W.J. Baars
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19 records found
1
Stratified Turbulent Flow Past a Monopile
Designing an Experimental Setup
Offshore Wind Farms (OWFs) are increasingly being placed in deeper, seasonally stratified parts of the sea, where the foundations supporting the turbines interact with density gradients in ways that aren't well understood.
The current knowledge primarily comes from Computational Fluid Dynamics (CFD) simulations and a limited number of field measurements.
Laboratory experiments, capable of providing high-resolution, repeatable data, remain scarce due to the complexity of the facilities required to create stable density stratification in a boundary layer.
This thesis addresses this knowledge-gap by developing and evaluating a new method for generating a two-layer, stably stratified turbulent boundary layer in an existing water tunnel.
This setup is then used to investigate the two-way interaction between such a boundary layer and the wake of a monopile.
A stratified turbulent boundary layer was created by combining Irwin spires, to thicken the boundary layer, with a gravity-driven injection system that introduced salt water into the flow through the sides of the spires.
Simultaneous velocity and density fields were measured using Particle Image Velocimetry (PIV) and Planar Laser Induced Fluorescence (PLIF), respectively, in the wake of a monopile model, over a range of bulk Richardson numbers.
Results show that the adapted facility successfully produced a stable density gradient while preserving most of the characteristic features of a turbulent boundary layer; the injection method did generate a gravity current and increased turbulence in the upper part of the flow, however.
This distinguishes the resulting boundary layer from the canonical flat-plate case.
In the wake of the monopile, stratification enlarged the recirculation region near the pycnocline, delayed the recovery of the streamwise velocity, suppressed vertical motion, and reduced the turbulent kinetic energy in the denser lower layer.
Only minor variations were observed across the tested Richardson numbers.
Conversely, the monopile wake was found to redistribute density, lowering it near the original pycnocline and raising it further up, while increasing both the turbulent buoyancy flux and the dissipation of scaled buoyancy variance, indicating increased irreversible mixing induced by the wake.
Although uncertainties in the upstream density evolution and in the Richardson number binning procedure limit a fully quantitative interpretation, the qualitative trends provide convincing evidence for a bidirectional interaction: stratification modifies the wake, and the wake in turn alters the local density distribution.
These findings demonstrate the feasibility of the developed method and provide a first idea of the wake dynamics, while also highlighting practical considerations for future facilities aiming to replicate environmentally relevant stratified conditions. ...
The current knowledge primarily comes from Computational Fluid Dynamics (CFD) simulations and a limited number of field measurements.
Laboratory experiments, capable of providing high-resolution, repeatable data, remain scarce due to the complexity of the facilities required to create stable density stratification in a boundary layer.
This thesis addresses this knowledge-gap by developing and evaluating a new method for generating a two-layer, stably stratified turbulent boundary layer in an existing water tunnel.
This setup is then used to investigate the two-way interaction between such a boundary layer and the wake of a monopile.
A stratified turbulent boundary layer was created by combining Irwin spires, to thicken the boundary layer, with a gravity-driven injection system that introduced salt water into the flow through the sides of the spires.
Simultaneous velocity and density fields were measured using Particle Image Velocimetry (PIV) and Planar Laser Induced Fluorescence (PLIF), respectively, in the wake of a monopile model, over a range of bulk Richardson numbers.
Results show that the adapted facility successfully produced a stable density gradient while preserving most of the characteristic features of a turbulent boundary layer; the injection method did generate a gravity current and increased turbulence in the upper part of the flow, however.
This distinguishes the resulting boundary layer from the canonical flat-plate case.
In the wake of the monopile, stratification enlarged the recirculation region near the pycnocline, delayed the recovery of the streamwise velocity, suppressed vertical motion, and reduced the turbulent kinetic energy in the denser lower layer.
Only minor variations were observed across the tested Richardson numbers.
Conversely, the monopile wake was found to redistribute density, lowering it near the original pycnocline and raising it further up, while increasing both the turbulent buoyancy flux and the dissipation of scaled buoyancy variance, indicating increased irreversible mixing induced by the wake.
Although uncertainties in the upstream density evolution and in the Richardson number binning procedure limit a fully quantitative interpretation, the qualitative trends provide convincing evidence for a bidirectional interaction: stratification modifies the wake, and the wake in turn alters the local density distribution.
These findings demonstrate the feasibility of the developed method and provide a first idea of the wake dynamics, while also highlighting practical considerations for future facilities aiming to replicate environmentally relevant stratified conditions. ...
Offshore Wind Farms (OWFs) are increasingly being placed in deeper, seasonally stratified parts of the sea, where the foundations supporting the turbines interact with density gradients in ways that aren't well understood.
The current knowledge primarily comes from Computational Fluid Dynamics (CFD) simulations and a limited number of field measurements.
Laboratory experiments, capable of providing high-resolution, repeatable data, remain scarce due to the complexity of the facilities required to create stable density stratification in a boundary layer.
This thesis addresses this knowledge-gap by developing and evaluating a new method for generating a two-layer, stably stratified turbulent boundary layer in an existing water tunnel.
This setup is then used to investigate the two-way interaction between such a boundary layer and the wake of a monopile.
A stratified turbulent boundary layer was created by combining Irwin spires, to thicken the boundary layer, with a gravity-driven injection system that introduced salt water into the flow through the sides of the spires.
Simultaneous velocity and density fields were measured using Particle Image Velocimetry (PIV) and Planar Laser Induced Fluorescence (PLIF), respectively, in the wake of a monopile model, over a range of bulk Richardson numbers.
Results show that the adapted facility successfully produced a stable density gradient while preserving most of the characteristic features of a turbulent boundary layer; the injection method did generate a gravity current and increased turbulence in the upper part of the flow, however.
This distinguishes the resulting boundary layer from the canonical flat-plate case.
In the wake of the monopile, stratification enlarged the recirculation region near the pycnocline, delayed the recovery of the streamwise velocity, suppressed vertical motion, and reduced the turbulent kinetic energy in the denser lower layer.
Only minor variations were observed across the tested Richardson numbers.
Conversely, the monopile wake was found to redistribute density, lowering it near the original pycnocline and raising it further up, while increasing both the turbulent buoyancy flux and the dissipation of scaled buoyancy variance, indicating increased irreversible mixing induced by the wake.
Although uncertainties in the upstream density evolution and in the Richardson number binning procedure limit a fully quantitative interpretation, the qualitative trends provide convincing evidence for a bidirectional interaction: stratification modifies the wake, and the wake in turn alters the local density distribution.
These findings demonstrate the feasibility of the developed method and provide a first idea of the wake dynamics, while also highlighting practical considerations for future facilities aiming to replicate environmentally relevant stratified conditions.
The current knowledge primarily comes from Computational Fluid Dynamics (CFD) simulations and a limited number of field measurements.
Laboratory experiments, capable of providing high-resolution, repeatable data, remain scarce due to the complexity of the facilities required to create stable density stratification in a boundary layer.
This thesis addresses this knowledge-gap by developing and evaluating a new method for generating a two-layer, stably stratified turbulent boundary layer in an existing water tunnel.
This setup is then used to investigate the two-way interaction between such a boundary layer and the wake of a monopile.
A stratified turbulent boundary layer was created by combining Irwin spires, to thicken the boundary layer, with a gravity-driven injection system that introduced salt water into the flow through the sides of the spires.
Simultaneous velocity and density fields were measured using Particle Image Velocimetry (PIV) and Planar Laser Induced Fluorescence (PLIF), respectively, in the wake of a monopile model, over a range of bulk Richardson numbers.
Results show that the adapted facility successfully produced a stable density gradient while preserving most of the characteristic features of a turbulent boundary layer; the injection method did generate a gravity current and increased turbulence in the upper part of the flow, however.
This distinguishes the resulting boundary layer from the canonical flat-plate case.
In the wake of the monopile, stratification enlarged the recirculation region near the pycnocline, delayed the recovery of the streamwise velocity, suppressed vertical motion, and reduced the turbulent kinetic energy in the denser lower layer.
Only minor variations were observed across the tested Richardson numbers.
Conversely, the monopile wake was found to redistribute density, lowering it near the original pycnocline and raising it further up, while increasing both the turbulent buoyancy flux and the dissipation of scaled buoyancy variance, indicating increased irreversible mixing induced by the wake.
Although uncertainties in the upstream density evolution and in the Richardson number binning procedure limit a fully quantitative interpretation, the qualitative trends provide convincing evidence for a bidirectional interaction: stratification modifies the wake, and the wake in turn alters the local density distribution.
These findings demonstrate the feasibility of the developed method and provide a first idea of the wake dynamics, while also highlighting practical considerations for future facilities aiming to replicate environmentally relevant stratified conditions.
Master thesis
(2026)
-
Vishwajeetsinh Jadhavrao, M.M. van Paassen, E. van Kampen, D.J. Juschus, E.J.J. Smeur, W.J. Baars
As aircraft propulsion is transitioning to hydrogen fuel-cells, ensuring safe and reliable operation remains a key challenge. This thesis develops a control methodology to automate the startup and shutdown of a multi-stack hydrogen fuel-cell propulsion system for aircraft, whilst minimizing stack degradation. The work was conducted in collaboration with DLR, with contributions from Airbus, ZAL and HSU, focusing on air-cooled, open-cathode PEM fuel cells for aerospace applications.
Control-oriented models were developed using Multi-level Flow Modelling, Finite State Machine, and subsystem physical models, which were validated against experimental data in Simulink. Based on these models, subsystem controllers were designed to operate under the supervisory Finite State Machine automatic controller. The simulation results verify the defined control safety and reliability requirements, demonstrating that the multi-level control methodology is robust in automating the startup and shutdown operations, highlighting the use case in future aircraft fuel-cell propulsion systems.
...
Control-oriented models were developed using Multi-level Flow Modelling, Finite State Machine, and subsystem physical models, which were validated against experimental data in Simulink. Based on these models, subsystem controllers were designed to operate under the supervisory Finite State Machine automatic controller. The simulation results verify the defined control safety and reliability requirements, demonstrating that the multi-level control methodology is robust in automating the startup and shutdown operations, highlighting the use case in future aircraft fuel-cell propulsion systems.
...
As aircraft propulsion is transitioning to hydrogen fuel-cells, ensuring safe and reliable operation remains a key challenge. This thesis develops a control methodology to automate the startup and shutdown of a multi-stack hydrogen fuel-cell propulsion system for aircraft, whilst minimizing stack degradation. The work was conducted in collaboration with DLR, with contributions from Airbus, ZAL and HSU, focusing on air-cooled, open-cathode PEM fuel cells for aerospace applications.
Control-oriented models were developed using Multi-level Flow Modelling, Finite State Machine, and subsystem physical models, which were validated against experimental data in Simulink. Based on these models, subsystem controllers were designed to operate under the supervisory Finite State Machine automatic controller. The simulation results verify the defined control safety and reliability requirements, demonstrating that the multi-level control methodology is robust in automating the startup and shutdown operations, highlighting the use case in future aircraft fuel-cell propulsion systems.
Control-oriented models were developed using Multi-level Flow Modelling, Finite State Machine, and subsystem physical models, which were validated against experimental data in Simulink. Based on these models, subsystem controllers were designed to operate under the supervisory Finite State Machine automatic controller. The simulation results verify the defined control safety and reliability requirements, demonstrating that the multi-level control methodology is robust in automating the startup and shutdown operations, highlighting the use case in future aircraft fuel-cell propulsion systems.
Smart Vortex Generators (SVGs) are used in a CFD study for control of loads on aeroelastic wings governed by mass-damper-spring structural model in the presence of unsteady wind gust disturbances, leveraging interpretable reduced-order system identification. The primary objective of the study is thus gust load alleviation (GLA). Unlike prior formulations, the dominant disturbance, such as a wind gust, is included explicitly as a control input within the model, allowing the MPC controller to respond adaptively and in anticipation to external system forcing. System dynamics are identified partly by Sparse Identification of Nonlinear Dynamics with Control (SINDYc) and Linear Parameter-Varying (LPV) system. SINDYc captures the wing-gust subsystem, while LPV offers physics-interpretable and sparse representation of the nonlinear aerodynamic effects associated with the different operating regimes of the SVGs—vortex generator and massive flow separation. Two types of nonlinearity are observed. Firstly, nonlinearities affecting amplitude and non-minimum-phase-like behaviour of the induced loads can be captured through one scheduling parameter—the SVG deflection angle. The second type of identified nonlinearity concerns the different GLA capability of SVGs in the presence of gusts with varying magnitude. This can be captured by second scheduling parameter—the SVG deflection angle multiplied by the effective angle of attack, identified by SINDYc. Once fully coupled reduced-order fluid-structure interaction dynamical model is constructed, model-predictive control (MPC) is deployed in order to study the maximum potential GLA capability of SVGs. It has been shown in previous studies that simple on/off control strategy induces secondary structural oscillations, resulting in more structural fatigue. It is shown here that optimal control strategy can mitigate these to an extent. MPC-controlled SVGs can best alleviate loads of low frequency high amplitude gusts, by up to ΔCL ≈ 0.23, while high frequency gusts require head start for SVG actuation for optimal GLA performance. In summary, the findings of this thesis demonstrate that SVGs constitute a viable candidate technology for active gust load alleviation when deployed together with optimal control framework.
...
...
Smart Vortex Generators (SVGs) are used in a CFD study for control of loads on aeroelastic wings governed by mass-damper-spring structural model in the presence of unsteady wind gust disturbances, leveraging interpretable reduced-order system identification. The primary objective of the study is thus gust load alleviation (GLA). Unlike prior formulations, the dominant disturbance, such as a wind gust, is included explicitly as a control input within the model, allowing the MPC controller to respond adaptively and in anticipation to external system forcing. System dynamics are identified partly by Sparse Identification of Nonlinear Dynamics with Control (SINDYc) and Linear Parameter-Varying (LPV) system. SINDYc captures the wing-gust subsystem, while LPV offers physics-interpretable and sparse representation of the nonlinear aerodynamic effects associated with the different operating regimes of the SVGs—vortex generator and massive flow separation. Two types of nonlinearity are observed. Firstly, nonlinearities affecting amplitude and non-minimum-phase-like behaviour of the induced loads can be captured through one scheduling parameter—the SVG deflection angle. The second type of identified nonlinearity concerns the different GLA capability of SVGs in the presence of gusts with varying magnitude. This can be captured by second scheduling parameter—the SVG deflection angle multiplied by the effective angle of attack, identified by SINDYc. Once fully coupled reduced-order fluid-structure interaction dynamical model is constructed, model-predictive control (MPC) is deployed in order to study the maximum potential GLA capability of SVGs. It has been shown in previous studies that simple on/off control strategy induces secondary structural oscillations, resulting in more structural fatigue. It is shown here that optimal control strategy can mitigate these to an extent. MPC-controlled SVGs can best alleviate loads of low frequency high amplitude gusts, by up to ΔCL ≈ 0.23, while high frequency gusts require head start for SVG actuation for optimal GLA performance. In summary, the findings of this thesis demonstrate that SVGs constitute a viable candidate technology for active gust load alleviation when deployed together with optimal control framework.
Master thesis
(2026)
-
S.M. Ribeiro Machado, M. Pini, W.J. Baars, C.M. de Servi, F.A. van Steen, Lorenzo Mazzei, Niccolò Casini
The increasing electrification of aircraft subsystems and the integration of high-power-density components are creating significant challenges for onboard thermal management systems. These key subsystems are reaching their performance limits, motivating the development of new advanced solutions. One promising option is the adoption of innovative heat exchanger concepts featuring complex internal geometries, such as strut-based lattice structures, and triply periodic minimal surfaces (TPMS), that are only feasible to manufacture through additive processes. The aim of this work is to develop a systematic methodology to characterize the thermo-hydraulic performance of advanced heat transfer structures based on periodic unit cell (Representative Volume Elements) simulations. The resulting numerical data are then used to calibrate reduced order models suitable for the preliminary design of heat transfer devices. More in detail, the study focuses on the characterization of four heat transfer topologies over a wide range of Reynolds and Prandtl numbers and structure porosities. The performance data from the simulations were used to derive empirical correlations for the friction factor and Nusselt number as a function of the topology porosity. These correlations were then implemented into an $\epsilon-NTU$-based model to support preliminary design activities. To verify this methodology, a cold plate was analyzed and the results were validated against CFD simulations. The methodology showed adequate accuracy for preliminary design, providing reliable predictions where CFD is computationally prohibitive. Future work will expand geometric inputs beyond porosity to broaden the design space of additively manufactured heat transfer devices.
...
The increasing electrification of aircraft subsystems and the integration of high-power-density components are creating significant challenges for onboard thermal management systems. These key subsystems are reaching their performance limits, motivating the development of new advanced solutions. One promising option is the adoption of innovative heat exchanger concepts featuring complex internal geometries, such as strut-based lattice structures, and triply periodic minimal surfaces (TPMS), that are only feasible to manufacture through additive processes. The aim of this work is to develop a systematic methodology to characterize the thermo-hydraulic performance of advanced heat transfer structures based on periodic unit cell (Representative Volume Elements) simulations. The resulting numerical data are then used to calibrate reduced order models suitable for the preliminary design of heat transfer devices. More in detail, the study focuses on the characterization of four heat transfer topologies over a wide range of Reynolds and Prandtl numbers and structure porosities. The performance data from the simulations were used to derive empirical correlations for the friction factor and Nusselt number as a function of the topology porosity. These correlations were then implemented into an $\epsilon-NTU$-based model to support preliminary design activities. To verify this methodology, a cold plate was analyzed and the results were validated against CFD simulations. The methodology showed adequate accuracy for preliminary design, providing reliable predictions where CFD is computationally prohibitive. Future work will expand geometric inputs beyond porosity to broaden the design space of additively manufactured heat transfer devices.
A Formula 1 car experiences dynamic aerodynamic conditions throughout a race, particularly during phases of acceleration, deceleration, and cornering. With the reintroduction of ground effect by Fédération Internationale de l’Automobile, FIA regulations in 2022, the aerodynamic sensitivity of the underfloor and front wing regions has become increasingly important. However, transient effects under such dynamic conditions are difficult to replicate in traditional wind tunnels or steady Computational Fluid Dynamics, CFD environments. This study numerically investigates the transient aerodynamic behavior of a front wing operating under ground effect, using a scaled model of the Tyrrell 026 Formula 1 front wing.
The research focuses on capturing transient aerodynamic behavior during straight-line motion, normal cornering, and yawed cornering configurations under both accelerating and decelerating conditions. 3D Reynolds Averaged Navier–Stokes (RANS) simulations using the SST 𝑘–𝜔 turbulence model are conducted in ANSYS Fluent. Detailed comparisons are made between steady state and unsteady results to understand how added mass effects, vortex shedding, and asymmetrical flow patterns influence force and moment coefficients.
Results show significant deviation in aerodynamic forces and moments during transient phases, particularly under cornering with yaw, where asymmetries in flow around the left and right endplates amplify aerodynamic imbalance. The study quantifies these effects using non-dimensional analysis and time-resolved post-processing, revealing critical dependencies between transient flow structures and aerodynamic response. These insights provide a foundation for improving the simulation fidelity of dynamic flow conditions and optimizing front wing setup for real-world race conditions. ...
The research focuses on capturing transient aerodynamic behavior during straight-line motion, normal cornering, and yawed cornering configurations under both accelerating and decelerating conditions. 3D Reynolds Averaged Navier–Stokes (RANS) simulations using the SST 𝑘–𝜔 turbulence model are conducted in ANSYS Fluent. Detailed comparisons are made between steady state and unsteady results to understand how added mass effects, vortex shedding, and asymmetrical flow patterns influence force and moment coefficients.
Results show significant deviation in aerodynamic forces and moments during transient phases, particularly under cornering with yaw, where asymmetries in flow around the left and right endplates amplify aerodynamic imbalance. The study quantifies these effects using non-dimensional analysis and time-resolved post-processing, revealing critical dependencies between transient flow structures and aerodynamic response. These insights provide a foundation for improving the simulation fidelity of dynamic flow conditions and optimizing front wing setup for real-world race conditions. ...
A Formula 1 car experiences dynamic aerodynamic conditions throughout a race, particularly during phases of acceleration, deceleration, and cornering. With the reintroduction of ground effect by Fédération Internationale de l’Automobile, FIA regulations in 2022, the aerodynamic sensitivity of the underfloor and front wing regions has become increasingly important. However, transient effects under such dynamic conditions are difficult to replicate in traditional wind tunnels or steady Computational Fluid Dynamics, CFD environments. This study numerically investigates the transient aerodynamic behavior of a front wing operating under ground effect, using a scaled model of the Tyrrell 026 Formula 1 front wing.
The research focuses on capturing transient aerodynamic behavior during straight-line motion, normal cornering, and yawed cornering configurations under both accelerating and decelerating conditions. 3D Reynolds Averaged Navier–Stokes (RANS) simulations using the SST 𝑘–𝜔 turbulence model are conducted in ANSYS Fluent. Detailed comparisons are made between steady state and unsteady results to understand how added mass effects, vortex shedding, and asymmetrical flow patterns influence force and moment coefficients.
Results show significant deviation in aerodynamic forces and moments during transient phases, particularly under cornering with yaw, where asymmetries in flow around the left and right endplates amplify aerodynamic imbalance. The study quantifies these effects using non-dimensional analysis and time-resolved post-processing, revealing critical dependencies between transient flow structures and aerodynamic response. These insights provide a foundation for improving the simulation fidelity of dynamic flow conditions and optimizing front wing setup for real-world race conditions.
The research focuses on capturing transient aerodynamic behavior during straight-line motion, normal cornering, and yawed cornering configurations under both accelerating and decelerating conditions. 3D Reynolds Averaged Navier–Stokes (RANS) simulations using the SST 𝑘–𝜔 turbulence model are conducted in ANSYS Fluent. Detailed comparisons are made between steady state and unsteady results to understand how added mass effects, vortex shedding, and asymmetrical flow patterns influence force and moment coefficients.
Results show significant deviation in aerodynamic forces and moments during transient phases, particularly under cornering with yaw, where asymmetries in flow around the left and right endplates amplify aerodynamic imbalance. The study quantifies these effects using non-dimensional analysis and time-resolved post-processing, revealing critical dependencies between transient flow structures and aerodynamic response. These insights provide a foundation for improving the simulation fidelity of dynamic flow conditions and optimizing front wing setup for real-world race conditions.
Mitigating Wall Interference in Transonic Wind Tunnel Testing of Swept Wings
A Numerical Study on Flowfield Correction Using Wall Deformation in the Cryogenic Ludwieg Tube Göttingen
Master thesis
(2025)
-
T. Dolipski, L.L.M. Veldhuis, W.J. Baars, M. Kotsonis, M. Barahona, A. Theiß, S. Hein
This thesis presents a comprehensive numerical investigation of wall interference effects in transonic wind tunnel testing of swept-wing configurations, by the example of the Cryogenic Ludwieg Tube Göttingen (KRG) at the German Aerospace Center (DLR). By employing Reynolds-Averaged Navier–Stokes (RANS) simulations using the DLR TAU code and the Spalart–Allmaras turbulence model, the study quantifies how test section walls affect the spanwise flow homogeneity around a swept wing at high Reynolds numbers. A detailed analysis was conducted to isolate the influence of key flow parameters, like angle of attack, sweep
angle, chord length, and Mach number, on wall-induced distortions.
The simulations reveal that vertical tunnel walls introduce significant spanwise pressure gradients, especially at higher sweep angles and larger chord lengths, i.e., aspect ratios. These effects are mitigated through a mesh deformation approach based on Radial Basis Functions, which enables the adaptive reshaping of tunnel walls to align with streamlines from an idealized, unbounded flowfield. The implementation of these deformed geometries resulted in substantial improvements in spanwise flow uniformity, reducing the root mean square errors in the pressure coefficient distribution by nearly 50% in critical test section regions. Not only is it possible to reduce spanwise gradients in the flowfield, but also to reduce the difference to infinitely swept conditions almost over the entire width of the test section.
Beyond identifying optimal strategies to reduce the influence of the walls on the flowfield, the thesis validates its numerical findings against existing experimental and numerical data. The proposed methodology offers a valuable toolset for pre-test planning and supports the development of more representative aerodynamic experiments, especially in the context of laminar-to-turbulent transition studies. These contributions provide a scientifically robust foundation for minimizing wall interference in transonic testing and support the advancement of research on laminar-to-turbulent transition. ...
angle, chord length, and Mach number, on wall-induced distortions.
The simulations reveal that vertical tunnel walls introduce significant spanwise pressure gradients, especially at higher sweep angles and larger chord lengths, i.e., aspect ratios. These effects are mitigated through a mesh deformation approach based on Radial Basis Functions, which enables the adaptive reshaping of tunnel walls to align with streamlines from an idealized, unbounded flowfield. The implementation of these deformed geometries resulted in substantial improvements in spanwise flow uniformity, reducing the root mean square errors in the pressure coefficient distribution by nearly 50% in critical test section regions. Not only is it possible to reduce spanwise gradients in the flowfield, but also to reduce the difference to infinitely swept conditions almost over the entire width of the test section.
Beyond identifying optimal strategies to reduce the influence of the walls on the flowfield, the thesis validates its numerical findings against existing experimental and numerical data. The proposed methodology offers a valuable toolset for pre-test planning and supports the development of more representative aerodynamic experiments, especially in the context of laminar-to-turbulent transition studies. These contributions provide a scientifically robust foundation for minimizing wall interference in transonic testing and support the advancement of research on laminar-to-turbulent transition. ...
This thesis presents a comprehensive numerical investigation of wall interference effects in transonic wind tunnel testing of swept-wing configurations, by the example of the Cryogenic Ludwieg Tube Göttingen (KRG) at the German Aerospace Center (DLR). By employing Reynolds-Averaged Navier–Stokes (RANS) simulations using the DLR TAU code and the Spalart–Allmaras turbulence model, the study quantifies how test section walls affect the spanwise flow homogeneity around a swept wing at high Reynolds numbers. A detailed analysis was conducted to isolate the influence of key flow parameters, like angle of attack, sweep
angle, chord length, and Mach number, on wall-induced distortions.
The simulations reveal that vertical tunnel walls introduce significant spanwise pressure gradients, especially at higher sweep angles and larger chord lengths, i.e., aspect ratios. These effects are mitigated through a mesh deformation approach based on Radial Basis Functions, which enables the adaptive reshaping of tunnel walls to align with streamlines from an idealized, unbounded flowfield. The implementation of these deformed geometries resulted in substantial improvements in spanwise flow uniformity, reducing the root mean square errors in the pressure coefficient distribution by nearly 50% in critical test section regions. Not only is it possible to reduce spanwise gradients in the flowfield, but also to reduce the difference to infinitely swept conditions almost over the entire width of the test section.
Beyond identifying optimal strategies to reduce the influence of the walls on the flowfield, the thesis validates its numerical findings against existing experimental and numerical data. The proposed methodology offers a valuable toolset for pre-test planning and supports the development of more representative aerodynamic experiments, especially in the context of laminar-to-turbulent transition studies. These contributions provide a scientifically robust foundation for minimizing wall interference in transonic testing and support the advancement of research on laminar-to-turbulent transition.
angle, chord length, and Mach number, on wall-induced distortions.
The simulations reveal that vertical tunnel walls introduce significant spanwise pressure gradients, especially at higher sweep angles and larger chord lengths, i.e., aspect ratios. These effects are mitigated through a mesh deformation approach based on Radial Basis Functions, which enables the adaptive reshaping of tunnel walls to align with streamlines from an idealized, unbounded flowfield. The implementation of these deformed geometries resulted in substantial improvements in spanwise flow uniformity, reducing the root mean square errors in the pressure coefficient distribution by nearly 50% in critical test section regions. Not only is it possible to reduce spanwise gradients in the flowfield, but also to reduce the difference to infinitely swept conditions almost over the entire width of the test section.
Beyond identifying optimal strategies to reduce the influence of the walls on the flowfield, the thesis validates its numerical findings against existing experimental and numerical data. The proposed methodology offers a valuable toolset for pre-test planning and supports the development of more representative aerodynamic experiments, especially in the context of laminar-to-turbulent transition studies. These contributions provide a scientifically robust foundation for minimizing wall interference in transonic testing and support the advancement of research on laminar-to-turbulent transition.
This thesis project examines the drag crisis trigger mechanism for double layer fabrics on cylindrical cross-flow. The primary goal is to discover why the double layer fabric is able to trigger the drag crisis much sooner than conventional surface roughness.
The methodology employed in this research follows an experimental approach, using balance measurements to determine the aerodynamic drag at varying Reynolds numbers for different configurations. Particle Image Velocimetry (PIV) measurements are performed to examine the boundary layer, flow separation point and other flow phenomena occurring near the cylinder surface.
Over the course of this research, eight different double layer configurations have been the subject of study, as well as six single fabric configurations and two reference configurations consisting of a bare cylinder and a cylinder with zigzag trips. Balance measurements have been performed on all of the configurations to determine which configurations are deemed relevant to be studied with PIV techniques. Thus, PIV measurements have been performed on two double layer configurations with a varying underlayer and the same overlayer, as well as the study of the individual fabrics employed to make up the two-fabric construction. That is to say, the two underlayers and one overlayer used have been studied on their own.
The balance results uncover that the minimum drag coefficient across all double layer configurations is achieved with the smallest rib spacing. Conversely, this minimum drag coefficient is located at the highest critical Reynolds number across all configurations. Additionally, a relationship between the critical Reynolds number and the underlayer rib spacing has been determined for the studied configurations. Furthermore, the PIV measurements provide insight into the normalized velocity fields and reconstructed pressure fields. With these results, the development of the boundary layer on the foreside of the cylinder with double layer fabrics can be studied. Examining the flow near the surface, it can be seen that the presence of the ribs results in a localized flow convergence (upstream of the rib) and divergence (downstream of the rib), these geometric effects accelerate and decelerate the flow locally, causing static pressure oscillations on the foreside of the cylinder. With the appearance of localized adverse pressure gradients on the foreside of the cylinder, flow instabilities are seeded eventually trigger the transition of the boundary layer to a turbulent state, thus allowing the flow to remain attached to the cylinder surface for longer, ultimately delaying separation and reducing pressure drag.
While the study has provided valuable insights regarding the trigger mechanism for the drag crisis on double layer fabrics on cylinders, it has also paved the way for further research regarding this topic and the specific effects of rib height, behavior and performance in unsteady flows and whether the two fabric construction is strictly necessary. These considerations are addressed at the end of the conclusions chapter. ...
The methodology employed in this research follows an experimental approach, using balance measurements to determine the aerodynamic drag at varying Reynolds numbers for different configurations. Particle Image Velocimetry (PIV) measurements are performed to examine the boundary layer, flow separation point and other flow phenomena occurring near the cylinder surface.
Over the course of this research, eight different double layer configurations have been the subject of study, as well as six single fabric configurations and two reference configurations consisting of a bare cylinder and a cylinder with zigzag trips. Balance measurements have been performed on all of the configurations to determine which configurations are deemed relevant to be studied with PIV techniques. Thus, PIV measurements have been performed on two double layer configurations with a varying underlayer and the same overlayer, as well as the study of the individual fabrics employed to make up the two-fabric construction. That is to say, the two underlayers and one overlayer used have been studied on their own.
The balance results uncover that the minimum drag coefficient across all double layer configurations is achieved with the smallest rib spacing. Conversely, this minimum drag coefficient is located at the highest critical Reynolds number across all configurations. Additionally, a relationship between the critical Reynolds number and the underlayer rib spacing has been determined for the studied configurations. Furthermore, the PIV measurements provide insight into the normalized velocity fields and reconstructed pressure fields. With these results, the development of the boundary layer on the foreside of the cylinder with double layer fabrics can be studied. Examining the flow near the surface, it can be seen that the presence of the ribs results in a localized flow convergence (upstream of the rib) and divergence (downstream of the rib), these geometric effects accelerate and decelerate the flow locally, causing static pressure oscillations on the foreside of the cylinder. With the appearance of localized adverse pressure gradients on the foreside of the cylinder, flow instabilities are seeded eventually trigger the transition of the boundary layer to a turbulent state, thus allowing the flow to remain attached to the cylinder surface for longer, ultimately delaying separation and reducing pressure drag.
While the study has provided valuable insights regarding the trigger mechanism for the drag crisis on double layer fabrics on cylinders, it has also paved the way for further research regarding this topic and the specific effects of rib height, behavior and performance in unsteady flows and whether the two fabric construction is strictly necessary. These considerations are addressed at the end of the conclusions chapter. ...
This thesis project examines the drag crisis trigger mechanism for double layer fabrics on cylindrical cross-flow. The primary goal is to discover why the double layer fabric is able to trigger the drag crisis much sooner than conventional surface roughness.
The methodology employed in this research follows an experimental approach, using balance measurements to determine the aerodynamic drag at varying Reynolds numbers for different configurations. Particle Image Velocimetry (PIV) measurements are performed to examine the boundary layer, flow separation point and other flow phenomena occurring near the cylinder surface.
Over the course of this research, eight different double layer configurations have been the subject of study, as well as six single fabric configurations and two reference configurations consisting of a bare cylinder and a cylinder with zigzag trips. Balance measurements have been performed on all of the configurations to determine which configurations are deemed relevant to be studied with PIV techniques. Thus, PIV measurements have been performed on two double layer configurations with a varying underlayer and the same overlayer, as well as the study of the individual fabrics employed to make up the two-fabric construction. That is to say, the two underlayers and one overlayer used have been studied on their own.
The balance results uncover that the minimum drag coefficient across all double layer configurations is achieved with the smallest rib spacing. Conversely, this minimum drag coefficient is located at the highest critical Reynolds number across all configurations. Additionally, a relationship between the critical Reynolds number and the underlayer rib spacing has been determined for the studied configurations. Furthermore, the PIV measurements provide insight into the normalized velocity fields and reconstructed pressure fields. With these results, the development of the boundary layer on the foreside of the cylinder with double layer fabrics can be studied. Examining the flow near the surface, it can be seen that the presence of the ribs results in a localized flow convergence (upstream of the rib) and divergence (downstream of the rib), these geometric effects accelerate and decelerate the flow locally, causing static pressure oscillations on the foreside of the cylinder. With the appearance of localized adverse pressure gradients on the foreside of the cylinder, flow instabilities are seeded eventually trigger the transition of the boundary layer to a turbulent state, thus allowing the flow to remain attached to the cylinder surface for longer, ultimately delaying separation and reducing pressure drag.
While the study has provided valuable insights regarding the trigger mechanism for the drag crisis on double layer fabrics on cylinders, it has also paved the way for further research regarding this topic and the specific effects of rib height, behavior and performance in unsteady flows and whether the two fabric construction is strictly necessary. These considerations are addressed at the end of the conclusions chapter.
The methodology employed in this research follows an experimental approach, using balance measurements to determine the aerodynamic drag at varying Reynolds numbers for different configurations. Particle Image Velocimetry (PIV) measurements are performed to examine the boundary layer, flow separation point and other flow phenomena occurring near the cylinder surface.
Over the course of this research, eight different double layer configurations have been the subject of study, as well as six single fabric configurations and two reference configurations consisting of a bare cylinder and a cylinder with zigzag trips. Balance measurements have been performed on all of the configurations to determine which configurations are deemed relevant to be studied with PIV techniques. Thus, PIV measurements have been performed on two double layer configurations with a varying underlayer and the same overlayer, as well as the study of the individual fabrics employed to make up the two-fabric construction. That is to say, the two underlayers and one overlayer used have been studied on their own.
The balance results uncover that the minimum drag coefficient across all double layer configurations is achieved with the smallest rib spacing. Conversely, this minimum drag coefficient is located at the highest critical Reynolds number across all configurations. Additionally, a relationship between the critical Reynolds number and the underlayer rib spacing has been determined for the studied configurations. Furthermore, the PIV measurements provide insight into the normalized velocity fields and reconstructed pressure fields. With these results, the development of the boundary layer on the foreside of the cylinder with double layer fabrics can be studied. Examining the flow near the surface, it can be seen that the presence of the ribs results in a localized flow convergence (upstream of the rib) and divergence (downstream of the rib), these geometric effects accelerate and decelerate the flow locally, causing static pressure oscillations on the foreside of the cylinder. With the appearance of localized adverse pressure gradients on the foreside of the cylinder, flow instabilities are seeded eventually trigger the transition of the boundary layer to a turbulent state, thus allowing the flow to remain attached to the cylinder surface for longer, ultimately delaying separation and reducing pressure drag.
While the study has provided valuable insights regarding the trigger mechanism for the drag crisis on double layer fabrics on cylinders, it has also paved the way for further research regarding this topic and the specific effects of rib height, behavior and performance in unsteady flows and whether the two fabric construction is strictly necessary. These considerations are addressed at the end of the conclusions chapter.
Jet Actuated Control of SWBLI
The Modulation of Separation Using Injection
Master thesis
(2025)
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K.P.M. Teunissen, S. Hickel, L. Laguarda Sanchez, B.W. van Oudheusden, W.J. Baars
This study investigated the control of Shock-Wave/Boundary-Layer Interactions (SWBLI) using jet actuators at several locations, both upstream and within the separation bubble. Such interactions are critical in high-speed aerodynamic applications, where flow separation can lead to performance losses. Using Large Eddy Simulation (LES), the effects of injection on the separation region are studied. Where both the time-averaged as unsteady effects are examined. The results show that size reduction of the separation bubble is not effective when the actuators are placed within the separation region. However the actuators were able to modulate the large scale motion of the separation independent of the injection frequency or location. Indicating that the instabilities of SWBLI are potentially inherent and can be modulated from within the bubble. The proposed driving mechanism behind this are the pressure disturbances created by the injection, which is supported by the results.
...
This study investigated the control of Shock-Wave/Boundary-Layer Interactions (SWBLI) using jet actuators at several locations, both upstream and within the separation bubble. Such interactions are critical in high-speed aerodynamic applications, where flow separation can lead to performance losses. Using Large Eddy Simulation (LES), the effects of injection on the separation region are studied. Where both the time-averaged as unsteady effects are examined. The results show that size reduction of the separation bubble is not effective when the actuators are placed within the separation region. However the actuators were able to modulate the large scale motion of the separation independent of the injection frequency or location. Indicating that the instabilities of SWBLI are potentially inherent and can be modulated from within the bubble. The proposed driving mechanism behind this are the pressure disturbances created by the injection, which is supported by the results.
This thesis investigates the aerodynamic performance of various Distributed Electric Propulsion (DEP) configurations using low-fidelity modeling tools. The study is motivated by the growing interest in sustainable aviation and the potential of DEP systems to offer improved aerodynamic efficiency, reduced noise, and greater design flexibility. Conducted in collaboration with Pipistrel Vertical Solutions, the research focuses on evaluating and optimizing different DEP layouts that combine minimum induced loss (MIL) propellers, dedicated for efficient cruise, and lift-augmenting (LA) propellers, designed to enhance lift during take-off and landing.
A key objective of the work is to develop a computationally efficient optimization framework suitable for early-stage design, capable of assessing multiple DEP configurations. To that end, the study integrates propeller design methodologies, slipstream modeling, and a DEP-specific lifting line solver into a single low-fidelity analysis tool. This framework is then coupled with multi-objective optimization algorithms, such as NSGA-II and SMPSO, to explore a wide design space and identify configurations that minimize take-off distance and cruise power requirements.
The thesis compares three main DEP arrangements, varying the number, size, and spanwise placement of MIL and LA propellers. The results show how propeller positioning and interaction effects influence overall aerodynamic performance, and they highlight the value and limitations of low-fidelity models in capturing these phenomena. The insights gained provide a foundation for future development of DEP systems and for refining low-fidelity tools for preliminary aircraft design. ...
A key objective of the work is to develop a computationally efficient optimization framework suitable for early-stage design, capable of assessing multiple DEP configurations. To that end, the study integrates propeller design methodologies, slipstream modeling, and a DEP-specific lifting line solver into a single low-fidelity analysis tool. This framework is then coupled with multi-objective optimization algorithms, such as NSGA-II and SMPSO, to explore a wide design space and identify configurations that minimize take-off distance and cruise power requirements.
The thesis compares three main DEP arrangements, varying the number, size, and spanwise placement of MIL and LA propellers. The results show how propeller positioning and interaction effects influence overall aerodynamic performance, and they highlight the value and limitations of low-fidelity models in capturing these phenomena. The insights gained provide a foundation for future development of DEP systems and for refining low-fidelity tools for preliminary aircraft design. ...
This thesis investigates the aerodynamic performance of various Distributed Electric Propulsion (DEP) configurations using low-fidelity modeling tools. The study is motivated by the growing interest in sustainable aviation and the potential of DEP systems to offer improved aerodynamic efficiency, reduced noise, and greater design flexibility. Conducted in collaboration with Pipistrel Vertical Solutions, the research focuses on evaluating and optimizing different DEP layouts that combine minimum induced loss (MIL) propellers, dedicated for efficient cruise, and lift-augmenting (LA) propellers, designed to enhance lift during take-off and landing.
A key objective of the work is to develop a computationally efficient optimization framework suitable for early-stage design, capable of assessing multiple DEP configurations. To that end, the study integrates propeller design methodologies, slipstream modeling, and a DEP-specific lifting line solver into a single low-fidelity analysis tool. This framework is then coupled with multi-objective optimization algorithms, such as NSGA-II and SMPSO, to explore a wide design space and identify configurations that minimize take-off distance and cruise power requirements.
The thesis compares three main DEP arrangements, varying the number, size, and spanwise placement of MIL and LA propellers. The results show how propeller positioning and interaction effects influence overall aerodynamic performance, and they highlight the value and limitations of low-fidelity models in capturing these phenomena. The insights gained provide a foundation for future development of DEP systems and for refining low-fidelity tools for preliminary aircraft design.
A key objective of the work is to develop a computationally efficient optimization framework suitable for early-stage design, capable of assessing multiple DEP configurations. To that end, the study integrates propeller design methodologies, slipstream modeling, and a DEP-specific lifting line solver into a single low-fidelity analysis tool. This framework is then coupled with multi-objective optimization algorithms, such as NSGA-II and SMPSO, to explore a wide design space and identify configurations that minimize take-off distance and cruise power requirements.
The thesis compares three main DEP arrangements, varying the number, size, and spanwise placement of MIL and LA propellers. The results show how propeller positioning and interaction effects influence overall aerodynamic performance, and they highlight the value and limitations of low-fidelity models in capturing these phenomena. The insights gained provide a foundation for future development of DEP systems and for refining low-fidelity tools for preliminary aircraft design.
Aerodynamics of wing-integrated ram-air duct for propeller aircraft
A numerical investigation into wing-integrated duct performance and wing-body junction flow
Steady Reynolds-Averaged Navier-Stokes (RANS) simulations utilizing the k-ω SST turbulence model are conducted to investigate the aerodynamic performance of a wing-integrated ram-air duct housing a heat exchanger for propeller-driven aircraft, including its impact on wing-body junction flow. The research is conducted in two stages: first, a 2D aerodynamic analysis employing a Design of Experiment (DoE) methodology to assess the sensitivity of key geometrical parameters–including stagger angle, leading-edge droop, duct gap, and heat exchanger characteristics–on lift, drag, and duct mass flow rate; and second, a 3D investigation of the junction flow behavior in the nacelle/ducted-wing configuration. The heat exchanger pressure drop is modeled as a porous media zone using the Darcy-Forchheimer quadratic drag law. Heat transfer is incorporated through a variable energy source term applied via a userdefined function (UDF) based on the ε-NTU correlation. Findings from the 2D aerodynamic analysis indicate that heat exchanger characteristics, particularly porosity and thickness, have a more pronounced impact on aerodynamic performance than external duct geometry. However, intake stagger angle and leading-edge droop play critical roles in mitigating flow separation and optimizing the wing pressure distribution. In addition, the redistribution of pressure due to flow restriction alters the stagnation point location, inlet-velocity ratio, and static pressure distributions, all of which influence the aerodynamic loading of the ducted wing. The optimal ducted airfoil configuration, featuring a lower-surface outlet aft of the maximum thickness and a thin heat exchanger, minimizes aerodynamic penalties while maximizing duct mass flow rate. However, thermal feasibility assessments reveal that meeting the cooling demands of fuelcell systems necessitates a thicker heat exchanger to accommodate sufficient heat transfer area within the constrained wing volume. This increase in thickness impairs aerodynamic performance through increased pressure drop and resultant drag. Although higher porosity mitigates flow resistance, the required thickness offsets this advantage, reinforcing the inherent trade-off between aero-thermal performance. In 3D, the presence of the heat exchanger inside the duct fundamentally alters the local aerodynamics by modifying boundary layer interactions at the wing-body junction. The flow resistance imposed by the heat exchanger directly affects the strength and topology of secondary flow structures, particularly the horseshoe vortex (HSV), which governs junction flow behavior and whose strength scales with the Reynolds number based on the momentum thickness of the incoming boundary layer. At low porosity levels, the stronger HSV , with an increased vertical extent above the wing, entrains high-momentum freestream flow into the chordwise and spanwise boundary layers, mitigating corner flow separation. Conversely, at high porosity levels, lower flow resistance alters HSV topology, reducing its vertical extent and allowing part of the vortex to enter the duct, inducing a secondary vortex at the lower lip. This weakens the HSV’s ability to stabilize the boundary layer, leading to earlier separation, increased pressure losses, and higher drag. A moderate porosity level provides an optimal balance between HSV strength, vertical positioning, and junction flow stability, reducing corner flow separation and associated pressure losses. Collectively, these findings yield critical insights into integrating ram-air cooling ducts within the wings of propeller-driven aircraft, offering a compelling approach to achieving efficient thermal management systems with minimal aerodynamic penalty. This investigation provides unprecedented detail in visualizing and understanding the intricate coupling between ducted wing aerodynamics and heat exchanger-induced flow interactions, while emphasizing the need for further research to validate and expand upon these findings
...
Steady Reynolds-Averaged Navier-Stokes (RANS) simulations utilizing the k-ω SST turbulence model are conducted to investigate the aerodynamic performance of a wing-integrated ram-air duct housing a heat exchanger for propeller-driven aircraft, including its impact on wing-body junction flow. The research is conducted in two stages: first, a 2D aerodynamic analysis employing a Design of Experiment (DoE) methodology to assess the sensitivity of key geometrical parameters–including stagger angle, leading-edge droop, duct gap, and heat exchanger characteristics–on lift, drag, and duct mass flow rate; and second, a 3D investigation of the junction flow behavior in the nacelle/ducted-wing configuration. The heat exchanger pressure drop is modeled as a porous media zone using the Darcy-Forchheimer quadratic drag law. Heat transfer is incorporated through a variable energy source term applied via a userdefined function (UDF) based on the ε-NTU correlation. Findings from the 2D aerodynamic analysis indicate that heat exchanger characteristics, particularly porosity and thickness, have a more pronounced impact on aerodynamic performance than external duct geometry. However, intake stagger angle and leading-edge droop play critical roles in mitigating flow separation and optimizing the wing pressure distribution. In addition, the redistribution of pressure due to flow restriction alters the stagnation point location, inlet-velocity ratio, and static pressure distributions, all of which influence the aerodynamic loading of the ducted wing. The optimal ducted airfoil configuration, featuring a lower-surface outlet aft of the maximum thickness and a thin heat exchanger, minimizes aerodynamic penalties while maximizing duct mass flow rate. However, thermal feasibility assessments reveal that meeting the cooling demands of fuelcell systems necessitates a thicker heat exchanger to accommodate sufficient heat transfer area within the constrained wing volume. This increase in thickness impairs aerodynamic performance through increased pressure drop and resultant drag. Although higher porosity mitigates flow resistance, the required thickness offsets this advantage, reinforcing the inherent trade-off between aero-thermal performance. In 3D, the presence of the heat exchanger inside the duct fundamentally alters the local aerodynamics by modifying boundary layer interactions at the wing-body junction. The flow resistance imposed by the heat exchanger directly affects the strength and topology of secondary flow structures, particularly the horseshoe vortex (HSV), which governs junction flow behavior and whose strength scales with the Reynolds number based on the momentum thickness of the incoming boundary layer. At low porosity levels, the stronger HSV , with an increased vertical extent above the wing, entrains high-momentum freestream flow into the chordwise and spanwise boundary layers, mitigating corner flow separation. Conversely, at high porosity levels, lower flow resistance alters HSV topology, reducing its vertical extent and allowing part of the vortex to enter the duct, inducing a secondary vortex at the lower lip. This weakens the HSV’s ability to stabilize the boundary layer, leading to earlier separation, increased pressure losses, and higher drag. A moderate porosity level provides an optimal balance between HSV strength, vertical positioning, and junction flow stability, reducing corner flow separation and associated pressure losses. Collectively, these findings yield critical insights into integrating ram-air cooling ducts within the wings of propeller-driven aircraft, offering a compelling approach to achieving efficient thermal management systems with minimal aerodynamic penalty. This investigation provides unprecedented detail in visualizing and understanding the intricate coupling between ducted wing aerodynamics and heat exchanger-induced flow interactions, while emphasizing the need for further research to validate and expand upon these findings
Master thesis
(2025)
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L.J. Stevense, B.W. van Oudheusden, F.F.J. Schrijer, J. Bulut, S. Hickel, W.J. Baars
An impinging Shock Wave-Turbulent Boundary Layer Interaction (SWTBLI) at Mach 2 was investigated experimentally while implementing two-dimensional Shock Control Bumps (SCBs). The aim was to investigate the changes in the unsteady dynamics while changing the bump ramp angle, tail angle, spanwise shape, and the impinging shock location. Schlieren and oil flow visualisations were used to identify these changes. Unsteadiness was quantified through a spectral analysis based on Welch’s method, and a Canny-based edge-detection algorithm was developed to track the impingement and separation location in the Schlieren images.
An uncontrolled SWTBLI could successfully be generated, and the implementation of the baseline bump showed a replacement of the unsteady separation shock by a steady compression ramp shock originating from the leading edge. Spectral analysis confirmed this behaviour since the characteristic low-frequencies of the uncontrolled separation shock seemed to be removed for this compression ramp shock. An increase in the bump ramp angle showed the progressive generation of a separation shock upstream of the bump with spectral content trending towards the low frequencies of the uncontrolled interaction. On the contrary, no alterations were observed for an increase in tail angle. An upstream impingement produced similar behaviour as the increase in ramp angle; a separation shock was generated upstream of the bump with increased low-frequency spectral content. The downstream impingement, however, did not show any alteration in the same region. Major factors influencing the unsteady dynamics are identified as the impingement location and the ramp shape of the bump.
The developed edge-detection algorithm proved unsuccessful in quantifying the unsteadiness although it could detect the impinging and reflection shock of the interactions. Spatial standard deviation distributions of the interaction revealed increased deviation values in the impinging shock suggesting that the impinging shock fluctuates. Rather, it is suspected to be a form of noise intrinsic to the experimental technique. Therefore, this affects the detection of the edges and the calculated impingement and separation location. Future research is suggested to improve the noise mitigation method in the algorithm. Additionally, the benefits associated with the 2D-SCB would be most noticeable in an integrated approach with an industrial application. Finally, numerical simulations and/or different experimental techniques are suggested for future research to obtain a better quantification of the interaction and unsteady dynamics.
...
An uncontrolled SWTBLI could successfully be generated, and the implementation of the baseline bump showed a replacement of the unsteady separation shock by a steady compression ramp shock originating from the leading edge. Spectral analysis confirmed this behaviour since the characteristic low-frequencies of the uncontrolled separation shock seemed to be removed for this compression ramp shock. An increase in the bump ramp angle showed the progressive generation of a separation shock upstream of the bump with spectral content trending towards the low frequencies of the uncontrolled interaction. On the contrary, no alterations were observed for an increase in tail angle. An upstream impingement produced similar behaviour as the increase in ramp angle; a separation shock was generated upstream of the bump with increased low-frequency spectral content. The downstream impingement, however, did not show any alteration in the same region. Major factors influencing the unsteady dynamics are identified as the impingement location and the ramp shape of the bump.
The developed edge-detection algorithm proved unsuccessful in quantifying the unsteadiness although it could detect the impinging and reflection shock of the interactions. Spatial standard deviation distributions of the interaction revealed increased deviation values in the impinging shock suggesting that the impinging shock fluctuates. Rather, it is suspected to be a form of noise intrinsic to the experimental technique. Therefore, this affects the detection of the edges and the calculated impingement and separation location. Future research is suggested to improve the noise mitigation method in the algorithm. Additionally, the benefits associated with the 2D-SCB would be most noticeable in an integrated approach with an industrial application. Finally, numerical simulations and/or different experimental techniques are suggested for future research to obtain a better quantification of the interaction and unsteady dynamics.
...
An impinging Shock Wave-Turbulent Boundary Layer Interaction (SWTBLI) at Mach 2 was investigated experimentally while implementing two-dimensional Shock Control Bumps (SCBs). The aim was to investigate the changes in the unsteady dynamics while changing the bump ramp angle, tail angle, spanwise shape, and the impinging shock location. Schlieren and oil flow visualisations were used to identify these changes. Unsteadiness was quantified through a spectral analysis based on Welch’s method, and a Canny-based edge-detection algorithm was developed to track the impingement and separation location in the Schlieren images.
An uncontrolled SWTBLI could successfully be generated, and the implementation of the baseline bump showed a replacement of the unsteady separation shock by a steady compression ramp shock originating from the leading edge. Spectral analysis confirmed this behaviour since the characteristic low-frequencies of the uncontrolled separation shock seemed to be removed for this compression ramp shock. An increase in the bump ramp angle showed the progressive generation of a separation shock upstream of the bump with spectral content trending towards the low frequencies of the uncontrolled interaction. On the contrary, no alterations were observed for an increase in tail angle. An upstream impingement produced similar behaviour as the increase in ramp angle; a separation shock was generated upstream of the bump with increased low-frequency spectral content. The downstream impingement, however, did not show any alteration in the same region. Major factors influencing the unsteady dynamics are identified as the impingement location and the ramp shape of the bump.
The developed edge-detection algorithm proved unsuccessful in quantifying the unsteadiness although it could detect the impinging and reflection shock of the interactions. Spatial standard deviation distributions of the interaction revealed increased deviation values in the impinging shock suggesting that the impinging shock fluctuates. Rather, it is suspected to be a form of noise intrinsic to the experimental technique. Therefore, this affects the detection of the edges and the calculated impingement and separation location. Future research is suggested to improve the noise mitigation method in the algorithm. Additionally, the benefits associated with the 2D-SCB would be most noticeable in an integrated approach with an industrial application. Finally, numerical simulations and/or different experimental techniques are suggested for future research to obtain a better quantification of the interaction and unsteady dynamics.
An uncontrolled SWTBLI could successfully be generated, and the implementation of the baseline bump showed a replacement of the unsteady separation shock by a steady compression ramp shock originating from the leading edge. Spectral analysis confirmed this behaviour since the characteristic low-frequencies of the uncontrolled separation shock seemed to be removed for this compression ramp shock. An increase in the bump ramp angle showed the progressive generation of a separation shock upstream of the bump with spectral content trending towards the low frequencies of the uncontrolled interaction. On the contrary, no alterations were observed for an increase in tail angle. An upstream impingement produced similar behaviour as the increase in ramp angle; a separation shock was generated upstream of the bump with increased low-frequency spectral content. The downstream impingement, however, did not show any alteration in the same region. Major factors influencing the unsteady dynamics are identified as the impingement location and the ramp shape of the bump.
The developed edge-detection algorithm proved unsuccessful in quantifying the unsteadiness although it could detect the impinging and reflection shock of the interactions. Spatial standard deviation distributions of the interaction revealed increased deviation values in the impinging shock suggesting that the impinging shock fluctuates. Rather, it is suspected to be a form of noise intrinsic to the experimental technique. Therefore, this affects the detection of the edges and the calculated impingement and separation location. Future research is suggested to improve the noise mitigation method in the algorithm. Additionally, the benefits associated with the 2D-SCB would be most noticeable in an integrated approach with an industrial application. Finally, numerical simulations and/or different experimental techniques are suggested for future research to obtain a better quantification of the interaction and unsteady dynamics.
Hot-wire anemometry (HWA) remains an essential diagnostic tool for high–frequency velocity measurements in aerodynamic research, yet its practical accuracy strongly depends on robust temperature correction and reliable multi–component probe calibration. Despite the maturity of the technique, the literature provides no consensus on how best to correct for fluid–temperature variations or how to calibrate X-wire probes with consistent accuracy across wide velocity and angular ranges. The situation is further complicated at the TU Delft Low-Speed Wind Tunnel Laboratory (LSL), where the absence of an in-situ multi-wire calibration system has made efficient and repeatable probe preparation difficult. This thesis addresses these issues by developing a new automated calibration framework for multi-wire CTA probes and performing a comprehensive evaluation of existing temperature-correction and X-wire calibration schemes.
The first part of the work concerns the separation of velocity and temperature effects in CTA measurements. Six temperature-correction models—ranging from simple Bearman [1]-type scaling ignoring the change in fluid properties with temperature, to more complex formulations incorporating fluid-property variations and the Collis & Williams factor [2] are applied to four independent datasets from literature, covering a wide range of velocities and temperature conditions. Each model is assessed by collapsing data acquired at varying temperatures onto a single fourth-degree polynomial calibration curve to quantify the residual error via normalized root mean square percentage error. Across all datasets, simple temperature corrections perform surprisingly well. Relying on resistance-based wire-temperature estimation suffers from uncertainty, with the best success when calculating the fluid properties at the film temperature and applying the Collis & Williams correction. Allowing the wire temperature to act as an optimization parameter can improve the fit of the curves, confirming earlier findings that resistance-derived temperatures can introduce systematic error.
The second part of the thesis focuses on the calibration of X-wires. A new miniature, high-precision, in-situ yaw calibrator is designed and implemented. The device integrates a compact high-torque servo motor and a magnetic rotary encoder with high accuracy, assembled using high-precision alignment procedures. With the calibration and controlled through a LabView interface, and the position feedback and servo control integrated via an Arduino implementation, the system enables automated yaw sweeps. Its compact form factor allows installation directly inside restricted windtunnel test sections.
Using this system, a dense X-wire calibration dataset was collected, spanning velocities from 5 to 30ms−1 in 1ms−1 increments and yaw angles from −40◦ to 40◦ in 1◦ steps. The dataset enabled comparison of calibration schemes, including interpolation-based indirect methods (Lueptow [3] and Tropea [4] variants), and direct polynomial surface ...
The first part of the work concerns the separation of velocity and temperature effects in CTA measurements. Six temperature-correction models—ranging from simple Bearman [1]-type scaling ignoring the change in fluid properties with temperature, to more complex formulations incorporating fluid-property variations and the Collis & Williams factor [2] are applied to four independent datasets from literature, covering a wide range of velocities and temperature conditions. Each model is assessed by collapsing data acquired at varying temperatures onto a single fourth-degree polynomial calibration curve to quantify the residual error via normalized root mean square percentage error. Across all datasets, simple temperature corrections perform surprisingly well. Relying on resistance-based wire-temperature estimation suffers from uncertainty, with the best success when calculating the fluid properties at the film temperature and applying the Collis & Williams correction. Allowing the wire temperature to act as an optimization parameter can improve the fit of the curves, confirming earlier findings that resistance-derived temperatures can introduce systematic error.
The second part of the thesis focuses on the calibration of X-wires. A new miniature, high-precision, in-situ yaw calibrator is designed and implemented. The device integrates a compact high-torque servo motor and a magnetic rotary encoder with high accuracy, assembled using high-precision alignment procedures. With the calibration and controlled through a LabView interface, and the position feedback and servo control integrated via an Arduino implementation, the system enables automated yaw sweeps. Its compact form factor allows installation directly inside restricted windtunnel test sections.
Using this system, a dense X-wire calibration dataset was collected, spanning velocities from 5 to 30ms−1 in 1ms−1 increments and yaw angles from −40◦ to 40◦ in 1◦ steps. The dataset enabled comparison of calibration schemes, including interpolation-based indirect methods (Lueptow [3] and Tropea [4] variants), and direct polynomial surface ...
Hot-wire anemometry (HWA) remains an essential diagnostic tool for high–frequency velocity measurements in aerodynamic research, yet its practical accuracy strongly depends on robust temperature correction and reliable multi–component probe calibration. Despite the maturity of the technique, the literature provides no consensus on how best to correct for fluid–temperature variations or how to calibrate X-wire probes with consistent accuracy across wide velocity and angular ranges. The situation is further complicated at the TU Delft Low-Speed Wind Tunnel Laboratory (LSL), where the absence of an in-situ multi-wire calibration system has made efficient and repeatable probe preparation difficult. This thesis addresses these issues by developing a new automated calibration framework for multi-wire CTA probes and performing a comprehensive evaluation of existing temperature-correction and X-wire calibration schemes.
The first part of the work concerns the separation of velocity and temperature effects in CTA measurements. Six temperature-correction models—ranging from simple Bearman [1]-type scaling ignoring the change in fluid properties with temperature, to more complex formulations incorporating fluid-property variations and the Collis & Williams factor [2] are applied to four independent datasets from literature, covering a wide range of velocities and temperature conditions. Each model is assessed by collapsing data acquired at varying temperatures onto a single fourth-degree polynomial calibration curve to quantify the residual error via normalized root mean square percentage error. Across all datasets, simple temperature corrections perform surprisingly well. Relying on resistance-based wire-temperature estimation suffers from uncertainty, with the best success when calculating the fluid properties at the film temperature and applying the Collis & Williams correction. Allowing the wire temperature to act as an optimization parameter can improve the fit of the curves, confirming earlier findings that resistance-derived temperatures can introduce systematic error.
The second part of the thesis focuses on the calibration of X-wires. A new miniature, high-precision, in-situ yaw calibrator is designed and implemented. The device integrates a compact high-torque servo motor and a magnetic rotary encoder with high accuracy, assembled using high-precision alignment procedures. With the calibration and controlled through a LabView interface, and the position feedback and servo control integrated via an Arduino implementation, the system enables automated yaw sweeps. Its compact form factor allows installation directly inside restricted windtunnel test sections.
Using this system, a dense X-wire calibration dataset was collected, spanning velocities from 5 to 30ms−1 in 1ms−1 increments and yaw angles from −40◦ to 40◦ in 1◦ steps. The dataset enabled comparison of calibration schemes, including interpolation-based indirect methods (Lueptow [3] and Tropea [4] variants), and direct polynomial surface
The first part of the work concerns the separation of velocity and temperature effects in CTA measurements. Six temperature-correction models—ranging from simple Bearman [1]-type scaling ignoring the change in fluid properties with temperature, to more complex formulations incorporating fluid-property variations and the Collis & Williams factor [2] are applied to four independent datasets from literature, covering a wide range of velocities and temperature conditions. Each model is assessed by collapsing data acquired at varying temperatures onto a single fourth-degree polynomial calibration curve to quantify the residual error via normalized root mean square percentage error. Across all datasets, simple temperature corrections perform surprisingly well. Relying on resistance-based wire-temperature estimation suffers from uncertainty, with the best success when calculating the fluid properties at the film temperature and applying the Collis & Williams correction. Allowing the wire temperature to act as an optimization parameter can improve the fit of the curves, confirming earlier findings that resistance-derived temperatures can introduce systematic error.
The second part of the thesis focuses on the calibration of X-wires. A new miniature, high-precision, in-situ yaw calibrator is designed and implemented. The device integrates a compact high-torque servo motor and a magnetic rotary encoder with high accuracy, assembled using high-precision alignment procedures. With the calibration and controlled through a LabView interface, and the position feedback and servo control integrated via an Arduino implementation, the system enables automated yaw sweeps. Its compact form factor allows installation directly inside restricted windtunnel test sections.
Using this system, a dense X-wire calibration dataset was collected, spanning velocities from 5 to 30ms−1 in 1ms−1 increments and yaw angles from −40◦ to 40◦ in 1◦ steps. The dataset enabled comparison of calibration schemes, including interpolation-based indirect methods (Lueptow [3] and Tropea [4] variants), and direct polynomial surface
The aviation sector faces growing pressure to reduce greenhouse gas emissions, while projections are suggesting that emissions could double by 2050 to account for increases in passengers. Liquid hydrogen fuel cell-electric (LH2FCE) propulsion offers a promising solution, with the potential to reduce climate impact by up to 90\% compared to conventional turboprop engines. However, significant challenges exist, including heat dissipation, increased drag, and increased weight penalties from heavy fuel cell stacks, cryogenic hydrogen storage and thermal management systems.
This thesis evaluates the impact on payload and range of retrofitting a regional turboprop aircraft with an LH2FCE propulsion system, by performing the preliminary design of the balance-of-plant systems, and assessing system performance by means of steady state analyses in take-off, top-of-climb and cruise conditions. A lumped parameter model was developed to simulate the fuel cell and balance-of-plant components, including an air supply and thermal management system and ram air ducts. Results show that payload reductions of approximately 58-77\% are expected for a 1500 km range compared to current turboprop aircraft, primarily due to increased mass and drag penalties, which reduce the propulsion system's specific power and lift-to-drag ratio. Sensitivity analyses were conducted, highlighting the effects of fuel cell operational parameters. It was revealed that adjustments in fuel cell temperature, pressure, and current density in the different operating conditions can enhance system performance. Additionally, it was found that most systems must be sized for top-of-climb, except for the ram air duct, which is constrained by take-off conditions. Incorporating a variable inlet design may eliminate ram air drag in cruise, although detailed drag analysis is recommended. Integration of a turbine and halving the rate-of-climb demonstrated that achieving a 1500 km range with a competitive payload (>3000 kg, 35 passengers) is feasible for retrofits, while reserving mass for non-modeled systems such as batteries. Moreover, projections for 2030 suggest that the performance of current turboprop aircraft could be matched by LH2FCE systems.
These findings highlight that LH2FCE propulsion is a viable and sustainable alternative for regional aviation, provided the reduction in payload is acceptable. With advancements in fuel cells, heat exchangers, electric motors, and liquid hydrogen storage, LH2FCE aircraft could achieve performance of current kerosene-powered turboprop aircraft by 2030. ...
This thesis evaluates the impact on payload and range of retrofitting a regional turboprop aircraft with an LH2FCE propulsion system, by performing the preliminary design of the balance-of-plant systems, and assessing system performance by means of steady state analyses in take-off, top-of-climb and cruise conditions. A lumped parameter model was developed to simulate the fuel cell and balance-of-plant components, including an air supply and thermal management system and ram air ducts. Results show that payload reductions of approximately 58-77\% are expected for a 1500 km range compared to current turboprop aircraft, primarily due to increased mass and drag penalties, which reduce the propulsion system's specific power and lift-to-drag ratio. Sensitivity analyses were conducted, highlighting the effects of fuel cell operational parameters. It was revealed that adjustments in fuel cell temperature, pressure, and current density in the different operating conditions can enhance system performance. Additionally, it was found that most systems must be sized for top-of-climb, except for the ram air duct, which is constrained by take-off conditions. Incorporating a variable inlet design may eliminate ram air drag in cruise, although detailed drag analysis is recommended. Integration of a turbine and halving the rate-of-climb demonstrated that achieving a 1500 km range with a competitive payload (>3000 kg, 35 passengers) is feasible for retrofits, while reserving mass for non-modeled systems such as batteries. Moreover, projections for 2030 suggest that the performance of current turboprop aircraft could be matched by LH2FCE systems.
These findings highlight that LH2FCE propulsion is a viable and sustainable alternative for regional aviation, provided the reduction in payload is acceptable. With advancements in fuel cells, heat exchangers, electric motors, and liquid hydrogen storage, LH2FCE aircraft could achieve performance of current kerosene-powered turboprop aircraft by 2030. ...
The aviation sector faces growing pressure to reduce greenhouse gas emissions, while projections are suggesting that emissions could double by 2050 to account for increases in passengers. Liquid hydrogen fuel cell-electric (LH2FCE) propulsion offers a promising solution, with the potential to reduce climate impact by up to 90\% compared to conventional turboprop engines. However, significant challenges exist, including heat dissipation, increased drag, and increased weight penalties from heavy fuel cell stacks, cryogenic hydrogen storage and thermal management systems.
This thesis evaluates the impact on payload and range of retrofitting a regional turboprop aircraft with an LH2FCE propulsion system, by performing the preliminary design of the balance-of-plant systems, and assessing system performance by means of steady state analyses in take-off, top-of-climb and cruise conditions. A lumped parameter model was developed to simulate the fuel cell and balance-of-plant components, including an air supply and thermal management system and ram air ducts. Results show that payload reductions of approximately 58-77\% are expected for a 1500 km range compared to current turboprop aircraft, primarily due to increased mass and drag penalties, which reduce the propulsion system's specific power and lift-to-drag ratio. Sensitivity analyses were conducted, highlighting the effects of fuel cell operational parameters. It was revealed that adjustments in fuel cell temperature, pressure, and current density in the different operating conditions can enhance system performance. Additionally, it was found that most systems must be sized for top-of-climb, except for the ram air duct, which is constrained by take-off conditions. Incorporating a variable inlet design may eliminate ram air drag in cruise, although detailed drag analysis is recommended. Integration of a turbine and halving the rate-of-climb demonstrated that achieving a 1500 km range with a competitive payload (>3000 kg, 35 passengers) is feasible for retrofits, while reserving mass for non-modeled systems such as batteries. Moreover, projections for 2030 suggest that the performance of current turboprop aircraft could be matched by LH2FCE systems.
These findings highlight that LH2FCE propulsion is a viable and sustainable alternative for regional aviation, provided the reduction in payload is acceptable. With advancements in fuel cells, heat exchangers, electric motors, and liquid hydrogen storage, LH2FCE aircraft could achieve performance of current kerosene-powered turboprop aircraft by 2030.
This thesis evaluates the impact on payload and range of retrofitting a regional turboprop aircraft with an LH2FCE propulsion system, by performing the preliminary design of the balance-of-plant systems, and assessing system performance by means of steady state analyses in take-off, top-of-climb and cruise conditions. A lumped parameter model was developed to simulate the fuel cell and balance-of-plant components, including an air supply and thermal management system and ram air ducts. Results show that payload reductions of approximately 58-77\% are expected for a 1500 km range compared to current turboprop aircraft, primarily due to increased mass and drag penalties, which reduce the propulsion system's specific power and lift-to-drag ratio. Sensitivity analyses were conducted, highlighting the effects of fuel cell operational parameters. It was revealed that adjustments in fuel cell temperature, pressure, and current density in the different operating conditions can enhance system performance. Additionally, it was found that most systems must be sized for top-of-climb, except for the ram air duct, which is constrained by take-off conditions. Incorporating a variable inlet design may eliminate ram air drag in cruise, although detailed drag analysis is recommended. Integration of a turbine and halving the rate-of-climb demonstrated that achieving a 1500 km range with a competitive payload (>3000 kg, 35 passengers) is feasible for retrofits, while reserving mass for non-modeled systems such as batteries. Moreover, projections for 2030 suggest that the performance of current turboprop aircraft could be matched by LH2FCE systems.
These findings highlight that LH2FCE propulsion is a viable and sustainable alternative for regional aviation, provided the reduction in payload is acceptable. With advancements in fuel cells, heat exchangers, electric motors, and liquid hydrogen storage, LH2FCE aircraft could achieve performance of current kerosene-powered turboprop aircraft by 2030.
Deployable Vortex Generators based on Shape Memory Alloy Actuation
Aerodynamic Design and Performance Analysis
This research addresses the challenge of improving aerodynamic performance by delaying flow separation using deployable vortex generators (VGs) actuated by shape memory alloys. These VGs combine passive and active elements to enhance mixing within the boundary layer, restoring momentum near the wall and minimizing separation in an Adverse Pressure Gradient (APG). A literature review identified a gap in understanding the dynamics of shape-adaptive VGs. Through Computational Fluid Dynamics (CFD) and experimental studies, the research optimized VG designs for aerodynamic efficiency. Results showed that counter-rotating vanes were most effective in controlling flow, while side load and yaw moment challenges necessitated robust actuation mechanisms. A novel design using rigid vane actuators and shape memory alloys for precise adjustments was proposed to address these issues. This study significantly advances the understanding and application of shape-adaptive VGs in flow control.
...
This research addresses the challenge of improving aerodynamic performance by delaying flow separation using deployable vortex generators (VGs) actuated by shape memory alloys. These VGs combine passive and active elements to enhance mixing within the boundary layer, restoring momentum near the wall and minimizing separation in an Adverse Pressure Gradient (APG). A literature review identified a gap in understanding the dynamics of shape-adaptive VGs. Through Computational Fluid Dynamics (CFD) and experimental studies, the research optimized VG designs for aerodynamic efficiency. Results showed that counter-rotating vanes were most effective in controlling flow, while side load and yaw moment challenges necessitated robust actuation mechanisms. A novel design using rigid vane actuators and shape memory alloys for precise adjustments was proposed to address these issues. This study significantly advances the understanding and application of shape-adaptive VGs in flow control.
An Experimental investigation into the drag performance of chevron-shaped protrusions in a turbulent channel flow
Experimental investigation into the drag and one dimensional velocity statistics
Contrary to conventional wisdom favouring smooth surfaces for drag reduction, the last decades have brought forth different textured surfaces showing drag reduction in turbulent boundary layers, of which riblets are the most studied.
Sirovich and Karlson (1997) introduced a different textured surface named chevron-shaped protrusions, indicating a drag reduction of 10% in turbulent channel flows.
Later studies questioned their efficacy as no credible reproduction of the results was obtained throughout the years.
The last credible reproduction study by Carrasco Grau et al. (2023) suggested that the reasons for this discrepancy could reside in the difference in model and test section size.
All reproduction studies were performed in facilities with external boundary layers instead of internal and with a covered area no longer than 0.8 meters, instead of the 8-meter-long channel flow which was fully covered in these chevrons from the original study.
This work will investigate just that, using a channel flow facility with dimensions more akin to that used by Sirovich and a model size of 2.4 meters.
The study assembles and characterises an improved channel flow facility with dimensions identical to those used by tay et al. (2011) at the National University of Singapore (NUS), measuring a total length of eight meters and a test section of 2.4 meters.
This facility used an array of 29 static pressure taps to determine the skin friction via the mean pressure gradient method, and utilised hot wire measurements in the midpoint of the test section for investigation of the flow mechanics of the chevrons.
The measurements performed on a flat plate were compared to the results obtained at the NUS to characterise the facility and validate its suitability for single-point drag measurements in the order of 5-10% increase or decrease.
Once this was confirmed the different configurations of chevron-shaped protrusions were investigated.
The key findings of this study are twofold.
Firstly, the developed channel flow facility at DUT proves proficient in generating canonical boundary layer profiles, enabling accurate skin friction measurements of textured surfaces.
The facility provides good hot-wire measurements without sensor vibrations capable of investigating the boundary layer characteristics in this facility.
Secondly, the study establishes that chevron-shaped protrusions are likely unsuited for reducing turbulent skin friction in turbulent channel flows.
Despite an inability to definitively disprove the working hypothesis, the observed increase in drag suggests that the technique's efficacy is not determined by the facility type and model size.
It is not considered worthwhile to conduct additional research into chevron-shaped protrusions as a potential technique for reducing drag in turbulent boundary layers.
However, the minimal drag penalty associated with this technique opens new possibilities for the application of this technique in the aviation sector, mainly in aiding with separation control and heat transfer. ...
Sirovich and Karlson (1997) introduced a different textured surface named chevron-shaped protrusions, indicating a drag reduction of 10% in turbulent channel flows.
Later studies questioned their efficacy as no credible reproduction of the results was obtained throughout the years.
The last credible reproduction study by Carrasco Grau et al. (2023) suggested that the reasons for this discrepancy could reside in the difference in model and test section size.
All reproduction studies were performed in facilities with external boundary layers instead of internal and with a covered area no longer than 0.8 meters, instead of the 8-meter-long channel flow which was fully covered in these chevrons from the original study.
This work will investigate just that, using a channel flow facility with dimensions more akin to that used by Sirovich and a model size of 2.4 meters.
The study assembles and characterises an improved channel flow facility with dimensions identical to those used by tay et al. (2011) at the National University of Singapore (NUS), measuring a total length of eight meters and a test section of 2.4 meters.
This facility used an array of 29 static pressure taps to determine the skin friction via the mean pressure gradient method, and utilised hot wire measurements in the midpoint of the test section for investigation of the flow mechanics of the chevrons.
The measurements performed on a flat plate were compared to the results obtained at the NUS to characterise the facility and validate its suitability for single-point drag measurements in the order of 5-10% increase or decrease.
Once this was confirmed the different configurations of chevron-shaped protrusions were investigated.
The key findings of this study are twofold.
Firstly, the developed channel flow facility at DUT proves proficient in generating canonical boundary layer profiles, enabling accurate skin friction measurements of textured surfaces.
The facility provides good hot-wire measurements without sensor vibrations capable of investigating the boundary layer characteristics in this facility.
Secondly, the study establishes that chevron-shaped protrusions are likely unsuited for reducing turbulent skin friction in turbulent channel flows.
Despite an inability to definitively disprove the working hypothesis, the observed increase in drag suggests that the technique's efficacy is not determined by the facility type and model size.
It is not considered worthwhile to conduct additional research into chevron-shaped protrusions as a potential technique for reducing drag in turbulent boundary layers.
However, the minimal drag penalty associated with this technique opens new possibilities for the application of this technique in the aviation sector, mainly in aiding with separation control and heat transfer. ...
Contrary to conventional wisdom favouring smooth surfaces for drag reduction, the last decades have brought forth different textured surfaces showing drag reduction in turbulent boundary layers, of which riblets are the most studied.
Sirovich and Karlson (1997) introduced a different textured surface named chevron-shaped protrusions, indicating a drag reduction of 10% in turbulent channel flows.
Later studies questioned their efficacy as no credible reproduction of the results was obtained throughout the years.
The last credible reproduction study by Carrasco Grau et al. (2023) suggested that the reasons for this discrepancy could reside in the difference in model and test section size.
All reproduction studies were performed in facilities with external boundary layers instead of internal and with a covered area no longer than 0.8 meters, instead of the 8-meter-long channel flow which was fully covered in these chevrons from the original study.
This work will investigate just that, using a channel flow facility with dimensions more akin to that used by Sirovich and a model size of 2.4 meters.
The study assembles and characterises an improved channel flow facility with dimensions identical to those used by tay et al. (2011) at the National University of Singapore (NUS), measuring a total length of eight meters and a test section of 2.4 meters.
This facility used an array of 29 static pressure taps to determine the skin friction via the mean pressure gradient method, and utilised hot wire measurements in the midpoint of the test section for investigation of the flow mechanics of the chevrons.
The measurements performed on a flat plate were compared to the results obtained at the NUS to characterise the facility and validate its suitability for single-point drag measurements in the order of 5-10% increase or decrease.
Once this was confirmed the different configurations of chevron-shaped protrusions were investigated.
The key findings of this study are twofold.
Firstly, the developed channel flow facility at DUT proves proficient in generating canonical boundary layer profiles, enabling accurate skin friction measurements of textured surfaces.
The facility provides good hot-wire measurements without sensor vibrations capable of investigating the boundary layer characteristics in this facility.
Secondly, the study establishes that chevron-shaped protrusions are likely unsuited for reducing turbulent skin friction in turbulent channel flows.
Despite an inability to definitively disprove the working hypothesis, the observed increase in drag suggests that the technique's efficacy is not determined by the facility type and model size.
It is not considered worthwhile to conduct additional research into chevron-shaped protrusions as a potential technique for reducing drag in turbulent boundary layers.
However, the minimal drag penalty associated with this technique opens new possibilities for the application of this technique in the aviation sector, mainly in aiding with separation control and heat transfer.
Sirovich and Karlson (1997) introduced a different textured surface named chevron-shaped protrusions, indicating a drag reduction of 10% in turbulent channel flows.
Later studies questioned their efficacy as no credible reproduction of the results was obtained throughout the years.
The last credible reproduction study by Carrasco Grau et al. (2023) suggested that the reasons for this discrepancy could reside in the difference in model and test section size.
All reproduction studies were performed in facilities with external boundary layers instead of internal and with a covered area no longer than 0.8 meters, instead of the 8-meter-long channel flow which was fully covered in these chevrons from the original study.
This work will investigate just that, using a channel flow facility with dimensions more akin to that used by Sirovich and a model size of 2.4 meters.
The study assembles and characterises an improved channel flow facility with dimensions identical to those used by tay et al. (2011) at the National University of Singapore (NUS), measuring a total length of eight meters and a test section of 2.4 meters.
This facility used an array of 29 static pressure taps to determine the skin friction via the mean pressure gradient method, and utilised hot wire measurements in the midpoint of the test section for investigation of the flow mechanics of the chevrons.
The measurements performed on a flat plate were compared to the results obtained at the NUS to characterise the facility and validate its suitability for single-point drag measurements in the order of 5-10% increase or decrease.
Once this was confirmed the different configurations of chevron-shaped protrusions were investigated.
The key findings of this study are twofold.
Firstly, the developed channel flow facility at DUT proves proficient in generating canonical boundary layer profiles, enabling accurate skin friction measurements of textured surfaces.
The facility provides good hot-wire measurements without sensor vibrations capable of investigating the boundary layer characteristics in this facility.
Secondly, the study establishes that chevron-shaped protrusions are likely unsuited for reducing turbulent skin friction in turbulent channel flows.
Despite an inability to definitively disprove the working hypothesis, the observed increase in drag suggests that the technique's efficacy is not determined by the facility type and model size.
It is not considered worthwhile to conduct additional research into chevron-shaped protrusions as a potential technique for reducing drag in turbulent boundary layers.
However, the minimal drag penalty associated with this technique opens new possibilities for the application of this technique in the aviation sector, mainly in aiding with separation control and heat transfer.
Master thesis
(2022)
-
A. Aditya, F.F.J. Schrijer, B.W. van Oudheusden, J. Bulut, F. Avallone, W.J. Baars
Supersonic flows cause thin panels to flutter, which is characterized by high-amplitude self-sustained oscillations, increasing the risk of fatigue failure. Flutter is known to be exacerbated when a shock wave impinges on the panel, creating a shock wave/boundary-layer interaction (SWBLI) which promotes separation of flow and leads to increased aerodynamic and thermal loading on the panel. This novel fluid-structure interaction (FSI), known as shock-induced panel flutter, poses a risk to the structural integrity of components used in high-speed aerial vehicles, such as rocket nozzles and supersonic engine inlets.
Due to the complex coupled interaction between flow and structural dynamics in flutter, experiments provide a better option to investigate the underlying physical mechanisms, rather than computational studies, which usually have to compromise between computational cost and resultant accuracy. The ST-15 supersonic wind tunnel facility at TU Delft is employed to conduct experimental measurements of shock-induced panel flutter, using high-speed Schlieren imaging to capture flow structures and stereographic Digital Image Correlation (DIC) to record panel displacements in separate campaigns. Tests are done at Mach 2 with fully-clamped thin panels, and the flutter response is excited using different shock strengths and impingement locations.
In both campaigns, accelerometers are used to measure spurious vibrations around the wind tunnel test section. This helps reveal the existence of external vibrations inherent to the facility, which are also found to drive the frequency of the panel flutter: at 760-770 Hz without an impinging shock, and 605-640 Hz with an impinging shock. The latter frequency is detected in both - shock motion and panel displacements – which establishes coupling between flow and structure despite measurements being non-simultaneous. Changing the shock impingement location does not have a significant effect on the degree of flow separation caused over the thin fluttering panel, which is always higher than the separation on a rigid plate at the same shock strength, thus proving that fluttering panels are not viable means of shock-induced separation control. A stronger impinging shock produces increased shock-induced flow separation but results in less energetic panel flutter, which is attributed to the higher post-shock pressure rise suppressing the panel. Flutter is found to be most energetic, and consequently, the panel is most susceptible to fatigue failure, when the shock impinges at 60% of the panel length. ...
Due to the complex coupled interaction between flow and structural dynamics in flutter, experiments provide a better option to investigate the underlying physical mechanisms, rather than computational studies, which usually have to compromise between computational cost and resultant accuracy. The ST-15 supersonic wind tunnel facility at TU Delft is employed to conduct experimental measurements of shock-induced panel flutter, using high-speed Schlieren imaging to capture flow structures and stereographic Digital Image Correlation (DIC) to record panel displacements in separate campaigns. Tests are done at Mach 2 with fully-clamped thin panels, and the flutter response is excited using different shock strengths and impingement locations.
In both campaigns, accelerometers are used to measure spurious vibrations around the wind tunnel test section. This helps reveal the existence of external vibrations inherent to the facility, which are also found to drive the frequency of the panel flutter: at 760-770 Hz without an impinging shock, and 605-640 Hz with an impinging shock. The latter frequency is detected in both - shock motion and panel displacements – which establishes coupling between flow and structure despite measurements being non-simultaneous. Changing the shock impingement location does not have a significant effect on the degree of flow separation caused over the thin fluttering panel, which is always higher than the separation on a rigid plate at the same shock strength, thus proving that fluttering panels are not viable means of shock-induced separation control. A stronger impinging shock produces increased shock-induced flow separation but results in less energetic panel flutter, which is attributed to the higher post-shock pressure rise suppressing the panel. Flutter is found to be most energetic, and consequently, the panel is most susceptible to fatigue failure, when the shock impinges at 60% of the panel length. ...
Supersonic flows cause thin panels to flutter, which is characterized by high-amplitude self-sustained oscillations, increasing the risk of fatigue failure. Flutter is known to be exacerbated when a shock wave impinges on the panel, creating a shock wave/boundary-layer interaction (SWBLI) which promotes separation of flow and leads to increased aerodynamic and thermal loading on the panel. This novel fluid-structure interaction (FSI), known as shock-induced panel flutter, poses a risk to the structural integrity of components used in high-speed aerial vehicles, such as rocket nozzles and supersonic engine inlets.
Due to the complex coupled interaction between flow and structural dynamics in flutter, experiments provide a better option to investigate the underlying physical mechanisms, rather than computational studies, which usually have to compromise between computational cost and resultant accuracy. The ST-15 supersonic wind tunnel facility at TU Delft is employed to conduct experimental measurements of shock-induced panel flutter, using high-speed Schlieren imaging to capture flow structures and stereographic Digital Image Correlation (DIC) to record panel displacements in separate campaigns. Tests are done at Mach 2 with fully-clamped thin panels, and the flutter response is excited using different shock strengths and impingement locations.
In both campaigns, accelerometers are used to measure spurious vibrations around the wind tunnel test section. This helps reveal the existence of external vibrations inherent to the facility, which are also found to drive the frequency of the panel flutter: at 760-770 Hz without an impinging shock, and 605-640 Hz with an impinging shock. The latter frequency is detected in both - shock motion and panel displacements – which establishes coupling between flow and structure despite measurements being non-simultaneous. Changing the shock impingement location does not have a significant effect on the degree of flow separation caused over the thin fluttering panel, which is always higher than the separation on a rigid plate at the same shock strength, thus proving that fluttering panels are not viable means of shock-induced separation control. A stronger impinging shock produces increased shock-induced flow separation but results in less energetic panel flutter, which is attributed to the higher post-shock pressure rise suppressing the panel. Flutter is found to be most energetic, and consequently, the panel is most susceptible to fatigue failure, when the shock impinges at 60% of the panel length.
Due to the complex coupled interaction between flow and structural dynamics in flutter, experiments provide a better option to investigate the underlying physical mechanisms, rather than computational studies, which usually have to compromise between computational cost and resultant accuracy. The ST-15 supersonic wind tunnel facility at TU Delft is employed to conduct experimental measurements of shock-induced panel flutter, using high-speed Schlieren imaging to capture flow structures and stereographic Digital Image Correlation (DIC) to record panel displacements in separate campaigns. Tests are done at Mach 2 with fully-clamped thin panels, and the flutter response is excited using different shock strengths and impingement locations.
In both campaigns, accelerometers are used to measure spurious vibrations around the wind tunnel test section. This helps reveal the existence of external vibrations inherent to the facility, which are also found to drive the frequency of the panel flutter: at 760-770 Hz without an impinging shock, and 605-640 Hz with an impinging shock. The latter frequency is detected in both - shock motion and panel displacements – which establishes coupling between flow and structure despite measurements being non-simultaneous. Changing the shock impingement location does not have a significant effect on the degree of flow separation caused over the thin fluttering panel, which is always higher than the separation on a rigid plate at the same shock strength, thus proving that fluttering panels are not viable means of shock-induced separation control. A stronger impinging shock produces increased shock-induced flow separation but results in less energetic panel flutter, which is attributed to the higher post-shock pressure rise suppressing the panel. Flutter is found to be most energetic, and consequently, the panel is most susceptible to fatigue failure, when the shock impinges at 60% of the panel length.
The turbulent boundary layer development under the influence of an air cavity is studied experimentally using planar PIV, with the aim of gaining insight and building upon the flow physics typically encountered in the application of air layer drag reduction. A detection technique based on correlation values is implemented to obtain an approximate shape of the air cavity and the location of the air-water interface. The technique was successful in identifying the maximum cavity thickness with sufficient accuracy. The leading and trailing edges of the cavity however, were harder to identify, the former owing to a limitation of the developed technique and the latter due to the dynamic nature of the flow and a slightly limited FOV. The ratio of the initial boundary layer thickness to the maximum thickness of the air cavity is 6.7, and as a consequence the boundary layer did not separate at the leeward side of the air cavity. The turbulent boundary layer is observed to feel the presence of the air cavity up to 8.5-9.5 cm upstream due to an adverse pressure gradient. Alternating streamwise pressure gradients are generated due to the curvature of the air cavity: from an adverse to favourable and back to adverse. Compared to solid bump studies in literature, additional perturbations due to a free-slip boundary condition and the unstable nature of air cavity increase the complexity of the current flow. The mean velocity profile and stresses are able to capture the effects of alternating streamwise pressure gradients and air injection, with variations mostly restricted to the inner region. Effects of streamline curvature in the outer region are found to be minimal, while potential effects of the free-slip condition were much harder to identify separately and further research would be needed to appropriately assess them. The mean velocity profile is found to deviate from the classic logarithmic behaviour at the apex of the air cavity, although the flow does not seem to relaminarise. Quadrant analysis shows differences in Reynolds stress producing events compared to the baseline turbulent boundary layer case hinting at possible alteration to coherent structuring of the turbulent boundary layer developing below the air cavity.
...
The turbulent boundary layer development under the influence of an air cavity is studied experimentally using planar PIV, with the aim of gaining insight and building upon the flow physics typically encountered in the application of air layer drag reduction. A detection technique based on correlation values is implemented to obtain an approximate shape of the air cavity and the location of the air-water interface. The technique was successful in identifying the maximum cavity thickness with sufficient accuracy. The leading and trailing edges of the cavity however, were harder to identify, the former owing to a limitation of the developed technique and the latter due to the dynamic nature of the flow and a slightly limited FOV. The ratio of the initial boundary layer thickness to the maximum thickness of the air cavity is 6.7, and as a consequence the boundary layer did not separate at the leeward side of the air cavity. The turbulent boundary layer is observed to feel the presence of the air cavity up to 8.5-9.5 cm upstream due to an adverse pressure gradient. Alternating streamwise pressure gradients are generated due to the curvature of the air cavity: from an adverse to favourable and back to adverse. Compared to solid bump studies in literature, additional perturbations due to a free-slip boundary condition and the unstable nature of air cavity increase the complexity of the current flow. The mean velocity profile and stresses are able to capture the effects of alternating streamwise pressure gradients and air injection, with variations mostly restricted to the inner region. Effects of streamline curvature in the outer region are found to be minimal, while potential effects of the free-slip condition were much harder to identify separately and further research would be needed to appropriately assess them. The mean velocity profile is found to deviate from the classic logarithmic behaviour at the apex of the air cavity, although the flow does not seem to relaminarise. Quadrant analysis shows differences in Reynolds stress producing events compared to the baseline turbulent boundary layer case hinting at possible alteration to coherent structuring of the turbulent boundary layer developing below the air cavity.
Master thesis
(2021)
-
Yu-Jui Lai, B.W. van Oudheusden, F.F.J. Schrijer, M. van Nesselrooij, R. Delfos, W.J. Baars
This research studies the fluid-structure interaction (FSI) of compliant surfaces in air flows, with an objective of finding possible turbulent viscous drag reduction. A compliant surface is a thin layer of viscoelastic material drawing inspiration from dolphin epidermis, which was thought to have drag-reducing capabilities by Gray (1936). Research into the drag reduction capabilities of compliant surfaces has been long going for more than five decades, starting with Kramer (1960), yet no firm conclusion has been reached. In addition, most of the experimental research has been focused on water flows, with air flows regarded as incapable of inertially forcing a compliant surface to deform. This research attempts to disprove this assumption by applying the proper inertial scaling to the FSI between the compliant surface and air flows. Compliant surfaces are characterised by the stiffness and thickness, which are presumed to be of first-order influence on the FSI. A parameter sweep of these compliant surface properties was conducted by drag delta measurements and flow visualisation by planar particle image velocimetry (PIV) in the M-Tunnel at the Low Speed Laboratory (LSL) Drag delta results confirm the possibility of turbulent viscous drag reduction by compliant surfaces, with a measured drag delta of -3.43 %. This result is further supported by the decrease in 1D turbulence intensity at the test plate's trailing edge from hot-wire measurements, and a smaller decay of the shape factor H from PIV. Quadrant analysis of the PIV data found evidence of a reduction in combined Q2 and Q4 events, further supporting the drag delta measured. Correlation between the compliant surface's viscoelastic properties and the drag delta found a negative correlation between the magnitude of complex shear modulus and a positive correlation between the loss tangent and drag reduction. The high loss tangent for the drag-reducing compliant surface and its subtle positive correlation with the drag reduction indicates that the viscoelasticity might have a greater influence on the FSI than expected.
...
This research studies the fluid-structure interaction (FSI) of compliant surfaces in air flows, with an objective of finding possible turbulent viscous drag reduction. A compliant surface is a thin layer of viscoelastic material drawing inspiration from dolphin epidermis, which was thought to have drag-reducing capabilities by Gray (1936). Research into the drag reduction capabilities of compliant surfaces has been long going for more than five decades, starting with Kramer (1960), yet no firm conclusion has been reached. In addition, most of the experimental research has been focused on water flows, with air flows regarded as incapable of inertially forcing a compliant surface to deform. This research attempts to disprove this assumption by applying the proper inertial scaling to the FSI between the compliant surface and air flows. Compliant surfaces are characterised by the stiffness and thickness, which are presumed to be of first-order influence on the FSI. A parameter sweep of these compliant surface properties was conducted by drag delta measurements and flow visualisation by planar particle image velocimetry (PIV) in the M-Tunnel at the Low Speed Laboratory (LSL) Drag delta results confirm the possibility of turbulent viscous drag reduction by compliant surfaces, with a measured drag delta of -3.43 %. This result is further supported by the decrease in 1D turbulence intensity at the test plate's trailing edge from hot-wire measurements, and a smaller decay of the shape factor H from PIV. Quadrant analysis of the PIV data found evidence of a reduction in combined Q2 and Q4 events, further supporting the drag delta measured. Correlation between the compliant surface's viscoelastic properties and the drag delta found a negative correlation between the magnitude of complex shear modulus and a positive correlation between the loss tangent and drag reduction. The high loss tangent for the drag-reducing compliant surface and its subtle positive correlation with the drag reduction indicates that the viscoelasticity might have a greater influence on the FSI than expected.
Attenuating Jet Installation Noise with a Lobed Nozzle
Master of Science Thesis
Over the last few decades, the bypass ratio and nacelle diameter of modern turbofans have been increasing to achieve higher efficiencies. This trend has forced the underwing-mounted engines to be coupled more closely to the wing surface and the flap system, which significantly enhances the low-frequency noise, known as jet installation noise (JIN).
In this project, lobed nozzles are proposed and tested as a passive flow control technique for the reduction of the JIN. Measurements were performed using hot-wire anemometry and a linear array of microphones to investigate the effects of lobed nozzles on the near-field structure of a jet. The lobed nozzles were found to alter the convection velocities of the pressure fluctuations and the transition to turbulence of the jet. Moreover, far-field pressure measurements were performed inside the anechoic wind tunnel to examine the effects of the near-field structure on the perceived far-field noise. These findings can be used to assess the effectiveness of lobed nozzles for the reduction of the JIN in specific jet installation configurations.
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In this project, lobed nozzles are proposed and tested as a passive flow control technique for the reduction of the JIN. Measurements were performed using hot-wire anemometry and a linear array of microphones to investigate the effects of lobed nozzles on the near-field structure of a jet. The lobed nozzles were found to alter the convection velocities of the pressure fluctuations and the transition to turbulence of the jet. Moreover, far-field pressure measurements were performed inside the anechoic wind tunnel to examine the effects of the near-field structure on the perceived far-field noise. These findings can be used to assess the effectiveness of lobed nozzles for the reduction of the JIN in specific jet installation configurations.
...
Over the last few decades, the bypass ratio and nacelle diameter of modern turbofans have been increasing to achieve higher efficiencies. This trend has forced the underwing-mounted engines to be coupled more closely to the wing surface and the flap system, which significantly enhances the low-frequency noise, known as jet installation noise (JIN).
In this project, lobed nozzles are proposed and tested as a passive flow control technique for the reduction of the JIN. Measurements were performed using hot-wire anemometry and a linear array of microphones to investigate the effects of lobed nozzles on the near-field structure of a jet. The lobed nozzles were found to alter the convection velocities of the pressure fluctuations and the transition to turbulence of the jet. Moreover, far-field pressure measurements were performed inside the anechoic wind tunnel to examine the effects of the near-field structure on the perceived far-field noise. These findings can be used to assess the effectiveness of lobed nozzles for the reduction of the JIN in specific jet installation configurations.
In this project, lobed nozzles are proposed and tested as a passive flow control technique for the reduction of the JIN. Measurements were performed using hot-wire anemometry and a linear array of microphones to investigate the effects of lobed nozzles on the near-field structure of a jet. The lobed nozzles were found to alter the convection velocities of the pressure fluctuations and the transition to turbulence of the jet. Moreover, far-field pressure measurements were performed inside the anechoic wind tunnel to examine the effects of the near-field structure on the perceived far-field noise. These findings can be used to assess the effectiveness of lobed nozzles for the reduction of the JIN in specific jet installation configurations.