A passive cooling system for an auxiliary power unit (APU) installation on an aircraft is provided. The system is for an auxiliary power unit having at least a compressor portion of a gas turbine engine and an oil cooler contained separately within a nacelle. The system includes the auxiliary power unit housed within the nacelle of the aircraft, an engine exhaust opening defined in the aft portion of the nacelle and communicating with the gas turbine engine, at least a first air inlet duct communicating with a second opening defined in said nacelle and with said compressor portion and the oil cooler is located within a second duct communicating with an opening other than the engine exhaust opening of said nacelle and with the engine exhaust opening. Exterior cooling air and engine exhaust ejected through said engine exhaust opening entrain cooling air through said second duct to said oil cooler, and thus provide engine oil cooling. An exhaust eductor is also provided.

Passive cooling system for auxiliary power unit installation

Disclosed is an exhaust mixer (50) including a passage (54) extending from an inlet (56) to an outlet (58) that is coincident with a centerline axis of mixer (50). Several ridges (68) are circumferentially disposed about the axis and each flare away from the centerline axis relative to a direction along the centerline axis from inlet (56) toward outlet (58). Ridges (68) each define a corresponding one of several inner channels (74) radially disposed about passage (54) that each intersect passage (54) between inlet (56) and outlet (58). Several outer channels (84) are also radially disposed about passage (54) and are each positioned between a corresponding pair of inner channels (74). Ridges are each shaped to turn inner channels (74) and outer channels (84) about the axis as ridges (68) extend along the indicated direction. Inner channels (74) diverge away from the axis and one another in this direction while outer channels (84) converge toward the axis and one another in this direction.

Exhaust mixer and apparatus using same

An aircraft engine thrust mount includes a pair of spaced apart thrust links pivotally joined at one end to an engine first frame and at opposite ends to a lever. A platform is fixedly joined to an aircraft pylon and includes first and second spaced apart clevises and a center clevis therebetween. The center clevis is pivotally joined to a center of the lever for carrying thrust force from the first frame through both thrust links and to the platform and pylon during normal operation. A pair of pins pivotally join the thrust links to the lever and extend through apertures in the platform first and second clevises to define predetermined gaps therebetween for allowing the lever to rotate relative to the platform center clevis up to a predetermined angular rotation. In a failure mode of operation the pins contact the platform first and second clevises for transmitting thereto either tension or compression thrust loads for providing an alternate failsafe loadpath. In a preferred embodiment, the platform further includes an aft mount joined to an engine second frame spaced aft of the engine first frame.

Aircraft engine thrust mount

A bypass gas turbine aeroengine has an exhaust mixer of the lobed type. The trailing edges of the confronting sides of each lobe are cut back to form notches or scallops in the sides of the lobe. To reduce aerodynamic losses and engine length and weight, the final turbine stage within the engine core is without outlet guide vanes so that the turbine exhaust gas stream issuing from the turbine has a substantial swirl component of velocity. Correction back to an axial direction of mean flow at the downstream end of the exhaust mixer nozzle is achieved by making the area of the scallop in one of the confronting sides of each lobe greater than the area of the scallop in the other side, the arrangement being such that sufficient of the turbine exhaust stream effuses from the larger scallops in directions contrary to the swirl component of velocity to produce the desired correction.

Bypass gas turbine aeroengines and exhaust mixers therefor

An aircraft bypass turbofan engine subassembly for ground deceleration (eg, thrust reversal) A fan nacelle has a through passageway connecting its inner exterior surface and outer exterior surface A passageway end is located on the inner exterior surface aft of the aft-most row of fan rotor blades A row of pivotable fan outlet guide vanes is located radially between the fan and core nacelles and longitudinally aft of the flow splitter and the passageway end Apparatus is provided to pivot the vanes to block passage of bypass air and to open the passageway for ground deceleration

Aircraft bypass turbofan engine thrust reverser

A gas turbine engine is provided with a long nacelle which forms both the inlet and exhaust of the engine. The nacelle is spaced apart from a core engine to define an annular bypass duct therebetween. The incoming air stream is pressurized by a front fan disposed in the inlet and then divided between the core engine and bypass duct. A lobed mixer device is provided downstream of the core engine to intermix the core engine and bypass duct exhaust streams. The mixed streams are discharged from a common nozzle formed at the downstream end of the nacelle. A non-redundant system of mounts are provided to secure the engine to the pylon of an aircraft in such a manner as to reduce engine thrust induced bending loads and insure symmetry of these thrust loads. The total reaction of the thrust loads at the mounting points is reduced by a unique interlocking system by which all of the major engine assemblies are joined. The engine is configured such that it may be easily removed from an aircraft and replaced with a spare engine in minimum time. This quick change capability is achieved by mounting the engine such that it may be removed from the aircraft by disconnecting the forward and rear mounts and an engine bleed duct. The mid portion of the engine nacelle and all accessories mounted thereon remain with the aircraft.

Long duct mixed flow gas turbine engine

In a mixed flow gas turbine engine of the type which includes an outer nacelle spaced apart from a core engine to define an annular bypass duct therebetween, a thrust reverse is incorporated into the outer nacelle immediately aft of a fan disposed in the inlet. The thrust reverser reverses the discharge direction of the airflow exhausted from the bypass duct thereby to develop reverse thrust. The unique design and location of the thrust reverser permits the use of an aerodynamically efficient and relatively easy to manufacture cylindrical structure immediately aft of the engine fan frame.

Thrust reverser for a gas turbofan engine

A variable cycle gas turbine engine for commercial supersonic aircraft is provided with variable geometry means to simultaneously control the bypass and pressure ratios for different engine flight conditions wherein the engine may be operated at supersonic speed with a lower bypass and higher pressure ratio than at subsonic speed. The relatively high noise levels generally associated with the supersonic mode of operation for conventional supersonic engines can be significantly reduced by controlling the variable geometry means in a manner which increases the bypass ratio and decreases the pressure ratio of the fan section of the engine. In this manner, supersonic aircraft may approach and take off from airports situated adjacent densely populated areas without unduly disturbing the local residents.

Variable cycle gas turbine engine

1,192,503. Cyclone separators. GENERAL ELECTRIC CO. 30 Aug., 1968 [29 Sept., 1967], No. 41489/68. Heading B2P. [Also in Division F1] The invention relates to air intake structure for a gas turbine engine which constitutes the power plant of a helicopter. A cowl defining an air intake is provided axially aligned with and forward of the engine intake, the cowl having a plurality of filters extending radially therethrough, the filters serving to remove dust and like particles and passing essentially clean air to the interior of the cowl when there is air flow therethrough. Closure doors are provided which are movable between a first position in which air may flow axially through the cowl to the engine intake, and a second position in which such air flow is prevented. A valve is associated with each filter element, the valve closing when air flows axially through the cowl when the closure doors are open, thus preventing flow of air through the filter elements, and opening when such axial air flow is prevented by closing of said doors. The helicopter shown is provided with two gas turbine engines E, each having a cowl 14 axially aligned with the engine intake. The cowl 14 comprises an outer screen 22, an outer panel 24 and an inner panel 26. A plurality of vortex type filters 28 are associated with the panels 24, 26 each filter comprising a vortex tube 30 and a frusto conical outlet 32. A flap valve 36 is pivotally mounted at 37 adjacent the outlet 32, being shown in the open position in Fig. 6 and in the closed position in Fig. 7. A pair of closure doors 34 is mounted at the forward open end of the cowl. In normal operation the doors 34 are open and air flows axially through the cowl thus acting to close the flap valves 36 so that no air enters through the filters 28. When operating in dust-laden conditions the doors 34 are closed so that no air flows axially into the cowl and the flap valves 36 now open, and air flows 'radially inwardly through the filters 28, dust particles being separated out and clean air passing into the engine intake 16. The dust particles separated out pass between the panels 24, 26 and are discharged through outlet 35, Fig. 2 not shown.

Helicopter engine air inlets

A variable geometry inlet system is provided for a subsonic/supersonic gas turbine engine wherein a secondary inlet airflow is selectively ejected into a primary inlet airflow in order to augment the overall mass airflow to the engine and increase inlet pressure recovery. A plurality of foils are cooperatively disposed in rotatable connection with respect to a nacelle in order to modulate the secondary inlet airflow and match the velocity of the secondary airstream to that of the primary airstream thus reducing the mixing loss which would otherwise occur if the two streams commingled at different velocities. The primary and secondary inlet airstreams may also be maintained at a sufficiently high velocity to reduce the forward propagation of engine generated noise. The foils may also be rotated to define an outlet whereby a portion of the primary inlet airflow may be ejected overboard during supersonic operation to reduce aerodynamic drag as might otherwise result from a malfunctioning engine.

Donald Farley Sargisson

Papers

Long duct mixed flow gas turbine engine

Thrust reverser for a gas turbofan engine

Variable cycle gas turbine engine

Helicopter engine air inlets

Variable air inlet system for a gas turbine engine