This paper presents the development of an integrated approach which targets the aerodynamic design of separate-jet exhaust systems for future gas-turbine aero-engines. The proposed framework comprises a series of fundamental modeling theories which are applicable to engine performance simulation, parametric geometry definition, viscous/compressible flow solution, and design space exploration (DSE). A mathematical method has been developed based on class-shape transformation (CST) functions for the geometric design of axisymmetric engines with separate-jet exhausts. Design is carried out based on a set of standard nozzle design parameters along with the flow capacities established from zero-dimensional (0D) cycle analysis. The developed approach has been coupled with an automatic mesh generation and a Reynolds averaged Navier–Stokes (RANS) flow-field solution method, thus forming a complete aerodynamic design tool for separate-jet exhaust systems. The employed aerodynamic method has initially been validated against experimental measurements conducted on a small-scale turbine powered simulator (TPS) nacelle. The developed tool has been subsequently coupled with a comprehensive DSE method based on Latin-hypercube sampling. The overall framework has been deployed to investigate the design space of two civil aero-engines with separate-jet exhausts, representative of current and future architectures, respectively. The inter-relationship between the exhaust systems' thrust and discharge coefficients has been thoroughly quantified. The dominant design variables that affect the aerodynamic performance of both investigated exhaust systems have been determined. A comparative evaluation has been carried out between the optimum exhaust design subdomains established for each engine. The proposed method enables the aerodynamic design of separate-jet exhaust systems for a designated engine cycle, using only a limited set of intuitive design variables. Furthermore, it enables the quantification and correlation of the aerodynamic behavior of separate-jet exhaust systems for designated civil aero-engine architectures. Therefore, it constitutes an enabling technology toward the identification of the fundamental aerodynamic mechanisms that govern the exhaust system performance for a user-specified engine cycle.

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The aerodynamic performance of the bypass exhaust system is key to the success of future civil turbofan engines. This is due to current design trends in civil aviation dictating continuous improvement in propulsive efficiency by reducing specific thrust and increasing bypass ratio (BPR). This paper aims to develop an integrated framework targeting the automatic design optimization of separate-jet exhaust systems for future aero-engine architectures. The core method of the proposed approach is based on a standalone exhaust design tool comprising modules for cycle analysis, geometry parameterization, mesh generation, and Reynolds-averaged Navier–Stokes (RANS) flow solution. A comprehensive optimization strategy has been structured comprising design space exploration (DSE), response surface modeling (RSM) algorithms, as well as state-of-the-art global/genetic optimization methods. The overall framework has been deployed to optimize the aerodynamic design of two civil aero-engines with separate-jet exhausts, representative of current and future engine architectures, respectively. A set of optimum exhaust designs have been obtained for each investigated engine and subsequently compared against their reciprocal baselines established using the current industry practice in terms of exhaust design. The obtained results indicate that the optimization could lead to designs with significant increase in net propulsive force, compared to their respective notional baselines. It is shown that the developed approach is implicitly able to identify and mitigate undesirable flow-features that may compromise the aerodynamic performance of the exhaust system. The proposed method enables the aerodynamic design of optimum separate-jet exhaust systems for a user-specified engine cycle, using only a limited set of standard nozzle design variables. Furthermore, it enables to quantify, correlate, and understand the aerodynamic behavior of any separate-jet exhaust system for any specified engine architecture. Hence, the overall framework constitutes an enabling technology toward the design of optimally configured exhaust systems, consequently leading to increased overall engine thrust and reduced specific fuel consumption (SFC).

Without careful consideration of aerodynamic installation eects on exhaust system performance the projected benets of high bypass ratio engines may not be achievable. This work presents a computational study of propulsion system integration in order to quantify the eect that aircraft installation has on the aerodynamic performance of separate-jet aero-engine exhaust systems. Within this study the sensitivity of exhaust nozzle performance metrics to aircraft incidence and under wing position were investigated for two engines of dierent specic thrust. Upon installation, thrust generation was found to be benecial or detrimental relative to an isolated engine depending on the position of the engine relative to the wing leading edge. The dominant installation eect was observed on the exhaust afterbodies and, over the range of engine positions investigated at cruise conditions, the installed modied velocity coecient was shown to vary up to 1 % relative to an isolated engine. Furthermore, due to variations in the core nozzle mass ow rate by

As the specific thrust of civil aero-engines reduces, the aerodynamic performance of the exhaust system will become of paramount importance in the drive to reduce engine fuel burn. This paper presents an aerodynamic analysis of civil aero-engine exhaust systems through the use of Reynolds Averaged Navier-Stokes Computational Fluid Dynamics . Two different numerical approaches were implemented and the numerical predictions were compared to measured data from an experimental high-bypass ratio separate-jet exhaust system. Over a fan nozzle pressure ratio range from 1.4 to 2.6, a comparison was drawn between values of thrust coefficient calculated numerically and values obtained from experimental measurements. In addition to the validation of numerical approaches, the effects of freestream Mach number and extraction ratio on the aerodynamic behavior of the exhaust system have been quantified and correlated to fundamental aerodynamic parameters.