Abstract: A nacelle and mounting arrangement for a high bypass ratio ducted fan aircraft engine mounted external to the aircraft main structure is disclosed. The nacelle and mounting arrangement isolate the engine from the adverse effects of certain aerodynamic forces acting on the nacelle by transferring substantially all of those forces directly from the nacelle to the aircraft and transferring substantially none of those forces to the engine. Various arrangements of the nacelle components which facilitate engine removal while avoiding the placement of nacelle component interfaces in highly stressed regions of the nacelle are also disclosed.

Abstract: A gas turbine fan engine with improved maintainability features is disclosed. Forward and rearward sections of the core of the engine are joined one to the other, and each are independently separable from the pylon on which the engine is mounted without the necessity of removing the other from the airplane.

Abstract: A segmented seal (36) is provided for sealing between two adjacent rotor disks (18, 20). The individual segments (36) includes sideplate members (42, 44) and integral hook members (50, 52) which engage the facing sides of the respective disks (18, 20). Strut members (60, 62) support the central portion (64) of the axially and circumferentially extending wall members (40) which collectively establish the radially inner gas boundary for an annular working fluid stream (30).

Abstract: The low pressure compressor and fan together with the intermediate case are removable as a unit from the core engine by removal of the nose cone and a stub shaft at the front of the low pressure turbine shaft and at the ID of the fan hub (50) to provide access to attachment bolts at the inner portion of the intermediate case which holds the bearing for the high pressure rotor in position within the intermediate case after assembly of the engine.

Abstract: An oil lubricating and cooling system is positioned in the engine cowl and includes a heat exchanger, a compressor for delivering air from air inlets in the surface of the cowl through the heat exchanger and hence to air outlets also in the cowl surface and oil pumps also located within the cowl for circulating oil from the sump in the cowl to the heat exchanger to the reduction gear within the cowl that serve to drive the fan also supported within the cowl and having the blades extending outwardly therefrom. The reduction gear, compressor and pumps are all driven from the engine shaft, the end of which is located in the cowl.

Abstract: In a gas turbine engine with a bypass duct, a rotating stage of airfoils in the bypass duct is supported through and driven by torque transmitting members extending across the core stream flow path, but which are sufficiently flexible in the radial direction to avoid passing centrifugal loads from the rotating stage across the core stream flow path. The rotating bypass stream stage is also supported through a bearing secured directly to the engine main support structure radially outwardly of the core stream flow path. The majority of aircraft maneuver loads are reacted at this bearing, and a minor portion of the maneuver loads pass across the core stream to a bearing located radially inwardly thereof. The airfoils may be either variable pitch airfols or fixed airfoils. If variable pitch, they may be used to reverse the flow in the bypass duct for thrust reversing.

Abstract: Thrust balance in a gas turbine engine of the thrust balance seal is modulated by judiciously locating a seal between the inner diffuser case and the rear of the high compressor rotor shaft and bleeding air out of the cavity between both seals. The bled air is utilized to cool a portion of the high pressure turbine rotor. Cool air bled from mid-stage compressor is fed downstream of the second seal to cool the bearing compartment.

Abstract: A nonrotating flow directing assembly 14 for a gas turbine engine is disclosed. The flow directing assembly is formed of a circumferentially segmented inner case 24 supported by an annular sleeve 22. The inner case 24 is formed of a plurality of arcuate segments 26 extending axially continuously through the engine. A method of assembling the circumferentially continuous annular sleeve about an axially continuous rotor is also disclosed. In an alternate embodiment the inner case 124 is formed of several pluralities of arcuate segments 126.

Abstract: A seal structure between a circumferentially extending array of stator vanes and an outer case is disclosed. Various construction details that allow the seal structure to accommodate differences in axial and radial growth between the outer case and the vane array are developed. A continuous ring disposed between the vane array and the outer case cooperates with a first seal member to block leakage of working medium between the ring and the vane array. The continuous ring cooperates with a second seal member to block leakage of the working medium between the ring and the outer case.

Abstract: An outer air seal circumscribing the tips of the rotor blades of a rotary machine is disclosed. Techniques for limiting the radial clearance between the tips of the blades and the outer air seal at steady state conditions are developed. Structure independent of the machine case for isolating the thermal response of the outer air seal from the thermal response of the machine case is discussed and illustrated.