the first housing end (86) and the second housing end (88). The pump housing (74) includes a plurality of housing body portions (90A-E) axially stacked along the rotational axis (82) to form the pump housing (74). The plurality of housing body portions (90A-E) collectively form an internal oil passage (92) through the pump housing (74).