Aeroengine Turbine-Induced Reverse Effects on Combustion and Aerothermal Features

Modern aeroengines require compact combustor/turbine integration with intensified cross-domain interactions. Current studies neglect turbine-induced reverse effects, contradicting physics. Conventional interface-based methods (sliding mesh/Multiple Reference Frame) enforce artificial pressure continuity at stationary/rotating interfaces, inducing numerical dissipation that attenuates reverse-propagating pressure waves and weakens turbine-to-combustor feedback. To address this, this study proposes an interface-free rotating framework with radiation modeling, capturing rotation-induced reverse-propagating pressure gradients through direct solution of global compressible flow equations. This approach simplifies mesh generation processes, achieving 16.7% computational time savings without compromising solution accuracy. The detailed analysis reveals that turbine structures constrain dilution-zone vortex development via adverse pressure gradients, driving hot streak expansion, and exacerbating temperature nonuniformity. Critically, rotation-induced pressure gradients propagate in reverse to the combustor, steepening existing pressure gradients and amplifying velocity fluctuation amplitudes. Low-speed fluctuations (less than or equal to 13,300 rpm) are dominated by swirl intensity; increasing speed attenuates swirl strength, progressively reducing fluctuations. At high speeds (greater than or equal to 19,950 rpm), turbine-induced reverse-propagating gradients constitute a dominant contribution, counteracting swirl decay. At this regime, the influence of this reverse effect traces back to the combustor front section. These findings provide critical guidance for suppressing hot streak diffusion and mitigating instability risks in combustor designs.