Abstract This study investigates the aerodynamic loads and structural implications for a future high bypass ratio (HBPR) aeroengine. Reynolds-averaged Navier–Stokes (RANS) computations are performed at top of climb (ToC), maximum takeoff (MTO), and maximum sideslip (MS) conditions over a broad range of angles of attack. The computational approach is validated against experimental data for both airframe and powerplant aerodynamics. The impact of airframe installation on the powerplant aerodynamic loads is quantified against an isolated configuration, and the underlying aerodynamic mechanisms are identified. Additionally, the importance of the aerodynamic terms on key structural aspects is quantified through finite element (FE) analyses. It is found that airframe installation modifies the powerplant aerodynamic loads through two fundamental mechanisms. First, at high-lift conditions, the increased airframe upwash changes the ingested streamtube and leads to pronounced flow expansion around the inlet lip. This results in increased inlet vertical loads by up to 9% of the net takeoff thrust, relative to the isolated geometry. Second, the elevated static-pressure field near the wing, combined with the swept-wing design, generates a strong inboard–outboard asymmetry on the powerplant. This leads to substantial changes in the lateral loading on the thrust reverser unit (TRU) relative to the isolated case, which are equal to 9% and 11% of the net thrust under takeoff and sideslip conditions, respectively. Additionally, the inclusion of aerodynamic loads from the isolated and installed powerplants increases the rear engine-to-pylon mount load by factors of 5.1 and 7.7, respectively, relative to a mechanical-only load reference. This study quantifies for the first time the effect of airframe installation on the powerplant aerodynamic loads, identifies the underlying driving mechanisms, and demonstrates the structural significance of the aeroloads for future HBPR powerplant design.

Understanding engine response to unsteady intake flow distortion is a crucial requirement to de-risk the development of novel aircraft configurations. This is more critical for configurations with highly embedded engines. Recent advances in non-intrusive, laser-based flow diagnostics demonstrated the ability to measure unsteady flows in convoluted intakes with high resolution in time and space. This work presents novel non-intrusive, unsteady flow measurements ahead of a fan rotor coupled to a convoluted diffusive intake. The fan rotor caused a local increase of the maximum levels of swirl intensity at the blade tip region, as well as flow re-distribution at the interface plane between the fan and the inlet duct compared to the baseline configuration with no fan in place. This contributed to the reduction of the overall swirl angle unsteadiness across the main flow distortion frequencies. This research presents a notable advance in unsteady fan-intake interaction characterisation. The work shows that high-resolution optical measurements offer notably better understanding of these complex aerodynamic interactions and have the potential to be part of larger scale, industrial testing programmes for future product development and certification.

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.

In-flight measurements of aerodynamic quantities are a requirement to ensure the correct scaling of Reynolds and Mach number and for the airworthiness certification of an aircraft. The ability to obtain such measurement is subject to several challenges such as instrument installation, environment, type of measurand, and spatial and temporal resolution. Given expected, more frequent use of embedded propulsion systems in the near future, the measurement technology needs to adapt for the characterization of multi-type flow distortion in complex flow, to assess the operability of air-breathing propulsion systems. To meet this increasing demand for high-fidelity experimental data, the Filtered Rayleigh Scattering (FRS) method is identified as a promising technology, as it can provide measurements of pressure, temperature and 3D velocities simultaneously, across a full Aerodynamic Interface Plane (AIP). Tau his work demonstrates the application of a novel FRS instrument, to assess the flow distortion in an S-duct diffuser, in a ground testing facility. A comparison of FRS results with Stereo-Particle Image Velocimetry (S-PIV) measurements reveals good agreement of the out of plane velocities, within 3.3 % at the AIP. Furthermore, the introduction of machine learning methods significantly accelerates the processing of the FRS data by up to 200 times, offering a substantial prospect towards real time data analysis. This study demonstrates the further development of the FRS technique, with the ultimate goal of inlet flow distortion measurements for in-flight environments.

A pathway to in-flight application of filtered Rayleigh scattering (FRS) is herein presented, including a viable concept, based on recently published related work. The proposed pathway considers the key technical, operational, and regulatory challenges to enable in-flight measurements using FRS for inlet flow distortion characterization ahead of the aeroengine. Solutions to these challenges are proposed, in particular methods for light delivery, flow imaging and integration of the measurement system in the in-flight environment. This complements the experimental lab-scale demonstration of an FRS concept for flow distortion measurements and provides a route for further exploitation as a diagnostic tool for next-gen aircraft.

More compact and short intakes can be a key enabler for the design of Ultra-High Bypass Ratio (UHBR) large civil aero engines. Under key design conditions such as crosswind significant flow distortion at the fan face can adversely affect the performance and engine compatibility. Shorter intakes may result in stronger intake-fan aerodynamic coupling, and the resulting unsteady interactions currently remain insufficiently understood. This study investigates how the intake length and crosswind direction can affect the intake-fan unsteady aerodynamics and distortion. Two computational models are used: a steady Reynolds-Averaged Navier-Stokes method with an Immersed Boundary Method with Smeared Geometry (RANS-IBMSG), and a time resolved fully coupled model (URANS-TRF). Swirl distortion is quantified in the relative frame of reference using a blade tracking method based on blade incidence angle. For the combination of research fan and intake analysed, the findings show that the fan proximity can reduce the onset of gross separation for shorter intake designs. This effect enables a reduction of intake length by about 25% while still meeting a crosswind operating condition requirement. In general, the RANS-IBMSG method predicts the main characteristics of the onset of gross intake separation and is in good agreement with the unsteady URANS-TRF simulations. Based on the unsteady simulations, a gross separation on the intake surface occurred at a lower crosswind velocity for the corotating configuration compared to the counter-rotating case for the research fan and intake studied. The RANS- IBMSG method does not capture some of the aerodynamic features responsible for differences between counter- and co-rotating configurations. Overall, while low-order fan models are useful for early-stage intake design evaluations, coupled unsteady simulations are required to fully capture the effects of crosswind direction and intake-fan interactions.

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).

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.

Highly integrated propulsion systems to achieve fuel savings and reduction of emissions in future aircrafts call for new measurement methods to assess inlet conditions at the engine fan face. Propulsion systems are expected to operate at higher levels of total pressure, total temperature, and swirl distortion due to flow interaction with aerodynamic surfaces and inherent flow distortion within convoluted intakes. Filtered Rayleigh Scattering (FRS) offers capability to assess all these quantities at once, and without the need of seeding particles which cannot be used for in-flight measurements. This paper aims at increasing the technology readiness level of this measurement technique through the application on a lab-scale S-duct diffuser tests and benchmark against Stereo-Particle Image Velocimetry (S-PIV) measurements. Methods to improve the optical integration and mitigate the effect of varying background conditions are hereby explored. Overall, this represents a step forward in the use of FRS as a turnkey solution for the testing and development phase of future propulsion systems.

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.

It is envisaged that future civil aero-engines will operate with ultra-high bypass ratios to reduce the specific fuel consumption. To achieve the expected benefits from the new engine cycles, these new powerplants may mount compact nacelles. For these new configurations the aerodynamic coupling between the powerplant and the airframe may increase. For this reason, it is required to quantify and further understand the effects of aircraft integration for compact aero-engine nacelles. This study provides an insight of the changes in flow aerodynamics as well as quantification of the most relevant performance metrics of the powerplant, airframe and the combined aircraft system across a range of different installation positions. Relative to a conventional architecture, there is an aerodynamic benefit in net vehicle force of about 1.2% for a compact powerplant when installed in forward positions. This is the same improvement that was identified when the aero-engine nacelles were in isolation. However, for close-coupled installation positions, the aerodynamic benefit in net vehicle force erodes to 0.44% due to the larger effects of aircraft integration on compact nacelles.

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.

Within the context of preliminary aerodynamic design with low order models, the methods have to meet requirements for rapid evaluations, accuracy and sometimes large design space bounds. This can be further compounded by the need to use geometric and aerodynamic degrees of freedom to build generalised models with enough flexibility across the design space. For transonic applications, this can be challenging due to the non-linearity of these flow regimes. This paper presents a nacelle design method with an artificial neural network (ANN) for preliminary aerodynamic design. The ANN uses six intuitive nacelle geometric design variables and the two key aerodynamic properties of Mach number and massflow capture ratio. The method was initially validated with an independent dataset in which the prediction error for the nacelle drag was 2.9% across the bounds of the metamodel. The ANN was also used for multi-point, multi-objective optimisation studies. Relative to computationally expensive CFD-based optimisations, it is demonstrated that the surrogate-based approach with ANN identifies similar nacelle shapes and drag changes across a design space that covers conventional and future civil aero-engine nacelles. The proposed method is an enabling and fast approach for preliminary nacelle design studies.

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.