The present invention relates to a tip turbine engine, and more particularly to mounting the tip turbine engine along a single mounting plane.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan and a low pressure compressor, a middle core engine including a combustor, and an aft low pressure turbine all located along a common longitudinal axis. Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure and therefore requires multiple mounting planes along the elongated engine structure to support the engine and mount the engine on an aircraft. Utilizing multiple mounting planes may complicate the mounting process, add weight, and make assembly laborious and expensive.
A recent development in gas turbine engines is the more longitudinally compact tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan, which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor located radially outward from the fan. The combustor ignites the fuel mixture to form a high energy gas stream which drives turbine blades that are integrated onto the tips of the hollow bypass fan blades for rotation therewith as disclosed in U.S. Patent Application Publication Nos.: 2003192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust to weight ratio equivalent to conventional turbofan engines of the same class within a package of significantly shorter longitudinal length.
Accordingly and because of the unique architecture and shorter longitudinal length of the tip turbine engine, it is desirable to provide more efficient and simplified mounting assemblies for the tip turbine engine.