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Understanding shock wave behavior in supersonic flow environments is critical for optimizing the aerodynamic performance of turbomachinery components. This study introduces a novel transonic linear cascade design, focusing on advanced blade manufacturing and experimental validation. Blades were 3D-printed using Inconel 625, enabling tight control over the geometry and surface quality, which were verified through extensive dimensional accuracy assessments and surface finish quality checks using coordinate measuring machines (CMMs). Numerical simulations were performed using Ansys CFX with an implicit pressure-based solver and high-order numerical schemes to accurately model the shock wave phenomena. To validate the simulations, experimental tests were conducted using Schlieren visualization, ensuring high fidelity in capturing the shock wave dynamics. A custom-designed test rig was commissioned to replicate the specific requirements of the cascade, enabling stable and repeatable testing conditions. Experiments were conducted at three different inlet pressures (0.7-bar, 0.8-bar, and 0.9-bar gauges) at a constant temperature of 21 °C. Results indicated that the shock wave intensity and position are highly sensitive to the inlet pressure, with higher pressures producing more intense and extensive shock waves. While the numerical simulations aligned broadly with the experimental observations, discrepancies at finer flow scales suggest the need for the further refinement of the computational models to capture detailed flow phenomena accurately. [ABSTRACT FROM AUTHOR]

Understanding shock wave behavior in supersonic flow environments is critical for optimizing the aerodynamic performance of turbomachinery components. This study introduces a novel transonic linear cascade design, focusing on advanced blade manufacturing and experimental validation. Blades were 3D-printed using Inconel 625, enabling tight control over the geometry and surface quality, which were verified through extensive dimensional accuracy assessments and surface finish quality checks using coordinate measuring machines (CMMs). Numerical simulations were performed using Ansys CFX with an implicit pressure-based solver and high-order numerical schemes to accurately model the shock wave phenomena. To validate the simulations, experimental tests were conducted using Schlieren visualization, ensuring high fidelity in capturing the shock wave dynamics. A custom-designed test rig was commissioned to replicate the specific requirements of the cascade, enabling stable and repeatable testing conditions. Experiments were conducted at three different inlet pressures (0.7-bar, 0.8-bar, and 0.9-bar gauges) at a constant temperature of 21 °C. Results indicated that the shock wave intensity and position are highly sensitive to the inlet pressure, with higher pressures producing more intense and extensive shock waves. While the numerical simulations aligned broadly with the experimental observations, discrepancies at finer flow scales suggest the need for the further refinement of the computational models to capture detailed flow phenomena accurately.

The aerodynamics of a high-speed low-pressure turbine (LPT) cascade were investigated under steady and unsteady inlet flows. The tests were performed at outlet Mach (M) and Reynolds numbers ( R e ) of 0.90 and 70k, respectively. Unsteady wakes were simulated by means of a wake generator equipped with bars. A bar reduced frequency ( f + ) of ∼0.95 was used for the unsteady case. The inlet flow field was characterized in terms of the total pressure profile and incidence. The blade aerodynamics at midspan and the secondary flow region were investigated by means of pneumatic taps and hot-film sensors. The latter provided a novel view into the impact of the secondary flows on the heat transfer topology on the blade suction side (SS). The cascade performance was quantified in terms of the outlet flow angle and losses by means of a directional multi-hole probe. The results report the phase-averaged impact of unsteady wakes on the secondary flow structures in an open test case high-speed LPT geometry. [ABSTRACT FROM AUTHOR]

The influence of unsteady wakes incoming from the upstream stages is of high relevance in modern high-speed, low-pressure turbines (LPT) operating at transonic exit Mach numbers and low Reynolds numbers for their potential to trigger transition and influence the separation of the boundary layer on the blade suction side. The aim of this paper is the experimental characterization of the influence of incoming wakes on the 2D aerodynamics of a high-speed LPT cascade operating at a low Reynolds number and transonic exit Mach number. A detailed analysis of the status of the flow along the blade under investigation and its impact on the profile loss are presented for a range of Mach numbers from 0.70 to 0.95 and Reynolds numbers from 70k to 120k under steady and unsteady inflow conditions. Tests were conducted at on- and off-design engine realistic conditions in the VKI S-1/C wind tunnel on the SPLEEN C1 transonic cascade. The wakes incoming from an upstream blade row have been replicated using a set of rotating bars, which shed wakes at an engine-representative reduced frequency ( f + = 0.95 ) and flow coefficient ( Φ = 0.80 ). A set of densely instrumented traversable blades were used to sample the surface pressure distributions. The development of the boundary layers along the blade suction side is examined through quasi-wall shear stress obtained with surface-mounted hot-film sensors. Wake traverses were carried out downstream of the cascade with a miniaturized L-shaped five-hole probe to characterize the blade losses. The introduction of periodic incoming wakes promotes variations in the flow topology over the blade. The effect on the suction side separation bubble is shown to depend on the exit flow conditions. At low Mach numbers, the incoming wakes determine a reduction in the size of the bubble; in contrast, this effect is not registered as the exit Mach number increases. Consistently, a high dependence of the unsteady wake effect on the profile loss on the exit Reynolds and Mach numbers is demonstrated. [ABSTRACT FROM AUTHOR]

Shrouded wind and hydrokinetic turbines are being widely studied nowadays. Although actuator disk studies have concluded that the addition of a diffuser around a rotor can improve the Power Coefficient, not many designs have realized those benefits when compared to standard high performance bare turbines, as shown by Nunes et al. (2020) [1]. One reason for low power coefficient on several designs is massive flow separation on the hub surface due to high adverse pressure gradient inside the diffuser, resulting on low mass flow capture and, hence, poor performance. This work presents a novel design methodology for shrouded rotors, which takes into consideration the influence of the entire linear cascade on each annular section. Also, the blade root is left unloaded to guarantee that no boundary layer separation occurs on the hub surface by allowing a layer of energized fluid to bypass the rotor. A turbine modeled by this method has been numerically studied and is shown herein to deliver a peak power coefficient of 0.415 normalized by the diffuser's largest cross sectional area. [Display omitted] [ABSTRACT FROM AUTHOR]

Experimental and numerical studies of a linear high-loaded turbine cascade with a dual−cavity tip structure are presented in this paper. The experimental conditions contained an increase in the outlet Mach number from 0.42 to 0.92, a change in the incidence angle from −15° to 15° and an increase in the relative clearance size from 0.36% to 1.4%. The ability of the dual−cavity tip to control leakage losses and vortices is assessed using the total pressure coefficient and the Q-criterion. This research indicates that the leakage vortex interacts strongly with the passage vortex, and the change in working conditions affects the balance between the two vortices and thus the flow field structure. The experimental and numerical results prove that the dual−cavity tip can reduce losses in all operating conditions, with the best control effect reduced by 0.025 in a large clearance size condition. In addition, the leakage control effect of the blade tip structure is more influenced by the incoming flow angle and clearance size than the Mach number. [ABSTRACT FROM AUTHOR]

Recent works have already demonstrated that placing a crescent-shaped block upstream of a cylindrical hole could enhance the cooling performance of flat-plate films. The flow and cooling performance of the crescent-shaped block applied over the pressure and suction sides of the blade is investigated in this article. The Reynolds-averaged Navier-Stokes equations are solved with the Shear Stress Transport model for turbulence closure. Two optimized blocks are obtained from the flat-plate film cooling in our previous work, and two positions on the pressure and suction sides are tested. The blowing ratio varies from 0.5 to 2.0. The results show that when the block is applied on the blade surface, it yields a different cooling performance compared with the flat plate due to different geometry curvature and pressure gradient. The cooling performance on the suction side is slightly higher than on that on the pressure side, while the aerodynamic loss on the suction side is much higher. For the different blocks, the qualitative change of cooling performance vs. blowing ratios held on turbine blades is quite close to that of flat plates. The optimized smaller block in the flat plate provides better cooling performance at lower blowing ratios, while the larger block is superior when the blowing ratios are higher. [ABSTRACT FROM AUTHOR]

The aerodynamics of a high-speed low-pressure turbine (LPT) cascade were investigated under steady and unsteady inlet flows. The tests were performed at outlet Mach (M) and Reynolds numbers (Re) of 0.90 and 70k, respectively. Unsteady wakes were simulated by means of a wake generator equipped with bars. A bar reduced frequency (f+) of ∼0.95 was used for the unsteady case. The inlet flow field was characterized in terms of the total pressure profile and incidence. The blade aerodynamics at midspan and the secondary flow region were investigated by means of pneumatic taps and hot-film sensors. The latter provided a novel view into the impact of the secondary flows on the heat transfer topology on the blade suction side (SS). The cascade performance was quantified in terms of the outlet flow angle and losses by means of a directional multi-hole probe. The results report the phase-averaged impact of unsteady wakes on the secondary flow structures in an open test case high-speed LPT geometry.

The influence of unsteady wakes incoming from the upstream stages is of high relevance in modern high-speed, low-pressure turbines (LPT) operating at transonic exit Mach numbers and low Reynolds numbers for their potential to trigger transition and influence the separation of the boundary layer on the blade suction side. The aim of this paper is the experimental characterization of the influence of incoming wakes on the 2D aerodynamics of a high-speed LPT cascade operating at a low Reynolds number and transonic exit Mach number. A detailed analysis of the status of the flow along the blade under investigation and its impact on the profile loss are presented for a range of Mach numbers from 0.70 to 0.95 and Reynolds numbers from 70k to 120k under steady and unsteady inflow conditions. Tests were conducted at on- and off-design engine realistic conditions in the VKI S-1/C wind tunnel on the SPLEEN C1 transonic cascade. The wakes incoming from an upstream blade row have been replicated using a set of rotating bars, which shed wakes at an engine-representative reduced frequency (f+=0.95) and flow coefficient (Φ=0.80). A set of densely instrumented traversable blades were used to sample the surface pressure distributions. The development of the boundary layers along the blade suction side is examined through quasi-wall shear stress obtained with surface-mounted hot-film sensors. Wake traverses were carried out downstream of the cascade with a miniaturized L-shaped five-hole probe to characterize the blade losses. The introduction of periodic incoming wakes promotes variations in the flow topology over the blade. The effect on the suction side separation bubble is shown to depend on the exit flow conditions. At low Mach numbers, the incoming wakes determine a reduction in the size of the bubble; in contrast, this effect is not registered as the exit Mach number increases. Consistently, a high dependence of the unsteady wake effect on the profile loss on the exit Reynolds and Mach numbers is demonstrated.

Modern aeroengine designs strive for peak specific fuel and thermal efficiency. To achieve these goals, engines have more highly loaded compressor stages, thinner aerofoils, and blended titanium integrated disks (blisks) to reduce weight. These configurations promote the occurrence of aeroelastic phenomena such as flutter. Two important parameters known to influence flutter stability are the reduced frequency and the ratio of plunge and pitch components in a combined flap mode shape. These are used as design criteria in the engine development process. However, the limit of these criteria is not fully understood. The following research aims to bridge the gap between semi-analytical models and modern compressors by systematically investigating the flutter stability of a linear compressor cascade. This paper introduces the plunge-to-pitch incidence ratio, which is defined as a function of reduced frequency and pitch axis setback for a first flap (1F) mode shape. Using numerical simulations, in addition to experimental validation, aerodynamic damping is computed for many modes to build stability maps. The results confirm the importance of these two parameters in compressor aeroelastic stability as well as demonstrate the significance of the plunge-to-pitch incidence ratio for predicting the flutter limit. [ABSTRACT FROM AUTHOR]

Experimental and numerical studies of a linear high-loaded turbine cascade with a dual−cavity tip structure are presented in this paper. The experimental conditions contained an increase in the outlet Mach number from 0.42 to 0.92, a change in the incidence angle from −15° to 15° and an increase in the relative clearance size from 0.36% to 1.4%. The ability of the dual−cavity tip to control leakage losses and vortices is assessed using the total pressure coefficient and the Q-criterion. This research indicates that the leakage vortex interacts strongly with the passage vortex, and the change in working conditions affects the balance between the two vortices and thus the flow field structure. The experimental and numerical results prove that the dual−cavity tip can reduce losses in all operating conditions, with the best control effect reduced by 0.025 in a large clearance size condition. In addition, the leakage control effect of the blade tip structure is more influenced by the incoming flow angle and clearance size than the Mach number.

Recent works have already demonstrated that placing a crescent-shaped block upstream of a cylindrical hole could enhance the cooling performance of flat-plate films. The flow and cooling performance of the crescent-shaped block applied over the pressure and suction sides of the blade is investigated in this article. The Reynolds-averaged Navier-Stokes equations are solved with the Shear Stress Transport model for turbulence closure. Two optimized blocks are obtained from the flat-plate film cooling in our previous work, and two positions on the pressure and suction sides are tested. The blowing ratio varies from 0.5 to 2.0. The results show that when the block is applied on the blade surface, it yields a different cooling performance compared with the flat plate due to different geometry curvature and pressure gradient. The cooling performance on the suction side is slightly higher than on that on the pressure side, while the aerodynamic loss on the suction side is much higher. For the different blocks, the qualitative change of cooling performance vs. blowing ratios held on turbine blades is quite close to that of flat plates. The optimized smaller block in the flat plate provides better cooling performance at lower blowing ratios, while the larger block is superior when the blowing ratios are higher.

The high-speed Over-Tip-Leakage (OTL) flow has a significant impact on the aerodynamic performance of the High-Pressure Turbine (HPT) passage and generates high thermal load for the blade tip. Different tip sealing and cooling design strategies are applied to reduce the OTL loss and help turbine survive in a high temperature environment. High-speed linear cascade experimental rigs play an important role in understanding the flow physics and evaluating their performance. Multiple blades and passages are often required to maintain a reasonable flow periodicity. To match the engine representative Reynolds number and Mach number, a high-speed multi-passage cascade design inevitably demands more compressed air supply. A very large amount of heating power is also required if the engine condition wall-to-gas temperature ratio needs to be matched. In this study, a simplified 2-Passage linear cascade rig for high-speed tip heat transfer research was developed. Both the design method and the rig performance are presented. Different from existing design method to match two-dimensional blade loading, this study shows there are other design flexibilities, such as assist blade tip gap, tailboard adjustment, and profiling adjustment, to match the periodic three-dimensional OTL flow structures. The design method was validated by experimental effort. High resolution tip heat transfer coefficient distribution at stationary and rotating conditions (Rotating Mach number = 0.35) are reported. The enlarged test model can offer much more improved resolution of optical measurement near the tip region.

Modern aeroengine designs strive for peak specific fuel and thermal efficiency. To achieve these goals, engines have more highly loaded compressor stages, thinner aerofoils, and blended titanium integrated disks (blisks) to reduce weight. These configurations promote the occurrence of aeroelastic phenomena such as flutter. Two important parameters known to influence flutter stability are the reduced frequency and the ratio of plunge and pitch components in a combined flap mode shape. These are used as design criteria in the engine development process. However, the limit of these criteria is not fully understood. The following research aims to bridge the gap between semi-analytical models and modern compressors by systematically investigating the flutter stability of a linear compressor cascade. This paper introduces the plunge-to-pitch incidence ratio, which is defined as a function of reduced frequency and pitch axis setback for a first flap (1F) mode shape. Using numerical simulations, in addition to experimental validation, aerodynamic damping is computed for many modes to build stability maps. The results confirm the importance of these two parameters in compressor aeroelastic stability as well as demonstrate the significance of the plunge-to-pitch incidence ratio for predicting the flutter limit.

Targeting higher efficiencies and lower fuel consumption of turbomachines, heat transfer and profile loss are research topics of particular interest. In contrast to that, the interaction of both was, so far, rarely investigated, but gains in importance in recent research activities. The profile loss of engine components can be characterised by the airfoil wakes at the blade rows utilising established measurement and evaluation methods for which an adiabatic flow is typically supposed. To enable the investigation of the influence of heat transfer at the blade on the loss characteristics, a novel evaluation procedure was set up. In addition to the pneumatic data, the total temperature in the airfoil wake at a linear cascade was measured by means of a five-hole probe with an integrated thermocouple. For the evaluation and analysis of these data, different definitions of the loss coefficient were investigated and, finally, extended to account for thermal aspects. Furthermore, established techniques to average the local wake data were applied and compared with special focus to their suitability for non-adiabatic cases. Moreover, an extended version of the mixed-out average as defined by Amecke was utilised applying not only a far-reaching consideration of a temperature gradient but also the inclusion of the third spatial dimension to enable the evaluation of field traverses in addition to single wake traverses. These techniques were applied to wake measurement data from a linear compressor cascade gained in a special test set-up in the high-speed cascade wind tunnel for different operating points and different blade temperatures. The suitability of the new methods could be proven, and initial steps of the aerodynamic analysis of the resulting data are presented. Thereby, the acquired techniques turned out as powerful methods for the evaluation of wake traverses on compressor and turbine cascades under non-adiabatic conditions. [ABSTRACT FROM AUTHOR]

Abstract: High-pressure turbine vanes and blades are subjected to a turbulent combustor flow affecting the heat transfer and boundary layer transition, hence, the temperature distribution. The accurate prediction of the temperature distribution is crucial for a reliable design and cooling implementation. Engine-representative measurements are hence mandatory for improving design tools. Recently, convective heat transfer measurements were conducted on a high-pressure turbine inlet guide vane (VKI LS89 airfoil) in the Isentropic Compression Tube (CT-2) facility at the von Karman Institute. This contribution focuses on the effect of high freestream turbulence generated by a new turbulence grid allowing a range of turbulence intensities in excess of 10% with representative length scales of the order of 1–2 cm. Three cases with varying turbulence levels are discussed in this paper. The different flow conditions are exit isentropic Mach numbers of 0.70–0.97, Reynolds numbers of 0.53 × 106 and 1.15 × 106 and a constant temperature ratio equal to 1.36. The heat transfer distributions along the vane suction side indicate a clear link between boundary layer transition and the stream-wise pressure gradients even at high levels of freestream turbulence intensity. Differences are put in evidence in the dynamics of the transition development. Future developments will focus also on the contribution of the other flow parameters under high turbulence. Heat transfer predictions from the boundary layer code TEXSTAN and Reynolds-Averaged Navier–Stokes code elsA (ensemble logiciel pour la simulation en Aérodynamique) are additionally compared to the experiments. Inherent difficulties associated with high turbulence modelling are clear from this early numerical work. [ABSTRACT FROM AUTHOR]

Since its first operation in 1956 at DFL Braunschweig and after its movement to Munich, the High-Speed Cascade Wind Tunnel (HGK) at Bundeswehr University Munich is intensively used for fundamental and application-oriented research on aero-thermodynamics of turbomachinery bladings. Numerous systematic airfoil design studies were performed over the last decades. Thanks to the HGK facility, which enables thorough and detailed cascade testing at turbomachinery-relevant conditions, many of those airfoils for different purposes finally made it into turbomachinery applications. Nowadays, the HGK still provides very useful contributions to the understanding of the complicated flow in compressor and turbine bladings, and thereby extends the knowledge on relevant physical phenomena. As a consequence of the intense usage, this unique test facility was subject to a major revision and upgrade. The performed changes are presented within this paper including an overview on new capabilities in terms of the extended operating range, the data acquisition system, and the recently available measurement equipment.

To better understand the wake effects at low Reynolds numbers, large-eddy simulations of a 50% reaction low-pressure turbine stage and a linear cascade with two different bar wake generators were carried out for a chord Reynolds number of 50,000. For the chosen front-loaded high-lift airfoil, the endwall structures are stronger than for more traditional mid-loaded moderate-lift airfoils. By comparing the 50% reaction stage results with the bar wake generator results, insight is gained into the effect of the three-dimensional wake components on the downstream flow field.For the cases with bar wake generator, the endwall boundary layer is growing faster because of the relative motion of the endwall with respect to the freestream. The half-width of the wake is approximately matched for the larger one of the two considered bar wake generators. To improve the quality of the phase-averaged flow fields, the proper orthogonal decomposition was employed as a filter to remove the low-energy unsteady flow field content. Both the mean flow and filtered phase-averaged flow fields were analyzed in detail. Visualizations of the phase-averaged flow field reveal a periodic suppression of the laminar suction side separation from the downstream airfoil even for the smaller bar wake generator. The passage vortex is entirely suppressed for the 50% reaction stage and for the larger bar wake generator. Furthermore, the phase-averaged data for the 50% reaction stage reveal a new longitudinal flow structure that is traced back to near-wall wake vorticity. This flow structure is missing for the bar wake generator cases.

This study proposes to attach stator vanes to PowerWindow, a linear cascade wind turbine, to improve the flow direction in the device. By controlling the angle of attack, the stator vanes increase the acting force and decrease the undesirable force on PowerWindow blades. An analytical model using blade element momentum theory is developed for the new configuration, referred to as stator‐augmented PowerWindow. The analytical model has been verified by a computational fluid dynamic simulation. This study shows that the stator vanes are able to minimize/neutralize the undesirable axial force on PowerWindow so that the thrust coefficient decreases from 0.035 in the original model to −0.005 in the stator‐augmented one. In addition, by increasing the acting force on the blades, the stator augmentation will simultaneously enhance the coefficient of performance by up to 10%. This study also shows that by using stator vanes to control the angle of attack, unlike in the original PowerWindow, the direction of rotation of the stator‐augmented PowerWindow will remain the same regardless of the wind direction, increasing the utility of the device in practice. [ABSTRACT FROM AUTHOR]

This manuscript presents a rod-linear cascade model for emulating rotor-stator interaction noise. The model is intended as a test platform for studying noise mitigation techniques for a turbofan fan stage, while it also extends the classical rod-airfoil configuration by considering a row of blades based on realistic geometrical details. The rod-linear cascade model consists of a rod positioned upstream of a 7-blade linear cascade, such that the rod wake impinges onto the central blade. The rod is scaled to obtain a fundamental shedding frequency equal to the first blade passing frequency of the NASA-Glenn Source Diagnostics Test (SDT) fan stage at approach condition. The cascade blade profile is also based on the OGV of the SDT sampled at 90% of the radial span. Subsequently, numerical simulations are performed using lattice-Boltzmann Method on a computational setup comprised of a contraction and a test section enclosing the rod-linear cascade model. The integral length scales of the rod wake and the mean loading of the central blade have been found to be in good agreement with the trends observed in the SDT fan stage. The primary noise sources are localized at the central blade leading edge, although noise propagation to the far-field is influenced by additional diffraction by the other blades. Furthermore, the acoustic-blade row interaction causes intense pressure fluctuation within the inter-blade channels, including in those that are not directly affected by the rod wake. [ABSTRACT FROM AUTHOR]

This paper develops an analytical model to investigate the aerodynamic of the power generation of a linear cascade wind turbine, referred to as PowerWindow. Based on blade element momentum theory, this analytical model elucidates some flow characteristics of PowerWindow such as axial induction factor and local instantaneous angle of attack along the blades. The model also provides an understanding of the effects of blade pitch angle, blade speed ratio, and solidity on the angle of attack, axial induction factor, and power generation of PowerWindow when operating in "elevated" position. In elevated position, PowerWindow is installed on a tower as opposed to directly mounted on a roof. The validity and accuracy of the analytical model are verified with computational fluid dynamic simulations. The results indicate that by doubling the solidity or increasing the blade pitch angle by 8°, the coefficient of performance of PowerWindow increases from 8% to 12% and 16%, respectively. The analytical model can serve as the tool to find the design parameters of PowerWindow and their influence on power generation for given life wind condition. [ABSTRACT FROM AUTHOR]

Experimental investigations of non-equilibrium spontaneous condensation in transonic steam flows in a linear blade cascade were carried out. The cascade consists of the rotor blades of the low-pressure (LP) steam turbine last stage. The experimental testing facility is a part of a small-scale steam power plant located at the Silesian University of Technology in Gliwice. The steam parameters at the testing facility inlet correspond to the wet steam conditions in a low-pressure steam turbine. Static pressure measurements on the blade surface as well as Schlieren images were used to assess the flow field in the steam turbine rotor tip blade linear cascade for different flow conditions. The experimental results were used to validate an in-house CFD code. [ABSTRACT FROM AUTHOR]

A description of non-equilibrium turbulence in compressors is introduced by using the equation of Lagrangian velocity gradient correlation. This description is consistent to other traditional definitions in homogeneous isotropic turbulence and can be easily extended to complex anisotropic flows in compressors. We then show by post-processing in a linear cascade, that the non-equilibrium phenomena are obvious in the regions of corner separation, wake and boundary layer. Theoretical predictions show good agreement in magnitude with numerical results. We also show that the non-equilibrium phenomena cannot be eliminated even when the flow is far downstream, which indicates the importance of non-equilibrium phenomena in compressors. The aim of the present contribution is to provide a method to quantitatively describe the non-equilibrium phenomena in compressors, and to inspire future improvement on turbulence models by considering these non-negligible non-equilibrium properties in the future. [ABSTRACT FROM AUTHOR]

High-pressure turbine vanes and blades are subjected to a turbulent combustor flow affecting the heat transfer and boundary layer transition, hence, the temperature distribution. The accurate prediction of the temperature distribution is crucial for a reliable design and cooling implementation. Engine-representative measurements are hence mandatory for improving design tools. Recently, convective heat transfer measurements were conducted on a high-pressure turbine inlet guide vane (VKI LS89 airfoil) in the Isentropic Compression Tube (CT-2) facility at the von Karman Institute. This contribution focuses on the effect of high freestream turbulence generated by a new turbulence grid allowing a range of turbulence intensities in excess of 10% with representative length scales of the order of 1−2 cm. Three cases with varying turbulence levels are discussed in this paper. The different flow conditions are exit isentropic Mach numbers of 0.70−0.97, Reynolds numbers of 0.53∙× 106 and 1.15∙× 106 and a constant temperature ratio equal to 1.36. The heat transfer distributions along the vane suction side indicate a clear link between boundary layer transition and the stream-wise pressure gradients even at high levels of freestream turbulence intensity. Differences are put in evidence in the dynamics of the transition development. Future developments will focus also on the contribution of the other flow parameters under high turbulence. Heat transfer predictions from the boundary layer code TEXSTAN and Reynolds-Averaged Navier−Stokes code elsA (ensemble logiciel pour la simulation en Aérodynamique) are additionally compared to the experiments. Inherent difficulties associated with high turbulence modelling are clear from this early numerical work.

Performance of a turbomachinery is strongly related to the flow structure in the tip gap. Flow in the tip gap is 3-D and highly complex. Pressure driven flow between the pressure side and suction side of the blade is a source of aerodynamic loss. Thus, in order to improve the performance flow structure in this small gap should be clarified. Studies indicate that interaction of tip leakage flow with main flow inside the turbomachinery results in a considerable loss. However, structure of leakage flow has not been understood completely. Small values of tip gap height creates difficulties for experimental measurements. In this study, flow structure through the tip gap will be investigated comprehensively by implementing Computational Fluid Dynamic (CFD) methods. Effects of tip clearance height (% 0.7, % 1, and % 1.5) and relative motion between blade and casing on formation of leakage flow and tip leakage vortex have been investigated. CFD results are consistent with studies in literature. It was observed that k-ω SST turbulence model was better to predict flow structures. Increasing tip clearance height results in higher aerodynamic losses and leakage flow. It was obtained that relative motion between blades and casing reduces the aerodynamic loss by weakening the tip leakage vortex. [ABSTRACT FROM AUTHOR]

Turbo-expanders have been used for energy conversion in many fields. They are sometimes operated in partial admission mode, depending on the stability of the energy source, such as solar energy or waste thermal energy, which can fluctuate depending on environmental or process condition. The ability to operate in partial admission can be advantageous because it allows continuous operation without interchanging components. However, this method also leads to lower turbo-expander efficiency. This experimental study was conducted to investigate losses at various partial admission rates. In addition, the effect of tip clearance was studied since it is another important parameter affecting turbo-expander performance. A linear cascade was fabricated with a nozzle that was set to 65° based on the axial direction. A blade row was designed to be movable along the pitchwise direction in order to investigate flow characteristics as a function of blade movement. Flow structures and pressure losses were measured for various partial admission rates and tip clearances. The experiment was conducted at a Reynolds number of 3 × 10 5 based on the chord. The experimental results showed that losses in the passages depended not only on partial admission rate but also location relative to the nozzle. [ABSTRACT FROM AUTHOR]

Flow measurement using a linear compressor or turbine cascade is a well-established technique to characterize the flow in turbomachines with a certain degree of abstraction. A common way to obtain a general characterization of the flow is to measure the flow downstream of the cascade with a five-hole probe, obtaining, e.g., total pressure losses and flow turning. Pneumatic five-hole probes are used to capture steady or time-averaged flow quantities, if not specified otherwise. In dependency of probe geometry, measurement set-up and flow properties, such measurements can be very time-consuming. Various techniques, in order to decrease the measurement time, are proposed in literature but for certain applications the efforts required to implement such techniques can outweigh the enhanced measurement speed. In this paper, methods proposed by other authors are combined and extended to allow for fast or transient five-hole probe measurements at strongly varying flow conditions. The effectiveness of this method is presented for flow measurements downstream of a compressor cascade with attached and stalled flow (by varying the Reynolds number) as well as with steady and periodically unsteady inflow. The new method allows to reduce the measurement time by up to 90 percent without compromising measurement accuracy. In fact, due to higher spatial resolution, the flow downstream of the cascade can be better resolved with the new method.

The present paper describes an experimental and numerical study of steam condensing flow in a linear cascade of turbine stator blades. The experimental research was performed on the facility of a small scale steam power plant located at Silesian University of Technology in Gliwice, Poland. The test rig of the facility allows us to perform the tests of steam transonic flows for the conditions corresponding to these which prevail in the low-pressure (LP) condensing steam turbine stages. The experimental data of steam condensing flow through the blade-to- blade stator channel were compared with numerical results obtained using the in-house CFD numerical code TraCoFlow. Obtained results confirmed a good quality of the performed experiment and numerical calculations.

A linear cascade has been geometrically scaled according to the similarity theory,aseries of three dimensional numerical simulationsare conducted on these geometric similarcascades withthe CFD software.The differences of the aerodynamic performance betweenthe original and the scaled linear cascades are investigated in details.The result showsthatwith the increasing of linear cascade size,the shapefactor decreases at the same relative arc length position and the locationof separation point moves towards trailing edge,resultinginthe total pressure loss coefficient decreasesand the static pressureraises coefficient increases. [ABSTRACT FROM AUTHOR]

Flow field in linear cascades under high subsonic/transonic conditions has been investigated numerically. Active boundary layer control was employed in the form of sources (jets) distributed across the upper surface of the profile in a cascade. The main goal of this analysis was to evaluate the possible increase in aerodynamic performance of the cascades and define the optimal combination of jets' parameters for each operating regime. The performed analyses were 2D, viscous and compressible. A 4-equation transition turbulence model was used to close the set of equations describing the fluid flow. Artificial neural networks were investigated as a means to decrease the total computation time by providing a prediction of the cascade performance for a specific regime. Results obtained by neural networks of three different architectures are presented and discussed. [ABSTRACT FROM AUTHOR]

Abstract: In a low-speed linear cascade wind tunnel, effects of boundary layer suction on the flow characteristic in a highly-loaded compressor cascade with a stator clearance are studied experimentally, and two types of clearance including “long clearance” and “short clearance” are considered separately. In particular, hole-type suctions with different locations and distributions on the endwall are applied to control clearance flow and loss. The nature of flow separation is described by analyzing the properties of the limiting streamlines, based on the ink-trace flow visualizations on the cascade surface and lower endwall. The configuration of suction holes was associated with the kinetic energy of a clearance flow and the interaction of tip clearance flow with the endwall boundary layer and the suction surface boundary layer. The separation line and reattachment line form the base of a separation structure for corners with clearance. The results show that the clearance flow can be greatly improved and the losses reduced to some extent when boundary layer suction is adopted appropriately. The maximum amplitude of the loss reduction in different schemes with clearance is 32%. [Copyright &y& Elsevier]

For convenience of both measurement and adjusting the clearance size and incidence, the current research is mainly conducted by experiments on an axial compressor linear cascade. The characteristics and the condition under which the unsteadiness of tip leakage flow would occur were investigated by dynamic measuring in different clearances, inlet velocities and incidences. From the experiment it is found that increasing tip clearance size or reducing rotor tip incidence can affect the strength of the tip clearance flow. Then the experimental results also indicate the tip leakage shows instability in certain conditions, and the frequency of unsteadiness is great influenced by inflow angle. The condition of occurrence of tip leakage flow unsteadiness is when the leakage flow is strong enough to reach the pressure side of the adjacent blade. The main cause of tip leakage flow unsteadiness is the tip blade loading. [ABSTRACT FROM AUTHOR]

Partial admission has advantages over full admission for some operating conditions of turbine. However, additional losses such as expansion, mixing, or pumping are generated in partial admission compared to full admission. Thus, an experiment was conducted in a linear cascade apparatus having a partial admission region in order to investigate the effect of partial admission on a blade row. The admission region was formed by a spouting nozzle installed at the inlet of the linear cascade apparatus. Its cross section was circular, and its diameter was 180 mm. The nozzle was set to have a flow angle of 65°. The tested blade was axial-type, and its chord was 200 mm. The experiment was conducted at a Reynolds number of 3 × 10 based on the chord. One blade row of nineteen identical blades was applied to the linear cascade for the partial admission experiment, and this blade row moved along the pitchwise direction in front of the admission region. The operating forces and surface pressures on a blade were measured along the direction at the steady state. For investigating the effect of solidity, three different solidities of 1.25, 1.38, and 1.67 were applied to the blade row. From the experimental results, the maximum rotational force and axial force increased for a larger solidity. [ABSTRACT FROM AUTHOR]