Natural Boundary Layer Transition Research Articles

Aerodynamic and Aeroacoustic Effects of Different Transition Mechanisms on an Airfoil

The paper delves into the detailed sound generation and propagation mechanisms associated with tripping-induced flow perturbations and trailing-edge scattering using wall-resolved large-eddy simulations. Two distinct boundary-layer tripping techniques, namely a geometrically resolved stair strip and an artificially modeled trip using suction and blowing, are investigated. To facilitate comparison, the natural boundary-layer transition is also simulated as a baseline scenario. The analysis takes into account a Reynolds number of 4×105, a Mach number of 0.058, and a nonzero angle of attack of 6.25° over a NACA 0012 airfoil configuration. The mechanisms for sound generation and propagation related to trailing-edge noise remain consistent across the three transition scenarios. However, boundary-layer tripping notably leads to intricate, scenario-specific noise generation: there is an interaction between the laminar separation bubble and tripping for the stair strip case, whereas laminar boundary-layer instability is evident for the suction and blowing scenario. Aerodynamic flowfields involving acoustic noise sources, their propagating natures near the wall, and far-field acoustics are cross-examined in detail. The comprehensive analysis of observed phenomena provides valuable insights for understanding nonlinear flow and acoustic interactions relevant to airfoil noise and designing new types of trips under adverse pressure gradient flows.

Aeroacoustic noise reduction by application of end plates on wall-mounted finite airfoils

One known method to reduce vortex shedding from the tip of a blade is the use of end plates or winglets. Although the aerodynamic impact of such end plates has been investigated in the past, no studies exist on the effect of such end plates on the far-field noise. The aeroacoustic noise reduction of three different end-plate geometries is experimentally investigated. The end plates are applied to the free end of a wall-mounted symmetric NACA 0012 airfoil and a cambered NACA 4412 airfoil with an aspect ratio of 2 and natural boundary layer transition. Microphone array measurements are taken in the aeroacoustic open-jet wind tunnel at BTU Cottbus-Senftenberg for chord-based Reynolds numbers between 75,000 and 225,000 and angles of attack from 0^circ to 30^circ. The obtained acoustic spectra show a broad frequency hump for the airfoil base configurations at higher angles of attack that is attributed to tip noise. Hot-wire measurements taken for one configuration show that the application of an end plate diffuses the vorticity at the tip. The aeroacoustic noise contribution of the tip can be reduced when the endplates are applied. This reduction is most effective for higher angles of attack, when the tip vortex is the dominant sound source.Graphic abstract

Influence of small wavy roughness on flatplate boundary layer natural transition

One of the effective ways to reduce viscous drag around an airfoil is by delaying the boundary layer transition. In this study, we analyzed the influence of a small wavy roughness on a two-dimensional, natural boundary layer transition, using direct numerical simulation that resolved each small roughness. A parametric study was conducted on the wavy roughness wavelength. Our results show that in some cases the transition delays whose characteristics depend on the roughness wavelength. In a detailed analysis, we found that the wavy roughness firstly affects the process of primary vortex growth, Tollmien–Schlichting (TS) instability. In addition, we found that the secondary vortex pairing also depended on it. In the most transition-delayed cases, the roughness wavelength was different far from the TS instability one, and the vortex pairing occurred firstly in upstream however not much in downstream, keeping the vortex size is kept small.

Experimental investigation of the effect of transpiration cooling on second mode instabilities in a hypersonic boundary layer

The influence of localized nitrogen transpiration on second mode instabilities in a hypersonic boundary layer is experimentally investigated. The study is conducted using a 7^circ half-angle cone with a length of 1100 mm and small nose bluntness at 0^circ angle-of-attack. Transpiration is realized through a porous Carbon/Carbon patch of 44 times 82 mm located near the expected boundary layer transition onset location. Transpiration mass flow rates in the range of 0.05–1% of the equivalent boundary layer edge mass flow rate were used. Experiments were conducted in the High Enthalpy Shock Tunnel Göttingen (HEG) at total enthalpies around 3 MJ/kg and unit Reynolds numbers in the range of 1.4 cdot 10^6 , to 6.4 cdot 10^6 , {text {m}}^{-1}. Measurements were conducted by means of coaxial thermocouples, Atomic Layer Thermopiles (ALTP), pressure transducers and high-speed schlieren. The present study shows that the most amplified second mode frequencies were shifted to lower values as nitrogen is transpired into the boundary layer. In some cases the instability amplitudes were found to be significantly reduced. The observed frequency reduction was verified to correlate with the change of the relative sonic line height in the boundary layer. The amplitude damping was observed to occur only until the most amplified frequencies were reduced to around 50% of their undisturbed values. When transpiration within this limit was performed shortly upstream of the natural boundary layer transition onset, a transition delay of approximately 17% could be observed.Graphic abstract

Boundary layer transition over a foil using direct numerical simulation and large eddy simulation

Transitional boundary layers over lifting bodies represent an important class of flows in many industrial applications, and accurately capturing the transition is crucial for the prediction of important phenomena such as lift, drag, and trailing-edge noise. In this study, we consider how large eddy simulations (LESs) can be used to capture the natural boundary layer transition and compare the results to fully resolved direct numerical simulations that provide a detailed picture of the transition and trailing edge flow. The ability of LES to capture the transition is considered by looking at different elements of the subfilter scale modeling and discretization. The behavior of the subfilter scale model is shown to be critical, and it must remain inactive during the early stages of transition to avoid erroneous predictions due to excessive dissipation. Dispersion errors, when present, can cause the natural transition mechanism to be bypassed at an earlier stage, which leads to higher levels of turbulent kinetic energy at the trailing edge.

Surface curvature effects on the tonal noise of a wall-mounted finite airfoil.

This paper is concerned with the influence of camber on the noise of a wall-mounted finite airfoil with natural boundary layer transition. Tonal noise measurements taken in an aeroacoustic wind tunnel are presented for airfoils with aspect ratio of 2, NACAxx12 profile and camber between 0 and 6% at 40% chord. The results show camber is an important parameter that determines the operating conditions for which acoustic tone generation occurs and the number and intensity of the tones produced. Airfoils with 0%-2% camber have an acoustic signature that is dominated by a high amplitude primary tone, whereas the spectra of airfoils with higher camber of 4%-6% feature a more pronounced side tone structure. Tonal noise production does not collapse with lift coefficient, demonstrating that the local flow conditions influence the noise source. Tonal noise production is explained in terms of changes to mean flow topology, namely the location of flow separation, which is linked to tonal noise generation. Scaling of airfoil tonal noise is found to vary with angle of attack and pressure gradient. Empirical scaling laws for the primary tone frequency dependence on velocity are also derived for the cambered airfoils.

Tonal noise production from a wall-mounted finite airfoil

This study is concerned with the flow-induced noise of a smooth wall-mounted finite airfoil with flat ended tip and natural boundary layer transition. Far-field noise measurements have been taken at a single observer location and with a microphone array in the Virginia Tech Stability Wind Tunnel for a wall-mounted finite airfoil with aspect ratios of L/C=1–3, at a range of Reynolds numbers (ReC=7.9×105–1.6×106, based on chord) and geometric angles of attack (α=0–6°). At these Reynolds numbers, the wall-mounted finite airfoil produces a broadband noise contribution with a number of discrete equispaced tones at non-zero angles of attack. Spectral data are also presented for the noise produced due to three-dimensional vortex flow near the airfoil tip and wall junction to show the contributions of these flow features to airfoil noise generation. Tonal noise production is linked to the presence of a transitional flow state to the trailing edge and an accompanying region of mildly separated flow on the pressure surface. The separated flow region and tonal noise source location shift along the airfoil trailing edge towards the free-end region with increasing geometric angle of attack due to the influence of the tip flow field over the airfoil span. Tonal envelopes defining the operating conditions for tonal noise production from a wall-mounted finite airfoil are derived and show that the domain of tonal noise production differs significantly from that of a two-dimensional airfoil. Tonal noise production shifts to lower Reynolds numbers and higher geometric angles of attack as airfoil aspect ratio is reduced.

Large Eddy Simulation of a Controlled Diffusion Compressor Cascade

In this research a Controlled Diffusion (CD) compressor cascade stator blade is simulated at a Reynolds number of ∼700,000, based on inflow velocity and chord length, using Large Eddy Simulation (LES). A wide range of flow inlet angles are computed, including conditions near the design angle, and at high negative and positive incidence. At all inlet angles the surface pressure distributions are well-predicted by the LES. Near the design angle the computed suction side boundary layer thickness agrees well with experimental data, whilst the pressure side boundary layer is poorly predicted due to the inability of LES to capture natural boundary layer transition on the present grid. A good estimation of the loss is computed near the design angle, whilst at both high positive and negative incidences the loss is less well predicted owing to discrepancies between the computed and experimental boundary layer thickness. At incidences above the design angle a laminar separation bubble forms near the leading edge of the suction surface, which undergoes a transition to turbulence. Similar behaviour is noted on the pressure surface at negative incidence. At high negative incidence contra-rotating vortex pairs are found to form around the leading edge in response to an unsteady stagnation line across the span of the blade. Such structures are not apparent in time-averaged statistical data due to their highly-transient nature.

Influence of the External Disturbances on Natural Boundary-Layer Transition in Rectangular Wing Flows

Influence of the External Disturbances on Natural Boundary-Layer Transition in Rectangular Wing Flows

Influence of the External Disturbances on Natural Boundary-Layer Transition in 2-D Wing Flows

Influence of the External Disturbances on Natural Boundary-Layer Transition in 2-D Wing Flows

Development of a Large-Scale Wind Tunnel for the Simulation of Turbomachinery Airfoil Boundary Layers

The procedures employed for the design of a closed-circuit, boundary layer wind tunnel are described. The tunnel was designed for the generation of large-scale, two-dimensional boundary layers on a heated flat surface with Reynolds numbers, pressure gradients, and free-stream turbulence levels typical of turbomachinery airfoils. The results of a series of detailed tests to evaluate the tunnel performance are also described. Testing was conducted for zero pressure gradient flow with natural boundary layer transition. Heat transfer data and boundary layer profiles are presented for a flow with 0.25 percent free-stream turbulence. The flow in the tunnel test-section was shown to be highly uniform and two-dimensional. Test boundary layer profile and convective heat transfer data were self-consistent and in excellent agreement with classic correlations. Test-section free-stream total pressure, multi-component turbulence intensity, longitudinal integral scale, and spectral distributions are presented for grid-generated turbulence levels ranging from 1 to 7 percent. The test-section free-stream turbulence was shown to be both homogeneous and nearly isotropic. Anticipated applications of the facility include studies of the heat transfer and aerodynamics for conditions typical of those existing on gas turbine airfoils.