| description abstract | Numerical experiments are carried out to investigate the tone noise radiated from a turbofan engine inlet under conditions at which the relative flow past the rotor tip is supersonic. Under these conditions, the inlet tone noise is generated by the upstream-propagating rotor-locked shock wave field. The spatial evolution of this shock system is studied numerically for flows through two basic hard-walled configurations: a slender nacelle with large throat area and a thick nacelle with reduced throat area. With the flight Mach number set to 0.25, the spatial evolution of the acoustic power through the two inlets reveals that the reduced throat area inlet provides superior attenuation. This is attributed to the greater mean flow acceleration through its throat and is qualitatively in accord with one-dimensional theory, which shows that shock dissipation is enhanced at high Mach numbers. The insertion of a uniform extension upstream of the fan is shown to yield greater attenuation for the inlet with large throat area, while the acoustic performance of the reduced throat area inlet is degraded. This occurs because the interaction of the nacelle and spinner potential fields is weakened, resulting in a lower throat Mach number. The effect of forward flight on the acoustic power radiated from the two inlets is also investigated by examining a simulated static condition. It is shown that the slender nacelle radiates significantly less power at the static condition than in flight, whereas the power levels at the two conditions are comparable for the thick nacelle. The reason for this behavior is revealed to be a drastic overspeed near the leading edge of the slender nacelle, which occurs to a lesser degree in the case of the thick inlet. This has implications for ground acoustic testing of aircraft engines, which are discussed. | |
