The development of aero engines with geared fans may require the use of an air/oil heat exchanger (AOHE) embedded within the bypass duct to dissipate heat from the power gearbox of the fan. It is important that the AOHE system is designed and installed to minimise any detrimental impacts on exhaust performance while meeting the heat exchanger (HEX) heat transfer requirements. This paper introduces and demonstrates the capabilities of a coupled mixed fidelity method to model a naturally ventilated AOHE embedded within the bypass duct. The method is demonstrated in this paper by using it to explore and quantify the trade-offs between the HEX design and performance and the impact on the exhaust system. Overall, for the specific example considered, the introduction of the HEX could reduce the cruise thrust by 0.7% while meeting the required heat transfer at maximum take-off. Overall, the work shows how a mixed fidelity method can be used for preliminary design assessments of the integration of the HEX with the bypass duct.

This work presents a numerical and experimental investigation of the nacelle aerodynamics for high-bypass- ratio aeroengines. A conventional nacelle that is representative of a current standard, and a compact design that is envisaged for future aeroengines, were optimized with an existing computational method. Both nacelles were tested in a large-scale transonic wind tunnel. For the first time, the aerodynamic benefits of compact nacelles are demonstrated through an experimental test campaign. Measurements and computational fluid dynamics (CFD) simulations confirmed the drag reduction of compact configurations across a wide range of operating points with different flight Mach numbers, mass-flow capture ratios, and angles of attack. For midcruise conditions with a Mach number of 0.85, this was a drag reduction of 8.5% and 8.8% for the measurements and CFD, respectively. These benefits are similar to an isolated optimization, that is, not installed in the wind tunnel, which confirmed the capabilities of the method to identify the drag benefit of compact designs. Relative to the measurements, the main aerodynamic characteristics on the nacelles were captured by CFD in terms of isentropic Mach number distributions and shock location. This work provides a quantitative evaluation for the use of CFD within an industrial setting for nacelle design and analysis.

Larger ultra-high bypass ratio (UHBR) aero-engines introduce an aerodynamic integration challenge. In close-coupled, podded underwing configurations, the aerodynamic interference between the propulsion system and the airframe could penalize the aircraft net vehicle force (NVF) and erode some of the novel cycle benefits and fuel burn reduction. Nonaxisymmetric designs of the bypass nozzle can improve the performance of the aircraft by mitigating some of the penalizing effects induced by the integration of the powerplant. However, due to the prohibitive computational cost of the design methods, only lower-fidelity design approaches have been feasible in an industrial time-scale. This work develops a relatively low-cost multi-fidelity design optimization methodology for non-axisymmetric exhausts where the effects of the propulsion system installation are considered. The methodology combines inviscid and viscous aerodynamic data to formulate multi-fidelity surrogate models which drive a genetic algorithm (GA) optimization. The method enabled the incorporation of the viscosity effects in the optimization process at a reasonable computational cost and led to better designs relative to a methodology based only on lower-fidelity data. Overall, the optimization of non-axisymmetric exhausts can benefit the net vehicle force of the complete engine-aircraft system in cruise by up to 0.9% of the engine standard net thrust which can reduce fuel burn by a similar amount. The optimization with multi-fidelity surrogate models reduced the computational time by a factor of four relative to a method based only on viscous aerodynamic data.

This work presents a combined experimental and numerical investigation of the effect of wing integration on the aerodynamic behaviour of a typical large civil aero-engine exhaust at wind-milling conditions. Engine performance simulations established estimates of Fan and Core Nozzle Pressure Ratios (FNPR and CNPR, respectively) for representative "engine-out" wind-milling scenarios. The experimental data and Reynolds Averaged Navier Stokes (RANS) Computational Fluid Dynamic (CFD) simulations encompassed End of Runway (EoR) take-off, diversion, and cruise wind-milling conditions for both isolated and installed configurations. The impact of FNPR, CNPR, free-stream Mach number (M infinity), and high-lift surfaces on the installed suppression effect were evaluated. The measured and CFD predicted fan and core nozzle maps were implemented into the engine performance model to estimate the engine re-matching characteristics due to the impact of the installation, and the effect on engine mass flow. The effect of installation can reduce the fan and core nozzle discharge coefficients by up to 13% and 26%, respectively, relative to the isolated configuration for representative EOR wind-milling conditions. RANS CFD captures the effect of suppression on both the fan and core with an accuracy between 0.1% and 1.2%, depending on Mach number, which is sufficient for industrial design and analysis purposes. The engine performance analyses showed that the installed suppression effect can result in a 10% reduction of engine mass flow at EOR wind-milling. Within the context of nacelle design under wind-milling, this effect of exhaust suppression must be considered in determining the intake Mass Flow Capture Ratio (MFCR).

Future short and sharp nacelles for ultra-high bypass ratio jet engines must maintain aerodynamic function during off-design conditions such as runway crosswind operations. An experimental rig is presented using a two-dimensional representation of the nacelle lip in a crosswind in a wind tunnel. The current paper describes how changing the incoming Mach number and engine mass-flow rate influences the flowfield around the lip, particularly the Mach number distribution and the boundary-layer state. Laser Doppler Velocimetry, Shadowgraph imaging, and pressure-sensitive paint were used to study the flowfields in detail. The target supersonic flowfields were achieved, with attached and separated lip boundary-layer states found when varying the isentropic Mach number distribution around the lip geometry; the flow reached a peak isentropic Mach number of 1.44 before boundary-layer separation.

Purpose Aerodynamic shape optimisation is a complex problem usually governed by transonic non-linear aerodynamics, a high dimensional design space and high computational cost. Consequently, the use of a numerical simulation approach can become prohibitive for some applications. This paper aims to propose a computationally efficient multi-fidelity method for the optimisation of two-dimensional axisymmetric aero-engine nacelles. Design/methodology/approach The nacelle optimisation approach combines a gradient-free algorithm with a multi-fidelity surrogate model. Machine learning based on artificial neural networks (ANN) is used as the modelling technique because of its ability to handle non-linear behaviour. The multi-fidelity method combines Reynolds-averaged Navier Stokes and Euler CFD calculations as high- and low-fidelity, respectively. Findings Ratios of low- and high-fidelity training samples to degrees of freedom of n LF /n DOFs = 50 and n HF /n DOFs = 12.5 provided a surrogate model with a root mean squared error less than 5% and a similar convergence to the optimal design space when compared with the equivalent CFD-in-the-loop optimisation. Similar nacelle geometries and aerodynamic flow topologies were obtained for down-selected designs with a reduction of 92% in the computational cost. This highlights the potential benefits of this multi-fidelity approach for aerodynamic optimisation within a preliminary design stage. Originality/value The application of a multi-fidelity technique based on ANN to the aerodynamic shape optimisation problem of isolated nacelles is the key novelty of this work. The multi-fidelity aspect of the method advances current practices based on single-fidelity surrogate models and offers further reductions in computational cost to meet industrial design timescales. Additionally, guidelines in terms of low- and high-fidelity sample sizes relative to the number of design variables have been established.

Purpose The decrease in specific thrust achieved by Ultra-High Bypass Ratio (UHBPR) aero-engines allows for a reduction in specific fuel consumption. However, the typical associated larger fan size might increase the nacelle drag, weight and the detrimental interference effects with the airframe. Consequently, the benefits from the new UHBPR aero-engine cycle may be eroded. This paper aims to evaluate the potential improvement in the aerodynamic performance of compact nacelles for installed aero-engine configuration. Design/methodology/approach Drooped and scarfed non-axisymmetric compact and conventional nacelle designs were down selected from a multi-point CFD-based optimisation. These were computationally assessed at a set of installation positions on a contemporary wide-body, twin-engine transonic aircraft. Both cruise and off-design conditions were evaluated. A thrust and drag accounting method was applied to evaluate different aircraft, powerplant and nacelle performance metrics. Findings The aircraft with the compact nacelle configuration installed at a typical installation position provided a reduction in aircraft cruise fuel consumption of 0.44% relative to the conventional architecture. However, at the same installation position, the compact design exhibits a large flow separation at windmilling conditions that is translated into an overall aircraft drag penalty of approximately 5.6% of the standard cruise net thrust. Additionally, the interference effects of a compact nacelle are more sensitive to deviations in mass flow capture ratio (MFCR) from the nominal windmilling diversion condition. Originality/value This work provides a comprehensive analysis of not only the performance but also the aerodynamics at an aircraft level of compact nacelles compared to conventional configurations for a range of installations positions at cruise. Additionally, the engine-airframe integration aerodynamics is assessed at an off-design windmilling condition which constitutes a key novelty of this paper.

When a civil aircraft engine is operated at windmill during the cruise flight phase, there is supersonic flow acceleration around the leading edge of the fan cowl toward the external surface. The terminating normal shock wave can separate the turbulent boundary layer developing on this external surface. A series of experiments at a flight-relevant Reynolds number (1.2 million based on lip thickness) are performed in a quasi-two-dimensional wind tunnel rig to investigate the underlying flow physics. At a nominal inflow Mach number of 0.65 and a nacelle incidence angle of 4.5 deg, as the equivalent engine mass-flow rate is reduced, an increase in shock strength results in flow separation when the shock exceeds Mach 1.4. Over a 10% range in the notional engine mass-flow rate, the boundary layer developing on the external fan cowl thickens by a factor of three on the onset of separation. A reduction in the incoming Mach number from 0.65 to 0.60 weakens the shock wave and thus delays separation. An increase in surface roughness has no significant effect in situations where the boundary layer remains attached. However, for separated cases, an increased local roughness height causes a greater separation extent and a thicker boundary layer downstream of the shock wave.

Under crosswind operating conditions, the flow field of an aero-engine intake can be characterized by notable unsteady flow distortion. These distortions are typically associated with flow separation within the intake as well as with the ingestion of the ground vortex. This unsteady flow distortion can have a detrimental effect on the intake performance and potentially on the operability of the downstream compression system. Measurements of the unsteady velocity field within a model-scale intake under crosswind conditions were acquired using stereo particle image velocimetry (S-PIV). This work analyzes the S-PIV data to quantify the unsteady flow distortion, as well as the characteristics of the ingested ground vortex, in a short intake under crosswind conditions. The swirl distortion metrics were calculated for a range of crosswind velocities and intake mass flow capture ratios (MFCRs). The conditions at which the intake flow separates depend on crosswind velocity, ground clearance, the design of the intake, and the MFCR. Flow characteristics of both low MFCR diffusion-driven and high MFCR shock-induced separation were identified. The circumferential extent and intensity of the swirl distortion are strongly dependent on the crosswind velocity and mass flow rate. The swirl distortion caused by the diffusion-driven separation is greater than that due to the shock-induced separation. The diffusion-driven separation affects a larger portion of the intake aerodynamic interface plane with greater time-averaged and peak distortion levels compared to shock-induced separation. The ground vortex characterization at the aerodynamic interface plane showed a decreasing level of unsteadiness in vortex meandering with increasing MFCR.

Aero-engine nacelles have to fulfill design requirements at both cruise and off-design conditions. Under engine windmilling conditions the ingested streamtube massflow is relatively low. A key off-design condition is take-off, which, in conjunction with an engine windmilling scenario, results in the stagnation point of the ingested streamtube being located significantly inside the intake. The combination of high angle of attack and low engine massflow rates leads to a strong flow acceleration and subsequent diffusion of the boundary layer on the upper quadrants of the external nacelle cowl, which can terminate with subsonic separation from the leading-edge. Under this condition, Reynolds number effects can play a dominant role on the separation onset and characteristics and 3D-annular wind tunnel tests cannot always achieve Reynolds’ number equivalent to full scale. A novel quasi-2D rig configuration representative of the aerodynamics of a full-size aero-engine nacelle under windmilling end of runway conditions examined in detail the characteristics of the boundary layer on the external cowl of a nacelle prior to diffusion-induced separation. Separation of the boundary layer was independently promoted through changes to represent different engine massflow rates and freestream Mach number on the rig to determine the limits of steady Reynolds Averaged Navier Stokes (RANS) methods to discern the onset of boundary layer separation. For the conditions and geometry investigated, the combined experimental and computational results showed that there was laminar to turbulent transition of the boundary layer ahead of the subsonic diffusion. The work showed that steady RANS can predict the onset of boundary layer separation with an uncertainty of approximately 10% on notional engine massflow rate and 0.05 on freestream Mach number relative to a nominal operating freestream Mach number of 0.25. This provides guidance for the industrial design and optimization of future windmilling-tolerant nacelles for large ultra-high bypass ratio turbofan engines.

In crosswind operating conditions, an aero-engine intake can be affected by notable unsteady flow distortions at the fan face. These distortions are typically associated with the ingestion of the ground vortex as well as with flow separation within the intake and can have a detrimental effect on the intake performance and therefore on the operability of the downstream compression system. Measurements of the unsteady velocity field within a modelscale intake under crosswind conditions were conducted using Stereo Particle Image Velocimetry (S-PIV) to characterize the velocity field and hence the intake flow distortion across the Aerodynamic Interface Plane (AIP) inside the duct. The intake distortion metrics were calculated for three operating conditions at a fixed crosswind velocity and increasing Mass Flow Capture Ratio (MFCR). The conditions at which the flow separates depend on crosswind velocity, ground clearance, the design of the intake and the MFCR. Flow characteristics of both low MFCR diffusion-driven, and high MFCR shock-induced separation were identified. The circumferential extent and intensity of the swirl distortion were found to be highly dependent on the crosswind velocity and MFCR. The swirl distortion caused by the diffusion-driven separation is greater than that due to shock-induced separation. The diffusion-driven separation was found to affect a bigger position of the intake AIP with higher time-average and peak values. An intermittent separation, that was observed for one value of MFCR in the range investigated, was found to cause peak levels of distortion twice the time-averaged values. Localized high swirl levels at a radial position near the intake surface correspondent to the tip region of a notional fan were observed. These can be expected to be detrimental to the operating stability of the downstream compression system.

This paper presents a numerical investigation of the effect of wing integration on the aerodynamic behaviour of a typical large civil aero-engine exhaust system at wind-milling flow conditions. The work is based on the Dual Stream Jet Propulsion (DSJP) test rig, as will be tested within the Transonic Wind Tunnel (TWT) located at the Aircraft Research Association (ARA) in the UK. The DSJP rig was designed to measure the impact of the installed pressure field due to the effect of the wing on the aerodynamic performance of separate-jet exhausts. The rig is equipped with the Dual Separate Flow Reference Nozzle (DSFRN), installed under a swept wing. Computational fluid dynamic simulations were carried out for representative ranges of fan and core nozzle pressure ratios for “engine-out” wind-milling scenarios at End of Runway (EOR) take-off, diversion, and cruise conditions. Analyses were done for both isolated and installed configurations to quantify the impact of the installed pressure field on the fan and core nozzle discharge coefficients. The impact of fan and core nozzle pressure ratios, as well as free-stream Mach number and high-lift surfaces on the installed suppression effect were also evaluated. It is shown that the installed pressure field can reduce the fan nozzle discharge coefficient by up to 16%, relative to the isolated configuration for EOR wind-milling conditions. The results were used to inform the design and set-up of the experimental activity which is planned for 2023.

Unsteady inlet flow distortion can influence the stability and performance of any propulsion system, in particular for more novel, short and slim intakes of future aero-engine configurations. As such, the requirement for measurement methods able to provide high spatial resolution data is important to aid the understanding of these flow fields. This work presents flow field characterisations at a crossflow plane within a short aeroengine intake using stereoscopic particle image velocimetry (SPIV). A series of tests were conducted across a range of crosswind and high angle of attack conditions for a representative short and slim aspirated intake configuration at two operating points in terms of mass flow rate. The velocity maps were measured at a crossflow plane within the intake at an axial position L/D = 0.058 from where a fan is expected to be installed. The diameter of the measurement plane was 250 mm, and the final spatial resolution of the velocity fields had a vector pitch of 1.5 mm which is at least two orders of magnitude richer than conventional pressure-based distortion measurements. The work demonstrates the ability to perform robust non-intrusive flow measurements within modern intake systems in an industrial wind tunnel environment across a wide range of operating conditions; hence, it is suggested that SPIV can potentially become part of standard industrial testing. The results provide rich datasets that can notably improve our understanding of unsteady distortions and influence the design of novel, closely coupled engine-intake systems.

The boundary layer on the external cowl of an aeroengine nacelle under windmilling diversion conditions is subjected to a notable adverse pressure gradient due to the interaction with a near-normal shock wave. Within the context of computational fluid dynamics (CFD) methods, the correct representation of the characteristics of the boundary layer is a major challenge in capturing the onset of the separation. This is important for the aerodynamic design of the nacelle, as it may assist in the characterization of candidate designs. This work uses experimental data obtained from a quasi-2D rig configuration to provide an assessment of the CFD methods typically used within an industrial context. A range of operating conditions are investigated to assess the sensitivity of the boundary layer to changes in inlet Mach number and mass flow through a notional windmilling engine. Fully turbulent and transitional boundary-layer computations are used to determine the characteristics of the boundary layer and the interaction with the shock on the nacelle cowl. The correlation between the onset of shock-induced boundary-layer separation and the preshock Mach number is assessed, and it was found that the CFD is able to discern the onset of boundary-layer separation.

The next generation of ultrahigh 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 (CFD) 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 preshock Mach number to mitigate the adverse unsteady interaction between the fan and the shock.

For Ultra-High Bypass Ratio aero-engines, the exhaust system is likely to play a significant role on the aerodynamics and performance of the aircraft.For this reason, relatively rapid methods for the aerodynamic design and optimisation of exhaust systems are required to inform design decisions at early stages of the design process.Previous exhaust optimisation works encompassed Response Surface Model (RSM) based optimisations of nozzle configurations that were parametrised with a significant number of design variables.The RSM were constructed with a large database of designs that were assessed with fine computational meshes and well resolved boundary layers.However, the large number of design variables and the computational cost required to evaluate each exhaust design limited the optimisation capabilities.This work develops a relatively more rapid exhaust optimisation method based on CFD in-the-loop and dimensionality reduction.The methodology is based on coarse meshes and wall functions to guide the optimisation process and is coupled with methods for the identification of the dominant design variables.For an UHBR aero-engine exhaust design space of 16 design variables, it was found that the velocity coefficient could be characterised with only seven parameters.Based on these results, various optimisation methods were developed and applied.These targeted the maximisation of the velocity coefficient by optimising just the 7 dominant design variables.With these approaches, a similar benefit in exhaust performance relative to the baseline optimisation method was obtained approximately 4 times faster.

Aero-engine nacelles are typically designed to fulfil both design and off-design aircraft manoeuvres. Under-off design conditions one of the objective is to avoid large flow separation either on the external cowl or within the intake that can influence aircraft and engine operability. One particular scenario is represented by a low engine mass flow regime associated with one inoperative engine, also known as a windmilling condition. Under windmilling, the boundary layer on the external cowl of the nacelle can separate either due to the interaction with shockwaves or due to notable adverse pressure gradient towards the trailing edge. Both mechanisms are computationally difficult to model and there is a need for more validation of computational fluid dynamics (CFD) methods. The aim of this work is to develop a rig configuration which will provide CFD validation data for the aerodynamics of a nacelle under representative windmilling conditions. Two flight regimes are considered, namely windmilling diversion and end-of-runway. CFD simulations of a 3D nacelle are used to determine primary aerodynamic mechanisms associated with boundary layer separation. Two rig configurations are developed and both 2D and 3D CFD analyses are used to achieve the design objectives. Overall, this work presents the design philosophy and methods that were pursued to develop a quasi-2D rig configuration representative of the aerodynamics of 3D-annular aero-engine nacelles under windmilling conditions.

The slim fan cowl profiles used for ultra-high bypass ratio aircraft engines are designed considering off-design operating conditions, such as engine windmilling during take-off climb out or during cruise. The current paper describes wind tunnel experiments studying how incoming Mach number and engine mass-flow rate influence the aerodynamics governing external fan cowl flow separation in both these windmilling scenarios. A transonic region may form on the forebody surface if the engine becomes inoperative during take-off climb out, with peak Mach number up to 1.2. The subsequent adverse pressure gradient can separate the local boundary layer, resulting in flow separation which originates near the highlight and a more uniform fan cowl pressure distribution. Meanwhile, engine shut down during cruise results in a large supersonic region on the external fan cowl surface which terminates in a normal shock wave. When the Mach number of this shock exceeds about 1.35, a closed separation bubble develops, which causes up to a four-fold increase in the boundary-layer thickness downstream of the shock wave.