Shock/boundary-layer interaction control with vortex generators and passive cavity

Tidal energy can play an important role in the Net Zero transition. Increasing tidal turbine performance through innovation is crucial if the cost of tidal energy is to become competitive compared to other sources of energy. The present investigation is a proof-of-concept study for the application of Vortex Generators (VGs) on tidal turbines in view of increasing their performance. The more mature wind energy industry uses passive VGs either as a retrofit or in the blade design process to reduce separation at the inboard part of wind turbine blades. Tidal turbine blades also experience flow separation and here we examine whether passive vane VGs can be used to reduce or suppress that separated flow. First, a wind tunnel investigation is performed to assess the performance of VGs on a 20% thick profile from the blade. Then, the VG effect on the 2D-profile is modelled in a Reynolds Averaged Navier-Stokes in-house solver. Results show that low profile VGs, i.e. VGs shorter than the local boundary layer, can increase the performance of the blade profile and successfully reduce flow separation. The VG effect on blade performance is examined in model scale and in full-size. VGs successfully suppress separation in both cases and it is shown that full-size information should be used for the placement of VGs. A maximum power coefficient increase of 1.05% is observed at a tip speed ratio of λ=3. The present proof-of-concept study demonstrates for the first time the potential of passive VGs to be included either in the design process of a tidal turbine blade or as a retrofit solution.

Experiments have been performed in a blowdown supersonic wind tunnel to investigate the effect of arrays of sub-boundary layer vortex generators placed upstream of a normal shock/ boundary layer interaction. The investigation makes use of a recovery shock wave and the naturally grown turbulent boundary layer on the wind tunnel floor. Experiments were performed at Mach numbers of 1.5 and 1.3 and a freestream Reynolds number of 28 × 106. Two types of vortex generators were investigated - wedge-shaped and arrays of counter-rotating vanes. It was found that at Mach 1.5 the vane-type VGs eliminated and the wedge-type VGs greatly reduced the separation bubble under the shock. When placed in the supersonic part of the flow both VGs caused a wave pattern consisting of a shock, re-expansion and shock. The re-expansion and double shocks are undesirable features since they equate to increased total pressure losses and hence increased -wave drag. Furthermore there are indications that the vortex intensity is reduced by the normal shock/ boundary layer interaction. When the shock was located directly over the VGs there was no re-expansion present, but the 'damping' effect of the shock on the vortex persisted. It appears that the vortices produced by the wedge-shaped VGs lift off the surface more rapidly. Similar results were observed at Mach 1.3, where the flow was unseparated.

Shock-boundary layer interaction control with low-profile vortex generators and passive cavity

In this paper, some preliminary experiments and the calculations were performed to clarify the flow field, in which the rods surrounded by a porous cavity were normally inserted into the main supersonic flow. As a result, it is found that the starting shock wave severely interacts with the rods, the bow shock wave, its reflections, and the porous wall, which are numerically well predicted under some conditions. In case of the single rod, the main supersonic flow is sucked in front of the rod and blown downstream of the rod. Furthermore, the suction mass flux through the porous holes depends on the number and the locations of the rods. On the other hand, the three rods aligned parallel to the flow make favorable suction area downstream of the rods and also generate as less total pressure loss as that for a single rod. In addition, the spanwised three rods cause the strongest bow shock wave making the large mass flux through the porous wall and also large total pressure loss downstream of the rods. The calculation also suggests that interaction between the bow shock wave and boundary layer along the upper wall results in the boundary layer separation.

Flow separation in overexpanded single expansion ramp nozzles (SERN) involves complex phenomena, such as shock waves, expansion waves, turbulent boundary layers, and shear layers. Computational fluid dynamics plays a crucial role in studying unsteady flow behaviour in supersonic nozzles, allowing for an investigation into the dynamic flow field characteristics. However, the application of OpenFOAM as a numerical tool for studying SERN in the field of compressible flows, particularly in the overexpansion state where the flow field characteristics are more complex, has received relatively less attention. In this study, the flow field characteristics of an overexpanded SERN under different turbulence models are investigated through a combination of experiments and numerical calculations. The qualitative and quantitative predictive performance of two compressible flow solvers in OpenFOAM, namely, rhoCentralFOAM and sonicFOAM, are compared in terms of flow separation pattern and separation pattern transitions within the overexpanded SERN. The ability of rhoCentralFOAM and sonicFOAM to accurately predict complex flow states is evaluated. Results indicate that the numerical simulations conducted using rhoCentralFOAM and sonicFOAM successfully capture flow separation, separated shock waves, separated bubbles and shear layers for two types of restricted shock separation patterns at the same nozzle pressure ratio (NPR), demonstrating agreement with experimental results. However, sonicFOAM initiates the transition in the separation pattern 0.0773 NPR earlier than rhoCentralFOAM during the whole separation pattern transition process of the SERN. The transition process in sonicFOAM lasts longer and exhibits a greater variation in NPR. SonicFOAM fails to accurately predict certain aspects, such as the pressure rise after the separation bubble, the reattachment shock wave, and tends to overestimat the length of the separation shock length. Consequently, sonicFOAM cannot be recommended as a suitable solver for accurately capturing the separation pattern of an overexpanded nozzle.

When non-equilibrium condensation occurs in a supersonic flow field, the flow is affected by the latent heat released. In the present study, in order to control the transonic flow field with shock wave, a condensing flow was produced by an expansion of moist air on a circular bump model and shock waves were occurred in the supersonic parts of the fields. Furthermore, the additional passive technique of shock / boundary layer interaction using the porous wall with a cavity underneath was adopted in this flow field. The effects of these methods on the shock wave characteristics were investigated numerically and experimentally. The result obtained showed that the total pressure loss in the flow fields might be effectively reduced by the suitable combination between non-equilibrium condensation and the position of porous wall.

A detailed numerical simulation is presented to investigate the new de-swirling methods and their effect on the mixing mechanisms of a turbofan mixer with 12 lobes. The numerical simulation employed a commercial solver, ANSYS CFX, using k-ω SST model. The core-to-bypass temperature ratio and pressure ratio were set to 2.59, and 0.97 respectively, giving the Mach number of 0.66 and bypass ratio of 2.65 at mixing nozzle outlet. The inlet swirl typically accelerates the jet-flow mixing by enhancing the vortices intensity and interaction, but leakage swirling flow can cause a three-dimensional separation bubble and the recirculation zone resulting in the dramatic increasing the total pressure loss and thrust loss. Removal of the leakage swirling flow between the lobes’ trough and centre-body was the key to limit the negative influence of inlet swirl. Two IGV design were investigated, DS1 and DS2. DS1 was installed at the upstream of the lobed mixer, could remove the negative effect of inlet swirl properly, but also inhibited the active role of the inlet swirl. The total pressure and thrust loss reduced by 0.31% and 3.8%, respectively, but the mixing efficiency also decreased by 1.72%. DS2, an integrated strut with the lobed mixer design, not only ensured the structure strength of the lobed mixer, but also reduced the length and weight of the exhaust system. This method suppressed the flow separation bubble on centre-body to some extent, and eliminated the recirculation zone downstream of the cenrebody, resulting in the total pressure loss decrease of 0.31% and thrust gain of 3.63%. On the other hand, the method DS2 also made full use of the inlet swirl to enhance the jet-flow mixing, resulting in the mixing efficiency increased 1.54% compared with that of the DS1 case. Under the off-design conditions with the incidence angle of ±10°, the aerodynamic performance of the DS2 cases didn’t changed too much such as the DS1 cases.

Numerical studies of shock wave interactions with a supersonic turbulent boundary layer in compression corner:Turning angle effects

Secondary flow effects like the corner stall between the wall and the vane in a compressor stage are responsible for a large part of total pressure losses. An extensive experimental study of flow control in a highly loaded compressor cascade was performed in order to decrease the separation and reduce the losses by means of vortex generators. The vortex generators were attached at the surface of the cascade side walls. These flow control devices produce strong vortices, which enhance the mixing between the main flow and the decelerated boundary layer at the side wall. Thus, the corner flow separation and the total pressure losses could be reduced. The experiments were carried out with a compressor cascade at a high-speed test facility at the DLR in Berlin at minimum loss (design point) and off-design of the cascade at Reynolds numbers up to Re = 0.6 × 106 (based on 40 mm chord) and Mach numbers up to M = 0.7. The cascade consisted of five vanes. The blade profiles are comparable to the hub section of the stator vanes used in the transonic compressor test rig running at Technische Universita¨t Darmstadt. In the range between −2° and +4° angle of incidence the total pressure losses of the cascade could be reduced up to 4.6% by means of vortex generators, whereas the static pressure rise was not influenced. Based on the results of the cascade measurements, the vortex generators were applied in front of the stator row of the single stage axial compressor at Technische Universita¨t Darmstadt. A numerical simulation of the compressor flow provided an indication for the adjustment of the vortex generators at the hub and casing. In the experiments the pressure rise and the efficiency of the axial compressor was measured and it could be shown that vortex generators partially improve the efficiency.

Flow control is required to suppress aero-optical effects for optical windows with backward-facing steps. This study investigates the impact of vortex generator (VG) spanwise width on vortical breakdown and aero-optical distortions in a Mach 3 flow field over a backward-facing step, using the nano-tracer planar laser scattering technique. Vortex evolution is quantified through fractal dimension analysis and two-point spatial correlation, while aero-optical effects are evaluated via gray-density calibration and ray-tracing methodology. Experimental results indicate that VG with larger spanwise widths facilitate accelerated vortical structure breakdown. However, excessive spanwise widths of VGs lead to a reduction in the global fractal dimension of the flow field and an increase in vortex size. Enhanced interaction between VG wakes and lower-wall turbulent boundary layers (TBLs) is observed for wider VGs. For spanwise widths of 24 and 18 mm, wake-induced TBL vortex size initially decreases before increasing along the streamwise direction. Time-averaged density fields further reveal a closer proximity of wake structures to the lower-wall TBL with increased spanwise widths of VGs. Optical path difference (OPD) in the computational domain initially increases, then decreases, and eventually stabilizes due to vortex breakdown and TBL/wake development. Notably, the 18 mm VG configuration exhibits a higher root mean square deviation of the OPD than other cases. This parametric study offers insights into optimizing optical window flow control strategies to mitigate aero-optical distortions.

Experimental and numerical simulation of the attenuation effect of blast shock waves in tunnels at different altitudes

Vortex generators (VGs) are applied to a two-stage highly-loaded transonic fan using the full blended blade and endwall (F-BBEW) method to assess their effects on fan performance. In the last stage stator, VGs were arranged on the hub and shroud within the blade passages. Numerical simulations are conducted to explore the effects of VGs on performance of the highly loaded transonic fan and the VGs action mechanism. The results demonstrate that the implementation of VGs led to an increase in fan aerodynamic performance, improving the fan pressure ratio margin by 2.25% while maintaining efficiency. The total pressure loss of the last stage stator decreased, with the loss coefficient reduced from 5.33% to 4.58% at 1.15 times the design pressure ratio. Furthermore, the VGs improve the flow field of the last stage stator. While the VGs on the hub suppressed local secondary flows, they do not provide significant benefits due to interaction with shock waves. In contrast, the VGs on the shroud reduce secondary flows and mitigate separation induced by upstream VGs, improving corner flow and overall flow quality of the last stage stator.

Effect of the flow rate of injected air on the temperature of porous wall in the supersonic flow

The performance of a compressor cascade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as the corner separation between the wall and the vane. An extensive experimental study of vortex generator application in a highly loaded compressor cascade was performed in order to control these effects and enhance the aerodynamic performance. The results of the study will be used in future projects as a basis for parameterization in the design and optimization process for compressors in order to develop novel nonaxisymmetric endwalls as well as for blade modifications. The study includes the investigation of two vortex generator types with different geometrical forms and their application on several positions in the compressor cascade. The investigation includes a detailed description of the secondary flow effects in the compressor cascade, which is based on numerical and experimental results. This gives the basis for a specific approach of influencing the cascade flow by means of vortex generators. Depending on the vortex generator type and position, there is an impact on the end wall cross flow, the development of the horse shoe vortex at the leading edge of the vane, and the extent of the corner separation achieved by improved mixing within the boundary layer. The experiments were carried out on a compressor cascade at a high-speed test facility at DLR in Berlin at minimum loss (design point) and off-design of the cascade at Reynolds numbers up to Re = 0.6 × 106 (based on 40-mm chord) and Mach numbers up to M = 0.7. At the cascade design point, the total pressure losses could be reduced by up to 9% with the vortex generator configuration, whereas the static pressure rise was nearly unaffected. Furthermore, the cascade deflection could be influenced considerably by vortex generators and also an enhancement of the cascade stall range could be achieved. All these results will be presented and discussed with respect to secondary flow mechanisms. Finally, the general application of vortex generators in axial compressors will be discussed.

An experimental study is conducted on the interaction of a swept shock wave with a turbulent boundary layer. The shock wave is generated by a sharp unswept fin in a Mach 2 flow, where the strength of the interaction is varied from weak to moderate by changing the angle of attack of the fin from 10 to 15 deg, which corresponds to a normal Mach number of 1.3 and 1.4, respectively. Surface oil-flow visualization is used to study the topography of the interaction where surface flow features such as the upstream influence line and the separation lines are identified. By taking advantage of the quasi-conical symmetry of the flowfield, velocity field measurements are acquired in the conical reference frame for the interaction at the moderate interaction strength of at two locations from the fin apex. Flowfield features such as the -shock structure, slip line, and the separation bubble with the reverse flow are clearly visible in the velocity field. These results are also examined in the spherical coordinate frame, and good agreement in the spatial location of the critical features is found, providing direct quantitative, experimental evidence of quasi-conical symmetry of the flowfield above the surface. An examination of the velocity field downstream of the rear-foot of the -shock shows a region (a “streamtube,” bounded on one side by the slip line emanating from the triple point) where the flow accelerates to transonic and supersonic speeds. This flow eventually turns toward and impinges upon the flat plate, a phenomenon referred to as an “impinging jet” in the literature. It is believed to be the principal cause of the high mean, unsteady pressures, very high heating, and skin friction coefficients that the present measurements confirm. An examination of the turbulence kinetic energy reveals a significant increase in amplitude in the separated flow region that is bounded by the flattened vortex under the -shock.