Suction Side Surface Research Articles

Phantom Cooling Effects on Rotor Blade Surface Heat Flux in a Transonic Full Scale 1 + 1/2 Stage Rotating Turbine

Abstract An experimental and numerical investigation of phantom cooling effects on cooled and uncooled rotating high pressure turbine blades in a full scale 1 + 1/2 stage turbine test is presented. Objectives set to capture, separate, and quantify the effects of upstream vane film-cooling and leakage flows on downstream rotor blade surface heat flux. Multiple series of 1 + 1/2 stage rotating high pressure turbine tests were carried out in the Air Force Research Laboratory, Turbine Research Facility, at Wright-Patterson Air Force Base, Ohio. A non-proprietary research turbine was instrumented with high frequency double-sided thin film heat flux gauges custom made at AFRL. High bandwidth, time-resolved surface heat flux is measured on multiple film-cooled and non-film-cooled HPT rotor blades downstream of both film-cooled and non-film-cooled vane sectors. Upstream wake passing and heat flux is characterized on both rotor pressure and suction side surfaces, along with quantifying rotor phantom cooling effects from non-uniform first-stage vane film cooling and leakage flows. Fast response heat flux measurements quantify how rotor phantom cooling impacts the blade pressure side greatest. Measurements show upstream vane film-cooling alone can account for 50% of the rotor blade cooling effect in some instances, and even outweigh the rotor blade film cooling effect far from the blade showerhead holes. Added unsteady aero numerical simulation demonstrates how variations in inlet total temperature and incidence angle also contribute to the circumferentially non-uniform rotor heat flux. Results from this investigation aid modeling and design efforts in optimizing film cooling performance in real high-pressure turbine flow fields. Understanding the behavior of such non-uniform circumferential rotor phantom cooling effects can be critical to optimizing the efficiency, fuel consumption, range, and durability of advanced turbomachines.

Transition mechanism and loss analysis of a separated flow over herringbone riblets in a compressor cascade using the lattice Boltzmann method

Herringbone riblets were regarded as a promising approach to control the separation bubble on the compressor blade. However, the underlying mechanism requires further elucidation. And numerical simulations with body-fitted meshes often face challenges in mesh generation due to the tiny and complex geometries involved. In the present research, high-fidelity simulations using the Lattice Boltzmann Method and Immersed Boundary Method were performed to investigate the effects of herringbone riblets on separated flow in a compressor cascade. At a low Reynolds number of 90 000, a separation bubble appears on the blade suction surface. The application of herringbone riblets on the suction side surface shows that it effectively reduces the bubble length from 0.24c to 0.12c and reduces the loss coefficient by 11%. A counter-rotating mode of secondary flow occurs before the separation, with a near-wall spanwise motion from the divergent region to the convergent region and a compensating flow from the convergent region to the divergent region in the outer layer of the boundary layer. Transition occurs earlier on the suction side surface due to the complex flow patterns. Four different mechanisms are responsible for the earlier transition. Over the divergent region, engulfing of a high momentum fluid from the outer layer to the inner layer of the boundary layer suppresses the separation bubble, forcing a high-momentum passage where an attached boundary layer is observed. This thinner boundary layer leads to an earlier natural transition. Second, the discharge of fluid from the herringbone cusp causes the overflow from the riblet channel beside the divergent line, i.e., overflow transition. Meanwhile, the transition over the converging region is attributed to the accumulation of disturbance. Finally, in the middle region with yawed riblets, transition in a separated shear layer occurs earlier under the influence of adjacent transition mechanisms over the divergent/convergent region. These mechanisms also bring about a serrated structure in the downstream wake. Overall, this research confirms the role of the counter-rotating mode produced by herringbone riblets in separation control and reveals the transition mechanisms for loss production. The findings suggest that proper utilization of herringbone riblets can provide significant improvement on the compressor blade performance.

Aerothermal Performance of Slashface Leakage With Double Surface Angle Under Realistic Swirling Inflow

Abstract The turbine vanes being manufactured and assembled piece by piece leads to an unavoidable gap on the vane endwall called slashface. Usually, the coolant leakage introduced from the slashface with a single surface angle aims to prevent hot gas ingression and protect the vane endwall. However, the cooling effectiveness is greatly reduced when considering the swirling flow from the combustor. In the current work, against strong swirling inflow, a new solution to improve endwall film cooling performance through applying the slashface leakage with the double surface angle is carried out. Through numerical method, the effects of the location of the angle transition region (D/L = 0%, 25%, 50%, 75%, 100%) and mass flow ratio (MFR = 0.5%, 1.0%, 1.5%) of the leakage on the turbine vane endwall aerothermal and film cooling performance are investigated. The results indicated that the coverage area of the coolant on the endwall was obviously enlarged when the slashface leakage with the double surface angle was employed. The film cooling level on both the endwall and suction side surface of the vane increased with increasing MFR. Moreover, the endwall thermal load was reduced. The designs of D/L = 25% and 50% are recommended for their high endwall film cooling level and low endwall thermal load. This paper provides turbine designers with a new idea to increase endwall film cooling performance by the slashface leakage with the double surface angle when considering aggressive swirling inflow.

Experimental and Numerical Investigation of Roughness Structure in Wind Turbine Airfoil at Low Reynolds Number

This paper experimentally and numerically investigates the effects of suction side surface roughness on the aerodynamic performances of the NACA 0015 turbine blade profile. Three different NACA 0015 turbine blade configurations, which are smooth (K0), single roughness (K1), and double roughness (K2), are considered. The experimental studies were conducted using the HM-170 GUNT open wind tunnel model. The aerodynamic characteristics of these three blade configurations are evaluated in terms of their lift coefficient (CL), drag coefficient (CD), and aerodynamic efficiency (CL/CD). The maximum CL (CL,max) for K0 was obtained at 25°, whereas the CL,max angles for the K1 and K2 roughness blade profiles were reduced to 22.5°, utilizing the rough surfaces on the suction side. The experimental analysis revealed that the K2 profile demonstrated a 21% and 19% enhancement in maximal CL over the K0 and K1 profiles, respectively. The highest CL/CD was observed with K1, except at low attack of angle (αoα), where the smooth blade profile resulted in slightly better performance. Experimental analysis showed peak CL/CD at αoα of 7.5° for K0, and 12.5° for both K1 and K2, with K1's optimal CL/CD being 2.85% and 8.5% higher than K0 and K2, respectively. Numerical analysis indicated that the CL/CD,avg for K1 was observed to be 11% and 8% higher than that of K0 across all αoα.

Film cooling effect of upstream jump coolant on turbine endwall with combustor-turbine interface misalignment

The investigation of jump cooling performance on turbine endwall independently causes its actual cooling effectiveness to deviate from the design values. In this paper, the endwall jump cooling structure with combustor-turbine interface cavity and combustor liners is modeled. Taking the realistic combustor-turbine interface features into account, the effect of jump coolant on turbine endwall film cooling and heat transfer characteristics is numerically studied. Under different endwall misalignment modes, the aerodynamic interaction mechanisms of realistic combustor outflow profile, jump coolant jet and cascade secondary flows at turbine endwall region are revealed. The modification effects of combustor-turbine interface features on jump cooling of the endwall are investigated. The results show that the combustor outflow severely ingests into the combustor-turbine interface cavity after flowing across several steps. Two branches of cavity vortex are subsequently generated in the middle-pitch region of cascade to affect the jump coolant jet. At low coolant blowing ratio, the attachment and separation sides of cavity vortex individually result in high and low cooling regions, while the horseshoe vortex leads to a large wedge-shaped cooling region. At z/Cax < 0, the backward step leads to a higher cooling effectiveness than forward step by 0.2. The jump coolant only leads to phantom cooling effect on downstream part of the vane suction side surface. The net heat flux reduction (NHFR) between two cavity vortexes and at pressure side junction are individually 0<NHFR<0.2 and −0.2<NHFR < −0.1. At high blowing ratio, the jump coolant can cover the pressure side junction with the highest cooling effectiveness of 0.5. The jump coolant can flow across the horseshoe vortex and directly upwash the vane surface. The weak protection regions between two cavity vortexes and near the attachment line of the pressure side horseshoe vortex enlarges. Compared with the forward step, the backward step achieves a lower NHFR by up to 0.1 and a higher Nusselt number. This paper provides an in-depth analysis of the aero-thermal physics of endwall jump cooling concerning realistic combustor-turbine interface conditions.

Noise generation mechanisms of a micro-tube porous trailing edge

The noise generation and flow characteristics of a forced-transitioned NACA 0012 airfoil with a micro-tube porous trailing edge and its solid counterpart have been numerically investigated. The near-field flow dynamics and far-field noise predictions are obtained using compressible large-eddy simulations and the Ffowcs-William and Hawkings acoustic analogy, respectively. The computed far-field noise levels agree with experimental data, showing that the micro-tube structure reduces low-frequency trailing-edge noise without noticeably altering the noise directivity but induces additional high-frequency noise along the direction perpendicular to the airfoil chord. An in-depth analysis of the trailing-edge flow and surface pressure fields reveals that the noise reduction is largely linked to the ‘cross-jet-like’ flow permeation through the micro-tube structures. The interaction between the flow permeation and the boundary layer turbulence leads to a reduction in the strength of the Reynolds-stress source term in the pressure-side flow field and a reduction in the magnitude of the source terms in Amiet’s trailing-edge noise model. Spectral proper orthogonal decomposition is performed to identify flow structures that are spatially and temporally coherent within the micro-tube geometries and to link the flow features to the attenuation in noise scattering efficiency. The effect of the micro-tube structure on the wavenumber–frequency spectra of the trailing-edge surface pressure field is quantitatively analysed. Additional high-energy regions exist in the wavenumber–frequency spectra of the porous trailing edge, resulting from the turbulent eddies that permeate through the micro-tubes. Further, the high-frequency noise increase mechanism is identified as the acoustic dipole noise arising from interactions of edge-induced flow separation with the suction-side surface.

Flow Control around NACA0015 Airfoil Using a Dielectric Barrier Discharge Plasma Actuator over a Wide Range of the Reynolds Number

In this study, an experimental investigation of separation control using a dielectric barrier discharge plasma actuator was performed on an NACA0015 airfoil over a wide range of Reynolds numbers, angles of attack, and nondimensional burst frequencies. The range of the Reynolds number was based on a chord length ranging from 2.52 × 105 to 1.008 × 106. A plasma actuator was installed at the leading edge and driven by AC voltage. Burst mode (duty-cycle) actuation was applied, with the nondimensional burst frequency ranging between 0.1–30. The control authority was evaluated using the time-averaged distribution of the pressure coefficient Cp and the calculated value of the lift coefficient Cl. The baseline flow fields were classified into three types: (1) leading-edge separation; (2) trailing-edge separation; and (3) the hysteresis between (1) and (2). The results of the actuated cases show that the control trends clearly depend on the differences in the separation conditions. In leading-edge separation, actuation with a burst frequency of approximately F+= 0.5 creates a wide negative pressure region on the suction-side surface, leading to an increase in the lift coefficient. In trailing-edge separation, several actuations alter the position of turbulent separation.

Suppression effect on TLV cavitation of overhanging grooves with different tab orientations: An experimental investigation

In this paper, a high speed camera and endoscopic particle image velocimetry system are utilized to investigate the influence of tab orientations of overhanging grooves on the suppression effect of tip-leakage vortex cavitation around a truncated NACA0012 hydrofoil. Four kinds of overhanging grooves (OHGs) with different tab orientations, namely OHGs_1–4, are tested, with the original hydrofoil as a comparison. The results suggest that compared with the baseline, all the overhanging grooves tested in our experiments show the suppression effect on TLV cavitation. However, the tab orientations of overhanging grooves leave significant influence on its effect. OHGs_2 with tabs extending beyond the pressure side surface shows the best suppression effect for all the measured gap sizes, followed by OHGs_1 with tabs extending beyond both the suction and pressure side surfaces. For OHGs_3 and OHGs_4, due to the lack of tabs extending beyond the pressure side surface, the suppression effect is not so good as OHGs_1 and OHGs_2. Further analysis indicates that the reason for the better suppression effect of OHGs_2 is the interference of tabs extending beyond the pressure side surface on the fusion of tip-leakage vortex and tip-separation vortex, which decreases the tip-leakage vortex circulation and increases its radius.

Heat transfer characteristics of a high-pressure turbine under combined distorted hot-streak and residual swirl: an unsteady computational study

• Unsteady CFD simulations were performed for a transonic turbine stage to reveal the effects of uniform and swirled flows on hot-streak transport. • Two ideal hot-streak maps, radial and rounded, were evaluated against a realistic distorted hot-streak map. • The residual swirl contributed to the homogenisation of the radial hot-streak and the dispersion of the rounded and distorted hot-streaks through the rotor. • The distorted hot-streak showed the most complex transport behaviour through the stage. • The residual swirl was found to intensify the leakage vortex and increase the heat transfer rates near the tip region. In new generation aero-engines, lean-burn combustors are equipped with swirl injectors in order to reduce pollutant emissions. At the exit of these combustors, the flow is dominated by temperature non-uniformity (hot-streak) and residual swirl. Available research has focused mainly on the isolated effects of the residual swirl only or the hot-streak only on the aerodynamic performance or thermal performance of the high-pressure turbine (HPT). Only few studies investigated the combined swirl and hot-streak on the aerothermal performance of the HPT using either idealized uniform or rounded hot-streak topologies. Realistic hot-streaks generated from lean burn combustors have more complex topologies that involves distortions. The present study investigates the effects of residual swirl and distorted hot-streak simultaneously on a first stage of a HPT. Other hot-streak types, uniform and rounded, have been also investigated with and without swirl to perform comparisons with the distorted hot-streak. Unsteady Reynolds-averaged Navier–Stokes (URANS) computations have been conducted to assess the aerothermal performance of a HPT under the influence of different hot-streaks. Results revealed that all hot-streaks without swirl were almost preserved as transported through the vane and altered as transported through the rotor due to the secondary flows. Under the residual swirl, all hot-streaks were remarkably altered at the vane exit and deformed more at the rotor exit. The uniform hot-streak was homogenised through the rotor and the rounded and distorted hot-streaks were dispersed. The distorted hot-streak showed the most complex transport behaviour through the stage. Results also revealed that the leakage flow through the rotor tip gap generated high heat transfer rates, in particular on the rotor blade suction side and tip surface. The leakage flow formed the known leakage vortex at the tip gap exit, which induced high heat transfer rates. The residual swirl was found to intensify the leakage vortex and consequently higher heat transfer rates are generated compared to the uniform flow condition.

Optimization of Turbine Blade Aerodynamic Designs Using CFD and Neural Network Models

Optimization methods have been widely applied to the aerodynamic design of gas turbine blades. While applying optimization to high-fidelity computational fluid dynamics (CFD) simulations has proven capable of improving engineering design performance, a challenge has been overcoming the prolonged run-time due to the computationally expensive CFD runs. Reduced-order models and, more recently, machine learning methods have been increasingly used in gas turbine studies to predict performance metrics and operational characteristics, model turbulence, and optimize designs. The application of machine learning methods allows for utilizing existing knowledge and datasets from different sources, such as previous experiments, CFD, low-fidelity simulations, 1D or system-level studies. The present study investigates inserting a machine learning model that utilizes such data into a high-fidelity CFD driven optimization process, and hence effectively reduces the number of required evaluations of the CFD model. Artificial Neural Network (ANN) models were trained on data from over three thousand two-dimensional (2D) CFD analyses of turbine blade cross-sections. The trained ANN models were then used as surrogates in a nested optimization process alongside a full three-dimensional Navier–Stokes CFD simulation. The much lower evaluation cost of the ANN model allows for tens of thousands of design evaluations to guide the search of the best blade profiles to be used in the more expensive, high-fidelity CFD runs, improving the progress of the optimization while reducing the required computation time. It is estimated that the current workflow achieves a five-fold reduction in computational time in comparison to an optimization process that is based on three-dimensional (3D) CFD simulations alone. The methodology is demonstrated on the NASA/General Electric Energy Efficient Engine (E3) high pressure turbine blade and found Pareto front designs with improved blade efficiency and power over the baseline. Quantitative analysis of the optimization data reveals that some design parameters in the present study are more influential than others, such as the lean angle and tip scaling factor. Examining the optimized designs also provides insight into the physics, showing that the optimized designs have a lower amount of pressure drop near the trailing edge, but have an earlier onset of pressure drop on the suction side surface when compared to the baseline design, contributing to the observed improvements in efficiency and power.

Flow characteristics and performance of the propulsion system with wavy duct

The present study applies the wavy shape on a wake equalizing duct before a propeller. The numerical simulations are performed to investigate the effect of wavy duct on a propulsion performance in the ranges of angles of attack (0°≤α≤15°) and the wavenumber of the duct ((4≤Wn≤32)). All ducts contribute to improving the propeller efficiency. The effect of the wavy duct on the force coefficients and their dependence on Wn become significant with increasing α. When α=10°, the wavy duct gives larger force coefficients and propeller efficiency with increasing Wn. The better performance of the propeller at the largest Wn is physically explained by the increase of the axial velocity in the slipstream region, more uniform distribution of the upstream velocity, and wider distributions of the low and high pressures in the suction and pressure side surfaces, respectively. The vortical structures of the wavy ducts and those size depend on the circumferential plan and Wn. The largest Wn breaks large vortical structures into smaller structures. The increase of the local momentum near the troughs contributes to the modification of vortical structures. Also, the appearance of additional axial vortices over the troughs is associated to modify the vortical structures with the largest Wn.

Detailed velocity and heat transfer measurements of an advanced insert for impingement cooling

This work reports the results of paired experiments for the aft section of a complex internal cooling flow within a gas turbine vane using Magnetic Resonance Velocimetry (MRV) and steady-state Infrared (IR) thermography. The aft cooling insert for the vane with two-pass impingement was designed at a scale five times larger than the original and built using stereolithography (SLA) fabrication methods. The MRV technique was used to measure the three-dimensional, three-component velocity field for a test case with a Reynolds number range of 2,600 to 7,600 based on diameter of the impingement holes. Flow distribution, impingement hole performance, and cross flow effects are discussed for the experiment in which a dilute aqueous copper sulfate solution was used as the working fluid. A paired experiment with a geometrically similar design employed electrical heating of a thin stainless steel shim to model a constant heat flux boundary condition of an interior wall of the turbine vane. The modeled vane insert was then operated with air as the working fluid at two test conditions with Reynolds numbers in the range of approximately 1,200 to 7,600 based on the diameter of the impingement holes. An IR camera was used to measure the surface temperature of the shim. Using energy balances and the known heat flux, the temperature data were used to determine heat transfer characteristics of the impinging jets for the pressure and suction side surfaces, including the Nusselt number. The MRV and IR data sets provide detailed insight into the surface effects of the flow distribution and the result on the local and area-averaged heat transfer performance. A strong coupling between the velocity field and temperature data provide insight into design feature performance, and serve as a validation data set for matched computational simulations. Finally, a comparison with internal heat transfer correlations is presented using the data from Florschuetz et al. [1]. The results showed a lack of agreement with Florschuetz, leading to the development of a novel methodology for estimating the heat transfer performance of an impinging hole with crossflow. Measurement uncertainty was estimated to be ±5% for velocity and ±5% for the spatially averaged Nusselt number distributions.

The effect of wavy surface on boundary layer instabilities of an airfoil

PurposeThe purpose of this paper is to investigate airfoil’s tonal noise reduction mechanism when deploying surface irregularities, such as surface waviness by means of spatial stability analyses.Design/methodology/approachFlow field calculations over smooth and wavy-surface NACA 0012 airfoils at 2° angle of attack and at Reynolds number of 200,000 are performed using the large eddy simulation (LES) approach. Three geometrical configurations are considered: a smooth NACA 0012 airfoil, wavy surface on the suction side (SS) and wavy surface on the pressure side (PS). The spatial stability analyses using the LES-generated flow fields are conducted and validated against the Orr-Sommerfeld stability analysis for the smooth airfoil configuration.FindingsThe spatial stability analyses show that inclusion of the wavy-type modification on the SS of the airfoil does not lead to altering of the acoustic feedback loop mechanism, with respect to the mechanism observed for the smooth airfoil configuration. In contrast, applying the surface modifications to the airfoil PS leads to a significant reduction of the amplification range of disturbances in the vicinity of the trailing edge for the frequency of the acoustic feedback loop mechanism.Practical implicationsThe spatial analyses using, for example, LES-generated flow fields can be widely used to determine acoustic sources and associated distributions of amplifications for a wide range of applications in the aeroacoustics.Originality/valueThe spatial stability analysis approach based on flow fields computed a priori using the LES method has been introduced, validated and used to determine behaviour of the acoustic feedback loop when accurate reconstruction of geometry effects is required.

환형 섹터 터빈 노즐 캐스케이드에서 팬 형상 막냉각 홀의 복합분사각도 적용에 따른 막냉각 효율 특성

The present study investigated the effect of compound angle injection of fan-shaped film cooling holes on film cooling performance on the nozzle vane surface in an annular sector cascade. Two different configurations were examined to investigate effect of compound angle injection. The baseline has streamwise injection holes on the pressure and suction side surfaces and normal injection holes in leading edge region. The other configuration has compound angle injection holes on the leading edge and the pressure side surface. An annular sector turbine cascade test facility was used and the measurements using an infrared thermography method were conducted at the exit Reynolds number of 2.2x10<SUP>6</SUP> and exit Mach number of 0.8. Total coolant mass flow rate ranges between 5 and 10% of mainstream. The results showed that the baseline configuration shows skewed trajectory of coolant downstream of holes on the pressure side surface, which is not observed in ‘flat-plate’ study and strongly related to mainstream flow direction near the nozzle surface. On the other hands, the coolant follows the mainstream direction and spread more uniformly on the pressure side surface with compound angle injection configuration. However as coolant mass flow rate increases, compound angle injection induces stronger interaction of mainstream and coolant and film cooling effectiveness drops quickly in the downstream region. Also, film cooling effectiveness in the suction side surface shows lower values for compound angle injection configuration since compound angle injection on the showerhead region causes non-uniform distributions of film cooling effectiveness due to stronger secondary flow.

Influence of slashface leakage coupled with quasi-labyrinth seal technique on gas turbine endwall aerothermal performance and blade suction side surface phantom cooling

Quasi-labyrinth seal technique was innovatively adopted to control the slashface flow, and the effect of slashface leakage on endwall aerothermal performance and blade suction side surface phantom cooling was numerically investigated. Simulations of five different labyrinth seal lengths ( Ls) and three coolant momentum flux ratios ( I) were conducted by computationally solving the Reynolds-averaged Navier–Stokes equations coupled with the shear stress transport k-ω turbulence model. The results show that the slashface coolant can be accelerated and conveyed downstream inside the slashface carried by the ingested mainstream and passage vortex. By setting up the quasi-labyrinth seal composed of a series of aperture cavities along the slashface, the coolant downstream transportation can be largely weakened, and the labyrinth seal with various lengths all have the potential to move the endwall coolant coverage toward upstream. When I = 0.48, the coolant downstream migration is severest and it has the best coverage on the endwall with seal. The laterally averaged cooling effectiveness is elevated for the fore part endwall but decreased for the back part endwall after sealing, and the highest cooling effectiveness is increased by 5% for 0.6 Ls seal. The 0.6–1.0 Ls seal has the approximate effect and is better than 0.4 Ls seal, and 0.6 Ls seal has the best outcome considering the countering effects of structural strength and endwall cooling performance. For 0.6 Ls sealed case, the high heat transfer level caused by coolant attachment is enhanced by 3.5% when I = 0.48 but decreased by 3.3% and 4.4% when I = 0.96 and 1.42 compared with baseline case. The low heat transfer caused by horseshoe vortex separation is enhanced when I = 1.42 for the sealed case. The quasi-labyrinth seal mainly weakens the back part suction side phantom cooling performance, and the well phantom cooling region moves upstream with the increase of I. The averaged phantom cooling effectiveness is decreased by 0.56%, 0.3%, and 1.44% when I = 0.48, 0.96, and 1.42 for the sealed cases. The results provide the gas turbine designers a better insight into improved slashface leakage control as well as its detailed surface cooling effects.

Metallurgical Failure Analysis of Cracked First Stage Vane of Heavy-Duty Industrial Gas Turbine Engine for Power Generation

Abstract A first stage turbine vane of a heavy-duty gas turbine engine, exhibiting an outward bulge of approximately 0.6 mm and a visible crack in the top coating, was delivered by the client for a detailed destructive failure investigation. The subject component was an experimental part used in a test field engine. The suction side wall of the casting was cut off to reveal a gaping crack on the inside of the aft cavity that was also visible by borescopic inspection. The cooling insert of the aft cavity was clean and did not have any blocked cooling holes. The cooling pattern on the inside wall of the aft cavity was unremarkable. No evidence whatever of impact damage was found on the suction side surface of the subject vane in the cracked region. The hypothesis as to the most probable root cause of the cracking that is still standing is a combination of low wall thickness, non-optimal process parameters for bond coat pre-heat, and potentially high thermal/mechanical loading during engine operation. It appears this is not a wide-spread phenomenon but rather an isolated occurrence.

Numerical investigation and optimization of airfoil flow control using passive air-jet

In this paper, a numerical investigation of airfoil flow control using passive air jet is presented. The study of generate a passive jet from pressure side to suction side was conducted in order to improve the airfoil characteristics. Such characteristics include boundary layer separation and stall inception. The numerical simulations conducted using CFDRC software. Different structured and unstructured finite volume technique is used to solve the steady compressible Navier—Stokes equations. The code results were validated with the experimental testing results for different turbulence model. Parametric study is performed for the location, slot width and angle of a synthetic jet on the suction side of a subsonic flow over a NACA 23012C airfoil. The maximum lift is achieved with the jet flow being normal to suction side surface but this come with penalty of the high drag. The maximum lift to drag ratio was accompanying with synthetic jet being located at 43% chord and 30° jet angle. Finally, optimum geometry of jet is obtained using simplex algorithm with initial shape obtained from the parametric study.

Comparison of low wind speed aerodynamics of unsymmetrical blade H-Darrieus rotors-blade camber and curvature signatures for performance improvement

Unsymmetrical blade Vertical axis wind turbine (VAWT) can be nicely adapted in the built environment because of its improved performance, self-starting ability, simple construction, easy maintenance and capability to adapt to wind direction changes. For proliferation of such VAWT installations, the challenge is to improve its efficiency in low wind speed condition (usually less than 10 m/s) of the built environment, which depends on its blade profile design. In an unsymmetrical blade, blade thickness-to-chord ratio, percentage camber, blade curvature around aerodynamic moment center are the controlling parameters for effective blade-fluid interactions for performance improvement, thus requiring a detailed investigation of the aerodynamics of unsymmetrical blade VAWT. The objective of this paper is to compare low wind speed aerodynamics of two different unsymmetrical airfoil blade profiles to obtain design information of blade camber and blade curvature around aerodynamic moment center for performance improvement of H-Darrieus VAWT, which is hardly available in the existing literature. Detailed blade-fluid interactions are analysed by Ansys Computational Fluid Dynamics (CFD) software for different low wind speeds between 4 m/s and 8 m/s. It is found that more rounded curvature of the suction side surface of advancing S815 blade around aerodynamic moment center and thicker blade profile (maximum thickness is 26.2% at 25.7% chord) has a distinct influence for performance improvement of the rotor in the power stroke. In the returning stroke, the VAWT is benefited from higher camber of EN0005 blade (maximum camber percentage of 10% at 55% chord) and lesser blade curvature of the blade suction side. Next, vorticity structures of each blade profile have also been monitored at different azimuthal positions and necessary performance insights with regards the development of static/deep stall vortices and other boundary layer features have been drawn. Further, blade polars, power performance, and blade loading have also been obtained and compared with published results. Overall better designs of unsymmetrical blade profiles have also been recommended. The present study leads to the understanding of the design blade camber and curvature signatures in both power and returning strokes for possible performance improvement of small-sized VAWT in low wind speed condition.

Effects of simulated swirl purge flow and mid-passage gap leakage on turbine blade platform cooling and suction surface phantom cooling performance

Effects of the coolant swirl ratio (SR) and density ratio (DR) of simulated upstream slot purge flow, which is used to simulate the rotation effect, and the mid-passage leakage blowing ratio M on the turbine blade platform cooling and suction side surface phantom cooling performance are numerically investigated using three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations coupled with the SST k-ω turbulence model. Simulations with four slot injection swirl ratios (SR) of 0.4, 0.6, 0.8 and 1.0, three platform blowing ratios M of 0.5, 1.0, 1.5 and three coolant density ratios (DR) of 1.0, 1.5, 2.0 under four different mass flow ratios (MFR) that range from 0.5% to 1.25% are conducted. The calculated results indicate that at low mass flow ratio (MFR = 0.5% and 1.0%) the best and worst platform cooling performance appear at SR = 0.4 and 0.8, respectively. But at MFR = 1.25%, the swirl ratio has positive effect on the cooling effectiveness after eliminating the mainstream ingestion. In addition, the phantom cooling effectiveness also increases with the increase of the relative motion between coolant and blade with the movement of separation vortex toward suction side. With the increase of DR (changes simultaneously in slot and midpassage gap), the uncooled area near the leading edge, which is usually referred to as the hot ring, is enlarged at high swirl ratio and the platform cooling effectiveness decreases for the fore portion. Similarly, the density ratio has negative effect on the phantom cooling effectiveness for all the cases. As for the mid-passage gap, its coolant blowing ratio M has a significant impact on the downstream portion platform cooling by acting partly as an endwall fence. With the increase of blowing ratio M, the leakage coverage on the platform increases firstly and then decreases at low density ratio (DR = 1.0 and 1.5) but increases continuously at DR = 2.0. As a result, the platform cooling effectiveness reaches its highest at M = 1.0 and 1.5 for DR = 1.0 and 2.0, respectively. The phantom cooling effectiveness increases slightly when increasing M from 1.0 to 1.5, but has nearly no difference for further increasing M.

Multiple control modes of nanosecond-pulse-driven plasma-actuator evaluated by forces, static pressure, and PIV measurements

The control authority of a nanosecond-pulse-driven dielectric-barrier discharge plasma actuator (ns-DBDPA) was evaluated via wind tunnel experiments with the simultaneous measurement of lift and drag forces, pressure on the airfoil surface, and particle image velocimetry (PIV) measurements. In these experiments, a Reynolds number of Re = 2.6 × 105 was applied with a freestream velocity of 40 m/s under atmospheric pressure. The force measurements revealed multiple peaks of lift force recovery and drag force modulation depending on the angle of attack, α, and non-dimensional frequency, F+. At the positive post-stall α close to stall α of approximately 16°, F+ values around 2.0 were effective for lift recovery and drag reduction. When the deep-stall angle α is larger than 20° (either positive or negative), relatively low F+ values around 0.25 were effective for lift recovery. When actuating at a deep-stall angle corresponding to F+ = 0.25, the surface pressure measurements showed that a near flat pressure distribution is formed on the suction side, and the PIV measurement showed that this near flat distribution is caused by the increase in backflow velocity near the surface of the airfoil. This backflow enhancement near the suction side surface leads to the reduction in pressure in separated flow, resulting in significant increases in the lift and drag coefficients. Thus, this simultaneous measurement of force, pressure, and PIV is capable of evaluating the multiple control modes underlying lift and drag control by ns-DBDPA.