Natural Laminar Flow Research Articles

Numerical study of natural laminar flow based on CRM NLF test

Numerical study of natural laminar flow based on CRM NLF test

Attenuation of boundary-layer instabilities for natural laminar flow design on supersonic highly swept wings

To meet the challenge of drag reduction for next-generation supersonic transport aircraft, increasing attention has been focused on Natural Laminar Flow (NLF) technology. However, the highly swept wings and high-Reynolds-number conditions of such aircraft dramatically amplify Crossflow (CF) instabilities inside boundary layers, making it difficult to maintain a large laminar flow region. To explore novel NLF designs on supersonic wings, this article investigates the mechanisms underlying the attenuation of Tollmien–Schlichting (TS) and CF instabilities by modifying pressure distributions. The evolution of TS and CF instabilities are evaluated under typical pressure distributions with different leading-edge flow acceleration region lengths, pressure coefficient slopes and pressure coefficient deviations. The results show that shortening the leading-edge flow acceleration region and using a flat pressure distribution are favorable for suppressing CF instabilities, and keeping a balance of disturbance growth between positive and negative wave angles is favorable for attenuating TS instabilities. Based on the uncovered mechanisms, a strategy of supersonic NLF design is proposed. Examination of the proposed strategy at a 60° sweep angle and Ma = 2 presents potential to exceed the conventional NLF limit and achieve a transition Reynolds number of 17.6 million, which can provide guidance for NLF design on supersonic highly swept wings.

Similar Solutions for Changing Characteristics of Natural Convection Laminar Flow around a Vertical Rectangular Inclined Plane Surface

The study investigates laminar natural convection flow around a vertical rectangular inclined plane surface, specifically focusing on the effects of viscous and pressure stress work. Previous research primarily addressed these effects in two-dimensional natural convection flow scenarios, while this study extends the analysis to encompass three exterior situations, accounting for variations in fluid properties outside compressible boundary layers. Employing numerical methods, the study aims to predict essential flow parameters including Prandtl’s number (Pr), controlling parameter (C), and additional parameters (A, B, D, E, ∅). The investigation yields velocity profiles (F'(0), S'(0)), temperature profiles , skin frictions (F''(0), S''(0)), and heat transfer coefficients θ'(0). Numerical results are presented to illustrate the impact of different parameters on these flow characteristics. Additionally, tabulated data in Tables-2 and Tables-3 offer further insights into the predicted flow parameters and their variations under varying conditions.

Conformally Decambered Natural Laminar Flow Blades for Vertical-Axis Wind Turbines

A two-element natural laminar flow airfoil—commonly used on medium-altitude, long-endurance, unmanned air vehicles—was conformally decambered for application on a two-bladed, H-rotor, vertical-axis wind turbine. Blade kinematics were used to determine the virtual camber line, which was then used to conformally map the original profile onto the chord line. Both decambered and original blade profiles were evaluated experimentally using large chord-radius ratios (0.6 and 0.75) that exploited dynamic stall to produce the driving torque. Decambered blades showed substantially greater power and torque coefficients than the original blades, up to 60 and 27%, respectively, which represents the first experimental validation of conformal decambering. Relatively large peak power coefficients of 0.28 were attained, despite maximum chord-based Reynolds numbers being less than 2×105. Depending upon the chord-radius ratio, either light or deep dynamic stall occurred in the second upstream quadrant, and the flap flow remained attached virtually throughout. In contrast, on the original profiles, massive separation was observed on the blades, and the flap flow remained separated due to assumed outer surface flow separation. Future research should consider surface pressure and flowfield measurements, significantly higher Reynolds numbers, and variable intracycle flap deflection mechanisms to optimize performance and minimize unsteady loads.

Natural laminar flow leading edge: Requirements, design, and experimental validation under operational loads

To unlock the potential of natural laminar flow aircraft wings, novel structural designs are necessary for the wing leading edge and its attachment. Those designs may not impinge on operability and maintenance of the aircraft. This paper presents a leading edge and attachment design for the natural laminar flow environment as well as the testing of this design on a large-scale ground-based demonstrator. The leading edge design is numerically verified by considering operational loads. For operational viability, the replacement of damaged leading edges without alterations to the spare part is desirable. In a series of tests, such interchange trials are made with two ▪ full complexity leading edge segments and the aerodynamic step height at the interface of leading edge and wing cover is assessed in both ground and cruise deformation. The leading edge design and its attachment concept were proven to support natural laminar flow step height requirements even under global part deformations of a multi-material structure under thermal loading both numerically and experimentally. Interchangeability of the leading edge is demonstrated with very low mean variation in step height between different installations.

Natural laminar flow airfoil design via adjoint-based transition onset delay

This work is focused on the aerodynamic shape optimization (ASO) of airfoils with the aim of attaining natural laminar flow (NLF) designs. The two-equation Amplification Factor Transport (AFT) model is utilized for an accurate prediction of the transitional boundary layer flows. The NLF designs are achieved by utilizing two different objective functions. In the first approach, the conventional drag minimization problem is solved with a target lift coefficient. In addition, a second approach is proposed in this work that directly aims at delaying the transition onset on both the pressure and suction sides, thus expanding the natural laminar flow region around the airfoil. It is shown that by utilizing a sigmoid fitting of the turbulence index profile, the transition onset locations can be accurately predicted in a differentiable and smooth fashion, which is essential to the adjoint-based sensitivity analysis of the RANS solver. The efficacy of the proposed technique for NLF design has been tested for NACA 0012 and RAE 2822 airfoils at moderate lift coefficients. Results have shown that a significantly lower drag count can be achieved by delaying the transition onset locations compared to the case where only a drag minimization is sought. Additionally, the transition onset locations on both sides of the airfoil are delayed using our proposed technique which results in a significant expansion of the natural laminar flow region.

An adjoint-based methodology for calculating manufacturing tolerances for natural laminar flow airfoils susceptible to smooth surface waviness

An adjoint-based method is presented for determining manufacturing tolerances for aerodynamic surfaces with natural laminar flow subjected to wavy excrescences. The growth of convective unstable disturbances is computed by solving Euler, boundary layer, and parabolized stability equations. The gradient of the kinetic energy of disturbances in the boundary layer (E) with respect to surface grid points is calculated by solving adjoints of the governing equations. The accuracy of approximations of ΔE\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\Delta E$$\\end{document}, using gradients obtained from adjoint, is investigated for several waviness heights. It is also shown how second-order derivatives increase the accuracy of approximations of ΔE\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\Delta E$$\\end{document} when surface deformations are large. Then, for specific flight conditions, using the steepest ascent and the sequential least squares programming methodologies, the waviness profile with minimum L2-\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$L2-$$\\end{document}norm that causes a specific increase in the maximum value of N- factor, ΔN\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\Delta N$$\\end{document}, is found. Finally, numerical tests are performed using the NLF(2)-0415 airfoil to specify tolerance levels for ΔN\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\Delta {N}$$\\end{document} up to 2.0 for different flight conditions. Most simulations are carried out for a Mach number and angle of attack equal to 0.5 and 1.25∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$1.25^{\\circ }$$\\end{document}, respectively, and with Reynolds numbers between 9×106\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$9\ imes 10^6$$\\end{document} and 15×106\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$15\ imes 10^6$$\\end{document} and for waviness profiles with different ranges of wavelengths. Finally, some additional studies are presented for different angles of attack and Mach numbers to show their effects on the computed tolerances.Graphic abstract

Aerostructural Optimization and Comparative Study of Twin-Fuselage and Strut-Braced-Wing Aircraft Configurations

The ultrahigh-aspect-ratio wing (UHARW) concept is a promising configuration to achieve future sustainable aviation goals. Twin-fuselage (TF) and strut-braced-wing (SBW) configurations are characterized by smaller structural bending moments and shear forces in the wing and are promising concepts for realizing UHARW designs. This paper addresses the aerostructural optimization problem of TF and SBW configurations with UHARW by using a coupled adjoint aerostructural optimization tool, which is composed of a geometrically nonlinear structural solver and a quasi-three-dimensional natural laminar flow (NLF) aerodynamic solver. The optimization results show significant improvements in fuel efficiency and performance for the TF and SBW aircraft, with fuel mass reductions of 13 and 10%, respectively, compared to the corresponding baseline aircraft designed in the conceptual design phase. In comparison to the original reference aircraft A320neo, the optimized TF and SBW have 48 and 31% lower fuel weights, respectively. The NLF range of both upper and lower wing surfaces is expanded during optimization. The optimized SBW configuration has a wing aspect ratio of 26.01, while the optimized TF has a wing aspect ratio of 20.74, indicating that the SBW concept is more conducive to realizing UHARW design compared with the TF configuration studied in this work. The optimized TF aircraft has a lighter fuel weight and gross weight compared to the optimized SBW aircraft, which is because the TF aircraft has a lighter operational empty weight, including a lighter fuselage structural weight, landing gear weight, etc., whereas the top-level aircraft requirements are the same for both aircraft, including range, payload, and cruise Mach.

Aerodynamic drag reduction through a hybrid laminar flow control and variable camber coupled wing

This paper presents computational fluid dynamics (CFD) results of a transonic transport aircraft wing, featuring a hybrid laminar flow control (HLFC) and variable camber (VC) technology coupling. The goal of the paper is a quantitative assessment of synergistic effects for aerodynamic drag reduction, when combining the possibility of actively shaping the surface pressure distribution of the wing through VC with the passive natural laminar flow aspect of HLFC. Modeling approaches for both VC integration, achieved through deflections of an Adaptive Dropped Hinge Flap (ADHF), and boundary layer suction into a CFD workflow are presented. Transition prediction is accomplished with the γ−Reθ+Cross-flow(CF) transition turbulence model, and linear stability theory (LST) coupled to a two-N-factor transition prediction method. Both methods reflect synergy-driven effects. While the extent of laminar flow predicted by the γ−Reθ+CF model is limited, both favorable ADHF settings and boundary layer suction lead to an overall drag reduction. This is correspondingly reflected through the analyses via LST. Furthermore, the latter predicts a pronounced downstream shift in transition position with increasing suction strength, where the inherent differentiation of transition mechanisms in the two-N-factor method between Tollmien-Schlichting (TSI) and cross-flow instabilities (CFI) reflects a marked reduction in TSI N-factors with increasing ADHF deflection angles. Even though N-factors associated to CFI partially increase through ADHF deflections, the intended synergistic effects are represented throughout the investigated parameter range, leading to an increase in the extent of laminar flow for adequate ADHF deflection angles.

Multidisciplinary Design Optimization of Transonic Wings with Boundary-Layer Suction

A quasi-three-dimensional aerodynamic solver is developed for the aerodynamic analysis of wings in a transonic regime that is able to capture the effect of BLS in hybrid laminar flow control (HLFC) application or transition to turbulent flow for natural laminar flow (NLF). The tool provides accurate results, but without the high computational cost of high-fidelity tools. The solver combines the use of an Euler flow solver characterized by an integral boundary-layer method and linear stability analysis using a approximation for transition prediction. In particular, a conical transformation is adopted, including the determination of the shock-wave position. The solver is implemented in a multidisciplinary design optimization (MDO) framework, including wing weight estimation and aircraft performance analysis. The framework consists of different modules: aerodynamics, structure, suction system analysis, and performance evaluation. Using a genetic algorithm and considering HLFC technology, wing MDO has been performed to find the optimum wing planform and airfoil shape. A backward-swept wing (BSW) aircraft, developed inside the Cluster of Excellence–Sustainable and Energy Efficient Aviation (A) is studied. Novel technologies such as active flow control, limited maximum load factors due to load alleviation, and novel materials allow a fuel weight reduction of 6%.

Investigation of the heat transfer coefficient for a red clay brick

AbstractThere is a need for the development of materials for thermal isolation. Heat energy is used in different ways, for example, for house warming or production of electricity from solar power plants. However, to use the heat energy efficiently isolating materials are needed. There are different materials available to isolate houses and heat storage containers. However, those materials are expensive and some are over engineered. The best available material for thermal isolation would be air that has the thermal coefficient . Thus, materials that trap a large amount of air are good isolators. On the other hand, some isolating materials are harmful for the environment, for example, glass wool, asbestos. An alternative are red clay bricks because red clay is a natural product that is present everywhere on the planet. But the thermal coefficient of red clay bricks is high . To reduce the thermal coefficient cavities can be added in the brick. The cavities trap air and reduce the thermal coefficient of the brick.We simulate the heat transfer through a hollow brick. The holes in the brick are cavities. In order to determine the heat transfer coefficient we solve the heat flux through the hollow brick. For that a model of the brick is formulated as sequence of a solid and a fluid part. In the solid part only heat conduction is present and in the fluid heat transfer is present in form of conduction, convection and radiation. The flow field in the fluid part is modelled with the velocity‐vorticity formulation of the incompressible Navier‐Stokes. It is assumed that the Rayleigh number is below 106, hence, natural convective laminar fluid flow is present in the cavities. The cavities are also heated by heat radiation of the surrounding walls.The numerical realisation is done with the Boundary‐Domain Integral Method (BDIM). To handle suitable geometries the complexity of the BDIM is reduced from quadratic to logarithmic or almost linear by applying the ‐methodology. M is the number of unknown domain nodes. Numerical studies will be presented.

Aerofoil Design and Boundary Layer Transition Flow Studies at Subsonic Speeds for a Typical Transport Aircraft

Reliable computational aerodynamic prediction of airfoil drag, maximum lift and the effect of Reynolds number continue to be a challenge to aerodynamicists, using large computer programs and memory resources even today. Design of an aerofoil for an efficient wing is presented here for the typical transport aircraft which is being developed. Flow around five digit NACA23015 aerofoil was investigated at three distinct Reynolds numbers viz, 2 x 106, 10 x 106 and 20 x 106 typical to small and medium size transport aircrafts for various angle of incidences up to stall. This paper provides insights into the geometry shape design of an aerofoil taking into consideration the effect of Reynolds number on overall efficiency i.e. the lift to drag ratio of an aerofoil. Further the paper provides some aspects of prediction and comparison of the movement of transition point for free transition flow at specific Reynolds numbers. Extensive use is made here of the open source XFOIL. With its analysis and excellent interactive inverse design capabilities for low speed single element aerofoil, the design of the aerofoil has been studied with respect to lifting characteristics, drag characteristics and pitching moment. The results are compared with those of the classical NACA 23015 aerofoil and another NASA Natural Laminar Flow (NLF) aerofoil. The comparative numerical studies using the XFOIL has shown some promising results of the modified aerofoil for use on the aircraft under design.

Natural convection laminar flow between two vertical walls with a nonlinear variation of density and viscosity with temperature

AbstractUnder the combined influence of buoyancy force and constant pressure gradient, the free convection flow and heat transmission inside a channel placed vertically and consisting of two parallel walls are examined. Through the application of the Runge–Kutta fourth‐order approach and the shooting procedure, the modeled equations of the problem including nonlinear density–temperature and quadratic viscosity–temperature change at constant but different wall temperatures are numerically solved. For distinct values of the embedded parameters, the fluctuations in fluid's velocity and temperature are examined. For various situations of linear, quadratic, and nonlinear dependence of density on temperature, numerical values for heat transmission rate, volume flow rate and skin friction are determined and tabulated. The fluctuations of the Nusselt number with the Grashof number corresponding to thermal expansion coefficients are depicted through graphical interpretations. It was found that the combined effect of thermal expansion coefficients raises the values of the physical entities, Nusselt number, volume flow rate, and heat transmission rate. For the lower Reynolds number values, the Nusselt number value stays around one, indicating the same impact of conduction and convection on heat transmission.

Hybrid Laminar Flow Control Optimizations for Infinite Swept Wings

Maintaining the laminar flow on surfaces through active control is a significantly promising technique for reducing fuel burn and alleviating environmental concerns in commercial aviation. However, there is a lack of systematic parameter studies for the hybrid laminar flow control (HLFC) together with natural laminar flow (NLF). To address this need, we optimize the infinite swept wings with different sweep angles and at various conditions, including different Mach numbers, Reynolds numbers, and lift coefficients. The Reynolds-averaged Navier-Stokes (RANS) solver coupled with the linear stability theory is applied for the laminar-turbulent transition prediction, and the traditional optimization method based on evolutionary algorithms is applied for laminar flow wing optimization. The optimization results found that HLFC is required when the NLF fails at a larger sweep angle (35°) and Reynolds number ( 20 × 10 6 ). The lower pressure peak with boundary-layer suction is found to delay the transition of the regional aviation condition. Besides, the pressure distribution of HLFC is similar to NLF results at the lower Reynolds number ( 10 × 10 6 ) or sweep angle (25°), i.e., a gentle negative pressure gradient near the leading edge and a small favorable pressure gradient behind it. Clarifying the characteristics of laminar flow wings will advance the application of the laminar flow technique within its field.

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.

Non-overlapping High-accuracy Parallel Closure for Compact Schemes: Application in Multiphysics and Complex Geometry

Compact schemes are often preferred in performing scientific computing for their superior spectral resolution. Error-free parallelization of a compact scheme is a challenging task due to the requirement of additional closures at the inter-processor boundaries. Here, sources of the error due to sub-domain boundary closures for the compact schemes are analyzed with global spectral analysis. A high-accuracy parallel computing strategy devised in “ A high-accuracy preserving parallel algorithm for compact schemes for DNS. ACM Trans. Parallel Comput. 7, 4, 1-32 (2020)” systematically eliminates error due to parallelization and does not require overlapping points at the sub-domain boundaries. This closure is applicable for any compact scheme and is termed here as non-overlapping high-accuracy parallel (NOHAP) sub-domain boundary closure. In the present work, the advantages of the NOHAP closure are shown with the model convection equation and by solving the compressible Navier–Stokes equation for three-dimensional Rayleigh–Taylor instability simulations involving multiphysics dynamics and high Reynolds number flow past a natural laminar flow airfoil using a body-conforming curvilinear coordinate system. Linear scalability of the NOHAP closure is shown for the large-scale simulations using up to 19,200 processors.

Adjoint-based robust optimization design of laminar flow wing under flight condition uncertainties

It is an inherent uncertainty problem that the application of laminar flow technology to the wing of large passenger aircraft is affected by flight conditions. In order to seek a more robust natural laminar flow control effect, it is necessary to develop an effective optimization design method. Meanwhile, attention must be given to the impact of crossflow (CF) instability brought on by the sweep angle. This paper constructs a robust optimization design framework based on discrete adjoint methods and non-intrusive polynomial chaos. Transition prediction is implemented by coupled Reynolds-Averaged Navier-Stokes (RANS) and simplified eN method, which can consider both Tollmien-Schlichting (TS) wave and crossflow vortex instability. We have performed gradient enhancement processing on the general Polynomial Chaos Expansion (PCE), which is advantageous to reduce the computational cost of single uncertainty propagation. This processing takes advantage of the gradient information obtained by solving the coupled adjoint equations considering transition. The statistical moment gradient solution used for the robust optimization design also uses the derivatives of coupled adjoint equations. The framework is applied to the robust design of a 25° swept wing with infinite span in transonic flow. The uncertainty quantification and sensitivity analysis on the baseline wing shows that the uncertainty quantification method in this paper has high accuracy, and qualitatively reveals the factors that dominate in different flow field regions. By the robust optimization design, the mean and standard deviation of the drag coefficient can be reduced by 29% and 45%, respectively, and compared with the deterministic optimization design results, there is less possibility of forming shock waves under flight condition uncertainties. Robust optimization results illustrate the trade-off between the transition delay and the wave drag reduction.

Design Exploration of Transonic Airfoils for Natural and Hybrid Laminar Flow Control Applications

Designing transonic laminar swept wing at high Reynolds numbers is challenging due to the premature flow transition from the prominent crossflow instabilities (CFIs). Hence, natural laminar flow (NLF) wings are limited to lower sweep angles. In this study, hybrid laminar flow control (HLFC) is used to extend this limit to design low-drag transonic infinite wing. A multi-objective genetic algorithm with competing drag components as objectives is used to design NLF and HLFC airfoils. Optimal design is a tradeoff between the drag reduction from skin friction drag, pressure drag, and suction drag, and does not resemble initial airfoils. Airfoil shape and suction distribution are coupled and optimized simultaneously. Euler flow solver with integral boundary-layer method is coupled with a higher-fidelity Linear Stability Analysis using 2.5D approximation for transition prediction. At wing sweep of 22.5°, Mach number of 0.78, and Reynolds number of 30 million, optimum NLF airfoil has 27% lower drag than an optimum turbulent airfoil. The optimum HLFC airfoil showed a 25% lower total drag than the NLF airfoil. A higher optimum sweep angle was observed for low-drag HLFC airfoil (20.5°) in comparison to the NLF airfoil (16.8°) and both favored an undercut on the lower surface to dampen CFI.

Numerical Investigations of Laminar Separation Bubbles Under the Influence of Periodic Gusts

Direct numerical simulations were performed to investigate the behavior of laminar separation bubbles subject to vertical gusts in an airfoil flow. Oscillating boundary-layer flows containing gusts—including unsteady pressure gradients—were prescribed via boundary conditions as well as forcing terms on a domain located in the upper rear section of a natural laminar flow airfoil. For this hybrid approach, unsteady Reynolds-averaged Navier-Stokes (URANS) simulations of the entire flowfield were carried out in conjunction with the so-called disturbance velocity approach, providing transient boundary conditions for the direct numerical simulation. A steady-state reference case with a Reynolds number of is compared and validated with results from wind tunnel experiments. Results of simulations at four different gust frequencies at the same amplitude , each with and without additional disturbances to the boundary layer introducing oblique resonance, are presented and discussed in this paper. The time-dependent behavior of convective instability modes is evaluated by using the continuous wavelet transform and linear stability theory with an unsteady extension. Furthermore, the contribution of the absolute instability—which appears to be larger in the case of oscillating flow compared to the steady-state case—is discussed. Lock-in effects are identified at high gust frequencies.