Boundary Layer Trip Research Articles

Characterization of a Glow Discharge in a Tripped Hypersonic Boundary Layer

A direct current glow discharge plasma was generated on a canonical 2.75° half-angle wedge test article. Experiments were conducted in the Texas A&M University Actively Controlled Expansion Tunnel at M=5.7 and Re=6×106/m. The effects of Reynolds number, polarity, and upstream perturbation were all independently studied via imaging, power measurements, and optical emission spectroscopy (OES). These measurements are the first to be conducted in a tripped hypersonic boundary layer and allow a deeper understanding of the interplay between the plasma and tripped flow. The plasma, formally characterized as a normal glow discharge, had negligible Joule and cathode heating effects on the boundary layer. Analogously, the installation of trips upstream of the electrodes changed only the appearance of the plasma, creating periodic streaks corresponding to the wakes and vortices produced by each trip but not altering the power across the electrodes. The OES indicates generation of nitric oxide (NO) in both the positive column and, mainly, at the cathode at the relatively low plasma power of ∼47W.

Effect of Tripping and Domain Width on Transonic Buffet on Periodic NASA-CRM Airfoils

Transonic buffet is an instability characterized by shock oscillations and separated boundary layers. High-fidelity simulations have typically been limited to narrow domains to be computationally feasible, overly constraining the flow and introducing modeling errors. Depending on the boundary-layer state upstream of the interaction, different buffet features are observed. High-fidelity simulations (implicit large-eddy simulation) were performed on the periodic (infinite) NASA-CRM wing at a moderate Reynolds number to assess the sensitivity of the two-dimensional transonic buffet to boundary-layer state and domain width. Simulations were cross-validated against low-fidelity Reynolds-averaged Navier–Stokes (RANS)/unsteady RANS and global stability analysis, and excellent agreement was found near the onset. By varying the boundary-layer tripping amplitude, laminar, transitional, and turbulent buffet interactions were obtained. All cases consisted of a single shock and low-frequency oscillations (St≈0.07). The transitional interaction also exhibited reduced shock movement, a 15% increase in CL¯, and energy content at higher frequencies (St≈1.3). Spanwise domain studies showed sensitivity at the shock location and near the trailing edge. We conclude that the span width must be greater than the trailing-edge boundary-layer thickness to obtain span-independent solutions. For largely separated cases, the sensitivity to span width increased, and variations across the span were observed. This was found to be associated with a loss of two-dimensionality in the flow.

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.

Characteristics of turbulent boundary layers generated by different tripping devices

In this study, we employ a high-fidelity flow visualisation technique to investigate the resemblance between a turbulent boundary layer generated in a lab environment with a limited development length by different tripping devices and a naturally developed canonical turbulent boundary layer. Furthermore, we aim to provide statistics of the wall-normal velocity fluctuation of the tripped boundary layers and analyse them compared with the canonical turbulent boundary layer. This is achieved by conducting two-dimensional particle image velocimetry measurements to acquire the velocity components in the stream-wise wall-normal plane downstream of the tripping location. Different criteria including integral parameters, diagnostic plot, stream-wise and wall-normal turbulence intensities, and mean velocity deficit profiles are employed to compare the boundary layer profiles generated by different trips with well-behaved turbulent boundary layers presented in the literature. The generation of Reynolds shear stress in tripped boundary layers is analysed using quadrant analysis, and suitably sized trips are shown to have contributions of sweeps and ejections in a similar manner to a well-behaved turbulent boundary layer. Furthermore, through the detection of uniform momentum zones, the presence of coherent structures in the logarithmic and outer layers of these well-behaved tripped boundary layers is shown. The findings of this study identify the blockage created by trips as the main factor affecting the turbulence statistics at a certain downstream distance. The results also suggest that the investigation of inner-scaled mean stream-wise velocity is recommended as a simple yet reliable and rapid assessment of the impact of a trip on the turbulent boundary layer development in short test sections.

Hypersonic FLEET velocimetry and uncertainty characterization in a tripped boundary layer

Femtosecond laser electronic excitation tagging (FLEET) velocimetry is applied in a hypersonic boundary layer behind an array of turbulence-inducing trips. One-dimensional mean velocity and root-mean-square (RMS) of velocity fluctuation profiles are extracted from FLEET emissions oriented across a 2.75∘ wedge and through a boundary layer above a flat plate in two test campaigns spanning 21 tunnel runs. The experiment was performed in the Texas A&M University Actively Controlled Expansion tunnel that operated near Mach 6.0 with a Reynolds number near 6 × 106 m−1 and a working fluid of air at a density near 2.5 × 10−2 kg m−3. Detailed analysis of random and systematic errors was performed using synthetic curves for error in the mean velocity due to emission decay and the error in the RMS velocity fluctuation due to random error. The boundary layer behind an array of turbulence-inducing trips is documented to show the breakdown of coherent structures. FLEET velocimetry is compared to the tunnel Data Acquisition System, Vibrationally Excited Nitric Oxide Monitoring results, and Reynolds-Averaged Navier–Stokes computational fluid dynamics to verify results.

Noise Reduction on a Low Reynolds Number Propeller

With the mass application of drones in daily work, more attention has been paid to the noise problem. The noise generation mechanism of low Reynolds number propellers is studied in this paper. It is found that there are laminar separation bubbles on the suction surface, causing an increase in broadband noise. Leading edge boundary layer trips are added to remove the bubbles. Finally, experimental measurements indicate that isolated propeller noise reduction is up to 5 dBA. It still has a 1.5 dBA reduction in the full-loaded hovering drone test, which brings many conveniences for daily work.

Large-eddy simulations of flow and aeroacoustics of twin square jets including turbulence tripping

In this study, we investigate the flow and aeroacoustics of twin square (i.e., aspect ratio of 1.0) jets by implicit large-eddy simulations (LESs) under a nozzle pressure ratio of 3.0 and a temperature ratio of 1.0 conditions. A second-order central scheme coupled with a modified Jameson's artificial dissipation is used to resolve acoustics as well as to capture discontinuous solutions, e.g., shock waves. The flow boundary layer inside of the nozzle is tripped, using a small step in the convergent section of the nozzle. The time-averaged axial velocity and turbulent kinetic energy of LES with boundary layer tripping approaches better to particle image velocimetry experimental data than the LES without turbulence tripping case. A two-point space–time cross-correlation analysis suggests that the twin jets are screeching and are coupled to each other in a symmetrical flapping mode. Intense pressure fluctuations and standing waves are observed between the jets. Spectral proper orthogonal decomposition (SPOD) confirms the determined mode and the relevant wave propagation. The upstream propagating mode associated with the shock-cell structures is confined inside jets. Far-field noise obtained by solving Ffowcs Williams and Hawkings equation is in good agreement with the measured acoustic data. The symmetrical flapping mode of twin jets yields different levels of the screech tone depending on observation planes. The tonalities—the fundamental tone, second and third harmonics—appear clearly in the far-field, showing different contributions at angles corresponding to the directivities revealed by SPOD.

The high-Reynolds-number stratified wake of a slender body and its comparison with a bluff-body wake

The high-Reynolds-number stratified wake of a slender body is studied using a high-resolution hybrid simulation. The wake generator is a 6 : 1 prolate spheroid with a tripped boundary layer, the diameter-based body Reynolds number is ${Re}= U_\infty D/\nu = 10^5$ , and the body Froude numbers are ${Fr}=U_\infty /ND=\{2,10,\infty \}$ . The wake defect velocity decays following three stages with different wake decay rates (Spedding, J. Fluid Mech., vol. 337, 1997, pp. 283–301) as for a bluff body. However, the transition points among stages do not follow the expected $Nt = Nx/U_\infty$ values. Comparison with the wake of a circular disk in similar conditions (Chongsiripinyo & Sarkar, J. Fluid Mech., vol. 885, 2020) quantifies the influence of the wake generator – bluff versus slender – in stratified flow. The strongly stratified ${Fr}=2$ wake is in a resonant state. The steady lee waves strongly modulate the mean flow, and relative to the disk, the 6 : 1 spheroid (a high-aspect-ratio shape) wake at ${Fr}=2$ shows an earlier transition from the non-equilibrium (NEQ) stage to the quasi-two-dimensional (Q2D) stage. The NEQ–Q2D transition is followed by a sharp increase in the turbulent kinetic energy and horizontal wake meanders. At ${Fr}=10$ , the start of the NEQ stage is delayed for the spheroid. Transfers between kinetic energy and potential energy reservoirs (both mean and turbulence) are analysed, and the flows are compared in phase space (with local Froude and Reynolds numbers as coordinates). Overall, the results of this study point to the difficulty of finding a universal framework for stratified wake evolution, independent of the features of the body, and provide insights into how buoyancy effects depend on the wake generator.

Boundary layer control for low Reynolds number fan rig testing

Ultra-high bypass ratio turbofans offer significant reductions in fuel and pollution due to their higher propulsive efficiency. Short intakes might lead to a stronger fan-intake interaction, which creates uncertainty in stability at off-design conditions. Due to the prohibitive cost of full-scale experimental testing, subscale testing in wind tunnels is used to understand this behaviour. The low Reynolds number of subscale models results in unrepresentative laminar shock-boundary layer interactions. The boundary layer state thus needs to be conditioned to better represent full-scale transonic fans. This paper proposes the use of an inexpensive and robust flow control method for the suction side of a fan blade. Design guidelines are given for the location and height of the discrete roughness elements used to control the boundary layer state. This paper also presents a rapid experimental validation methodology to ensure and de-risk the application of the boundary layer trip to 3D rig blades. The experimental methodology is applied to a generic aerofoil representative of a fan tip section. The experimental method proves that it is possible to reproduce boundary layers and pressure distributions of a full-scale fan blade on a 1/10 subscale model. The results obtained confirm that the boundary layer trip method successfully promotes transition at the location representative of full-scale blades, avoiding unrepresentative laminar shock wave boundary layer interactions. This highlights the importance of conditioning boundary layers in low Reynolds number fan rig testing.

Computational study of boundary layer effects on stochastic rotor blade vortex shedding noise

This work entailed the use of lattice-Boltzmann simulations to investigate the effects of different sUAS rotor blade surface treatments, volumetric grid resolutions, and surface roughness heights on broadband noise prediction. Aerodynamic results showed trailing edge flow separation and a mispredicted laminar separation bubble for the laminar-to-turbulent transitional wall-function case and the fully-turbulent wall-function case as well as a transition front for the laminar-to-turbulent transitional wall-function case. It was seen that the current version of PowerFLOW (6-2021) cannot predict turbulent structures associated with turbulent wall-bounded flow unless a triggering mechanism (e.g., boundary layer trip) is employed, possibly explaining aerodynamic force and broadband noise underprediction. Tonal noise results showed thickness (i.e., geometry dependent) noise dominance at out-of-plane observers located above the rotor plane for the fundamental blade passage frequency and at all observer locations for the second harmonic of the blade passage frequency. Extrapolation techniques were also successfully used over broadband noise results to separate regions of underpredicted broadband noise caused by inadequate spatial resolution from other underpredicted regions caused by the misprediction of flow physics using the current computational strategy.

Experimental investigation on the unsteady surface pressure fluctuation patterns over an airfoil

This article presents a comprehensive mapping of wall-pressure fluctuations over an airfoil under three different inflow conditions to shed light on some basic assumptions taken for granted for the recent aeroacoustic and aerodynamics experimental studies and in the noise prediction models. Unsteady and steady pressure measurements were performed over a heavily instrumented airfoil, which was exposed to smooth inflow, grid-generated turbulent inflow, and a smooth inflow with a tripping tape over the airfoil to explore the unsteady response of the airfoil for a broad range of angles of attack, 0°≤α≤20°. The results are presented in terms of non-dimensional pressure coefficient, root mean square non-dimensional pressure coefficient, frequency-energy content pattern map at isolated frequencies for the entire airfoil, and spectra of frequency-energy content at selected transducer locations. The results show that the unsteady airfoil response patterns for the tripped boundary layer and turbulence ingestion cases show a dramatic difference compared to the airfoil response patterns of the smooth inflow conditions. The response patterns differ across angles of attack, frequency, and between both sides of the airfoil. The results suggest a three-region pattern for the smooth inflow case, a two-region pattern for the tripped boundary layer case, and a two-region pattern for the turbulence ingestion case. Moreover, the results indicate that the presence of tripping may provide a flow with necessary statistical characteristics for the experimental rigs representing the full-scale application; however, it may misrepresent the frequency-dependent nature of the boundary layer.

Trailing-Edge Noise Comparability in Open, Closed, and Hybrid Wind Tunnel Test Sections

Open-jet, hard-wall, and hybrid test section configurations are typically used in wind tunnel tests, aiming to represent the ideal free-flight condition. Each test section type requires special adaptations of the experimental hardware and specific correction methodologies that account for systematic measurement errors. Differently from previous research, this paper investigates the comparability of measurements performed in three test section configurations in the same wind tunnel. With this approach, systematic errors related to the facility or boundary-layer tripping methodology are minimized. Basic 2D aerodynamic boundary corrections are evaluated using a DU97W300 airfoil. The corrected lift curve collapsed well while differences in stall behavior and distribution at larger angles of attack are nonnegligible. The trailing-edge noise produced by a NACA-0012, NACA-0018, and NACA-63018 served as reference aeroacoustic sources in this comparability analysis. Moreover, the noise reduction from trailing-edge serrations served as an additional reference for the comparability of relative noise levels. The chord-based Reynolds number of the experiments ranged from 260,000 to 660,000. Unsteady wall pressure measurements performed near the trailing-edge provided a reference for the aeroacoustic noise source terms, demonstrating a negligible change between the different test section types. The applied experimental hardware and acoustic corrections yielded aeroacoustic sources of comparable absolute far-field noise level in the three test section types within . The relative noise levels are comparable within . These results show that aeroacoustic measurements of trailing-edge noise performed in different test section configurations are equivalent, provided that adequate experimental and postprocessing methodologies are systematically implemented.

Experimental investigation of supersonic boundary-layer tripping with a spanwise pulsed spark discharge array

In this paper, a pulsed spark discharge plasma actuator array is deployed to control laminar–turbulent transition in a Mach 3.0 flat-plate boundary layer, and the subtle flow structures are visualized by nanoparticle planar laser scattering (NPLS) technique. Results show that the onset location of turbulence can be brought upstream by plasma actuation, corresponding to forced boundary-layer transition. Hairpin vortex packets evolved from the thermal bulbs play a vital role in the breakdown of laminar flow. With the help of a machine learning tool, all the relevant structures induced by plasma actuation are extracted from NPLS images, and a conceptual model of the hairpin vortex generation is proposed, including three stages: production and lift-up of the high-vorticity region, formation of the $\varLambda$ vortex and evolution of the hairpin vortex.

Aerodynamic performance investigation of sweptback wings with bio-inspired leading-edge tubercles

The aerodynamic behavior of sweptback wing configurations with bio-inspired humpback whale (HW) leading-edge (LE) tubercles has been investigated through computational and experimental techniques. Specifically, the aerodynamic performance of tubercled wings with symmetric (NACA 0015) and cambered (NACA 4415) airfoils is validated against the baseline model at various angles of attack ([Formula: see text]. The [Formula: see text]/[Formula: see text] ratio of the HW flipper is strategically reduced to 0.15 for ascertaining the flow control potential of the bio-inspired wings with sweptback configuration. It is a novel effort to quantify the effect of the leading-edge protuberances on stall delay, flow separation control and distribution of streamline vortices at unique [Formula: see text]/[Formula: see text] ratio outside the thickness range of HW flipper morphology. Four tapered sweptback wing models (Baseline A, Baseline B, HUMP 0015, HUMP 4415) are used with the amplitude-to-wavelength ([Formula: see text] ratio of 0.24 and Reynolds number about [Formula: see text]. The chordwise pressure distributions are recorded at the peak, mid and trough regions of the tubercled wings through a detailed wind tunnel testing and validated with numerical analysis. Additionally, the flow characteristics over the bio-inspired surfaces have been qualitatively analyzed through the laser flow visualization (LFV) technique to reveal the influence of laminar separation bubbles (LSBs). The essential aerodynamic characteristics such as boundary layer trip delay, vortex mixing, stall delay, and flow control at different AoA are addressed through consistent experimental data. As the sweptback configuration is a primary choice for airplane wings, the improved aerodynamic characteristics of the tubercled wings can be effectively utilized for the design of novel lifting surfaces, hydroplanes and wind turbines in the near future.

Noise from a Model-Scale Vertical-Axis Wind Turbine

Vertical-axis wind turbines are an attractive option for small-scale wind-power installations. In this application, noise is crucially important. To investigate the problem, measurements taken from a model-scale turbine in a wind tunnel are considered. The sound radiated by the model is clearly evident in the cross spectra from a pair of flush-mounted microphones. Parameter studies show that, in the normal operating range, speed is the dominant factor. The main source appears to be blade self-noise, although there are also indications of blade–wake interaction in the downstream half of the rotation path. Boundary-layer trips have a significant impact, showing that laminar instability is an important contributor to the self-noise. It will remain a risk at full scale, and its absence should be ensured at the design stage. To minimize the remaining sound radiation (at a given wind speed), configurations with lower tip-speed ratio at maximum power output should be preferred to ones having higher values of this parameter.

Measurements and analysis of hypersonic tripped boundary layer turbulence

Measurements and analysis of hypersonic tripped boundary layer turbulence

Tunnel Noise Effects on Hypersonic Cylinder-Induced Transitional Shock-Wave/Boundary-Layer Interactions

Experiments were conducted on a cylinder-induced shock-wave/boundary-layer interaction (SBLI) at Mach 6. The state of the interaction was brought through transition using a combination of boundary-layer trips and Reynolds number sweeps. The unit Reynolds number was varied over a range of 2 to 8 million per meter (M/m). The Actively Controlled Expansion wind-tunnel nozzle boundary layers transitioned near , which resulted in a threefold increase in the freestream disturbances. Trends related to SBLI properties (including the size of the separation region, peak surface heating, and surface pressure fluctuation spectra) correlated with the freestream disturbance signature. Separation distances on the centerline were bounded by laminar and turbulent conditions at and , respectively. The shape and extent of the separation region exhibited sensitivity to the transition process over a unit Reynolds number range of 3 to . Heat transfer from the reattachment shock was monitored for untripped and tripped configurations. The former demonstrated a near-linear relationship between heat transfer coefficient and unit Reynolds number, whereas the transitional SBLI configuration revealed a sudden rise in heat transfer coefficient that mirrored that of the freestream disturbances near . PCB® and Kulite® pressure transducers measured increased surface pressure fluctuations near for all model configurations, with the cylinder exciting a 40 kHz instability that led to SBLI transition for the tripped configuration.

Aerodynamic investigation of a linear cascade with tip gap using large-eddy simulation

The flow in a linear compressor cascade with tip gap is simulated using a wall-resolved compressible Large-Eddy Simulation. The cascade is based on the Virginia Tech Low Speed Cascade Wind Tunnel. The Reynolds number based on the chord is 3.88 x 10⁵ and the Mach number is 0.07. The gap considered in this study is 4.0 mm (2.9% of axial chord). An aerodynamic analysis of the tip-leakage flow allow us identifying the main mechanisms responsible for the development and the convection of the tip-leakage vortex downstream of the cascade. A region of high turbulence and vorticity levels is located along an ellipse that borders the top of the tip-leakage vortex. The influence of the airfoil suction side boundary layer development on the tip-leakage vortex is highlighted by tripping the flow. A tripped boundary layer induces a stronger and larger tip-leakage vortex that tends to move further away from the airfoil suction side and from the endwall compared with an untripped flow. The boundary layer turbulent state influences the tip-leakage flow development.

Comparative study on the aerodynamic performance of airfoil with boundary layer trip of various geometrical shapes

Performance of small scale wind turbine (SSWT) and miniature aerial vehicles (MAV) is always effected with Laminar separation bubble. The problem of a laminar separation bubble on the upper surface of an E216 airfoil operated at low Reynolds number (Re=100000) is investigated numerically. Incompressible steady two-dimensional simulation is carried out with Transition γ − Reθ turbulence model on the airfoil with a boundary layer trip (BLT). The performance of two different types of trips, namely, isosceles triangular (IT) and right-angled triangular (RA) is compared with that of the airfoil with a rectangular (RT) trip. The trip locations used are, 17% of the chord for location-1 and 10% of the chord for location-2 from the leading edge of the airfoil, while the trip heights selected are 0.3 mm, 0.5 mm, 0.7 mm, and 1 mm. Results showed that the boundary layer trip significantly affected the laminar separation and improved the aerodynamic performance of the airfoil. Maximum improvement in drag by 17.41% and corresponding lift to drag ratio by 10.86% are obtained for the isosceles trip at location-2 for the angle of attack of 6°. There is no observable advantage for isosceles and right-angled triangular trips over rectangular trips. Considering the geometrical complexity in fabrication, the rectangular trip is recommended.

Mitigation of the aerodynamic noise of small axial wind turbines - methods and experimental validation

Although the characteristic Reynolds number of small horizontal axis wind turbines is only in the order of 100,000, their noise emissions can be annoying. Three different modifications of the blades aiming at attenuating the aerodynamic noise mechanisms are investigated: (i) Optimization of the blade profile, (ii) application of boundary layer tripping and (iii) addition of serrations to the trailing edges of the blades. All modifications were designed using either a novel optimization method or state-of-the-art design rules. They were tested on a three-bladed research wind turbine with 3 m diameter, which was placed on a large flat roof of an isolated tall building. Shaft and sound power were measured during substantial time periods, statistically analyzed and compared. Depending on the type and thickness of the blade profile, both, tripping and trailing edge serrations, were found to reduce noise but always degraded the turbine shaft power. The model-based optimization of the 15 % thick baseline profile S834 resulted in a new 10 % thick profile with a novel shape of the pressure and suction sides. Blades made from this profile, without trips but with trailing edge serrations, provided the best compromise between noise reduction and degradation of the aerodynamic turbine performance.