Abstract
A hybrid methodology combining a detailed Large Eddy Simulation of a combustion chamber sector, an analytical propagation model of the extracted acoustic and entropy waves at the combustor exit through the turbine stages, and a far-field acoustic propagation through a variable exhaust temperature field was shown to predict far-field combustion noise from helicopter and aircraft propulsion systems accurately for the first time. For the single-stream turboshaft engine, the validation was achieved from engine core to the turbine exit. Propagation to the far field was then performed through a modeled axisymmetric jet. Its temperature modified the acoustic propagation of combustion noise significantly and a simple analytical model based on the Snell–Descarte law was shown to predict the directivity for axisymmetric single jet exhaust accurately. Good agreement with measured far-field spectra for all turboshaft-engine regimes below 2 kHz stresses that combustion noise is most likely the dominant noise source at low frequencies in such engines. For the more complex dual-stream turbofan engine, two regime computations showed that direct noise is mostly generated by the unsteady flame dynamics and the indirect combustion noise by the temperature stratification induced by the dilution holes in the combustion chamber, as found previously in the turboshaft case. However, in the turboengine, direct noise was found dominant at the combustor exit for the low power case and equivalent contributions of both combustion noise sources for the high power case. The propagation to the far-field was achieved through the temperature field provided by a Reynolds-Averaged Navier–Stokes simulation. Good agreement with measured spectra was also found at low frequencies for the low power turboengine case. At high power, however, turboengine jet noise overcomes combustion noise at low frequencies.
Highlights
Future aeroengines must be designed to limit cruising speed consumption, pollutant emissions and noise at landing and takeoff phases
Using a three-sensor technique, Liverbardon et al showed that a significant narrow component of direct combustion noise was found around 200 Hz generated in the combustion chamber, and that a broadband hump up to 1200 Hz could be attributed to broadband noise generated by the high pressure turbine (HPT), which could be indirect combustion noise [5]
It is worthy noting that a specific decorrelation filter is used to account for the interaction between the multiple engine burners, and an attenuation function through each turbine row is introduced to properly simulate the effect of the mean flow on the entropy waves
Summary
Future aeroengines must be designed to limit cruising speed consumption, pollutant emissions and noise at landing and takeoff phases. A detailed turbulent reactive LES is used only for the noise sources in the chamber It provides waves which are propagated through the turbine stages with a simplified acoustically compact method taking into account that combustion noise is in the low frequency range (0–1500 Hz). These methods are analytical extensions of the methods of Cumpsty and Marble [9], the limits of which were extensively studied by Leyko et al [22,23] through a stator vane, Wang et al [24] through a rotor blade, and more recently by Bauerheim et al [25] through a two-dimensional stage.
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