Jet engine inlet-fan system and design method
AssigneeMassachusetts Institute of Technology
InventorAli Merchant
Life of this patent
Estimating the term…
Abstract
A supersonic aircraft engine axial fan that includes a rotating blade row having blades that receive a supersonic entry flow in the absolute frame and decelerate the flow to a lower supersonic or subsonic velocity exit flow while adding work to the flow to increase stagnation pressure. It is preferred that the lower velocity be subsonic. When the fan is combined with a suitably designed inlet, the propulsion system is compact and lighter in weight than conventional engines for supersonic aircraft.
Claims9
An axial fan comprising:
- a rotating blade row including blades that receive an absolute supersonic entry flow in an absolute frame over some part or all of the span and decelerate the flow to a lower supersonic or subsonic velocity exit flow while adding work to the flow to increase stagnation pressure.
Integrated inlet/fan engine system for a supersonic aircraft comprising:
- an inlet having length to decelerate the flow from flight freestream supersonic velocity down to a velocity in the range of about Mach 1.2 to about Mach 1.3:
- a fan having a rotating blade row forming a fan face for receiving the decelerated supersonic flow at an end of the inlet, wherein blades in the blade row decelerate the flow to a lower supersonic or subsonic exit flow while adding work to the flow to increase stagnation pressure.
Method for designing a fan for a turbofan engine for use in a supersonic aircraft, the fan having a supersonic axial inflow and supersonic/subsonic axial outflow comprising:
- using a two-dimensional axisymmetric flow solver to calculate a base flow field through the fan;
- post-processing the base flow field calculation to obtain stream surface positions and flow conditions upstream and downstream of a blade of the fan to serve as boundary conditions;
- using the boundary conditions in a quasi-three-dimensional solver to calculate a flow field on an unwrapped stream surface to provide blade surface pressure distribution and boundary layer development on the blade;
- shaping an inner and outer annulus to achieve a uniform and efficient deceleration of the flow with flow across the fan;
- aligning and shaping the blade section with respect to the incoming flow such that the flow acceleration around the blade is minimized;
- aligning the blade section with respect to the incoming flow such that the absolute Mach number of about 1.2 to about 1.3 can be achieved at the fan entry face;
- shaping the front portion of the blade to achieve a pre-shock Mach number of about 1.4 to about 1.5;
- shaping the aft portion of the blade to achieve a decreasing pressure gradient near the trailing edge;
- shaping the blade to achieve a desired pressure ratio;
- and
- placing a boundary layer control actuator near the adverse pressure gradient or shock waves on the blade.
Dependencies are inferred from the claim text.