AIAA Journal Research Articles

Fluid–structural coupling of an impinging shock–turbulent boundary layer interaction at Mach 3 over a flexible panel

We present high-fidelity numerical simulations of the interaction of an oblique shock impinging on the turbulent boundary layer developed over a rectangular flexible panel, replicating wind tunnel experiments by Daub et al. (AIAA Journal, vol. 54, 2016, pp. 670–678). The incoming free-stream Mach and unit Reynolds numbers are $M_{\infty } = 3$ and $Re_{\infty }=49.4\times 10^6 {\rm m}^{-1}$ , respectively. The reference boundary layer thickness upstream of the interaction with the shock is $\delta _0 = 4$ mm. The oblique shock is generated with a rotating wedge initially parallel to the flow that increases the deflection angle up to $\theta _{{max}} = 17.5^{\circ }$ within approximately $15$ ms. A loosely coupled partitioned flow–structure interaction simulation methodology is used, combining a finite-volume flow solver of the compressible wall-modelled large-eddy simulation equations, an isoparametric finite-element solid mechanics solver and a spring-system-based mesh deformation solver. Simulations are conducted with rigid and flexible panels, and the results compared to elucidate the effects of panel flexibility on the interaction. Three-dimensional effects are evaluated by conducting simulations with both full ( $50 \delta _0$ ) and reduced ( $5\delta _0$ ) spanwise panel width, the latter enforcing spanwise periodicity. Panel flexibility is found to increase the separation bubble size and modify its spectral dynamics. Time- and spanwise-averaged streamwise profiles of the wall pressure exhibit a drop over the flexible panel prior to the interaction and a reduced peak pressure in comparison with the rigid case. Spectral analyses of wall pressure data indicate that the low-frequency motions have a similar spectral distribution for the rigid and flexible cases, but the flexible case shows a wider region dominated by low-frequency motions and traces of the panel vibration on the wall pressure signal. The sensitivity of the interaction to small variations in the wedge extent and incoming boundary layer thickness is evaluated. Predictions obtained from lower-fidelity modelling simplifications are also assessed.

Drag Decomposition Using Partial-Pressure Fields: ONERA M6 Wing

Drag decomposition has classically been undertaken using a far-field approach both from a wind-tunnel testing and computational fluid dynamics standpoint. It has been suggested by Schmitz (“Drag Decomposition Using Partial-Pressure Fields in the Compressible Navier-Stokes Equations,” AIAA Journal, Vol. 57, No. 5, 2019, pp. 2030–2038) that a decomposition of static pressure into partial pressure fields can allow one to decompose a computational fluids dynamics solution in the near field. This paper covers the implementation process of partial pressure fields on the ONERA M6 wing at conditions relevant to commercial transport aircraft. Relations between partial pressure field decomposition, the classical far-field theory, and physical drag sources are discussed. It is demonstrated that the near-field partial pressure field decomposition is equivalent to classical far-field analysis. In transonic flow, a new method to compute wave drag using a combination of both partial pressure fields and the far-field method is demonstrated.

New Sensitivity Analysis Methods for Complex Modal Parameters in Asymmetrical Damped System

Most of the sensitivity algorithms with respect to the design parameters adopted the normalization relationship with system property matrices weighted when analyzing an asymmetrical damping linear discrete dynamic system. In this paper, two new methods (an algebraic method and a direct method) based on the Euclidean norm as the normalization conditions are proposed. Both proposed methods are developed to calculate the first- and second-order sensitivities of the complex modal parameters for an asymmetric damped system. The precisions of the Taylor approximations of the second degree are also discussed in the proposed methods. Like other algebraic methods, this paper outlines the proof of numerical stability. Different from Nelson’s direct method (“Simplified Calculation of Eigenvector Derivatives,” AIAA Journal, Vol. 14, No. 9, 1976, pp. 1201–1205), this paper considers complexity of the solution space in complex vector space, and the partitioning scheme of the proposed direct algorithm is more practical. The usefulness and correctness are demonstrated by a one-parameter numerical experiment and a two-parameter numerical experiment.

Computational Study of Wing-Tip Effect for Flow-Control Authority of Dielectric-Barrier-Discharge Plasma Actuator

In the present study, large-eddy simulations of a three-dimensional flow around the NACA 0015 wing controlled by a dielectric-barrier-discharge (DBD) plasma actuator (PA) were conducted, and the influence of the wing tip on the flow was investigated. A half-span rectangular wing with an aspect ratio (AR) of two was selected. The Reynolds number was set to 63,000, and the angles of attack were 12 and 14 deg as the poststall condition. In both cases, the overall significantly increased by the DBD-PA. The increment of the was small around the wing tip, owing to the wing-tip vortex; whereas it increased substantially in the area from the wing root to the midspan, as reported in the previous study (Sato, M., Aono, H., Yakeno, A., Nonomura, T., Fujii, K., Okada, K., and Asada, K., “Multifactorial Effects of Operating Conditions of Dielectric-Barrier-Discharge Plasma Actuator on Laminar-Separated-Flow Control,” AIAA Journal, Vol. 53, No. 9, 2015, pp. 2544–2559) of a two-dimensional airfoil (that is, three-dimensional simulation with a two-dimensional airfoil without tip and root). This study also focused on the effect of the spanwise length of the DBD-PA and compared aerodynamic performance between four DBD-PA lengths to investigate an effective installation length in the current setting. Through the results, the factors that affect flow-control capability around a three-dimensional (3-D) wing (both those common in a two-dimensional airfoil and those peculiar in a 3-D wing) were discussed.

Journal Vol – 15 No -7, July 2020 Journal > Journal > Journal Vol – 15 No -7, July 2020 > Page 6 PERFORMANCE AND EMISSION CHARACTERISTICS OF GASOLINE-ETHANOL BLENDS ON PFI-SI ENGINE Authors: D.Vinay Kumar ,G.Samhita Priyadarsini,V.Jagadeesh Babu,Y.Sai Varun Teja, DOI NO: https://doi.org/10.26782/jmcms.2020.07.00051 admin July 26, 2020 Abstract: Alcohol based fuels can be produced from renewable energy sources and

Journal Vol – 15 No -7, July 2020 Journal > Journal > Journal Vol – 15 No -7, July 2020 > Page 6 PERFORMANCE AND EMISSION CHARACTERISTICS OF GASOLINE-ETHANOL BLENDS ON PFI-SI ENGINE Authors: D.Vinay Kumar ,G.Samhita Priyadarsini,V.Jagadeesh Babu,Y.Sai Varun Teja, DOI NO: https://doi.org/10.26782/jmcms.2020.07.00051 admin July 26, 2020 Abstract: Alcohol based fuels can be produced from renewable energy sources and

Optimized geometrical metrics satisfying free-stream preservation

Computational fluid dynamics and aerodynamics, which complement more expensive empirical approaches, are critical for developing aerospace vehicles. During the past three decades, computational aerodynamics capability has improved remarkably, following advances in computer hardware and algorithm development. However, for complex applications, the demands on computational fluid dynamics continue to increase in a quest to gain a few percent improvements in accuracy. Herein, we numerically demonstrate, in the context of tensor-product discretizations on hexahedral elements, that computing the metric terms with an optimization-based approach leads to a solution whose accuracy is overall on par and often better than the one obtained using the widely adopted Thomas and Lombard metric terms computation (Geometric conservation law and its application to flow computations on moving grids, AIAA Journal, 1979). We show the efficacy of the proposed technique in the context of low and high-order accurate nonlinearly stable (entropy stable) schemes on distorted, high-order tensor product elements, considering smooth three-dimensional inviscid and viscous compressible test cases for which an analytical solution is known. The methodology, originally developed by Crean et al. (2018) in the context of triangular/tetrahedral grids, is not limited to tensor-product cells and it can be applied to other cell-based diagonal-norm summation-by-parts discretizations, including spectral differences, discontinuous Galerkin finite elements, and flux reconstruction schemes.

Koopman-Based Approach to Nonintrusive Reduced Order Modeling: Application to Aerodynamic Shape Optimization and Uncertainty Propagation

A methodology for nonintrusive projection-based nonlinear model reduction originally presented by Renganathan et al. (“Koopman-Based Approach to Nonintrusive Projection-Based Reduced-Order Modeling with Black-Box High-Fidelity Models,” AIAA Journal, Vol. 56, No. 10, 2018, pp. 4087–4111) is further extended toward parametric systems with a focus on application to aerospace design. Specifically, the method is extended to address static systems with parametric geometry (that deforms the mesh) in addition to parametric freestream boundary conditions. The main idea is to first perform a transformation on the governing equations such that it is lifted to a higher-dimensional but linear underdetermined system. This enables one to extract the system matrices easily as compared to that of the original nonlinear system. The underdetermined system is closed with a set of model-dependent nonlinear constraints upon which the model reduction is finally performed. The methodology is validated on the subsonic and transonic inviscid flows past the NACA0012 and the RAE2822 airfoils with parametrized shapes. The utility of the approach is further demonstrated by applying it to two common problems in aerospace design, namely, derivative-free global optimization and parametric uncertainty quantification with Monte Carlo sampling. Overall, the methodology is shown to achieve accuracy up to 5% and computational speedup of two to three orders of magnitude relative to the full order model. Comparison against another nonintrusive model reduction method revealed that the proposed approach is more robust, accurate, and retains the consistency between the state variables.

Model of Distributed Receptivity to Kinetic Fluctuations in High-Speed Boundary Layers

The receptivity of high-speed compressible boundary layers to kinetic fluctuations is considered within the framework of fluctuating hydrodynamics. The formulation is based on the idea that KF-induced dissipative fluxes may lead to the generation of unstable modes in the boundary layer. Fedorov and Tumin (“Receptivity of High-Speed Boundary Layers to Kinetic Fluctuations,” AIAA Journal, Vol. 55, No. 7, 2017, pp. 2335–2348) solved the receptivity problem using an asymptotic matching approach, which used a resonant inner solution in the vicinity of the neutral point. This study adopts a method of multiple scales approach, which requires fewer assumptions about the locus of primary excitation. The approach is modeled after the one taken by Luchini (“Receptivity to Thermal Noise of the Boundary Layer over a Swept Wing,” AIAA Journal, Vol. 55, No. 1, 2017, pp. 121–130) to study low-speed incompressible boundary layers over a swept wing. The new framework is used to study examples of high-speed flat plate boundary layers whose spectra exhibit nuanced behavior near the generation point, such as first-mode instabilities and near-neutral evolution over moderate length scales.

Wall-Resolved Large-Eddy Simulations of Transonic Shock-Induced Flow Separation

Wall-resolved implicit large-eddy simulations based on spatial filtering are performed for shock-induced flow separation over an axisymmetric bump at a Mach number of 0.875 and a chord-based Reynolds number of 2.763 million. The incoming boundary layer has a momentum thickness Reynolds number of 6600 at 1.5 chords upstream of the bump. The calculations, which employ up to 24 billion grid points, simulate the experimental model of Bachalo and Johnson (“Transonic, Turbulent Boundary-Layer Separation Generated on an Axisymmetric Flow Model,” AIAA Journal, Vol. 24, No. 3, 1986, pp. 437–443), except that the tunnel walls are ignored and free air is assumed. The effects of domain span and grid resolution are examined along with main flowfield features. The predicted shock position as well as separation and reattachment locations agree well with experiment. Grid convergence is observed in the attached region well upstream of separation. Two-point azimuthal correlations suggest that a span of at least 20 deg is needed in the attached region, whereas a 120 deg span appears insufficient for correlations to completely decay to zero by midspan in the separation/reattachment region. Computed rms surface pressure fluctuations display a sharp primary peak at the shock foot and a broader secondary peak near the reattachment location. Simulations reveal evidence of low-frequency shock unsteadiness; however, a longer statistical sample is needed to investigate this phenomenon.

Dynamic Shear-Lag Model for Stress Transfer in Piezoelectric Transducer Bonded to Plate

Crawley and Luis (“Use of Piezoelectric Actuators as Elements of Intelligent Structures,” AIAA Journal, Vol. 25, No. 10, 1987, pp. 1373–1385) presented in the year 1987 a classic one-dimensional (1-D) quasi-static shear-lag model for the stress transfer between piezoelectric actuators and the host beam to which they are bonded through adhesive layers. Since then, it has been the backbone in modeling piezoelectric actuation and sensing. However, even after 30 years of its existence, the model remains to be extended to incorporate the inertia terms of the system, which is important for high-frequency applications, such as structural health monitoring. This work presents a consistent and generic extension of the 1-D shear-lag model for the dynamic stress transfer of a piezoelectric wafer transducer bonded to a thin isotropic plate, incorporating the inertia of the transducer, host plate, and the adhesive. The transducer and the plate were considered to be under plane strain condition. The plate was modeled using the classical Kirchhoff’s plate theory, whereas the wafer transducer was considered to undergo only extensional deformation. The model led to a simple closed-form solution for the shear stress induced at the transducer–plate interface under harmonic electric potential excitation. The results were compared with detailed two-dimensional finite element dynamic simulations. A detailed numerical study was conducted to illustrate the influence of various transducer, plate, and adhesive parameters, as well as the excitation frequency on the inertia effect on the interfacial shear stress.

Drag Decomposition Using Partial-Pressure Fields in the Compressible Navier–Stokes Equations

The static-pressure field in the steady and compressible Navier–Stokes equations is decomposed into Euler (inviscid) and dissipative (viscous) partial-pressure fields as a generalization of the incompressible pressure decomposition previously reported by Schmitz and Coder (“Inviscid Circulatory-Pressure Field Derived from the Incompressible Navier–Stokes Equations,” AIAA Journal, Vol. 53, No. 1, 2015, pp. 33–41). The primary purpose of partial-pressure fields is to provide a means of dissecting local drag contributors over a lifting body. The analysis shows that the integral of the Euler pressure over the surface of a lifting body of thickness recovers the Kutta–Joukowski theorem for lift, and results in Maskell’s formula for the vortex-induced drag in the limit of high Reynolds number; the combined integral of the dissipative pressure and wall shear stress results in a generalized form of Oswatitsch’s formula for entropy-flux drag. Transport equations are derived with well-posed boundary conditions for both the Euler and dissipative partial-pressure fields for implementation in computational fluid dynamics codes as a complement to far-field and volumetric methods for drag decomposition of complex aircraft configurations.

Analysis of Second-Mode Amplitudes on Sharp Cones in Hypersonic Wind Tunnels

This research effort coordinated by the NATO Science and Technology Organization Applied Vehicle Technology 240 specialists’ group compiles and analyzes second-mode amplitudes on sharp slender cones. The analysis focuses on pressure fluctuations measured with piezoelectric sensors in 11 hypersonic wind tunnels operated by nine organizations located in three NATO countries (Belgium, Germany, and the United States) and Japan. The measurements are at freestream Mach numbers between 5 and 14, unit Reynolds numbers between 1.5 and , and wall-to-total temperature ratios between 0.1 and 0.8. The study shows that second-mode growth rates can be predicted with parabolized stability equations (PSEs) over the wide range of conditions. The maximum second-mode amplitudes vary weakly at edge Mach number greater than but significantly decrease at lower . The maximum factor envelope from PSE and the measured amplitudes are used to estimate the initial amplitudes . At each Mach number, varies approximately as Reynolds number. This leads to transition factors that increase with Reynolds number. This behavior is consistent with the results from Marineau (“Prediction Methodology for Second-Mode-Dominated Boundary-Layer Transition in Wind Tunnels,” AIAA Journal, Vol. 55, No. 2, 2017, pp. 484–499). This study indicates that amplitude-based methods for second-mode-dominated hypersonic boundary-layer transition predictions should be applicable across a wide range of conditions.

Including Flow–Acoustic Interactions in the Helmholtz Computations of Industrial Combustors

A methodology to include flow–acoustic interactions in a Helmholtz solver is proposed and used to quantify their impact on the thermoacoustic modes of a real industrial combustor. This methodology takes advantage of the compactness of flow–acoustic interactions, modeling the regions of interest by matrices included in the Helmholtz computation thanks to a dedicated boundary condition. The one-dimensional formulation introduced in Ni et al. (“Accounting for Acoustic Damping in a Helmholtz Solver,” AIAA Journal, Vol. 55, No. 4, 2017, pp. 1205–1220) is extended to three-dimensional configurations so that the dissipation of dilution holes, swirlers, and perforated plates can be accounted for. The potential of the method is illustrated by performing the thermoacoustic analysis of an annular combustor from SAFRAN. The Helmholtz solution corresponding to an azimuthal mode is analyzed and shows that flow–acoustic interactions dissipate an important part of the acoustic energy but also limit the amplification created by the flame.

Scattering of Tollmien-Schlichting waves as they pass over forward-/backward-facing steps

Forward-/backward-facing steps in boundary-layer flows are often seen in engineering applications, and they have potential impacts on laminar-turbulent transition through scattering of the oncoming instability modes (e.g., Tollmien-Schlichting (T-S) waves). This issue is studied in the present paper by applying a local scattering framework, which is a rather generic mathematical framework on describing the mode scattering process. In this framework, a high-Reynolds-number triple-deck formalism is employed, and a transmission coefficient, defined as the ratio of the asymptotic amplitude of the instability mode downstream of the step to that upstream, is introduced. Through the systematical study, it has been found that both the forward- and backward-facing steps have a destabilizing effect on the oncoming T-S waves in subsonic boundary layers, this effect increases with the height of the step and/or the frequency of the T-S wave, and a backward-facing step (BFS) always has a greater impact than a forward-facing step (FFS). These facts agree with most of the previous investigations. However, one numerical study (WORNER, A., RIST, U., andWAGNER, S. Humps/steps influence on stability characteristics of two-dimensional laminar boundary layer. AIAA Journal, 41, 192–197 (2003)), which was based on an ad-hoc configuration, showed an opposite impact of an FFS. Through the investigation on the specific configuration, it is revealed that the wrong conclusion was drawn by misinterpreting the numerical results.

Unsteady Simulations of a Fan/Outlet-Guide-Vane System: Tone–Noise Computation

Tone noise from a rotor–stator system representing a turbofan is studied based on high-fidelity unsteady simulations. The baseline configuration of the NASA source-diagnostic-test geometry with the entire hard-wall nacelle is solved by using the unsteady Reynolds-averaged Navier–Stokes equations and the improved delayed detached-eddy simulation at approach and takeoff conditions. Following a paper by Shur et al. (“Unsteady Simulations of a Fan/Outlet-Guide-Vane System: Aerodynamics and Turbulence,” AIAA Journal, 2018), which focused on the simulation methodology for the turbulence treatment and the aerodynamics, this paper investigates noise-generation/propagation phenomena of the first and second blade-passing-frequency tones: The analyses include mode decomposition of the tone power levels, their comparison with the NASA source diagnostic test data, nonlinear propagation of rotor-locked tones in the inlet, and the unsteady response of the stator against the rotor viscous wakes. With a duct-mode-matching method, the improved delayed detached-eddy simulation can predict tone levels within 6 dB and usually less, except for the rotor-locked tones. Although both simulations capture nonlinear decay in the inlet, “identical blade” simulations are unable to redistribute the energy into multiple pure tones at the takeoff speed. In the viscous-wake/outlet-guide-vane interaction, unsteady stator responses computed by both simulations seem not to be in accordance with classical models; moreover, the unsteady Reynolds-averaged Navier–Stokes simulation can predict only cut-on tone levels due to the lack of turbulence contents.

Parametric Investigation of Stall Cell Formation on a NACA 0015 Airfoil

An experimental parametric exploration of stall cells on a NACA 0015 airfoil with aspect ratio was conducted. First, flowfield measurements were performed using stereoscopic particle image velocimetry at conditions where a stall cell was present. This enabled some of the properties associated with stall cell(s) and stall cell transition to be defined. For the parametric study, angles of attack from to 21.5 deg and Reynolds numbers from to were observed using oil-flow visualizations, demonstrating the transition from full span separation to fully formed single or dual stall cells. At and , full span separation with recirculation regions was present, whereas at midrange of Reynolds numbers tested, three-dimensional separation was observed with undulations across the span. At , a single stall cell formed when the Reynolds number exceeded a critical Reynolds number. At the highest angles of attack tested, a dual stall cell configuration was present at high Reynolds numbers. In addition, the evolution of a stall cell was observed to be reminiscent of the mechanism proposed by Weihs and Katz (“Cellular Patterns in Poststall Flow over Unswept Wings,” AIAA Journal, Vol. 21, No. 12, 1983, pp. 1757–1759). It was also shown that, under specific conditions, a bistable condition was present, where two different patterns were obtained for identical flow conditions. Finally, a less-detailed parametric investigation was also conducted for a NACA 0015 airfoil with , which suggested a relationship between the critical Reynolds number and the aspect ratio.

Effect of square dimples on the temperature distribution and heat transfer coefficient of an NACA0012 airfoil

ABSTRACTThe thermal boundary layer is the barrier for convective heat transfer, since the temperature within the thermal boundary layer is lower than the free stream temperature; therefore, it is preferable to maintain a small thermal boundary layer thickness (δt). Among the various categories of dimples like square, rectangle and triangle we considered square dimples for the upper surface of the NACA0012 wing throughout the span. The dimples increase the coefficient of lift due to delayed flow separation. The skin friction coefficient improvement reduces the δh which in turn reduces the (δt), therefore the rate of heat transfer is increased and surface temperature distribution becomes smooth. Dimples also act as fins, reducing the upper surface temperature to lower than the lower surface, which in turn increases the lift particularly in low-speed and micro-scale airfoils [Kim, Jongmin, Zvi Rusak, and Nikhil Koratkar. 2003. ‘Small-Scale Airfoil Aerodynamic Efficiency Improvement by Surface Temperature and Heat Transfer.’ AIAA Journal 41 (11): 2105–2113].

Large-Eddy Simulation of Flow over a Wall-Mounted Hump with Separation and Reattachment

Wall-resolved large-eddy simulation of a model turbulent flow involving favorable and adverse pressure gradients, imposed by surface curvature of a wall-mounted hump, is performed for a spanwise-periodic computational domain. The flow acceleration over the front portion of the hump is strong enough to exceed the relaminarization criterion of Narasimha and Sreenivasan (“Relaminarization in Highly Accelerated Turbulent Boundary Layers,” Journal of Fluid Mechanics, Vol. 61, No. 3, 1973, pp. 417–447) but only over a relatively short streamwise extent. Although the flow does not relaminarize, turbulent skin-friction variation exhibits a plateau, also observed in the experiment of Greenblatt et al. (“A Separation Control CFD Validation Test Case, Part 1: Baseline and Steady Suction,” AIAA Journal, Vol. 44, No. 12, 2006, pp. 2820–2830), before it again continues to rise. The subsequent adverse pressure gradient is strong enough to cause flow separation. The location and extent of flow separation, including skin-friction distributions, compare reasonably well with experimental data. Computed velocity and Reynolds stress profiles are also compared with the experimental results. A systematic study of the effect of computational domain span, subgrid-scale model, tunnel backpressure, incoming boundary layer, grid refinement, Mach number, and top tunnel wall contour (which models the blockage effect of the experimental setup) is carried out, and sensitivity of the results to these parameters is discussed.

Aeroacoustic Prediction of a Multi-Element Airfoil Using Wall-Modeled Large-Eddy Simulation

Wall-modeled large-eddy simulation is employed in combination with the Ffowcs Williams–Hawkings equation to predict and analyze the aeroacoustics of a 30P30N three-element airfoil. The results are compared with experimental measurements, and quantitative agreement is obtained in terms of flow statistics, frequency spectra of pressure fluctuations on the slat surface, and the far-field noise spectra, thus demonstrating the feasibility and accuracy of the wall-modeled large-eddy simulation approach. The slat cove is confirmed to be the primary sound source and emits narrowband tonal noise at several resonance frequencies in addition to broadband noise. The main element is the largest contributor to low-frequency noise, and the flap is a negligible noise source at all frequencies. Computational results for different freestream Mach numbers indicate that the resonance frequencies are well predicted by the semi-empirical models of Block (“Noise Response of Cavities of Varying Dimensions at Subsonic Speeds,” NASA Technical Note D-8351, 1976) and Terracol et al. (“Investigation of the Unsteady Flow and Noise Generation in a Slat Cove,” AIAA Journal, Vol. 54, No. 2, 2016). The intensity of flyover noise is shown to largely scale with the fifth-power of Mach number, as suggested by Guo and Joshi (“Noise Characteristics of Aircraft High Lift Systems,” AIAA Journal, Vol. 41, No. 7, 2003), but in most other directions, the Mach number scaling of sound intensity is higher than the fifth power.

Wind-Tunnel and CFD Investigations Focused on Transition and Performance Predictions of Laminar Wings

This paper is a continuation of the paper recently published by Hue et al. (“Experimental and Numerical Methods for Transition and Drag Predictions of Laminar Airfoils,” AIAA Journal, Vol. 53, No. 9, Sept. 2015, pp. 2694–2712) that focused on transition and drag predictions of laminar airfoils. The extension of such studies to three-dimensional configurations representative of modern civil aircraft is a further step toward the implementation of natural laminar-flow technologies. The present work, therefore, focuses on validating the laminar design of a low-swept wing for business jet applications. In 2015, an experimental campaign was carried out in the European Transonic Windtunnel, and included laminar/turbulent transition measurements with temperature-sensitive paint at Mach and Reynolds numbers typical of cruise flight conditions. Subsequently, fluid dynamics computations were performed on this aircraft geometry either with a Reynolds-averaged Navier–Stokes solver using both Tollmien–Schlichting and crossflow transition criteria, or with a boundary-layer code using either database methods for transition location or exact stability analyses. In this paper, experimental and numerical transition predictions are compared for three representative cases corresponding to different angles of attack. The agreement that is achieved is satisfactory, and extended regions of laminar flow are observed on the wing at cruise lift levels. In these conditions, the drag reduction can account for 10–15% of the aircraft drag.