Unsteady Boundary Layer Transition Research Articles

Direct Numerical Simulation of a High-Pressure Turbine Stage: Unsteady Boundary Layer Transition and the Resulting Flow Structures

Abstract In the present study, we investigate the unsteady boundary layer transition based on the direct numerical simulation database of a high-pressure turbine (HPT) stage (Zhao and Sandberg, 2021, High-Fidelity Simulations of a High-Pressure Turbine Stage: Effects of Reynolds Number and Inlet Turbulence, ASME Turbo Expo 2021, Paper No. GT2021-58995), focusing on the transition mechanisms on the rotor blade, affected by the incoming periodic wakes and the background freestream turbulence (FST). On the basis of the fully resolved flow fields, we provide detailed analysis of the flow structures responsible for the transition, and two distinctive transition paths have been identified. The first path is the typical bypass transition via the instability of Klebanoff streaks, which happens when the transition region is not directly affected by the wake. The suction-side boundary layer is disturbed at the leading edge, resulting in the formation of streamwise streaks. These streaky structures endure varicose instability in the region with adverse pressure gradient (APG), then quickly break down into turbulent spots, which then evolve into fully turbulent flow. The other transition path is a consequence of the direct interaction between the wake structures and the blade boundary layer, when the wake apex starts to affect the transitional region. To be specific, the wake structures directly interact with the separation bubble in the APG region, causing sudden breakdown into turbulence. A calmed region is found to follow the wake-induced turbulent boundary layer. It is observed that the recovery to a calmed region can be impacted by the FST, as the calmed region in case with no FST is much longer compared to cases with stronger FST.

Numerical Investigation of Unsteady Boundary Layer Transition on a Dynamically Pitching Rotor

Predictions of unsteady boundary layer transition are performed on a four-bladed rotor in axial inflow using a computational fluid dynamics approach. The configuration is based on experiments performed at the German Aerospace Center (DLR) in the 1.6-m × 3.4-m wind tunnel in the rotor test facility (RTG). Simulations are performed using the NASA OVERFLOW 2.3 solver with hybrid RANS/LES and laminar turbulent transition modeling. Solutions are based on a hover tip Mach number of 0.143 with prescribed cyclic pitching conditions. Computational methods and grid generation are described. The rotor flow field is analyzed, and the effect of transition modeling on unsteady boundary layer transition prediction is assessed. Laminar-turbulent transition predictions and rotor performance are compared to experimental measurements obtained at the DLR RTG. A study of sensitivity was performed on freestream turbulence intensity to investigate its effect on predicted rotor transition.

Effect of Crossflow Transition on the Pressure-Sensitive-Paint Rotor in Hover

The amplification factor transport transition model has been extended to include crossflow transition and applied to a rotor in hover. Validation of the crossflow model implementation has been performed using standard test cases. The pressure-sensitive-paint rotor is simulated in hover for various collective pitch angles as a test case for the transition model. The OVERFLOW 2.2n solver is used with laminar–turbulent transition modeling and hybrid Reynolds–averaged Navier–Stokes/large-eddy simulation turbulence modeling enabled. Grid generation, numerical methods, and transition modeling are described. The effects of crossflow transition are qualitatively and quantitatively assessed through comparisons to experiments, to a transition model without crossflow, and to a fully turbulent model. Additionally, several techniques for evaluating unsteady boundary-layer transition are described and are applied to solutions for the pressure-sensitive-paint rotor in hover.

Unsteady boundary-layer transition measurements and computations on a rotating blade under cyclic pitch conditions

The presented work tackles the lack of experimental investigations of unsteady laminar-turbulent boundary-layer transition on rotor blades at cyclic pitch actuation, which are important for accurate performance predictions of helicopters in forward flight. Unsteady transition positions were measured on the blade suction side of a four-bladed subscale rotor by means of non-intrusive differential infrared thermography (DIT). Experiments were conducted at different rotation rates corresponding to Mach and Reynolds numbers at 75% rotor radius of up to {M}_{75} = 0.21 and {mathrm {Re}}_{75} = 3.3 times 10^5 and with varying cyclic blade pitch settings. The setup allowed transition to be measured across the outer 54% of the rotor radius. For comparison, transition was also measured using conventional infrared thermography for steady cases with collective pitch settings only. The study is complemented by numerical simulations including boundary-layer transition modeling based on semi-empirical criteria. DIT results reveal the upstream and downstream motion of boundary-layer transition during upstroke and downstroke, a reasonable comparison to experimental results obtained using the already established sigma c_p method, and noticeable agreement with numerical simulations. The result is the first systematic study of unsteady boundary-layer transition on a rotor suction side by means of DIT including a comparison to numerical computations.

Application of hot-film anemometry to resolve the unsteady boundary layer transition of a laminar airfoil experiencing limit cycle oscillations

The unsteady boundary layer behavior of a supercritical laminar airfoil model, which undergoes limit cycle oscillations in pitch, is investigated by the application of hot-film anemometry. The data basis is a 2D flutter experiment under transonic flow conditions. The laminar airfoil model was elastically mounted with a single degree of freedom in pitch and performed self-excited limit cycle oscillations at a Mach number of {text {Ma}}=0.73 and a Reynolds number of {text {Re}}approx 2times 10^6. An analysis of the hot-film signals on the basis of the quasi-wall shear stress is carried out, with which the boundary layer state for steady and unsteady flow is resolved. An algorithm is presented which allows an automated detection of the transition position, so that a correlation between airfoil motion and transition location movement can be quantified. A sudden movement of the boundary layer transition is observed at the upper and lower reversal points of the limit cycle oscillation, while in parts of the up- and downstroke of the laminar airfoil a shock-induced transition occurs. During the limit cycle oscillation, a delayed response of the boundary layer occurs, resulting in a significant phase lag between the movement of the boundary layer transition and the motion of the laminar airfoil.

Advanced infrared thermography data analysis for unsteady boundary layer transition detection

Advanced data processing methods for detecting unsteady boundary layer transition in periodic aerodynamic processes by means of infrared thermography measurements are presented. The thermal radiation emitted from the heated suction surface of a pitching airfoil model in subsonic flow is measured with an infrared camera. The unsteady boundary layer transition location is detected by analyzing the difference in the infrared radiation signal over short periods of time with differential infrared thermography (DIT). The DIT method is optimized and automated in the present study, which facilitates the extension of the part of the motion period where valid DIT transition measurements are produced. Additionally, a new infrared thermography data processing method is introduced in this study. The extraction of the extrema of the measured radiation signal at fixed locations on the model surface yields instants of the motion period that relate to the occurrence of boundary layer transition. The local infrared thermography (LIT) approach can be extended to measuring the two-dimensional unsteady boundary layer transition front.

Optimization of differential infrared thermography for unsteady boundary layer transition measurement

Differential infrared thermography (DIT) is a method of analyzing infrared images to measure the unsteady motion of the laminar–turbulent transition of a boundary layer. It uses the subtraction of two infrared images taken with a short-time delay. DIT is a new technique which already demonstrated its validity in applications related to the unsteady aerodynamics of helicopter rotors in forward flight. The current study investigates a pitch-oscillating airfoil and proposes several optimizations of the original concept. These include the extension of DIT to steady test cases, a temperature compensation for long-term measurements, and a discussion of the proper infrared image separation distance. The current results also provide a deeper insight into the working principles of the technique. The results compare well to reference data acquired by unsteady pressure transducers, but at least for the current setup DIT results in an additional measurement-related lag for relevant pitching frequencies.

Synchronization of multi-plane measurement data by means of POD: application to unsteady boundary layer transition

Time-resolved particle image velocimetry measurements have been performed on the suction side of a low-pressure turbine cascade perturbed by periodic unsteady incoming wakes. The flow field has been investigated in two orthogonal planes. A wall-normal and a wall-parallel plane have been considered to identify and characterize the dimensions of the coherent structures developing within the boundary layer at the different wake passing phases. A new proper orthogonal decomposition (POD)-based procedure has been developed to sort the data acquired in the two planes. Particularly, the temporal coefficients of the POD modes related to wake migration have been used to provide phase-averaged data and synchronize the information obtained in the two measurement planes. It is shown that this technique does not require an external trigger, typically requested by previous available multi-plane phase-locking strategies. Moreover, POD has been applied to sub-ensembles of data at the same relative phase within the wake passing cycle. Thanks to the proposed procedure, the characteristic dimension of the eddies caused by the wake–boundary layer interaction has been investigated: at each phase of the wake passing cycle, the POD provides the representation of the most relevant flow features. It has been possible to observe the presence of boundary layer streaks and when they first occur within the wake passing period. One-dimensional auto-correlations of the velocity fields in the wall-parallel plane have been performed following the periodic perturbation induced by the passing wake in space and time. The auto-correlation provides the spacing between streaks, that, once scaled with the local boundary layer displacement thickness measured at the same phase in the wall-normal plane, is consistent with the available literature data for steady flows.

Measurement of unsteady transition on a pitching airfoil using dynamic pressure sensors

Unsteady boundary layer transition on a pitching OA209 airfoil in a wind tunnel was detected by using pressure fluctuation measurement at different oscillation frequency. Thirty Kulite dynamic pressure transducers flush-mounted on the airfoil surface recorded pressure signatures, and root mean square of pressure signatures were calculated. Results indicated that the criterion of transition for static airfoil defined as the peak of root mean square of pressure fluctuation was still suitable for detection of transition on a pitching airfoil. Fixed transition experiment for pitching airfoil was performed to validate the conclusion. Effect of oscillation frequency on transition was investigated. For small reduced frequency, the hysteresis loop is larger near leading edge. With increasing in the oscillation frequencies, the transition was promoted and relaminarization was enhanced.

Differential infrared thermography for boundary layer transition detection on pitching rotor blade models

Differential infrared thermography (DIT) was investigated and applied for the detection of unsteady boundary layer transition locations on a pitching airfoil and on a rotating blade under cyclic pitch. DIT is based on image intensity differences between two successively recorded infrared images. The images were recorded with a high framing rate infrared camera. A pitching NACA0012 airfoil served as the first test object. The recorded images were used in order to investigate and to further improve evaluation strategies for periodically moving boundary layer transition lines. The measurement results are compared with the results of unsteady CFD simulations based on the DLR-TAU code. DIT was then used for the first time for the optical measurement of unsteady transition locations on helicopter rotor blade models under cyclic pitch and rotation. Image de-rotation for tracking the blade was employed using a rotating mirror to increase exposure time without causing motion blur. The paper describes the challenges that occurred during the recording and evaluation of the data in detail. However, the results were found to be encouraging to further improve the method toward the measurement of unsteady boundary layer transition lines on helicopter rotor models in forward flight.

Unsteady boundary-layer transition in low-pressure turbines

This paper examines the transition process in a boundary layer similar to that present over the suction surfaces of aero-engine low-pressure (LP) turbine blades. This transition process is of significant practical interest since the behaviour of this boundary layer largely determines the overall efficiency of the LP turbine. Modern ‘high-lift’ blade designs typically feature a closed laminar separation bubble on the aft portion of the suction surface. The size of this bubble and hence the inefficiency it generates is controlled by the transition between laminar and turbulent flow in the boundary layer and separated shear layer. The transition process is complicated by the inherent unsteadiness of the multi-stage machine: the wakes shed by one blade row convect through the downstream blade passages, periodically disturbing the boundary layers. As a consequence, the transition to turbulence is multi-modal by nature, being promoted by periodic and turbulent fluctuations in the free stream and the inherent instabilities of the boundary layer. Despite many studies examining the flow behaviour, the detailed physics of the unsteady transition phenomena are not yet fully understood. The boundary-layer transition process has been studied experimentally on a flat plate. The opposing test-section wall was curved to impose a streamwise pressure distribution typical of modern high-lift LP turbines over the flat plate. The presence of an upstream blade row has been simulated by a set of moving bars, which shed wakes across the test section inlet. Further upstream, a grid has been installed to elevate the free-stream turbulence to a level believed to be representative of multi-stage LP turbines. Extensive particle imaging velocimetry (PIV) measurements have been performed on the flat-plate boundary layer to examine the flow behaviour. In the absence of the incoming bar wakes, the grid-generated free-stream turbulence induces relatively weak Klebanoff streaks in the boundary layer which are evident as streamwise streaks of low-velocity fluid. Transition is promoted by the streaks and by the inherent inflectional (Kelvin–Helmholtz (KH)) instability of the separation bubble. In unsteady flow, the incoming bar wakes generate stronger Klebanoff streaks as they pass over the leading edge, which convect downstream at a fraction of the free-stream velocity and spread in the streamwise direction. The region of amplified streaks convects in a similar manner to a classical turbulent spot: the leading and trailing edges travel at around 88% and 50% of the free-stream velocity, respectively. The strongest disturbances travel at around 70% of the free-stream velocity. The wakes induce a second type of disturbance as they pass over the separation bubble, in the form of short-span KH structures. Both the streaks and the KH structures contribute to the early wake-induced transition. The KH structures are similar to those observed in the simulation of separated flow transition with high free-stream turbulence by McAuliffe & Yaras (ASME J. Turbomach., vol. 132, no. 1, 2010, 011004), who observed that these structures originated from localised instabilities of the shear layer induced by Klebanoff streaks. In the current measurements, KH structures are frequently observed directly under the path of the wake. The wake-amplified Klebanoff streaks cannot affect the generation of these structures since they do not arrive at the bubble until later in the wake cycle. Rather, the KH structures arise from an interaction between the flow disturbances in the wake and localised instabilities in the shear layer, which are caused by the weak Klebanoff streaks induced by the grid turbulence. The breakdown of the KH structures to small-scale turbulence occurs a short time after the wake has passed over the bubble, and is largely driven by the arrival of the wake-amplified Klebanoff streaks from the leading edge. During this process, the re-attachment location moves rapidly upstream. The minimum length of the bubble occurs when the strongest wake-amplified Klebanoff streaks arrive from the leading edge; these structures travel at around 70% of the free-stream velocity. The bubble remains shorter than its steady-flow length until the trailing edge of the wake-amplified Klebanoff streaks, travelling at 50% of the free-stream velocity, convect past. After this time, the reattachment location moves aft on the surface as a consequence of a calmed flow region which follows behind the wake-induced turbulence.

Effects of Reynolds Number and Freestream Turbulence Intensity on the Unsteady Boundary Layer Development on an Ultra-High-Lift Low Pressure Turbine Airfoil

The effects of Reynolds numbers and the freestream turbulence intensities (FSTIs) on the unsteady boundary layer development on an ultra-high-lift low-pressure turbine airfoil, so-called T106C, are investigated. The measurements were carried out at both Tu=0.5% and 4.0% within a range of Reynolds numbers, based on the blade chord and the isentropic exit velocity, between 100,000 and 260,000. The interaction between the unsteady wake and the boundary layer depends on both the strength of the wake and the status of the boundary layer. At Tu=0.5%, both the wake’s high turbulence and the negative jet behavior of the wake dominate the interaction between the unsteady wake and the separated boundary layer on the suction surface of the airfoil. Since the wake turbulence cannot induce transition before separation on this ultra-high-lift blade, the negative jet of the wake has the opportunity to induce a rollup vortex. At Tu=4.0%, the time-mean separation on the suction surface is much smaller. With elevated FSTI, the turbulence in the wake just above the boundary layer is no longer distinguishable from the background turbulence level. The unsteady boundary layer transition is dominated by the wake’s negative jet induced boundary layer variation.

Computational Fluid Dynamics Study of Wake-Induced Transition on a Compressor-Like Flat Plate

Recent experimental work has documented the importance of wake passing on the behavior of transitional boundary layers on the suction surface of axial compressor blades. This paper documents computational fluid dynamics (CFD) simulations using a commercially available general-purpose CFD solver, performed on a representative case with unsteady transitional behavior. The study implements an advanced version of a three-equation eddy-viscosity model previously developed and documented by the authors, which is capable of resolving boundary layer transition. It is applied to the test cases of steady and unsteady boundary layer transition on a two-dimensional flat plate geometry with a freestream velocity distribution representative of the suction side of a compressor airfoil. The CFD results are analyzed and compared to a similar experimental test case from the open literature. Results with the model show a dramatic improvement over more typical Reynolds-averaged Navier–Stokes (RANS)-based modeling approaches, and highlight the importance of resolving transition in both steady and unsteady compressor aerosimulations.

On injection-ejection fluid influence through different accelerating porous surfaces on unsteady 2d incompressible boundary layer characteristics

Through the porous contour in perpendicular direction, the fluid of the same properties as incompressible fluid in basic flow, has been injected or ejected with velocity who is a function of the contour longitudinal coordinate and time. The corresponding equations of unsteady boundary layer, by introducing the appropriate variable transformations, momentum and energy equations and two similarity parameters sets, are transformed into generalized form. These parameters are expressing the influence of the outer flow velocity, the injection or ejection velocity and the flow history in boundary layer, on the boundary layer characteristics. Obtained generalized solutions are used to calculate the distributions of velocity, and shear stress in laminar-turbulent transition of unsteady incompressible boundary layer on different porous contours: circular cylinder, thin elliptical cylinder and aerofoil, whose centers velocities changes in time as a degree functions. The ejection of fluid postpones the boundary layer separation, i.e. laminar-turbulent transition, and vice versa the injection of fluid favors the separation. Boundary layer characteristics are found directly, no further numerical integration of momentum equation.

Advances in Unsteady Boundary Layer Transition Research, Part II: Experimental Verification

This two‐part article presents recent advances in boundary layer research into the unsteady boundary layer transition modeling and its validation. This, Part II, deals with the results of an inductive approach based on comprehensive experimental and theoretical studies of unsteady wake flow and unsteady boundary layer flow. The experiments were performed on a curved plate at a zero streamwise pressure gradient under periodic unsteady wake flow, in which the frequency of the periodic unsteady flow was varied. To validate the model, systematic experimental investigations were performed on the suction and pressure surfaces of turbine blades integrated into a high‐subsonic cascade test facility, which was designed for unsteady boundary layer investigations. The analysis of the experiment′s results and comparison with the model′s prediction confirm the validity of the model and its ability to predict accurately the unsteady boundary layer transition.

Advances in Unsteady Boundary Layer Transition Research, Part I: Theory and Modeling

This two‐part article presents recent advances in boundary layer research that deal with the unsteady boundary layer transition modeling and its validation. A new unsteady boundary layer transition model was developed based on a universal unsteady intermittency function. It accounts for the effects of periodic unsteady wake flow on the boundary layer transition. To establish the transition model, an inductive approach was implemented; the approach was based on the results of comprehensive experimental and theoretical studies of unsteady wake flow and unsteady boundary layer flow. The experiments were performed on a curved plate at a zero streamwise pressure gradient under a periodic unsteady wake flow, where the frequency of the periodic unsteady flow was varied. To validate the model, systematic experimental investigations were performed on the suction and pressure surfaces of turbine blades integrated into a high‐subsonic cascade test facility, which was designed for unsteady boundary layer investigations. The analysis of the experiment′s results and comparison with the model′s prediction confirm the validity of the model and its ability to predict accurately the unsteady boundary layer transition.

Advances in Unsteady Boundary Layer Transition Research, Part I: Theory and Modeling

Advances in Unsteady Boundary Layer Transition Research, Part I: Theory and Modeling

Advances in Unsteady Boundary Layer Transition Research, Part II: Experimental Verification

Advances in Unsteady Boundary Layer Transition Research, Part II: Experimental Verification

주기적 통과 후류의 방향과 주파수가 익형 위 비정상 천이경계층에 미치는 영향

주기적 통과 후류의 방향과 주파수가 익형 위 비정상 천이경계층에 미치는 영향

Unsteady incompressible MHD boundary layer on porous aerofoil in high accelerating fluid flow

The fluid, flowing past the surface, is incompressible and its electro-conductivity is constant. The present magnetic field is homogenous and perpendicular to the surface and through the porous contour the fluid has been injected or ejected. In order to study this problem, a polyparametric method known as generalized similarity method has been established. The corresponding equations of unsteady boundary layer, by introducing the appropriate variable transformations, momentum and energy equations and three similarity parameters sets, being transformed into generalized form. The numerical integration of the generalized equation with boundary conditions has been performed by means of the difference schemes and by using Tridiagonal Algorithm Method with iterations in the four parametric and twice localized approximation. So obtained generalized solutions are used to calculate the shear stress distribution in laminar-turbulent transition of unsteady boundary layer on porous high accelerating aerofoil. It's shown that for both in confuser and in diffuser regions the ejection of fluid postpones the boundary layer separation, and vice versa the fluid injection favors the separation. For both injection and ejection of fluid, the magnetic field increases the friction and postpones the laminar-turbulent transition.