Boundary Layer Transition Process Research Articles

An experimental investigation of supersonic conical cooling films with angles of attack

While the flow mechanisms of two-dimensional supersonic cooling films have been studied in-depth, this paper used the nanoparticle planar laser scattering and particle image velocimetry techniques to investigate the flow of supersonic conical cooling films at different angles of attack (AOAs). The mainstream Mach number was Ma∞=3.8, and the supersonic conical cooling film was tangentially injected via a precisely calibrated Maj=2.8 annular nozzle. Initially, the streamwise boundary layer transition process without cooling film injection was analyzed. The boundary layer transition on the leeward side occurred prematurely, whereas on the windward side, the transition process was notably delayed. Subsequently, the supersonic conical cooling film flow was qualitatively and quantitatively evaluated from the perspectives of turbulent structures, and the time-averaged and statistical characteristics of the velocity field. On the windward side, as the ratio of static pressure decreased, the effective cooling length also decreased with an increase in AOA. On the leeward side, at a small positive AOA, the supersonic conical cooling film mixed with the low-energy fluid within the thickened inner layer of the mainstream boundary layer, which mitigated the growth rate of the mixing layer and ultimately enhanced the effective cooling length. With a further increase in AOA, the supersonic conical cooling film experienced the three-dimensional detrimental effects of crossflow-separation vortices and downwash mainstream on the leeward surface, resulting in a decrease in the effective cooling length.

Large eddy simulations of periodic wake effects on boundary-layer transition of low-pressure turbine cascades

The periodic wake effect is one of the most important sources of unsteady disturbance in turbines. Its influence on the boundary layer transition process of the downstream blade suction surface is an important factor determining the turbine loss and aerodynamic performance, and also an effective potential approach of turbine loss control. In this paper, the high-load low-pressure turbine (LPT) cascade is taken as the research object, and the large eddy simulation based on the inhouse coed Multiblock Parallel Large-eddy Simulation is used to study the periodic influence of upstream wake. The unsteady transition process of the boundary layer on the suction surface of the turbine cascade and the spatial–temporal evolution of the vortex are discussed in detail. It is shown that there are three modes of boundary layer transition on the suction surface of the LPT cascade under the effect of wake, occurring alternately during the wake passing period. Each mode of transition has different characteristics in vortex structures, as well as boundary-layer separation and reattachment, thereby makes different losses. Although the transition mechanism and evolution process of the three modes are different, the calming regions exist in all three modes, which is important for the control of the boundary layer. This study gives an important reference for reducing the flow loss in high-load turbines by means of periodic wakes.

Impact of Wavy Wall Surface on Hypersonic Boundary-Layer Instability of Sharp Cone Model

A comprehensive investigation of the impact of wavy wall on hypersonic boundary layers is carried out in a Mach 6 Ludwieg tube based on a 7 deg half-angle sharp cone. Linear stability theory is applied on the cone base flow with both smooth and wavy surfaces first. The evaluation of amplification rate as well as the frequency of instability waves reveals that the second mode instability is suppressed downstream of the wavy wall region. The development of instabilities is then characterized using surface flush-mounted pressure sensors and focused laser differential interferometer (FLDI). The FLDI measurement is conducted within the wavy wall zone both at troughs and at crests. Good agreement is found between the surface pressure measurement and computation, except that instability waves are invisible shortly downstream of the wavy wall. Additionally, bispectral analysis indicates that the nonlinear interaction of instability waves downstream of the wavy wall region is suppressed. FLDI measurement along the streamwise direction shows that the second mode instability waves disappear in the wavy wall region, and the boundary-layer transition process is delayed. Spatial measurement with FLDI of boundary layer identifies various types of disturbance at the troughs and crests; moreover, the co-existence of low-frequency and high-frequency disturbances is observed off the wall surface downstream of the wavy wall.

A three-equation transition model with mechanical considerations

A three-equation transition model based on the transition V-model is proposed for subsonic flows in this study. Considering the mechanical approximation of the generation process of the pre-transitional vorticities, the value of laminar Reynolds shear stress related to the mean shear deformation was calculated in the original transition V-model. Then a new transition model, named V-SA model, was proposed, which considered the phenomenological process of transition and presented great results for flows with and without pressure gradient. It is well-known that the baseline Shear Stress Transport (SST) turbulence model shows excellent performance of accuracy and robustness in plentiful flow cases, but it is important to predict boundary layer transition. The current model (V-SST) successfully couples the V-model to the SST turbulence model by introducing the effective turbulent viscosity and additional correction terms into the transport equations. A thorough evaluation of its ability to predict transition features is performed versus the well-documented flat plate of ERCOFTAC, including T3A and T3B without pressure gradient, T3L2 and T3L3 with semi-circular leading edge, the three-dimensional 6:1 prolate-spheroid under two angles of attack, and the NLR-7301 airfoil under different Mach numbers. Numerical results show that the current model has an attractive and superior performance in the simulation of boundary layer transition processes

Hypersonic boundary layer transitions over a yawed, blunt cone

The boundary layer transition process for a straight circular cone with angle of attack (AoA) of 6° at Mach 6 was studied via high-resolution direct numerical simulations and stability analyses. A pair of strong streamwise vortices simultaneously induced a low-speed mushroom structure to emerge near the leeward plane. This structure was subject to low-frequency sinuous and high-frequency varicose instabilities. The transition between windward and leeward rays likely occurred due to the interactions between low-frequency traveling and stationary crossflow vortices. The lower-amplitude high-frequency secondary crossflow instability mode could be traced back to the upstream Mack mode. Although the axial and crossflow vortices were fundamentally different, there was a striking similarity between them since both high-frequency perturbations were driven by azimuthal shear and resulted in staggered positive and negative axial vortices along the streamwise direction.

Numerical study of the suction flow control of the supersonic boundary layer transition in a framework of gas-kinetic scheme

In this paper, we numerically study the supersonic boundary layer transition process and the suction flow control method by employing the DNS technology. By adopting the high resolution GKS method, the mechanism of the boundary transition and the statistical characteristics of the compressible turbulent boundary layer are investigated. Results indicate that the streamwise vortex and the three-dimensional shear layer with high instability are the main causes of the boundary layer transition. Suction control cases show that the laminar boundary layer suction can weaken the development of the unstable disturbance waves in the boundary layer, and delay the streamwise position of the local disturbance. Also, the space distributions of the vortex structure are weakened to delay the process of the transition.

Investigation on the boundary layer transition with the effects of periodic passing wakes

The boundary layer transition caused by wake is a problem related to basic fluid mechanics and engineering applications. In this paper, the interaction between the periodic passing wakes induced by moving cylinders and the boundary layer of the plate is investigated by large eddy simulation, and experiments in a low-speed water tunnel are designed to verify the simulations. The flow field velocity is measured by high-resolution pressure sensors and two-dimensional particle image velocimetry. The effects of wake passing frequencies fN = 0.63, 1, and 1.26 on the time average and statistical average characteristics of the boundary layer transition and the instantaneous flow structure are studied. The influence of large-scale wakes on the integral parameters of boundary layer thickness, skin friction coefficients, time-averaged velocity distributions, and velocity fluctuations is addressed. The results show that the onset of transition moves to the leading edge of the plate as the wake passing frequency increases, while the location of transition completion moves to the end of the plate. The specific boundary layer transition process is analyzed through the spanwise pocket and streamwise streaky structure propagation. The vortex structures in the boundary layer are extracted by the Q criterion, and the results show that the spanwise secondary vortex on the pressure side induced by the large-scale wake gradually loses its stability and results in transition. Moreover, the hairpin vortex in the suction surface continually lifts up the wall-normal location of the breakdown event. Thus, it turns the turbulence spot arrowhead pointing downstream into pointing upstream.

Computational aeroacoustic prediction of trailing edge noise for small wind turbines

The study of aeroacoustic noise generated by small wind turbines is important to increase acceptance and implementation of the technology. Small wind turbines have unique challenges due to the low Reynolds number (Re) flow the blades experience, which introduces a potential for tonal noise. Computational aeroacoustics can be applied during the design stage of the turbine blades to improve acoustic performance. This work aims to validate a fully analytical aeroacoustic model by analyzing a SD 7037 blade segment at static angles of attack, with comparison to experimental flow and acoustic data. The Ffowcs-Williams and Hawkings (FW-H) acoustic model is used in combination with Large Eddy Simulation (LES). The following simulation parameters were examined: mesh quality, mesh density, inlet turbulence and spanwise boundary condition. These parameters change the boundary layer (BL) transition process and the formation of the laminar separation bubble on the suction side of the blade segment, both of which impact the aeroacoustic noise prediction. It was found that improvement in mesh quality and density on the surface of the blade segment resulted in improved BL simulation and tonal noise prediction. Alteration of the inlet turbulence and spanwise boundary condition did not have as large of an effect.

An improved local correlation-based intermittency transition model appropriate for high-speed flow heat transfer

An improved local correlation-based intermittency transition model is developed for high-speed flow heat transfer in this paper. Based on the intermittency transition model, two improvements have been proposed. One is introducing a new correlation of the momentum thickness Reynolds number for high-speed boundary layer into the transition onset function. In this new correlation, the free stream Mach number and the free stream Reynolds number are used to reflect the compressible effects. The other improvement is the modification of the intermittency factor transport equation to include physical information regarding the transition processes. Several test cases containing supersonic and hypersonic flat plates, a slender cone at different Reynolds numbers, the X-51A forebody, and the Hypersonic Inflatable Aerodynamic Decelerator are conducted to assess the applicability of the improved intermittency transition model. The numerical results show the ability of the improved model to describe the high-speed boundary layer transition processes. However, further research is needed to reproduce the overshoot phenomena and verify the capacity of the improved model on complex cases.

Evaluation of Turbulent Spot Production Rate in Boundary Layers Under Variable Pressure Gradients for Gas Turbine Applications

Abstract The paper presents several results from an experimental data base on transitional boundary layers developing on a flat plate installed within a variable area opening endwall channel. Measurements have been carried out by means of time-resolved particle image velocimetry (PIV). The overall test matrix spans three Reynolds numbers, four freestream turbulence intensity levels, and four different flow pressure gradients. For each condition, 16,000 instantaneous flow fields have been acquired in order to obtain high statistical accuracy. The flow parameters have been varied in order to provide a gradual shift of the mode of transition from a by-pass process to separated flow transition. In order to quantify the influence of the flow parameter variation on the boundary layer transition process, the transition onset and end positions, and the turbulent spot production rate have been evaluated with a wavelet-based intermittency detection technique for every condition exhibiting a complete transition process. The by-pass transition mode has the longest transition length that decreases with increasing the Reynolds number. The transition length of the separated flow case is smaller than the by-pass one, and the variation of the flow parameters has a similar impact. The variation of the inlet turbulence intensity has a small influence on this parameter except for the condition at the highest turbulence intensity that always shows the lowest turbulent spot production rate because a by-pass type transition occurs. This large amount of data has been used to develop new correlations used to predict the spot production rate and the transition length in attached and separated flows.

Numerical study on supersonic boundary-layer transition and wall skin friction reduction induced by fuel wall-jet combustion

In this study, the effect of wall-jet combustion on boundary layer transition and skin friction reduction was numerically investigated. To effectively capture the characteristics of boundary layer flow, the transition k−kl−w model was employed as the turbulence model and laminar finite-rate model was chosen as the combustion model. This numerical method was firstly validated by two sets of experimental results in open domain. After that, the research on wall-jet combustion was conducted and the numerical results showed that when injecting hydrogen in different directions, both the hydrogen self-ignition location and the skin friction on the wall were not altered significantly. When the injection angle to the airflow direction was increased to 30°, the intensity of combustion was insufficient and the skin-friction coefficient would be increased. Meanwhile, boundary layer transition occurred in a relatively smaller Reynolds number at this condition. The variation of the wall-jet height would have a greater impact on both boundary layer transition and the skin friction. More components can diffuse to the lower wall as the height of the wall-jet was enlarged, which can make the process of boundary layer transition be postponed and the skin friction be reduced as well. Furthermore, greater skin-friction reduction would be achieved downstream the self-ignition location when the wall was adiabatic, while the original Reynolds numbers for boundary layer transition is the smallest if the wall temperature is set to 600 K. Finally, in order to simulate the effect of the back pressure on the combustion flow field, another injector is added near the exit at the upper wall to produce air-throttling flow jet. The results showed that altering back pressure nearly has little influence on boundary layer transition and the skin friction decreases with more air-throttling flow rate.

Description of the Flow in a Two-Stage Low-Pressure Turbine With Hub Cavities

Abstract In gas turbine, multi-stage row blading and technological effects can exhibit significant differences for the flow compared with isolated smooth blade rows. Upstream stages promote a non-uniform flow field at the inlet of the downstream rows that may have large effects on mixing or boundary layer transition processes. The rows of current turbines (and compressors) are already very closely spaced. Axial gaps between adjacent rows of approximately 1/4 to 1/2 of the axial blade chord are common practice. Future designs with higher loading and lower aspect ratios, i.e., fewer and bigger blades, and the ever present aim at minimizing engine length or compactness, will aggravate this condition even further. Interaction between cascade rows will therefore keep increasing and need to be taken into account in loss generation estimation. Also the cavities at hub platform induce purge flow blowing into main annulus and additional losses for the turbine. A robust method to account for the loss generated due to these different phenomena needs to be used. The notion of exergy (energy in the purpose to generate work) provides a general framework to deal with the different transfers of energy between the flow and the gas turbine. This study investigates the flow in a two-stage configuration representative of a low-pressure turbine including hub cavities based on large eddy simulation (LES). A description of the flow in the cavities, the main annulus, and at rim seal interface is proposed. The assessment of loss generated in the configuration is proposed based on an exergy analysis. The study of losses restricted to boundary layer contributions and secondary flows show the interaction processes of secondary vortices and wake generated in upstream rows on the flow in downstream rows.

Local-Variable-Based Model for Hypersonic Boundary Layer Transition

Based on the numerical study of the existing model, certain inconsistencies of the model were first allocated. It was found that the second-mode instability dominating the transition process in hypersonic boundary layers was not simulated properly by the model, and a single transport equation for total fluctuating kinetic energy would make the model fail to be self-consistent. To eliminate these discrepancies, a laminar kinetic energy transport equation was developed in terms of local variables. Then, a time scale correction function capable of reproducing the hypersonic transition process dominated by the second-mode instability was constructed. In addition, the transport equation for the intermittency factor, in which the self-consistent production and dissipation terms were constructed, was also revised. On this basis, the transport equations for the laminar kinetic energy and intermittency factor were coupled with the shear stress transport model through the concept of effective turbulent eddy viscosity to form a local-variable-based model for hypersonic boundary layer transition. Finally, hypersonic transition flows over a flat plate and a straight cone are employed to test and verify the model. Numerical results illustrate that both the transition onset and the length of the transition region predicted by the model agree well with the experimental data.

Multi-scale entropy analysis and conditional sampling of the velocity increment in a transitional boundary layer

Boundary layer transition induced by surface imperfections has a great impact on the performance of the long endurance vehicles. A better understanding of the physical mechanism of the transition process would be much helpful in developing new models for transition prediction. The velocity increment is important in the theory of turbulence, but it is rarely discussed in a transitional boundary layer. In this paper, the multi-scale entropy and the conditional sampling are used to study the velocity increment in the boundary layer transition process. Similarities in the distribution of velocity increment in each part (laminar/turbulent) of the intermittent flow are observed. It is also found that there is a process for both the laminar and turbulent part of the fluid motions to evolve into the mature state before the fully developed turbulence is reached.

Large eddy simulation of boundary layer transition flow around NACA0009 blunt trailing edge hydrofoil at high Reynolds number

The process of boundary layer transition is needed to be taken into account for better understanding of the flow around a hydrofoil. Based on the WALE model, this paper presents a large eddy simulation of the flow around NACA0009 blunt trailing edge hydrofoil to investigate the characteristics of the boundary layer transition at high Reynolds number. The accuracy of the WALE model in predicting the transition process is validated by comparing the calculated boundary layer velocity profiles, thicknesses and shape factors with experimental data. Moreover, the spatial development laws of velocity fluctuation in the transitional boundary layer are revealed by analyzing the profiles of resolved normal and shear Reynolds stress in both streamwise and transverse directions

Free-stream turbulence effects on the boundary layer of a high-lift low-pressure-turbine blade

The suction side boundary layer evolution of a high-lift low-pressure turbine cascade has been experimentally investigated at low and high free-stream turbulence intensity conditions. Measurements have been carried out in order to analyze the boundary layer transition and separation processes at a low Reynolds number, under both steady and unsteady inflows. Static pressure distributions along the blade surfaces as well as total pressure distributions in a downstream tangential plane have been measured to evaluate the overall aerodynamic efficiency of the blade for the different conditions. Particle Image Velocimetry has been adopted to analyze the time-mean and time-varying velocity fields. The flow field has been surveyed in two orthogonal planes (a blade-to-blade plane and a wall-parallel one). These measurements allow the identification of the Kelvin-Helmholtz large scale coherent structures shed as a consequence of the boundary layer laminar separation under steady inflow, as well as the investigation of the three-dimensional effects induced by the intermittent passage of low and high speed streaks. A close inspection of the time-mean velocity profiles as well as of the boundary layer integral parameters helps to characterize the suction side boundary layer state, thus justifying the influence of free-stream turbulence intensity on the blade aerodynamic losses measured under steady and unsteady inflows.

DNS study on the formation of Lambda rotational core and the role of TS wave in boundary layer transition

ABSTRACTΛ-vortex plays a pivotal role in the boundary layer transition process. In this paper, the generation of Lambda rotational core (Λ-vortex) in boundary layer transition and the role of TS (Tollmien–Schlichting) waves performed are studied by direct numerical simulation (DNS) in detail. It shows that the growth of TS wave under non-linear NS equations differs from the one under linearised NS (Navier–Stokes) equations at very early stage. With the evolution of TS waves, a newly found evolution procedure under non-linear NS equations helps to form the perturbation structures and lead them to evolve to Λ-vortices. The transition process undergoing the non-linear stage is also found to be much earlier than the previous studies had suggested.

Real-Gas and Surface-Ablation Effects on Hypersonic Boundary-Layer Instability over a Blunt Cone

There has been little research into surface-ablation effects on hypersonic boundary-layer instability, and the current understanding of real-gas effects on hypersonic boundary-layer instability still contains uncertainties. The objective of the current work was to analyze the hypersonic boundary-layer transition process using linear-stability theory, in which surface ablation, as well as real-gas effects, is present, and the second mode is the dominant instability mode. A thermochemical nonequilibrium linear-stability-theory code with a gas-phase model that includes multiple carbon species, as well as a linearized surface graphite-pseudoablation model, is developed and validated. It is validated with previously published linear-stability analysis and direct-numerical-simulation results. A high-order method for discretizing the linear-stability equations is used, which can easily include high-order boundary conditions. The developed linear-stability code, with mean-flow solutions produced from a high-order shock-fitting direct-numerical-simulation method for hypersonic flows with thermochemical nonequilibrium and surface-chemistry boundary conditions for graphite pseudoablation, is used to study hypersonic boundary-layer instability for a 7 deg half-angle blunt cone at Mach 15.99 and the Reentry-F experiment at 100 kft. Multiple mean-flow simulations are obtained with the same geometry and freestream conditions to help separate real-gas, blowing, and carbon-species effects on hypersonic boundary-layer instability. For the case at Mach 15.99, an factor comparison shows that real-gas effects significantly destabilize the flow when compared to an ideal gas. Blowing is destabilizing for the real-gas simulation, and has a negligible effect for the ideal-gas simulation due to the different locations of instability onset. Notably, carbon species resulting from ablation are shown to slightly stabilize the flow for both cases. It is also found that ablating nose-cone effects may safely be excluded when determining the factor at transition for Reentry-F at 100 kft.

Numerical investigation of the development of three-dimensional wavepackets in a sharp cone boundary layer at Mach 6

AbstractDirect numerical simulations were performed to investigate wavepackets in a sharp cone boundary layer at Mach 6. In order to understand the natural transition process in hypersonic cone boundary layers, the flow was forced by a short-duration (localized) pulse. The pulse disturbance developed into a three-dimensional wavepacket, which consisted of a wide range of disturbance frequencies and wavenumbers. First, the linear development of the wavepacket was studied by forcing the flow with a low-amplitude pulse (0.001 % of the free-stream velocity). The dominant waves within the resulting wavepacket were identified as the second-mode axisymmetric disturbance waves. In addition, weaker first-mode oblique disturbance waves were also observed on the lateral sides of the wavepacket. In order to investigate the nonlinear transition regime, large-amplitude pulse disturbances (0.5 % of the free-stream velocity) were introduced. The response of the flow to the large-amplitude pulse disturbances indicated the presence of a fundamental resonance mechanism. Lower secondary peaks in the disturbance wave spectrum were identified at approximately half the frequency of the high-amplitude frequency band, suggesting the possibility of a subharmonic resonance mechanism. However, the spectrum also indicated that the fundamental resonance was much stronger than the subharmonic resonance. A secondary stability investigation using controlled disturbances confirmed that fundamental resonance is indeed a dominant mechanism compared to subharmonic resonance. Furthermore, strong peaks in the disturbance wave spectrum were also observed for low-azimuthal-wavenumber second-mode oblique waves, hinting at a possible oblique breakdown mechanism. Thus, the wavepacket simulations indicate that the second-mode fundamental resonance and oblique breakdown mechanisms are the strongest for the investigated flow. Hence, both mechanisms are likely to be relevant in the natural transition process for a cone boundary layer at Mach 6.

Boundary-Layer Stability Analysis of the High Enthalpy Shock Tunnel Transition Experiments

Recent experiments focused on the hypersonic boundary-layer transition process have been conducted at Japan Aerospace Exploration Agency’s free-piston High Enthalpy Shock Tunnel facility. Numerical simulations of the experiments were conducted to generate mean-flow solutions upon which boundary-layer stability analyses were performed. These computations included the effects of chemical reactions and internal energy relaxation. The disturbances in the boundary layer were calculated by solving the linear parabolized stability equations. Comparisons of transition location, disturbance amplification rates, and disturbance frequencies were made between those measured experimentally and those calculated from the stability analysis. The results suggest that an factor of 8.0 is the value that indicates the onset of transition for this facility. A disturbance frequency comparison between the computations and experiment showed a reasonable agreement for the locations at which the flow remained laminar. Additional theoretical operational conditions were examined to investigate effects of finite-rate chemistry and vibrational excitation. For the conditions examined, the inclusion or exclusion of chemical kinetics modeling in the mean-flow solution influenced boundary-layer stability to a greater extent than the inclusion or exclusion of chemical kinetics in the stability analysis alone. In general, the inclusion of these effects increased disturbance frequencies, peak amplification rates, and the distance over which the disturbance frequencies were amplified.