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.

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).

ABSTRACT:The performance prediction of axial flow compressors and turbines still relies on the stationary testing of blade cascades. Most of the blade testing studies are done for operating conditions close to the design point or in off-design areas not too far from it. However, blade-and consequently engine-performance remain unexplored at relatively far off-design conditions, such as windmilling or sub-idle. Such regimes are dominated by blade operation under extremely low mass flows and rotational speeds that imply highly negative values of incidence angle, thus totally separated flows on the pressure side of the blades. Those flow patterns are difficult to be measured and even more difficult to be numerically predicted as the current modelling capability of separated internal flows is of limited reliability. In this paper, the performance of a 3-dimensional linear compressor cascade at highly negative incidence angle is initially experimentally investigated. The main objective of the study is to derive the total pressure loss and outlet flow angle through the blades and use the data for the validation-calibration of a numerical solver enhancing its capability to predict highly separated flows. The development of the CFD model and the simulation strategy followedare also presented.The numerical results are compared against the derived test data demonstrating a good agreement. In addition, most trends of the properties of interest have been captured sufficiently, therefore the physical phenomena are considered to be well captured, allowing the numerical tool to be used for further studies on similar test cases.

ABSTRACT:One of the certification requirements that a jet engine has to fulfil is its altitude relight capability. Relighting an aero gas turbine engine at high altitudes is more challenging than at sea level conditions. The pressure, air velocity, and temperature within the combustor at such conditions are very low, hindering the fuel atomization and evaporation process. After ignition, combustion efficiency can be relatively low due to the poor fuel atomization quality, leading to slow shaft acceleration rates. Further studies in this field can help determine and predict the fuel spray characteristics, which limit the relight and pull-away capability of the engine at these sub-idle conditions. Reported in this paper is the CFD analysis of a typical airblast atomizer, simulated at different sub-idle operating conditions. Two sets of models were used; one with a simple liner-only combustor with co-flow air, and the other with a more detailed geometry, including wall cooling slots, primary and secondary dilution holes, and co-flow air. For the simpler model, three different inner and outer liner wall spacing were modelled to examine the effect of the chamber volume on the fuel spray behaviour. The CFD models were then run at a chamber pressure of 101, 41 and 31kPa, typical of sub-idle conditions. The effect of such conditions on the atomization quality of the fuel spray was analysed. The study carried out indicates how the chamber pressure, chamber volume and AFR (through amount of co-flow air introduced) significantly affect the resulting spray characteristics. A parametric analysis was performed to extract a correlation between the spray SMD (Sauter Mean Diameter) and fuel flow rate.

With stricter regulations on engine altitude relight capability, the understanding of low-speed axial compressor performance is becoming increasingly important. At such far off-design conditions, compressors behave differently from design point, with large changes in the flow phenomena and reduced reliability on the established empirical equations and assumptions. This work focuses on the aerodynamics of a locked-rotor axial compressor at high inlet Mach number conditions, using a validated numerical simulation approach. In a locked-rotor compressor there is very little compression of the inflow. The air is forced to accelerate, with the rear stages seeing the highest velocities. Depending on the inlet Mach number, the velocity at the rear stages can be close to sonic, until choking conditions are reached. To predict accurately the zero-speed compressor performance close to the choking point, the corresponding blade aerodynamic coefficients are evaluated as a function of the blade’s physical parameters and the inlet Mach number. In addition, the blockage due to the separated flow as a result of the high negative incidences is investigated as a function of inlet Mach number, incidence, solidity and stagger angle. Models that predict the characteristics and choking mass flow of the compressor, require such data. This work offers a better insight into the low-speed and locked rotor characteristics of the compressor. The zero-speed line can be calculated through a stage-stacking technique using the aerodynamic coefficients and flow blockage derived from the numerical simulations. Low-speed lines between the zero and idle-speed line can subsequently be created through interpolation. Using this methodology, it is possible to generate a complete sub-idle map for a multi-stage axial compressor, enhancing the predictive capability of whole engine performance solvers.