Abstract: Air captured from the tips of a compressor of a gas turbine engine is diverted from the engine core and can be collected for auxiliary uses. Gas path separation can be achieved using part-span shrouded compressor blades or using blade tip cut-outs conforming to an airflow dividing annular shroud. In a preferred application for the present invention, the gas turbine engine is the auxiliary power unit of an aircraft. This permits compressed air generation for Auxiliary Power Unit oil cooling and for compartment pressurization without loosing significant mass flow to the engine core, while eliminating the need to provide a separate active cooling system such as an engine driven fan.

Abstract: The present invention is addressed to a multiple layer blade burst control system for use in turbine engines. The system utilizes a first inner containment ring and a second outer containment ring separated by a layer of thin ductile material. The inner containment ring and outer containment ring are each mounted to an engine housing by the use of a plurality of supporting legs. The first inner containment ring and the second outer containment ring each have complementary ribs and grooves on their respective inner and outer surfaces, so that when broken blades or blade fragments strike the inner surface of the inner containment ring, this ring absorbs the transfer of energy and moves towards the second containment ring. The movement of the inner containment ring also deforms the thin ductile material which creates an additional energy transfer. The system may also be utilized without the thin ductile layer, and a set of intermeshing protrusions and recesses.

Abstract: Gas turbine blades 14 are slid axially into disc 12 with a retention tang 22 on each abutting the disc. An axially extending space 28 between blade platforms 26 and the disc receives elongated strip 30. With prebent end 34 resiliently held against the disc the opposite end 36 is bent radially outward against the roots of adjacent blades. A bow 38 biases the blades outwardly, deterring vibration.

Abstract: A gas turbine engine rotor (18) and shaft (28) are secured against radial loosening between the shaft (28) and axially spaced rotor spigots (24, 26) by annular spacer rings (42, 43) which are fabricated from a material having a thermal coefficient of expansion significantly higher or lower than the thermal coefficients of the shaft (28) or rotor (18). The ring (42, 43) is pre-assembled during production of the engine in an interference fit with either the rotor hub (22) or the shaft (28), being sized to remain tight with the pre-assembled member over a range of operating temperatures, and becoming more tightly fit with the other member due to differential thermal expansion therebetween.