Natural Laminar Flow Airfoil Research Articles

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.

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

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.

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.

Numerical Investigation on the Transition Flow around NLF Airfoil

A natural laminar flow (NLF) airfoil is designed to reduce drag by expanding laminar flow areas. In-depth knowledge of transition performance is essential for its aerodynamic design. The k-ω-γ-Reθ framework, which consists of the SST k-ω turbulence model and γ-Reθ transition model, is employed to simulate transitional flows around an NLF wing RAE5243 airfoil. The transition performances of the RAE5243 airfoil under various values of turbulent intensity, temperature, angle of attack, and Mach number are simulated and compared. The results show that the rise of inflow turbulent intensity will promote an earlier transition on both the suction and pressure sides. The influence of wall temperature on transition is limited. The rise of angle of attack will lead to an earlier transition on the pressure side but a later transition on the suction side. With the rise of Mach number, the transition happens earlier under a zero and positive angle of attack but later under a negative angle of attack. In addition, the correlation of transition onset locations with respect to turbulent intensity, surface temperature, angle of attack, and Mach number is established based on numerical results.

In-flight measurement of free-stream turbulence in the convective boundary layer

Natural laminar flow airfoils have achieved such a level of refinement that further optimisation and subsequent wind tunnel testing need to regard the specific free-stream turbulence to be expected during operation. This requires the characterisation of this turbulence in terms of those properties which are relevant for boundary layer receptivity and subsequent transition. These parameters of turbulence change with environmental conditions and, in case of aircraft, along the flight profile. This study investigates the free-stream turbulence relevant for the case of sailplane airfoils. In-flight measurements with a constant temperature anemometer x-wire probe were conducted during cross-country flights in Central Europe and provided 22 h of flight data, covering thermalling phases as well as straight flight legs. Longitudinal and transversal velocity fluctuations were recorded well into the dissipation range. The special challenges of operating a constant temperature anemometer probe continuously for several hours are addressed. The permanent unsteadiness of the inflow poses challenges for the evaluation, but also provides a broad database of measured turbulence levels. The quality of the measurements is shown by verifying some of the predictions of Kolmogorov's inertial range theories. Free-stream turbulence in thermalling phases is sufficiently homogeneous to be described accurately, as the dissipation range fluctuates only in a limited range and follows a log-normal distribution. On the straight flight legs, the turbulence depends on the convective activity along the flight path. In general, within the convective part of the atmosphere, turbulence levels are found to be significantly larger than in low-turbulence wind tunnels.Graphical abstract

Robust Optimization of Natural Laminar Flow Airfoil Based on Random Surface Contamination

Natural laminar-flow (NLF) airfoils are one of the most promising technologies for extending the range and endurance of aircrafts. However, there is a lack of methods for the optimization of airfoils based on the surface contamination that destroys the laminar flow. In order to solve this problem, a robust optimization process is proposed using the Non-dominated Sorting genetic algorithm- II (NSGA-II) evolutionary algorithm, and Monte Carlo simulation combined with an aerodynamic calculation software Xfoil. Firstly, the airfoil is optimized normally and the aerodynamic performance of optimized airfoil under surface contamination is analyzed. Then, the original airfoil is robustly optimized under random surface contamination based on the assumption that its locations follow triangular and uniform probability distributions. Finally, all the optimized results and original airfoil are compared. It is found that robust optimization reduces the sensitivity of the airfoil to random surface contamination, hence, improving the robustness of the airfoil. The proposed methods make it possible to improve the aerodynamic performance of NLF airfoils considering surface contamination.

Airfoil Morphed Leading-Edge Design for High-Lift Applications Using Genetic Algorithm

A genetic algorithm was developed for the optimal design of morphing airfoil leading-edge shapes to maximize airfoil , while retaining a constant leading-edge arc length. Using a candidate slotted, natural laminar flow airfoil, an inviscid–viscous flow simulation program suite was used to evaluate a randomized population of airfoils during the design process at low-speed, high-lift flight conditions of a representative single-aisle commercial transport aircraft. The use of a continuous morphing strategy to perform treatment of the leading-edge geometry in high-lift scenarios, as opposed to leading-edge flap or slat elements, is envisaged as a means to mitigate surface discontinuities and retain laminar flow qualities of future airfoil sections in cruise conditions. By using different definitions of the objective cost function, multiple morphed airfoil designs were found through the genetic algorithm to create a library of possible designs aimed to purposefully alleviate the leading-edge stall tendencies of the baseline airfoil. Three morphed designs were validated through experimentation in a low-speed, open-return wind tunnel. Resulting data were used to verify that the leading-edge separation property of the baseline airfoil during stall was effectively alleviated with use of the morphed leading-edge geometries.

Climate Impact Mitigation Potential of Novel Aircraft Features

This work presents a transpacific airliner designed for minimal climate impact, incorporating several novel design features. These include open rotor engines, sustainable aviation fuels, natural laminar flow airfoils, and riblets. The design’s configuration and mission have been optimised simultaneously using a combination of standard preliminary techniques, experimental data, a multi-point mission analysis, and a model of average temperature response. It is demonstrated that, on an 8000 km mission, the design offers an 89.8% reduction in average temperature response relative to an Airbus A330-200, at the expense of a 7.3% increase in direct operating cost. The sensitivity of these results is investigated by comparing the performance over a range of operating conditions. In addition, several alternative designs incorporating only some of the above-mentioned features are analysed, allowing for an assessment of their individual contribution. Finally, a life-cycle average temperature response analysis is presented to place the climate impact of operation, manufacturing and end-of-life procedures in context.

Controlling transonic shock–boundary layer interactions over a natural laminar flow airfoil by vortical and thermal excitation

Investigations have been performed via implicit large eddy simulations to study the overall effects of exciting a flow field by thermal (wall-heating and wall-cooling) and vortical (with high and low frequencies) actuation. The actuator is placed on the suction surface of a natural laminar flow (SHM-1) airfoil having an angle of attack of α=0.38° (cruise setting). Oncoming flow has a Mach number of 0.72, and a Reynolds number based on a chord of Re=16.2×106, for which a complex shock system is formed on the suction surface. Vorticity dynamics of the flow is studied using time series of vorticity at different locations above the suction surface and instantaneous contour plots of vorticity in the domain. An inspection of the flow using snapshots of ∇ρ and ∇(ρT) is done to characterize the numerical schlieren. The comparative effects of the various forms of excitation on the shock–boundary layer interactions (SBLI) have been analyzed using time series of the magnitudes of ∇ρ across the identified shock structures from numerical schlieren snapshots. Also, the role of the frequency of imposed vortical actuation has been studied using vorticity and Mach contours for a comparative understanding of the control of the SBLI.

Partial Pressure Field for Airfoil Wave Drag

Far-field drag prediction methods have proven to be effective at capturing drag sources in both computational fluid dynamics and wind-tunnel testing analyses. Near-field decomposition, on the other hand, has been neglected due to the failure to divide pressure drag directly into its contributions to both form and wave drag. The work presented in this paper is an extension of previous research conducted into partial pressure fields that allow such a near-field drag decomposition. In particular, the current paper proposes a new partial pressure field that captures wave drag sources in two-dimensional compressible viscous flow. Pressure decompositions are performed on the conventional NACA0012 and RAE2822 airfoils, and the results are compared to those obtained by classical far-field decomposition, with reference data from previous studies providing a means of verification. Finally, the slotted natural laminar-flow airfoil, the S207, is analyzed to test the applicability of the decomposition method on multielement airfoils that are considered for next-generation aircraft design. It is demonstrated that an adjusted compressible pressure field captures wave drag as a reliable near-field method that correlates closely with the traditional far-field methods, particularly for typical transonic cases.

Characterization of low levels of turbulence generated by grids in the settling chamber of a laminar wind tunnel

Wind tunnel investigations of how Natural Laminar Flow (NLF) airfoils respond to atmospheric turbulence require the generation of turbulence, whose relevant characteristics resemble those in the atmosphere. The lower, convective part of the atmospheric boundary layer is characterized by low to medium levels of turbulence. The current study focuses on the small scales of this turbulence. Detailed hot-wire measurements have been performed to characterize the properties of the turbulence generated by grids mounted in the settling chamber of the Laminar Wind Tunnel (LWT). In the test section, the very low base turbulence level of Tuu ≅ 0.02% (10 ≤ f ≤ 5000 Hz) is incrementally increased by the grids up to Tuu ≅ 0.5%. The turbulence spectrum in the u-direction shows the typical suppression of larger scales due to the contraction between grids and test section. Still, the generated turbulence provides a good mapping of the spectrum measured in flight for most of the frequency range 500 ≤ f ≤ 3000 Hz, where Tollmien-Schlichting (TS)-amplification occurs for typical NLF airfoils. The spectra in v and w-direction exhibit distinct inertial subranges with slopes being less steep compared to the − 5/3 slope of the Kolmogorov spectrum. The normalized spectra in u-direction collapse together well for all grids, whereas in v- and w-directions the inertial- and dissipative subranges are more clearly distinguished for the coarser grids. It is demonstrated that the dissipation rate ε is a suitable parameter for comparing the wind tunnel turbulence with the atmospheric turbulence in the frequency range of interest. By employing the grids, turbulence in the range 4.4 × 10–7 ≤ ε ≤ 0.40 m2/s3 at free-stream velocity U∞ = 40 m/s can be generated in the LWT, which covers representative dissipation rates of free flight NLF applications. In the x-direction, the spectra of the v and w-components develop progressively more pronounced inertial- and dissipative subranges, and the energy below f ≈ 400 Hz decreases. In contrast, the spectral energy of the u-component increases across the whole frequency range, when moving downstream. This behavior can be explained by the combination of energy transport along the Kolmogorov cascade and the incipient return to an isotropic state.Graphic

Comparative study of transonic shock–boundary layer interactions due to surface heating and cooling on an airfoil

Implicit large eddy simulation results are compared to investigate the effects of wall-heating and wall-cooling on shock–boundary layer interaction over an airfoil. Heat flux is provided on the suction surface of the airfoil from x/c=0.40 to x/c=0.50 for a Mach number of 0.72 and a Reynolds number based on chord of Re=16.2×106. Flow quantities are compared for the effects of heating and cooling. Numerical Schlieren snapshots reveal an oscillation of the shock wave and its interaction with upstream propagating Kutta waves generated from the trailing edge of the airfoil. Quantitative data obtained from these Schlieren snapshots and the mean aerodynamic load values indicate a reduction in frequency of oscillation of shock wave and a decrease in shock strength for the case of heating. Flow control by heating shows higher fluctuations in flow features evident from instantaneous quantities. Both imposed excitations lead to a marginal increase in aerodynamic efficiency (lift/drag). We also compare the integral aerodynamic parameters, such as lift and drag coefficients, and their ratio, Cl, Cd,and Cl/Cd. The simulations reported here follow the techniques used in Sengupta et al. [“Thermal control of transonic shock–boundary layer interaction over a natural laminar flow airfoil,” Phys. Fluids 33(12), 126110 (2021)].

Thermal control of transonic shock-boundary layer interaction over a natural laminar flow airfoil

Implicit large eddy simulations are performed to show control of shock and boundary layer interaction over a natural laminar flow airfoil by harmonic heat transfer on the suction surface at a free-stream Mach number M∞=0.72 and angle of attack of α=0.38°, for which experimental/flight test results exist. Surface heat flux is added in a time-periodic manner with the exciter strip located at different locations in the vicinity of the shock location at x/c≈0.49 for the flow without any control. Four cases of localized excitation with Gaussian distribution centered at x/c=0.45, 0.48, 0.50, and 0.55 on the airfoil surface are considered, with a width of 10% chord. The effects of unsteady heat flux on shock strength, location, and modification of its structure are demonstrated by instantaneous and time-averaged flow quantities. Detailed vortical and entropy contour plots and numerical Schlieren indicate flow features, such as creation and propagation of Kutta waves and their interactions with the shock wave. Time-averaged load distributions reveal a shift in the location and strength of the shock, with altered lift and drag time histories. The Fourier transforms of lift and drag coefficients help explain the alteration of the strength and structures of the shock-induced unsteadiness. The imposed excitation results in an improvement of aerodynamic efficiency (lift to drag ratio). The numerical simulations follow the algorithm developed in Sengupta et al. [Comput. Fluids 88, 19–37 (2013)].

Coupled Boundary-Layer Suction and Airfoil Optimization for Hybrid Laminar Flow Control

The present work incorporates the effects of boundary-layer suction in designing airfoils for hybrid laminar flow control application. The overall airfoil shape, the suction velocity distribution, and its location are optimized for minimum total drag while sustaining laminar flow over 80% of the airfoil top surface. Nondominated Sorting Genetic Algorithm II is used with the aerodynamic load computations from XFOILSUC and transition prediction from the improved Van Ingen method. The airfoil is designed for short-range subsonic conditions for three design lift-coefficient scenarios: 1) 0.4, 2) in the range of 0.3–0.7, and 3) in the range of 0.3–0.7 with suction onset at midchord. The optimized airfoil for scenario 3, with suction onset at midchord, results in a thicker airfoil with natural laminar flow (NLF) on the upper surface comparable to NLF airfoils. The total drag reduces by one-third due to the effects of suction. On the other hand, without limiting the suction onset (scenario 2), a thinner airfoil with longer NLF on the bottom surface and lower total drag is obtained. Optimum suction onset location is predicted to be at 30% of the chord. Suction is also observed to improve the drag bucket at off-design conditions.

A Surrogate-Based Method for Compressible Boundary-Layer Transition Prediction

To improve robustness and efficiency of automatic transition prediction in aerodynamic design, a reduced model of linear stability analysis is usually adopted, such as -envelope or -database method. Nevertheless, building such a model is challenging when it comes to compressible flows, as the transition mechanism is more complex and multiple flow parameters should be taken into consideration. To address this problem, this paper proposes an efficient surrogate-based method for compressible boundary layers that uses pretrained surrogate models to substitute linear stability analysis, concerning stability analysis of both Tollmien–Schlichting and Mack modes, as well as transition prediction of flow over arbitrary-shaped airfoils. The proposed method is demonstrated by stability analysis of compressible flat-plate boundary layers at a wide range of Mach numbers of . It is also validated by transition prediction of flow over a low-speed natural-laminar-flow (NLF) airfoil NLF-0416 and a transonic NLF airfoil NPU-LSC-72613. Besides, a sample partitioning method is presented to accelerate surrogate-model training with large samples. Results show that the predicted growth rates of perturbations, factors, and corresponding transition locations by our method of using surrogate-based stability analysis agree well with those by a standard method of solving full linear stability equations.

Unsteady Response of Natural Laminar Flow Airfoil Undergoing Small-Amplitude Pitch Oscillations

Large-eddy simulations are performed to investigate the dynamic response of a natural laminar flow airfoil undergoing harmonic small-amplitude pitch oscillations at a chord based Reynolds number of . Large changes in the transition location as well as trailing-edge separation are observed throughout the pitch cycles, which leads to a nonlinear response of the aerodynamic forces. Despite the highly nonlinear nature of the flow, the evolution of the boundary layer over the airfoil can be modeled by using a simple phase-lag concept, which suggests a quasi-steady evolution of the boundary layer. A simple empirical model is developed based on this phase-lag assumption, which fits very well with the measured experimental data and identifies the primary source of non-linearities in the unsteady aerodynamic forces.

Tollmien–Schlichting Wave Control on an Airfoil Using Dynamic Surface Modification

The present study demonstrates experimental mitigation of Tollmien–Schlichting waves on an unswept natural laminar flow airfoil at a chord-based Reynolds number of . This was facilitated by dynamic surface modifications using piezoelectrically driven oscillating surface actuators. The actuators were located at three streamwise locations on the airfoil’s suction side. The upstream actuator introduced a disturbance to the flow, which phase locked the Tollmien–Schlichting waves. The downstream actuators were used to mitigate the induced Tollmien–Schlichting waves by introducing an antiphase disturbance with the proper amplitude. The disturbances included either a single-frequency or multiple-frequency input signals. The Tollmien–Schlichting waves were mitigated in both open- and closed-loop control schemes. In open-loop control, the results demonstrated that dynamic surface modification can be used to mitigate the waves, even in the presence of a large adverse gradient for varying frequency bandwidths. In addition, the closed-loop scheme, using an iterative learning control algorithm, succeeded in reducing the induced disturbance amplitudes to 11%. In addition, the experiments demonstrated the ability to not only mitigate the wave but also amplify it if transition to turbulence is desired.

A robust method for reliable transition detection in temperature-sensitive paint data

This work presents a method to automatically detect boundary-layer transition on thermographic images of aerodynamic surfaces, acquired using the temperature-sensitive paint measurement technique. It is applied to surface temperature distributions mapped onto three-dimensional grids representing the surfaces of interest. Motivated and substantiated by earlier work, the transition location is determined at the location corresponding to the maximal value of the streamwise gradient of the surface temperature in the transitional region. The robustness of the transition detection method was proven for a challenging case with low signal-to-noise ratio and moderate streamwise temperature gradients. Moreover, the present method was validated against transition data obtained using an uncorrelated transition detection method based on the measurement of surface pressure fluctuations. The developed transition detection algorithm enabled the accurate, reliable, and consistent detection of the transition location under a variety of test conditions, including experiments on natural laminar flow airfoil and wing models at high Reynolds numbers and at both subsonic and transonic Mach numbers.