Normal-shock/boundary-layer interactions in transonic intakes at high incidence

During high-incidence manoeuvres, shock-wave boundary layer interactions can develop over transonic inlet lower lips, significantly impacting aerodynamic performance. Here, a novel experimental rig is used to investigate the nature and severity of these interactions for a typical high incidence scenario. Furthermore, we explore the sensitivity to changes in angle of incidence and mass flow rate, as potentially experienced across off-design operations. The reference flow-field, informed by typical climb conditions, is defined by an incidence of 23° and a free stream Mach number M=0.435. The lower lip flow is characterised by a rapid acceleration around the leading-edge and a M=1.4 shock ahead of the intake diffuser. Overall, this flow-field is found to be relatively benign, with minimal shock-induced separation. Downstream of the interaction, the boundary layer recovers a healthy profile ahead of the nominal fan location. Increasing incidence by 2°, the separation becomes noticeably larger and unsteadiness develops. Detrimental effects are exacerbated at an even higher incidence of 26°. Increasing the mass flow rate over the lip by up to 15% of the initial value has minor effects on performance and is not found to inhibit the boundary layer profile recovery

A semi-empirical model for streamwise vortex intensification

Vortex intensification plays an important role in a wide range of flows of engineering interest. One scenario of interest is when a streamwise vortex passes through the contracting streamtube of an aircraft intake. There is, however, limited experimental data of flows of this type to reveal the dominant flow physics and to guide the development of vortex models. To this end, the evolution of wing-tip vortices inside a range of streamtube contractions has been measured using stereoscopic particle image velocimetry. A semi-empirical model has been applied to provide new insight on the role of vorticity diffusion during the intensification process. The analysis demonstrates that for mild flow contractions, vorticity diffusion has a negligible influence due to the low rates of diffusion in the vortex flow prior to intensification and the short convective times associated with the streamtube contraction. As the contraction levels increase, there is a substantial increase in the rates of diffusion which is driven by the greater levels of vorticity in the vortex core. A new semi-empirical relationship, as a function of the local streamtube contraction levels and vortex Reynolds number, has been developed. The model comprises a simple correction to vortex filament theory and provides a significant improvement in the estimation of vortex characteristics in contracting flows. For the range of contractions investigated, errors in the estimation of vortex core radius, peak tangential velocity and vorticity are reduced by an order of magnitude. The model can be applied to estimate the change in vortex characteristics for a range of flows with intense axial strain, such as contracting intake streamtubes and swirling flows in turbomachinery

Aspects of aero-engine nacelle drag

To address the need for accurate nacelle drag estimation, an assessment has been made of different nacelle configurations used for drag evaluation. These include a sting mounted nacelle, a nacelle in free flow with an idealised, freestream pressure matched, efflux and a nacelle with a full exhaust system and representative nozzle pressure ratio. An aerodynamic analysis using numerical methods has been carried out on four nacelles to assess a near field drag extraction method using computational fluid dynamics. The nacelles were modelled at a range of aerodynamic conditions and three were compared against wind tunnel data. A comparison is made between the drag extraction methods used in the wind tunnel analysis and the chosen computational fluid dynamics approach which utilised the modified near-field method for evaluation of drag coefficients and trends with Mach number and mass flow. The effect of sting mounting is quantified and its influence on the drag measured by the wind tunnel methodology determined. This highlights notable differences in the rate of change of drag with free stream Mach number, and also the flow over the nacelle. A post exit stream tube was also found to create a large additional interference term acting on the nacelle. This term typically accounts for 50% of the modified nacelle drag and its inclusion increased the drag rise Mach number by around ΔM = 0.026 from M=0.849 M=0.849 to M=0.875 M=0.875 for the examples considered

The aerodynamic effects of VHBR engine installation to the Common Research Model

This work describes the assessment of the effect of engine installation parameters such as engine position, size and power setting on the performance of a typical 300 seater aircraft at cruise condition. Two engines with very-high by-pass ratio and with different fan diameters and specific thrusts are initially simulated in isolation to determine the thrust and drag forces for an isolated configuration. The two engines are then assessed in an engine-airframe configuration to determine the sensitivity of the overall installation penalty to the vertical and axial engine location. The breakdown of the interference force is investigated to determine the aerodynamic origins of beneficial or penalising forces. To complete the cruise study a range of engine power settings were considered to determine the installation penalty at different phases of cruise. This work concludes with the preliminary assessment of cruise fuel burn for two engines. For the baseline engine, across the range of installed positions the resultant thrust requirement varied by 1.7% of standard net thrust. The larger engine was less sensitive with a variation of 1.3%. For an assessment over a 10000km cruise flight the overall effect of the lower specific thrust engine showed that the cycle benefits of –5.8% in specific fuel consumption was supplemented by a relatively beneficial aerodynamic installation effect but offset by the additional weight to give a -4.8% fuel burn reduction

Aerodynamic effects of propulsion integration for high bypass ratio engines

This work describes the assessment of the effect of engine installation parameters such as engine position, size, and power setting on the performance of a typical 300-seater aircraft at cruise condition. Two engines with very high bypass ratio and with different fan diameters and specific thrusts are initially simulated in isolation to determine the thrust and drag forces for an isolated configuration. The two engines are then assessed in an engine–airframe configuration to determine the sensitivity of the overall installation penalty to the vertical and axial engine location. The breakdown of the interference force is investigated to determine the aerodynamic origins of beneficial or penalizing forces. To complete the cruise study, a range of engine power settings is considered to determine the installation penalty at different phases of cruise. This work concludes with the preliminary assessment of cruise fuel burn for two engines. For the baseline engine, across the range of installed positions, the resultant thrust requirement varies by 1.7% of standard net thrust. The larger engine is less sensitive with a variation of 1.3%. For an assessment over a 10,000 km cruise flight, the overall effect of the lower specific thrust engine shows that the cycle benefits of −5.8% −5.8% in specific fuel consumption are supplemented by a relatively beneficial aerodynamic installation effect but offset by the additional weight to give a −4.8% −4.8% fuel-burn reduction

Design optimisation of separate-jet exhausts for the next generation of civil aero-engines

This paper presents the development and application of a computational framework for the aerodynamic design of separate-jet exhaust systems for Very-High-Bypass-Ratio (VHBR) gas-turbine aero-engines. An analytical approach is synthesised comprising a series of fundamental modelling methods. These address the aspects of engine performance simulation, parametric geometry definition, viscous/compressible flow solution, design space exploration, and genetic optimisation. Parametric design is carried out based on minimal user-input combined with the cycle data established using a zero-dimensional (0D) engine analysis method. A mathematical approach is developed based on Class-Shape Transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts, nozzles, after-bodies, and plugs. This proposed geometry formulation is coupled with an automated mesh generation approach and a Reynolds Averaged Navier–Stokes (RANS) flow-field solution method, thus forming an integrated aerodynamic design tool. A cost-e ective Design Space Exploration (DSE) and optimisation strategy has been structured comprising methods for Design of Experiment (DOE), Response Surface Modelling (RSM), as well as genetic optimisation. The integrated framework has been deployed to optimise the aerodynamic performance of a separate-jet exhaust system for a large civil turbofan engine representative of future architectures. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared to a baseline architecture, through optimum re-design of the exhaust system. Furthermore, the developed approach is shown to be able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system

Multi-objective optimisation of short nacelles for high bypass ratio engines

Future turbo-fan engines are expected to operate at low specific thrust with high bypass ratios to improve propulsive efficiency. Typically, this can result in an increase in fan diameter and nacelle size with the associated drag and weight penalties. Therefore, relative to current designs, there is a need to develop more compact, shorter nacelles to reduce drag and weight. These designs are inherently more challenging and a system is required to explore and define the viable design space. Due to the range of operating conditions, nacelle aerodynamic design poses a significant challenge. This work presents a multi-objective optimisation approach using an evolutionary genetic algorithm for the design of new aero-engine nacelles. The novel framework includes a set of geometry definitions using Class Shape Transformations, automated aerodynamic simulation and analysis, a genetic algorithm, evaluations at various nacelle operating conditions and the inclusion of additional aerodynamic constraints. This framework has been applied to investigate the design space of nacelles for high bypass ratio aero-engines. The multi-objective optimisation was successfully demonstrated for the new nacelle design challenge and the overall system was shown to enable the identification of the viable nacelle design space

Effect of lip shape on shock wave-boundary layer interactions in transonic intakes at incidence

The flow field around five transonic inlet lips at high incidence is investigated for a variety of flow conditions around a design point representative of high incidence manoeuvring. Changes to the operating point are simulated by varying the angle of incidence as well as changing the mass flow rate over the lip, intended to mimic the effect of an increase in engine flow. For these inflow conditions, the flow on the surface of the lip is characterised by a supersonic region, terminated by a near-normal shock wave. Of particular interest is the effect of lip geometry and operating point on the boundary layer at the equivalent fan location. The parametric investigation revealed a significant effect of lip shape on the position and severity of the shock wave-boundary layer interaction. From correlation studies based on the parametric investigation, it appears that the extent of shock-induced separation is the main factor affecting the boundary layer state downstream of the normal shock wave-boundary layer interaction. Somewhat surprisingly, this was found to be independent of shock strength but potentially related to the severity of the diffusion downstream of the shock. Alongside delaying flowreattachment, this diffusion is also likely to have a direct detrimental effect on the boundary layer development close to the engine fan

Neural network-based multi-point, multi-objective optimisation for transonic applications

In the context of aircraft applications, the overall design process can be challenging due to the different aerodynamic requirements at several operating conditions and the total associated computational overhead. For this reason, the use of low order models for the optimisation of complex non-linear problems is sometimes used. This paper addresses the challenge of transonic aerodynamic design optimisation through the integration of a set of neural networks for the prediction of integral values, the classification of flow features and the estimation of flow field characteristics. The design method improves the computational efficiency relative to an expensive design process driven by Computational Fluid Dynamics (CFD) evaluations. The approach is used for the multi-point, multi-objective optimisation of a compact aero-engine nacelle in which the design outcomes are validated using a CFD in-the-loop optimisation strategy. It is demonstrated that the method based on the neural network capability identifies similar nacelle designs at a 75% reduction in the overall computational cost, a drag uncertainty prediction within 2.8%, and a predictive accuracy for the classification metric of 98%. For downselected configurations, the main flow characteristics in terms of peak Mach number, pre-shock Mach number and shock location are well predicted by the neural network models compared with the CFD-based evaluations

Effect of unsteady fan-intake interaction on short intake design

The next generation of ultra-high bypass ratio civil aero-engines promises notable engine cycle benefits. However, these benefits can be significantly eroded by a possible increase in nacelle weight and drag due to the typical larger fan diameters. More compact nacelles, with shorter intakes, may be required to enable a net reduction in aero-engine fuel burn. The aim of this paper is to assess the influence of the design style of short intakes on the unsteady interaction under crosswind conditions between fan and intake, with a focus on the separation onset and characteristics of the boundary layer within the intake. Three intake designs were assessed and a hierarchical computational fluid dynamics approach was used to determine and quantify primary aerodynamic interactions between the fan and the intake design. Similar to previous findings for a specific intake configuration, both intake flow unsteadiness and the unsteady upstream perturbations from the fan have a detrimental effect on the separation onset for the range of intake designs. The separation of the boundary layer within the intake was shock driven for the three different design styles. The simulations also quantified the unsteady intake flows with an emphasis on the spectral characteristics and engine-order signatures of the flow distortion. Overall, this work showed that is beneficial for the intake boundary layer to delay the diffusion closer to the fan and reduce the pre-shock Mach number to mitigate the adverse unsteady interaction between the fan and the shock