Pratt & Whitney Canada Patente

An annular containment structure (34) for a rotor stage (32) of a gas turbine engine (20) is provided that includes an outer radial panel (38), a plurality of contact members (40), and a deformable structure (42). The outer radial panel (38) has an inner radial surface (36A) and an outer radial surface (36B). The contact members (40) are distributed around a circumference of the containment structure (34). The deformable structure (42) is engaged with the inner radial surface (36A) of the outer radial panel (38) and the plurality of contact members (40). The deformable structure (42) is configured to dissipate energy by plastically deforming. Each contact member (40) includes an interior segment (40A) and an inner perimeter segment (40B) that are connected to one another. The inner perimeter segments (40B) form at least a portion of an inner radial surface (36A) of the containment structure (34).

An engine (22) for an aircraft (1000) includes a gear box (68), an engine oil system (70), a starter generator (72), and a starter generator cooling circuit (92). The gear box (68) includes a gear assembly (74). The engine oil system (70) includes an oil pump (76), a heat exchanger (80), and an oil tank (78) forming an oil flow path (84) of the engine oil system (70). The oil pump (76) is coupled with the gear assembly (74). The oil pump (76) is operable to pump an engine oil along the oil flow path (84) through the heat exchanger (80) to the oil tank (78). The heat exchanger (80) includes a primary side (88) and a secondary side (90). The primary side (88) forms the oil flow path (84). The starter generator (72) is coupled with the gear assembly (74). The starter generator cooling circuit (92) has a coolant flow path (94) formed by the secondary side (90) and the starter generator (72).

A method includes flowing a fuel to a primary fuel manifold (128A) of a gas turbine engine (100), flowing the fuel to a secondary fuel manifold (128B) of the gas turbine engine (100) at a fuel flow rate, monitoring a compressor discharge pressure of the gas turbine engine (100), determining a ratio of the fuel flow rate of the secondary fuel manifold (128B) and the compressor discharge pressure, and adjusting the fuel flow rate of the fuel to the secondary fuel manifold (128B) based on the ratio.

A method includes receiving one or more flight condition inputs (202; 402) at a controller (152; 204; 404), determining a viability of a plurality of fuel mix ratios based on the one or more flight condition inputs (202; 402), selecting a fuel mix ratio of the plurality of fuel mix ratios, dispensing two or more fuels or solutions to a gas turbine engine (100; 218; 418) based on the selected fuel mix ratio, and monitoring performance parameters of the gas turbine engine (100; 218; 418) using a plurality of sensors (144; 220; 420) coupled to the controller (152; 204; 404). The two or more fuels or solutions may include water.

A method of balancing a rotational spool (50) within an assembled gas turbine engine (26) is provided. The rotational spool (50) includes turbine and compressor sections, and a shaft (46) that connects the turbine and compressor sections. The method includes: disposing a circumferentially extending track component (68) to rotate with the rotational spool (50); engaging a plurality of balance weights (70) with the track component (68) prior to assembly of the gas turbine engine (26); rotating the rotational spool (50) of the assembled gas turbine engine (26) and determining a presence or absence of an imbalance condition in the rotational spool (50); when an imbalance condition is determined: rotating the rotational spool (50) until the first balance weight (70) is circumferentially positioned relative to an access port (74); circumferentially moving the first balance weight (70) relative to the track component (68), and subsequently actuating the first balance weight (70) to the secured configuration.

A method includes obtaining at a first sampling time, a first set of n sampled values of n control parameters for one or more control systems of an aircraft. The method includes performing, in a higher power mode, a first sequential lookup and interpolation operation of a table maintained in the memory (110) to obtain a first output control value, and sending, the first output control value to the one or more control systems of the aircraft. Determining that at least one condition associated with a first steady state has been satisfied and switching to a first lower power mode of the device (100). In the first lower power mode, performing a second sequential lookup and interpolation operation to obtain a second output control value and sending the second output control value to the one or more control systems of the aircraft.

An aircraft engine (10) includes a gas path annularly extending around a central axis in serial flow communication across a compressor (14) and turbine (18). The turbine (18) has a casing (26), a shroud (32) secured to the casing (26), and blades (28) radially inward of the shroud (32), with a clearance defined between blade tips and the shroud (32). An active tip clearance control device (36) for directing a flow of cooling air to the casing (26) includes a first cooling air source (40) at a first temperature, a second cooling air source (42) at a second temperature than the first temperature, and a valve (46) having a first inlet fluidly connected to the first air source (40), a second inlet fluidly connected to the second air source (42), and an outlet selectively connectible to the first and second inlets. The outlet is fluidly connected to the casing (26) to selectively direct the flow of air from the first or second air source (40, 42) to the casing (26).

A turbine exhaust duct (144) for a gas turbine engine is provided that includes an exhaust manifold (50), an outer duct (52), first and second connecting panels, a first stiffener panel (64A), and a second stiffener panel. The exhaust manifold (50) includes an annular inlet and first and second exhaust ports (58). The outer duct (52) includes a wall panel, an axial end aperture, a first lateral side aperture, and a second lateral side aperture. The first connecting panel includes a first exhaust port connecting flange and a first lateral side outer duct flange. The first stiffener panel (64A) has a periphery flange that is attached to the first connecting panel and to the outer duct wall panel. The second stiffener panel has a periphery flange that is attached to the first connecting panel and to the outer duct wall panel.

A gas turbine engine (26) is provided that includes a combustor (38), a rotational spool (50), a spool balancing system, and at least one access port (74). The rotational spool (48) includes turbine and compressor sections, and a shaft that connects the turbine and compressor sections. The spool balancing system includes a circumferentially extending track component (68) and a plurality of balance weights (70) that are engaged with the track component (68). The balance weights (70) are disposable in secured and unsecured configurations. The track component (68) and the balance weights (70) are disposed to rotate with the rotational spool (50). The access port (74) is disposed within a non-rotational member of the gas turbine engine (26). The access port (74) is configured to receive a tool for engagement with the balance weights (70).

An apparatus comprising a propeller (102;304) and an engine (104;302;504) configured to drive the propeller (102;304). A controller (108;306;502) is configured to control fuel flow to the engine (104;302;504) to decelerate the propeller (102;304) responsive to detection of a propeller speed exceeding a predetermined threshold.

Apparatus and methods of monitoring a lubrication system (12) of an aircraft power plant are provided. The apparatus and method can detect a human error associated with servicing an aircraft engine. The method includes sensing an actual oil level (34) in an oil tank (26) of the lubrication system (12) of the aircraft power plant, and determining that the actual oil level (34) is below a habitual replenish level (54) that is higher than a minimum operating oil level for the oil tank (26). When the actual oil level (34) is below the habitual replenish level (54) an alert is generated.

An apparatus includes a propeller (102;304), an engine (104;302;504) configured to drive the propeller (102;304). The apparatus also includes a controller (108;306;502) configured to control fuel flow to the engine (104;302;504) responsive to the propeller (102;304) operating in a first mode of operation and to control fuel flow to the engine (104;302;504) in a second mode of operation responsive to the propeller (102;304) operating in an uncontrollable high thrust condition.

A de-aerator (103) for an aircraft power plant (10), has: a housing (108) defining an air-oil inlet (109), an oil outlet (110), and an air outlet (111), the housing (108) having a forward end (108A) and a rearward end (108B) opposite the forward end (108A), the air-oil inlet (109) defined at the forward end (108A); a rotor (115) received within the housing (108) and rotatable relative to the housing (108) about a rotation axis (A); and a shaft (116) rotatable about the rotation axis (A), the shaft (116) extending axially from a fore end (116A) at the forward end (108A) of the housing (108) to a rear end (116B) at the rearward end (108B) of the housing (108) relative to the rotation axis (A), the rotor (115) drivingly engaged to the shaft (116) at the fore end (116A) of the shaft (116), the rear end (116B) of the shaft (116) defining a coupler (116C) drivingly engageable to a power source for rotation of the shaft (116) and of the rotor (115) mounted thereto.

A de-aerator (103) has: a housing (108) defining an air-oil inlet (109), an oil outlet (110), and an air outlet (111); a rotor (115) within the housing (108) and rotatable about a rotation axis (A) and having: inner and outer rims (123A, 123B) and blades (123), passages between the blades (123) and communicating with the air-oil inlet (109); a conduit (127) extending around the rotation axis (A) and defining a central passage (125B) extending axially and communicating with the air outlet (111), the conduit (127) being radially spaced apart from the outer rim (123B) by radial passages (126) communicating with the oil outlet (110); and openings (124) located inwardly of the blades (123) and communicating with the central passage (125B), wherein one or more of: a cross-sectional area of the central passage (1258) taken on a plane transverse to the rotation axis (A) increasing in an axial direction away from the air-oil inlet (109); and the openings (124) located at least partially radially between an inner face of the conduit (127) and an outer face of the inner rim (123A).

A method for controlling a variable-pitch propeller (24) of an aircraft (12) includes using a feedback controller (62, 76), operating the variable-pitch propeller (24) at a current operating condition including a current rotational speed (NPf), determining a speed error (E) between the current rotational speed (NPf) of the variable-pitch propeller (24) and a set point rotational speed (NPref) for the variable-pitch propeller (24), determining a gain (Kp, Kd) for the feedback controller (62, 76) based on a power coefficient (Cp) and an advance ratio (J) of the variable-pitch propeller (24). Using the feedback controller (62, 76), a propeller pitch command (Co) is determined based on the speed error (E) and using the gain (Kp, Kd). The pitch of the variable-pitch propeller (24) is adjusted according to the propeller pitch command (Co) to reduce the speed error (E).

A method for testing a material strength of a bearing inner ring (104) includes mounting the bearing inner ring (104) on a test shaft (156) to form a ring test assembly (162). The bearing inner ring (104) includes a ring material having a first coefficient of thermal expansion. The test shaft (156) includes a test shaft (156) material having a second coefficient of thermal expansion. The second coefficient of thermal expansion is greater than the first coefficient of thermal expansion. The method further includes heating the ring test assembly (162), cooling the ring test assembly (162), and testing the material strength of the bearing inner ring (104). The testing includes performing a plurality of cycles of the steps of heating the ring test assembly and cooling the ring test assembly.

A hybrid-electric propulsion, HEP, system (20) for powering an aircraft is provided. The system (20) includes a gearbox (26), a gas turbine engine (22), and an electric motor (24). The gearbox (26) includes an input shaft (46), an output shaft (54), a drive gear (52), a first layshaft assembly (48), a second layshaft assembly (50), and a first gearbox accessory mount (56A). The gas turbine engine (22) is in drive communication with the input shaft (46) of the gearbox (26). The electric motor (24) is engaged with the first gearbox accessory mount (56A). The first layshaft, FL, assembly (48) includes a first layshaft (68), a first FL gear (64), and a second FL gear (66). The first and second FL gears (64, 66) are attached to the first layshaft (68). The first FL gear (64) is engaged with the input shaft (46) and the second FL gear (66) is engaged with the drive gear (52). The first gearbox accessory mount (56A) is in geared communication with the drive gear (52) independent of the first layshaft assembly (48).

An aircraft propulsion system is provided that includes a thermal engine (20) and a gearbox (36). The gearbox (36) includes an input shaft (46), an output shaft (48), a first epicyclic gear arrangement (42), a drive gear (44), and a first layshaft assembly (50). The first epicyclic gear arrangement (42) includes a sun gear (54), a plurality of planet gears (56), and a ring gear (60). The sun gear (54) is connected with the drive gear (44). The first layshaft assembly (50) includes a first layshaft input gear (70), a first layshaft ring engagement gear (72), and a first clutch (74). The first layshaft ring engagement gear (72) is engaged with the ring gear (60), and the first layshaft input gear (70) is engaged with the drive gear (44) and the input shaft (46). The first clutch (74) is disposable in a clutch-disengaged configuration and in a clutch-engaged configuration.

Torque measurement systems for aircraft propulsion systems include a gear assembly (200) having a shaft (202) and a bearing assembly (212) configured to rotationally support the shaft (202) within a housing (208). The bearing assembly (212) includes an inner bearing raceway (218) arranged on the shaft (202), an outer bearing raceway (220) arranged on the housing (208), a rolling element (216) arranged between the bearing raceways (218) and configured to travel along a path defined between the bearing raceways (218), and a bearing cage (228) arranged between the bearing raceways (218). The rolling element (216) is configured drive rotational movement of the bearing cage (228). An axial load member (222) is arranged to apply an axial load to the outer bearing raceway (220), a rotational velocity sensor (232) is configured to measure a rotational velocity of the bearing cage (228), and a controller (242) is configured to calculate a transferred toque of the gear assembly (200) based on a detected rotational velocity of the bearing cage (228).

A maintenance system (100) for performing maintenance of an aircraft engine (10) while the aircraft engine is still mounted to an aircraft (1), the system (100) having: a support fixture assembly (110) having a static member (111) removably securable to a first case (18C) of the aircraft engine (10) and a mobile member (112) removably securable to a second case (18D) of the aircraft engine (10), the static member (111) and the mobile member (112) being slidably engaged to one another; a brace fixture (120) secured to the static member (111); a shaft supporting rod (130) removably securable to the brace fixture (120), the shaft supporting rod (130) being centered relative to the case assembly (18) when supporting a shaft (22) of the aircraft engine (10), the shaft supporting rod (130) having a shaft-engageable end (131) removably securable to an end of the shaft (22) of the aircraft engine (10); and an intermediate support member (140) removably secured to the static member (111) and extending transversally therefrom, the intermediate support member (140) having a remote end (141) engageable to the shaft (22) of the aircraft engine (10).

A method of servicing an aircraft engine (10) mounted to an aircraft, the aircraft engine having a central axis (17) and a case assembly (18) including a first case (18D) secured to a second case (18C) and being axially offset from one another, the method includes: while the aircraft engine remains mounted to the aircraft: radially supporting, relative to the central axis, an end of a shaft (22) of the aircraft engine through a central passage (18F) extending axially through the first case; creating a gap (G) between the first case and the second case by axially moving the first case away from the second case and relative to the shaft while preventing axial and radial movements of the shaft relative to the central axis; radially and axially supporting the shaft at the gap; and removing the first case by disconnecting the end of the shaft while maintaining the shaft radially and axially supported at the gap.

A powerplant (301) of an aircraft (201) is provided and includes a primary tank (310) fluidly connected to a motor (202) via a lubrication circuit (330) that is operable to provide lubricant to a part of the motor (202) when the aircraft (201) is in use, the primary tank (310) having pre-defined maximum and minimum lubricant levels, a secondary tank (320) disposed relative to the primary tank (310) such that a portion of the secondary tank (320) is disposed height-wise between the maximum level and the minimum level of the primary tank (310) when the aircraft (201) is stationary on horizontal terrain and air and lubricant conduits (332, 331)).

A centrifugal compressor includes an impeller (18a) rotatable about a central axis (26), and a scroll housing (28) enclosing the impeller (18a). The scroll housing (28) includes a housing inlet (30) configured to direct an airflow (32) to the impeller (18a) for compression, and a scroll (36) positioned downstream of an impeller outlet (46). The scroll (36) is configured to direct a compressed airflow (32) toward a housing outlet (34). A plurality of diffuser pipes (38) are positioned in the scroll housing (28) and are configured to direct the compressed airflow (32) from the impeller outlet (46), through the plurality of diffuser pipes (38) and into the scroll (36). An aerodynamic sleeve (48) encloses at least one diffuser pipe (38; 38a, 38b) of the plurality of diffuser pipes (38).

An aircraft engine includes a scroll case (20) extending around a central axis (A) and having an inlet (22) fluidly connected to a source of combustion gases and an outlet (23), and a conduit (21) extending around the central axis (A) from the inlet (22) to the outlet (23). The conduit (21) includes an outer wall (26) and a radially inward inner wall (25), the outlet (23) defined radially between the inner wall (25) and the outer wall (26). A turbine (15) is downstream of the outlet (23) of the scroll case (20) relative to a flow of the combustion gases. Vanes (61) are circumferentially distributed around the central axis (A), the vanes (61) being removably mounted to the scroll case (20) proximate to the outlet (23) of the conduit (21) and radially between the inner wall (25) and the outer wall (26).

A hybrid-electric propulsion system (24) for an aircraft is provided that includes a thermal engine, an electric motor (28), a propulsion unit (32), a gearbox (22), and an electric motor disconnect system (52). The gearbox has a housing (46) and is configured to receive a first rotational drive input from the thermal engine and a second rotational drive input from the electric motor. The gearbox is also configured to transfer an output rotational drive to the propulsion unit. The electric motor disconnect system is disposed within the housing of the gearbox. The electric motor disconnect system is disposable in a connected state wherein the electric motor is engaged with the gearbox to provide the second rotational drive input to the gearbox, and disposable in a disconnected state wherein the electric motor is disengaged from providing the second rotational drive input to the gearbox.

A method for inspecting an acoustic liner (30) positioned in an inner cavity (34) of a component (22) of an aircraft engine (10) includes: emitting, through a drain hole (22a) of the component (22) and through a section (35; 35') of the acoustic liner (30) to be inspected, a sound signal (41a) having an initial intensity; capturing, via a sound capture device (42) positioned inside the inner cavity (34), a transmitted intensity of the sound signal (41b) following its transmission through the section (35; 35') of the acoustic liner (30) to be inspected; and determining a loss in signal intensity through the section (35; 35') of the acoustic liner (30) to be inspected by comparing the transmitted intensity to the initial intensity. Upon detecting that the loss in signal intensity is within a predetermined loss in signal intensity envelope for the acoustic liner (30), an indication that the acoustic liner (30) is acceptable is emitted.

A method for improvement of T6 probes accuracy based on swirl angle includes predicting a swirl angle of a gas turbine engine (102a, 102b) based on a power of the gas turbine engine (102a, 102b) and a delta value. The method includes determining an error value of an exhaust temperature (T6) of the gas turbine engine (102a, 102b) based on the predicted swirl angle and a correlation function (400) of swirl angle to error. Also, the method includes compensating a measured value of the T6 based on the error value.

A casing (20) for a gas turbine engine (10), including: a containment fabric layer (32) wrapped around a structural shell (28); and a bi-material shell (80) located around the containment fabric layer (32), the bi-material shell (80) formed from a plurality of bi-material shell segments (81), each of the plurality of bi-material shell segments (81) include an inner layer (82) and an outer layer (84), the outer layer (84) comprises a plurality of individual segments (85) that form the outer layer (84) and a plurality of embedded fuse lines (86) defined by the plurality of individual segments (85) of the outer layer (84), wherein an overall structure of the bi-material shell (80) does not impede a rapid expansion of the containment fabric layer (32).

An aircraft powerplant (20) apparatus includes a rotating structure (70) configured to rotate about an axis (72). The rotating structure (70) includes a bladed rotor, a first component, a second component, a first balancing mass (150A), a second balancing mass (150B) and a spring element (152). The first balancing mass (150A) and the second balancing mass (150B) are configured to rotationally balance the rotating structure (70) about the axis (72). The first component include a plurality of first component teeth (160). The second component includes a plurality of second component teeth (226). The first balancing mass (150A) includes a first mass ring (174A) and a plurality of first mass teeth (178A). The first mass teeth (178A) are meshed with the first component teeth (160). The second balancing mass (150B) includes a second mass ring (174B) and a plurality of second mass teeth (178B). The second mass teeth (178B) are meshed with the second component teeth (226). The spring element (152) is engaged with the first balancing mass (150A) and the second balancing mass (150B).

An exhaust system for an aircraft engine has: a turbine exhaust duct, TED, (30) having an annular inlet conduit extending around a central axis for directing combustion gases generally in an axial direction, and outlet conduits branching off from the annular inlet conduit; and an exhaust case (18D) surrounding the TED, the exhaust case being frustoconical and having outlet openings (18E) each communicating with a respective one of the outlet conduits, a probe opening extending through the exhaust case and bounded by a periphery, and a probe interface (40) mounted to the periphery of the probe opening, the probe interface including an interfacing plate secured to the periphery of the probe opening and defining an aperture sized to receive a probe therethrough, the aperture located radially inwardly of the probe opening.

A gas turbine engine (20) is provided that includes compressor, combustor, and turbine sections. The turbine section (28) includes a rotor stage (42) having a bladed rotor (48) and a shroud (50). The bladed rotor (48) includes a disk (54) and a plurality of rotor blades (52) extending radially outward from the disk (54). Each rotor blade (52) has a leading edge (58) and a blade tip surface (66). The shroud (50) is disposed radially outside of the bladed rotor (48), and circumferentially around the bladed rotor (48). The shroud (50) includes a primary segment (PS) (50A) and a forward segment (FS) (50B). The primary segment (50A) has a PS forward end surface (72) and a PS inner radial surface (78). The forward segment (50B) is disposed forward of the rotor blades (52) and is configured to deflect core gas flow away from a blade tip gap (86) disposed between the PS inner radial surface (78) and the blade tip surface (66) of each said rotor blade (52) of the plurality of rotor blades.

Gas turbine engines include an engine core defining an engine axis and having compressor section (300), combustor, and turbine sections arranged along the engine axis. A core flow path is defined through the sections of the engine core. A casing (302) of the compressor section (300) defines, in part, a portion of the core flow path, with a bleed air cavity (310) defined between the casing (302) of the compressor section (300) and an engine housing (312). An impeller (308) is arranged between the compressor (300) and combustor sections, the impeller (308) include an impeller shroud (316). A bleed aperture (314) is defined axially between the casing (302) and the impeller shroud (316) with a bleed flow extension (324) extending from the casing (302) and into the bleed air cavity (310) and a diffuser channel (322) is defined between a surface of the bleed flow extension (324) and a surface of the impeller shroud (316), with the diffuser channel (322) having an inlet area that is less than an outlet area.

A system (30) for use in the manufacturing of an aircraft engine (10) includes an engine component (20) to be machined. The component (20) has an opening (23) leading to an inner cavity (22) in an axial direction. The inner cavity (22) has an inner wall (25) with a tapered portion (25c) extending axially and a surface to be clamped (25d) extending radially outwardly from the opening (23) to the tapered portion (25c). The system (30) further includes a clamping device (40), for clamping the component (20), having a base (41), a shaft (42) displaceable relative to the base (41) and inserted through the opening (23), and a clamp (44) coupled to the shaft (42). The clamp (44) is movable, based on a position of the shaft (42) relative to the base (41), between an undeployed state in which the clamp (44) is free of contact with the inner wall (25) of the component (20) and a deployed state in which the clamp (44) engages the surface to be clamped (25d) to axially retain the component (20).

The tool (30) can be used for manipulating a tang (22) of a helical insert (20) during installation. The tool (30) can have a guide member (32) having a working tip (38), and a holder (40). The tool (30) can have a stem member (42) slidingly engaged with the guide member (32), the stem member (42) having a tang catch (44) adjacent the first end (34), and an actuator (46) protruding from the second end (36); and a biasing member (48) coupled between the guide member (32) and the stem member (42), the biasing member (48) biasing the stem member (42) to a rest position relative the guide member (32), in which rest position the tang catch (44) is retracted into the working tip (38).

A system for detecting a shaft shear comprises a gas turbine engine (102) having a turbine (208), a high inertia rotating component (210) and a driveshaft (212). The turbine (208) drives the high inertia rotating component (210) using the driveshaft (212). A first speed probe (202) is configured to detect a first speed of the driveshaft (212) at a first location associated with the driveshaft (212). A second speed probe (204) is configured to detect a second speed of the driveshaft (212) at a second location associated with the driveshaft (212). Control circuitry (302, 304) is configured to compare the first speed and second speed, determine if the shaft shear has occurred responsive to the comparison and generate a control signal to the gas turbine engine (102) responsive to the determination.

A double row stator assembly (200) for a gas turbine engine (100) includes a first vane row stator assembly (202) that includes a plurality of first-row vanes, a second vane row stator assembly (204) that includes a plurality of second-row vanes, and a wavy-edged flange (206) configured to secure the first vane row stator assembly to the second vane row stator assembly when the first vane row stator assembly and second vane row stator assembly are assembled to form the double row stator assembly. The second vane row stator assembly is configured to be positioned to be sequentially downstream of the first vane row stator assembly when the first vane row stator assembly and the vane row second stator assembly are assembled to form the double row stator assembly such that trailing edges (202c) of the first-row vanes (202a) overlap leading edges (204b) of the second-row vanes (204a).

A system for detecting shaft shear comprises a gas turbine engine (102) including at least one turbine (114, 118) configured to drive a fan (108) using a driveshaft (116, 120). A plurality of frangible plugs (104) is located downstream of the at least one turbine (114, 118) of the gas turbine engine (102). The plurality of frangible plugs (104) is configured to provide an indication of shearing of the driveshaft (116, 120) responsive to breakage of at least a portion of the plurality of frangible plugs (104) caused by impact of the at least one turbine (114, 118) with the portion of the plurality of frangible plugs (104).

A combustor (28) and an insert (58) for providing cooling fluid along a wall (38, 40) of the combustor (28) includes a body (64), a passage (48), and an outlet (66). The body (64) extends from a proximal end (64A) to a distal end (64B) that protrudes from the wall (38, 40). The passage (48) extends at least partially through the body (64) towards the distal end (64B). The outlet (66) extends through the body (64) at the distal end (64B) to fluidly connect the passage (48). The passage (48) and the outlet (66) form a fluid path through the insert (58) and the wall (38, 40).

A powerplant assembly (24) for an aircraft (20) includes a first driven rotor, a second driven rotor, a gas turbine engine (28), and an auxiliary engine (30A). The gas turbine engine (28) includes a compressor section (36), a combustor section (38), and a turbine section (40) forming a core gas flow path (88). The gas turbine engine (28) further includes a rotational assembly (46) including a bladed turbine rotor (56) and a shaft (52). The shaft (52) couples the bladed turbine rotor (56) with the first driven rotor. The auxiliary engine (30A) is mechanically independent of the gas turbine engine (28). The auxiliary engine (30A) includes an engine output shaft (66), an air inlet (68), and an exhaust outlet (70). The engine output shaft (66) is coupled with the second driven rotor. The auxiliary engine (30A) further forms the core gas flow path (88) with the air inlet (68) connected in fluid communication with the compressor section (36) and the exhaust outlet (70) connected in fluid communication with the combustor section (38).

A louver segment (44) includes an arcuate member (58) and a rib (74). The arcuate member (58) extends circumferentially partially about an axis (C) from a first side (62A) to a second side (62B) and extends axially parallel to the axis (C) from an upstream end (64A) to a downstream end (64B). The arcuate member (58) includes an inner surface (66) facing radially inward relative to axis (C) and an outer surface (68) facing radially outward relative to the axis (C). The rib (74) extends circumferentially along the outer surface (68) between the first (62A) and the second (62B) side. A combustor (28) includes a plurality of louver segments (44) circumscribing the axis (C).