Modular tip turbine engine
A tip turbine engine assembly includes a compressor module ( 12 ) and a fan module ( 14 ) located aft of the compressor module ( 12 ). The compressor module ( 12 ) and fan module ( 14 ) are fastened together along a mating portion with seals that generally prevent airflow from escaping through the mating portion. The compressor module ( 12 ) and fan module ( 14 ) are independently attachable to each other such that the compressor module ( 12 ) may be attached or detached to or from the fan module ( 14 ) without having to significantly disassemble the fan module ( 14 ), and verse visa.
1. A tip turbine engine assembly comprising:
a compressor module having an axial compressor and an annular combustor located radially outwards from said axial compressor; and
a fan module aft of said compressor module and including at least a rotor and a plurality of hollow fan blades attached to said rotor, wherein said fan module and said compressor module are removably attached together such that, when attached, said combustor is located axially forward of said plurality of hollow fan blades.
2. An assembly as recited in claim 1 , wherein each of said plurality of hollow fan blades defines a core airflow passage therethrough.
3. An assembly as recited in claim 1 , wherein said fan module comprises a planetary gearbox driven by said hollow fan blades and a turbine mounted on an outer tip of said plurality of hollow fan blades for rotation therewith.
4. An assembly as recited in claim 1 , wherein said compressor module is removable from an aircraft without removing said fan module.
5. An assembly as recited in claim 1 , wherein said fan module is removable from an aircraft without removing said compressor module.
6. An assembly as recited in claim 1 , further comprising a seal between said compressor module and said fan module.
7. An assembly as recited in claim 1 , wherein each of said plurality of hollow fan blades defines a core airflow passage therethrough, each core airflow passage having a turn directed axially forward into said annular combustor.
8. An assembly as recited in claim 1 , wherein said fan module includes a turbine located radially outward from said plurality of hollow fan blades.
9. An assembly as recited in claim 1 , wherein said compressor module and said fan module include a common axially facing mating surface at least between said annular combustor and said plurality of hollow fan blades.
10. A tip turbine engine assembly comprising:
a first jet engine module including at least a rotor and a plurality of hollow fan blades attached to said rotor; and
a second jet engine module having an axial compressor and an annular combustor located radially outwards from said axial compressor and axially forward of said plurality of hollow fan blades, wherein said first jet engine module is independently detachable from a said second jet engine module without having to disassemble said second jet engine module.
11. An assembly as recited in claim 10 , wherein at least one of said first jet engine module or said second jet engine module is independently attachable to an aircraft.
12. An assembly as recited in claim 10 , wherein each of said plurality of hollow fan blades defines a core airflow passage therethrough.
13. An assembly as recited in claim 12 , wherein said first jet engine module comprises a turbine located radially outward from said plurality of hollow fan blades.
14. An assembly as recited in claim 10 , wherein said first jet engine module comprises a gearbox.
15. A method of assembling a tip turbine engine comprising:
(a) aligning a central axis of a compressor module having an axial compressor and an annular combustor located radially outwards from said axial compressor with the central axis of a fan module which is aft of said compressor module, said fan module including at least a rotor and a plurality of hollow blades attached to said rotor; and
(b) removably attaching the compressor module to the fan module such that, when attached, said annular combustor is located axially forward of said plurality of hollow fan blades.
16. A method as recited in claim 15 , wherein said step (a) further comprises positioning the compressor module upstream from said fan module.
17. A method as recited in claim 15 , wherein said step (b) further comprises forming a seal between the compressor module and the fan module.
18. A method as recited in claim 15 , wherein said step (b) further comprises securing the compressor module and the fan module with a plurality of fasteners.
19. A method as recited in claim 18 , wherein said step (b) further comprises securing the compressor module and the fan module at an outer diameter and an inner diameter.