초록 ▼

A turboengine that reduces rear noise emissions has a thickness of a rear part of a sound-attenuating coating, which is borne internally by an external fan cowl, that is increased toward a front part of the sound-attenuating coating. An increased-thickness zone is connected to a pan of the coating,

A turboengine that reduces rear noise emissions has a thickness of a rear part of a sound-attenuating coating, which is borne internally by an external fan cowl, that is increased toward a front part of the sound-attenuating coating. An increased-thickness zone is connected to a pan of the coating, located in a critical area of a jet nozzle, by a surface having a curved profile.

대표청구항 ▼

The invention claimed is: 1. A bypass turboengine of an aircraft, the turboengine comprising: a nacelle that delimits the front an air inlet and includes a cold stream fan, a hot stream central generator, and a fan channel with an annular section having a nozzle for the cold stream, wherein: the fa

The invention claimed is: 1. A bypass turboengine of an aircraft, the turboengine comprising: a nacelle that delimits the front an air inlet and includes a cold stream fan, a hot stream central generator, and a fan channel with an annular section having a nozzle for the cold stream, wherein: the fan channel is formed between an internal cowl surrounding the hot stream central generator and an internal tubular face of an acoustic attenuation coating with the annular section carried internally by an external fan cowl forming the rear of the nacelle, the coating comprises: a front part, disposed upstream of the nozzle, having an optimal thickness for the acoustic attenuation of the noise produced by the fan and conveyed by the cold stream, and a rear part, contiguous with the front part and disposed on either side of the throat of the nozzle, the rear coating part having a thickness that decreases towards the rear edge of the external fan cowl delimiting an annular ejection orifice for the cold stream, the front coating part has, in the vicinity of its junction with the rear coating part, a convergent zone in which its internal tubular face begins to converge towards the nozzle, a critical zone of the fan channel is defined, beginning at the nozzle throat and extending frontward, in which any possible geometric modification of the fan channel and the internal tubular face of the rear coating part is impossible without a modification of parameters of the nozzle, in the convergent zone of the front coating part, the internal tubular face has a progressive increase in the thickness of the coating towards the rear coating part and this progressive modification in the internal tubular face of the contiguous rear coating part is continued until a zone of the latter with increased thickness is obtained in which the thickness is at least equal to the optimal thickness, and the rear end of the zone with increased thickness is linked to the front end of the critical zone by an internal tubular face with an inflection profile. 2. The bypass turboengine of claim 1, wherein the front boundary of the critical zone is defined such that the Mach number of the cold stream thereat is at least approximately equal to 0.8. 3. The bypass turboengine of claim 1, wherein in the convergent zone of the front part of the coating, the progressive shape modification begins when the Mach number of the cold stream lies between about 0.4 and about 0.55. 4. The bypass turboengine of claim 1, wherein the inflection profile is configured not to produce any detachment of the boundary layer. 5. The bypass turboengine of claim 4, wherein the shape parameter of the inflection profile is at most equal to 1.6. 6. The bypass turboengine of claim 1, wherein the acoustic attenuation coating with the annular section carried internally by the external fan cowl comprises an inflection profile disposed between an upstream zone, in which the thickness of the coating is at least approximately equal to an optimal thickness, and the critical zone of the nozzle.