Nacelle Design Research Articles

Numerical and experimental investigations of diffusion-induced boundary layer separation on aero-engine nacelles

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.

Installed nacelle aerodynamics at cruise and windmilling conditions

PurposeThe 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/approachDrooped 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.FindingsThe 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/valueThis 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.

Effect of U-shaped notches on nozzle performance and jet flow structures

In nacelle design, optimization of nozzle shape is crucial for flow control. To understand the effect of nozzle shape on its aerodynamic performance and jet flow structures, the numerical experiment is performed using the dual separate flow reference (DSFR) nozzle released by the Aero Systems Engineering (ASE) FluidDyne Laboratory. A pair of U-shaped notches are made on the nozzle. The results show that the discharge coefficient and thrust angle increase with the size of the notches in all the simulated operating conditions. The notches induce a pair of streamwise vortices, which enhance turbulent mixing and decrease turbulent kinetic energy in the mixing layer.

Artificial neural network for preliminary design and optimisation of civil aero-engine nacelles

Abstract 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.

Multi-fidelity and multi-objective aerodynamic short nacelle shape optimisation under different flight conditions

AbstractThroughout the course of a flight mission, a range of aerodynamic conditions, including design-point conditions and off-design conditions, are encountered. As the bypass ratio increases and the fan-pressure ratio decreases to reduce the engine’s specific fuel consumption, the engine diameters increase, which results in an increase in the nacelle weight and overall drag. To reduce its weight and drag, a shorter nacelle with a length-to-diameter ratio $L/D = 0.35$ is investigated. In this study, an adaptive cokriging-based multi-objective optimisation method is applied to the design of a short aero-engine nacelle. Two nacelle performance metrics were employed as the objective functions for the optimisation routine. The cruise drag coefficient is evaluated under cruise conditions, whereas the intake pressure recovery is evaluated under takeoff conditions. The cokriging metamodel are refined using an effective infilling strategy, where high-fidelity samples are infilled via the modified Pareto fitness, and low-fidelity samples are infilled via the Pareto front. By combining parameterised geometry generation, automated mesh generation, numerical simulations, surrogate model construction, Pareto front exploration based on the non-dominated sorting genetic algorithm-II and sample infilling, an integrated multi-objective optimisation framework for short aero-engine nacelles is developed. Two-objective and three-objective test functions are used to validate the effectiveness of the proposed framework. After the optimisation process, a set of non-dominated nacelle designs is obtained with better aerodynamic performance than the original design, demonstrating the effectiveness of the optimisation framework. Compared with the kriging-based optimisation framework, the cokriging-based optimisation framework outperforms the single-fidelity method with a higher hypervolume value at the same number of iteration loops.

Aerodynamic Optimization Framework for a Three-Dimensional Nacelle Based on Deep Manifold Learning-Assisted Geometric Multiple Dimensionality Reduction

As a core component of an aero-engine, the aerodynamic performance of the nacelle is essential for the overall performance of an aircraft. However, the direct design of a three-dimensional (3D) nacelle is limited by the complex design space consisting of different cross-section profiles and irregular circumferential curves. The deep manifold learning-assisted geometric multiple dimensionality reduction method combines autoencoders (AE) with strong capabilities for non-linear data dimensionality reduction and class function/shape function transformation (CST). A novel geometric dimensionality reduction method is developed to address the typical constraints of nacelle parameterization. Low-dimensional latent variables are extracted from the high-dimensional design space to achieve a parametric representation of 3D nacelle manifolds. Compared with traditional parametric methods, the proposed geometric dimensionality reduction method improves the accuracy and efficiency of geometric reconstruction and aerodynamic evaluation. A multi-objective optimization framework is proposed based on deep manifold learning to increase the efficiency of 3D nacelle design. The Pareto front curves under drag divergence constraints reveal the correlation between the geometry distribution and the surface isentropic Mach number distribution of 3D nacelles. This paper demonstrates the feasibility of the proposed geometric dimensionality reduction method for direct multi-objective optimization of 3D nacelles.

Geometrically compatible integrated design method for conformal rotor and nacelle of distributed propulsion tilt-wing UAV

The inner rotors of distributed propulsion tilt-wing Unmanned Aerial Vehicles (UAVs) are often folded in the cruising state and deployed in vertical take-off and landing to cope with the huge difference in thrust requirements. However, the blades of the conventional rotor have poor conformality with the nacelle profile, which will greatly increase the drag of the UAV after folding. This paper proposes an integrated method for the design of rotor and nacelle considering geometric compatibility to reduce the drag of the folded rotor and nacelle, so as to further improve the aerodynamic efficiency in cruise while ensuring the rotor efficiency in the vertical flight mode. A geometric mapping model based on nacelle design parameters and rotor design parameters is established, and a parametric model and aerodynamic optimization model of the outer arc airfoil family are developed. In addition, a rotor performance analysis model and a neural network response surface model for nacelle drag prediction that meet the requirements of confidence level are established. Based on the oblique inflow blade element momentum theory method, numerical simulation method, and genetic algorithm, an integrated optimization framework of the design of the conformal rotor and nacelle is built. Then, a geometrically compatible integrated optimization for the rotor and nacelle is carried out with the objective of maximizing energy efficiency in the full mission profile. Finally, a conformal rotor and nacelle design solution is obtained, which satisfies geometric compatibility and thrust constraints while providing high thrust efficiency and low cruising drag. A comparison of the results of the integrated design and the conventional rotor optimization design shows that the drag of the conventional rotor is 3.45 times that of the conformal integrated design in the cruising state, which proves the effectiveness and necessity of the proposed method.

Aerodynamic Investigation of the High-Lift Performance of a Propeller-Driven Regional Transport Aircraft with Distributed Propulsion

Increased high-lift capabilities due to propeller slipstream, i.e. slipstream deflection, is seen to be one of the main benefits of distributed propulsion, as it may lead to a reduction in the main wing size and thus to reduced drag in cruise flight and/or reduced system weight and complexity. The presented work assesses the potential of distributed propulsion on the high-lift capabilities of a novel transport aircraft design from an aerodynamic point of view. The assessment is based on a regional propeller-driven transport aircraft designed within the European IMOTHEP project. Based on the initial aircraft design, a sensitivity study on the number of propellers and propeller positions with regards to the maximum lift coefficient under take-off conditions has been performed. Moreover, adjustments to the nacelle design and the propulsor integration have been investigated. The study indicates significant increases in the maximum effective lift coefficient in take-off of up to +42% due to slipstream deflection. The increase is thereby strongly dependent on the number of propellers and the propeller positions.

Over-wing integration of ultra-high bypass ratio engines: A coupled wing redesign and engine position study

The integration of next-generation high-bypass turbofan engines poses a major challenge to the aeronautical industry due to the larger fans necessary to achieve more fuel-efficient engines. The limited space underneath the wings and the strict ground clearance constraints bring the necessity to investigate solutions other than the conventional under-wing mounted engines. Over-wing installed nacelles have the potential to solve the ground clearance issue and, in addition, might reduce ground noise due to acoustic shielding from the wing. Nevertheless, a strong and complex coupling between aerodynamics and propulsion is the result of such integration choice, and traditional design practices may result in configurations with prohibitively high drag penalties. This paper presents a novel wing redesign method, specifically developed for over-wing mounted engines. The wing is reshaped to recover the spanwise lift distribution of the clean airframe (wing-body) configuration, for a single aisle airliner at a cruise condition. The wing redesign is conducted along with an engine position sensitivity study, in which the wing is reshaped for different engine axial and vertical positions. The coupling between propulsion and aerodynamics is thoroughly investigated, as well as the interaction and interference effects between the wing, pylon, and nacelle. Moreover, the best over-wing solution is compared to a baseline under-wing mounted nacelle. Results show that, by applying the developed method, an overall drag reduction of 17.65 counts, or 6.4%, was obtained, compared to the initial over-wing configuration, comprising the original wing and baseline engine position. Nonetheless, the best over-wing nacelle design is still 5.58 counts, or 2%, higher in overall drag compared to the baseline under-wing mounted nacelle case.

Double-decoupled inverse design of natural laminar flow nacelle under transonic conditions

The inverse design based on the pressure distribution is an essential approach to realize the improvement of Natural Laminar Flow (NLF) performance for nacelles. However, the direct definition of target pressure distribution at design point is challenging for the dilemma to consider the constraints of shock wave and laminar flow at the same time. In addition, the universality of method will be limited when the inverse design is strongly coupled with the solver. Thus, a double-decoupled methodology based on the relationship of pressure distributions between design and off-design points is proposed in this paper, which realizes the decoupling of constraints in shock wave and laminar flow on target pressure distribution as well as the decoupling of flow field solution and inverse design method. Aimed at an isolated flow-through-nacelle of high bypass ratio, the target pressure distribution with appropriate favorable gradient and shock-free feature is defined according to physical principles at the off-design point of Ma = 0.80 while the transonic and laminar performance are examined at the design point of Ma = 0.85. The solution of flow field is based on γ-Reθ transition model and the inverse design is based on residual-correction method. With the inverse design starting from off-design point, the performance of shock wave and laminar flow at design point are both improved. The local shock wave after the lip of nacelle is eliminated effectively while the streamwise length of laminar flow region is doubled and exceeds to 30% of the chord length. The percentage of drag reduction for outboard surface is 12.7% for friction drag, 7.8% for pressure drag and 10.5% for total drag. The effects of inverse design on the process of transition are analyzed with detailed flow features. The robustness of laminar flow is examined under different variation factors of freestream which are deviated from the design point.

Numerical optimization of transonic natural laminar flow nacelles

Natural laminar flow nacelle is a promising technology for drag reduction. In this paper, an optimization platform is established for the design of transonic axisymmetric and three-dimensional natural laminar flow nacelles for large civil aircraft. The platform adopts the class/shape transformation method for geometric parameterization, a four-equation transition model for transition prediction, and the differential evolution algorithm combined with the radial basis function surrogate model as the optimization algorithm. The optimized axisymmetric nacelle demonstrates approximately 31% chord length of laminar flow, with the drag reduction of 13.3%. The influence of the Reynolds number and inlet mass flow rate on the optimization result is also investigated. The axis-symmetric nacelle optimization method is further used for the section profile design of a non-axisymmetric nacelle. An equivalent method is used to simulate the different local flow angles at different sections in the circumferential direction of the non-axisymmetric nacelle by using different inlet mass flow rates of the axisymmetric nacelle. The optimized natural laminar flow nacelle maintains over 30% chord length of laminar flow with robustness to the change of the freestream angle of attack. The total drag of the non-axisymmetric nacelle is reduced by 5.4% under cruise conditions.

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.

Structural optimization of engine nacelle fan cowl door hinge

To increase propulsive efficiency, future nacelle engines are projected to run at low specific thrust with large bypass ratios. Typically, this entails a larger fan diameter and nacelle, as well as increased drag and weight penalties. As a result, there is a need to construct more compact, nacelles’ components in comparison to present designs in order to minimize drag and weight. These designs are naturally more difficult, necessitating the use of a system to explore and identify the design area that is possible. Designing a nacelle is difficult due to the wide range of operational circumstances. This paper provides a multi-objective optimization technique for fan cowl door component of nacelle design utilizing an evolving genetic algorithm. A collection of geometry definitions utilizing Class Shape Transformations, automated simulation and analysis, a generic algorithm, assessments at various nacelle operating circumstances, and the addition of extra aerodynamic restrictions are all part of the unique framework. This framework was used to look into the nacelle design space for nacelle components. For the new nacelle component design challenge, multi-objective optimization was successfully shown, and the entire system was proved to allow the discovery of the feasible nacelle design space. Moreover, in this paper a series of fan cowl door hinge is designed from the scratch and optimized step by step for reducing the weight of the component without compromising for the stress capacity and load transformations within the allowable limit.

Fluid mechanism analysis on the interaction between natural laminar flow nacelle and wingbody on a transonic aircraft

A transonic natural laminar flow (NLF) nacelle, which is a streamlined fairing used to contain a turbofan engine and mounted under the wing of civil aircraft, can reduce friction drag. Because the fluid mechanism of the interaction between an NLF nacelle and a wingbody is not clear, laminar flow at high Reynolds numbers in the transonic regime is maintained difficultly. In this work, such interaction on a civil aircraft is investigated. Three NLF nacelles with different pressure distribution characteristics and a baseline nacelle with the turbulent flow are examined. These are installed under the wing of a widebody aircraft to investigate the fluid mechanism between a natural laminar flow nacelle and wingbody. The results show that the influence of the wingbody on the fluid characteristics of the nacelle should be considered in the NLF nacelle design. A well-designed isolated NLF nacelle is different from the one that considers the effect of the wingbody. A favorable pressure gradient in the front part of the nacelle is a key factor in drag reduction. An installed NLF nacelle owning a large pressure peak with a weak favorable pressure gradient and a strong shock wave in front of the nacelle is recommended to be applied in civil aircraft.

Acoustic Mode Attenuation in Ducts (Using CFD) with Time-Domain Impedance Boundary Condition

In the framework of improving aircraft nacelle design to comply international regulation on aircraft noise, a time-domain impedance boundary condition (TDIBC) based on the oscillo-diffusive representation is implemented in an industrial computational fluid dynamics (CFD) solver under a characteristic formalism, in order to take into account acoustic liner attenuation. A validation of such TDIBC is carried out, first on a two-dimensional benchmark case, then on a more realistic three-dimensional cylindrical configuration, with focus on the acoustic modal content and its attenuation with and without flow. Unsteady Reynolds-averaged Navier–Stokes computations are performed, with the grazing flow being considered laminar. Good agreement is found between the CFD results and the references (computational aeroacoustics results and experiments), with differences in the order of for most cases. The formulation chosen for the TDIBC proved to have no significant impact on numerical stability or computational time.

Inverse design of aerodynamic configuration using generative topographic mapping

The inverse design method of aerodynamic configuration is hard to give a reasonable pressure distribution, and strongly rely on experience of designers. The method has been difficult to adapt to the needs of modern aircraft design. Aiming at the shortcoming of the method, an efficient and robust aerodynamic configuration inverse design method is developed, employing knowledge of machine learning methods and optimization methods. The present method establishes the mapping between the high dimensional data obtained from aerodynamic shape and pressure distribution and the variables in the latent space. Then, the global optimization is carried out in the latent space by using the genetic algorithm. The optimum pressure distribution and the corresponding shape can be obtained. Through the the GTM model with high precision, there is not necessary for the flow solver in the whole iterative process, thus the design efficiency can be enhanced. Besides, by taking the advantage of optimization method, the target pressure distribution can be given in a very flexible way, and does not need to be physically meaningful. This feature can reduce reliance on the design experience. Airfoils in low speed and transonic flow and a three-dimensional laminar nacelle design cases are carried out. It is shown that the method robustly and efficiently converges to the target pressure, and has good engineering application potential.

Aero-propulsive interaction model for conceptual distributed propulsion aircraft design

PurposeThis research aims to present an aero-propulsive interaction model applied to conceptual aircraft design with distributed electric propulsion (DeP). The developed model includes a series of electric ducted fans integrated into the wing upper trailing edge, taking into account the effect of boundary layer ingestion (BLI). The developed model aims to estimate the aerodynamic performance of the wing with DeP using an accurate low-order computational model, which can be easily used in the overall aircraft design's optimization process.Design/methodology/approachFirst, the ducted fan aerodynamic performance is investigated using a low-order computational model over a range of angle of attack required for conventional flight based on ducted fan design code program and analytical models. Subsequently, the aero-propulsive coupling with the wing is introduced. The DeP location chordwise is placed at the wing's trailing edge to have the full benefits of the BLI. After that, the propulsion integration process is introduced. The nacelle design's primary function is to minimize the losses due to distortion. Finally, the aerodynamic forces of the overall configuration are estimated based on Athena Vortex Lattice program and the developed ducted fan model.FindingsThe ducted fan model is validated with experimental measurements from the literature. Subsequently, the overall model, the wing with DeP, is validated with experimental measurements and computational fluid dynamics, both from the literature. The results reveal that the currently developed model successfully estimates the aerodynamic performance of DeP located at the wing trailing edge.Originality/valueThe developed model's value is to capture the aero-propulsive coupling accurately and fast enough to execute multiple times in the overall aircraft design's optimization loop without increasing runtime substantially.

Optimal shape design and transition uncertainty analysis of transonic axisymmetric natural laminar flow nacelle at high Reynolds number

Laminar flow drag reduction is an important way to improve the cost effectiveness of civil aircraft. In this paper, a hybrid game based optimization method is applied to optimize an axisymmetric natural laminar flow nacelle with inflow and geometric constraints at high Reynolds number. The effectiveness of the optimization algorithm and the ability to explore the solution of multi-objective optimization problem from engineering science are confirmed. In order to maintain the stable laminar flow performance and to achieve robust aerodynamic drag reduction in changing flight conditions, the global sensitivity and uncertainty analysis of the designed axisymmetric laminar nacelle with respect to the coupling of geometric parameters and the coupling of Mach number and turbulence intensity are finally performed, and the effects of the input parameters on the laminar flow range and the total drag coefficient are analyzed. The sensitivities of each parameter and its contribution to the laminar flow region were obtained. This lays the foundation for the subsequent development of a robust optimization modeling and the specification of its effective search space for the three-dimensional natural laminar flow nacelle.

Multipoint Aerodynamic Design of Ultrashort Nacelles for Ultrahigh-Bypass-Ratio Engines

This paper presents a newly developed methodology for multipoint aerodynamic design of ultrashort nacelles for ultrahigh-bypass-ratio turbofan engines. An integrated aerodynamic framework, based on parametric geometry generation and flowfield solution via three-dimensional Reynolds-averaged Navier–Stokes equations, was built and used for designing several ultrashort nacelle shapes and to evaluate their aerodynamic performance. An approach for modeling the inlet–fan coupling is presented and validated. A design strategy is introduced, and various test cases are evaluated under the following critical operating conditions: midcruise, low speed/high angle of attack, and pure crosswind. The major design parameters are highlighted and their influence in the flowfield is discussed in detail for all the chosen flight conditions. Performance was evaluated by assessing inlet flow distortion and by bookkeeping of thrust and drag. The framework has proven to be suitable for designing high-performance nacelles capable of operating under critical flight conditions, without flow separation or high levels of distortion. Drooping the inlet by 4 deg is shown to reduce the drag at cruise by 1.9%, which also has a large beneficial impact on internal lip separation at high-incidence conditions. Furthermore, crosswind was identified as the most severe of the conditions, requiring a drastic reshaping of the nacelle to avoid internal lip separation. Two final nacelle designs were compared: the first allowed inlet separation under a 90 deg crosswind condition, whereas the second was reshaped to be separation-free under all operating conditions. Reshaping to avoid separation has increased drag by 5.1% at cruise.

Computational fluid dynamics for turbomachinery technologies with a focus on high-speed rotor blades

NASA's SR3 8-bladed rotor with a rotational speed of 6350 rpm, an advance ratio of 3.6, and a transonic Mach Number of 0.8 is utilized to evaluate the aerodynamic and rotor effectiveness predictive ability of a CFD simulation using the Ansys commercial software. The experimentally tested model [1] included eight 45° swept blades that were tested at a design operating altitude of 10.68 km (35,000 ft.) and it is the one with the most data in the public domain. In the model, blade sweep and the distinct contour of the spinner and nacelle design were implemented to reduce compressibility losses. The primary goal of this research is to compare performance predictions with the experimental results for power coefficient and efficiency evaluation. Analyses are carried out on a single rotating zone separated by a sliding mesh boundary. We also look at the accuracy of recording downward velocities and swirling angles in the propeller's wake, with the goal to use the RANS technique to investigate how these wake parameters react with aerodynamic surface and the results are compared with the experimental results. Initial computational findings demonstrate increased correlation with wind tunnel measurements at 63.3° degrees pitch.