Simulation of Turbofan Engine Flow Suppression Effects

Abstract

When turbofan engines are installed under the wing, the static pressures at the engine nozzles are often locally higher than those of the undisturbed environment. An increased static pressure downstream of an unchoked propelling nozzle leads to a reduction in nozzle pressure ratio, nozzle mass flow rate and gross thrust – referred to as flow suppression. The conventional industry turbofan thrust accounting method for flight and wind tunnel testing is based on a mass flow and momentum consideration and does not take the local static pressure at the nozzle of the installed engine into account. Neglecting flow suppression, the determined thrust is often overestimated. While this thrust accounting simplification was permissible for previous engine generations, the mass flow rate and thrust bias increases significantly for designs with increasing bypass and decreasing fan pressure ratios. Aerodynamic data is corrected for nacelle external flow induced suppression effects dominating in high speed conditions according to standard industry practice but no such corrections are applied to wing induced flow suppression effects which dominate in low speed conditions. NASA’s Common Research Model (CRM) configuration has been used for a parametric study based on full configuration Reynolds Averaged Navier-Stokes (RANS) computations complemented by rapid CFD results for the low speed high-lift flap effects. The impact of the nozzle back-pressure effect on the determination of mass flow rates and thrusts depending on aircraft layout, engine/nacelle design features, high-lift configuration, operational and flight condition parameters is explained. A straight forward data reduction approach avoiding systematic bias errors due to nozzle backpressure effects based on the continuity equation is proposed. Difficulties in measuring static pressures at the nozzle exit station are avoided by moving the mass flow evaluation station forward into the duct where measurements of representative static pressures are more robust. The suggested approach can not only be employed for wind tunnel tests utilizing Turbofan Propulsion Simulators (TPS) and Through Flow Nacelles (TFN) but also for full scale in-flight thrust determination.