Abstract: An aircraft engine mount includes a mounting platform fixedly joinable to an aircraft pylon, and a cooperating mounting ring fixedly joinable to an aircraft engine. The platform includes a pair of spaced apart lugs defining therebetween a failsafe slot. The mounting ring includes an inner flange fixedly joinable to the engine, and an outer flange extending from the inner flange into the failsafe slot. A pair of side links are pivotally joined to the outer flange and at least one of the lugs to carry loads therebetween. A failsafe pin extends across the failsafe slot and is fixedly joined to the lugs. A clearance hole is provided in the outer flange for receiving the failsafe pin and providing a failsafe alternate loadpath between the mounting ring and platform upon failure of the links to carry loads.

Abstract: An aircraft engine mount includes a mounting platform fixedly joinable to an aircraft pylon, and a cooperating mounting ring fixedly joinable to an aircraft engine. The platform includes four lugs arranged in forward and aft pairs. Four links are arranged in forward and aft pairs pivotally joined at opposite ends thereof to respective ones of the four lugs and to the mounting ring. The four links are symmetrically arranged both circumferentially and axially to carry both thrust loads and in-plane vertical and side loads between the mounting ring and mounting platform with failsafe redundancy.

Abstract: A turbofan engine bypass and exhaust system has a double-hoop configuration in which a fan exhaust passageway comprises shells of a constant circular contour to resolve air pressure induced stresses and increase in bypass passageway area asymmetrically in the horizontal direction without substantially increasing the overall engine height.

Abstract: A gas turbine engine having a compressor and an air-cooled turbine is provided with a cooling system for decreasing the temperature of the turbine cooling air. A heat exchanger, mounted on the compressor casing, receives a portion of the pressurized air which is bled from the compressor. A heat sink medium is pumped through the heat exchanger into heat exchange relationship with the pressurized air, thereby cooling the air. The cooled air is then further pressurized and routed to and circulated through the turbine rotor blades to provide improved cooling thereof. The intercooling of the compressor bleed air permits a reduction in the quantity of compressor air required for turbine rotor blade cooling or allows an increase in turbine entry temperature and thus provides an improvement in engine power and performance. In the case where the heat sink medium is engine fuel, the heat extracted from the compressor bleed air is returned to the engine operating cycle in the form of heated engine fuel.

Abstract: The invention concerns reduction of tip vortices in aircraft propellers, wherein a fence having anhedral or dihedral is attached to the blades of the propeller. Another form of the invention is applicable to counterrotating propeller systems, wherein a fence is attached to the blades of the forward propeller of the system, and no fences are attached to the aft propeller.

Abstract: A gas turbine engine comprising a core gas generator engine for generating combustion gases, an additional power turbine, an additional fan section, and a booster compressor is disclosed. The power turbine includes first and second counterrotatable turbine blade rows effective for rotating first and second drive shafts, respectively. The fan section includes a first fan blade row connected to the first drive shaft and a second fan blade row connected to the second drive shaft. The booster compressor includes a first compressor blade row connected to the first drive shaft and a second compressor blade row connected to the second drive shaft.

Abstract: An improved counterrotating power turbine is disclosed. The power turbine includes a support having an annular axial cross-section and forward and aft portions. The forward portion of the support has a diameter greater than the diameter of the aft portion. Also included are a first rotor and a second rotor. The first rotor is supported on the first portion of the support by first bearing means and extends to the aft portion forming a first interface therewith. The second rotor is supported on the first rotor by second bearing means and extends to the aft portion of the support forming a second interface therewith. In this manner, it is possible to have low diameter slip rings at the first and second interfaces while maintaining relative high diameters for the rotor bearing means.

Abstract: Turbine blade shrouds are sealed by injecting air on to the shroud immediately upstream of a shroud seal. This air serves two purposes: it creates a turbulent flow immediately upstream of the shroud seal, thus spoiling and reducing the leakage flow over the seal, and the hot gases which would otherwise leak over the seal are replaced by the cooler air, thus reducing energy wastage and improving S.F.C. (specific fuel consumption).