Terms Of Aerodynamic Performance Research Articles

Experimental and numerical investigation on the unsteady interaction in longitudinal stage separation for parallel-staged two-body configuration

Hypersonic stage separation is a significant process, probably involving complex aerodynamic interaction, which determines the survival of two-stage-to-orbit (TSTO) vehicles. The longitudinal stage separation (LSS) scheme is proposed to improve the safety of the parallel-staged TSTO vehicle, where a small interstage gap may result in weak interference. Therefore, an experimental and numerical study of LSS for the parallel-staged TSTO vehicle at Mach 7 with different angles of attack (AoA) is carried out. The dynamic interaction, including variations in the shock structure, wall pressure distribution, and unsteady aerodynamics, is investigated by testing and numerical simulation. The LSS experiments for the TSTO vehicle were performed using a high-speed pneumatic ejection launch system in the JF-12 shock tunnel, and the method was developed using high-speed visualization and image processing techniques to capture the separating trajectory. The numerical simulations were carried out using the overset grid method and solving the Navier–Stokes equations coupled with the rigid body dynamics equations to obtain the laminar flows over the TSTO vehicle during LSS. The qualitative and quantitative comparison of the test and numerical results showed good agreement in terms of aerodynamic performance, flowfield pattern, wall pressure, and separation trajectory. They show that the small interstage gap of the LSS leads to weak type I and VI shock–shock interactions, with short-duration weak shock reflection at a higher AoA. Furthermore, no shock reflection or interstage gap is observed at lower AoA. Moreover, no stage recontact is observed, and the safety and feasibility of LSS for parallel-staged TSTO vehicles are demonstrated.

Low-Engine-Order Forced Response Analysis of a Turbine Stage with Damaged Stator Vane.

A damaged stator vane can disrupt the circumferential symmetry of the design flow for turbine assemblies, which can lead to a low-engine-order (LEO) forced response of rotor blades. To help engineers be able to better address sudden vane damage failures, this paper conducts a mechanism analysis of the LEO forced response of rotor blades induced by a single damaged vane using an in-house computational fluid dynamic code (Hybrid Grid Aeroelasticity Environment). Firstly, it is found that the damaged vane introduces a family of LEO aerodynamic excitations with high amplitudes by full-annulus unsteady aeroelastic simulations of a single-stage turbine. In particular, the LEO forced response of the rotor blades excited by 3EO is 2.01 times higher than the resonance response excited by vane passing frequency, and the LEO resonance risk of the rotor blades is greatly increased. Then, by analyzing the flow characteristics of the wake and potential field of the stator row with a damaged vane, the localized high transient pressure in the notch cavity and the radial redistribution of the secondary vortex at the stator exit are the main sources of the low-order harmonic components in the flow field. Importantly, the interaction mechanisms in two regions with high LEO excitation amplitude on the rotor blade surface are revealed separately. Finally, an evaluation and comparison of a single damaged vane in terms of aerodynamic performance and LEO forced response was carried out. The results of this paper provide a good theoretical basis for engineers to effectively control the resonance response of rotor blades caused by a damaged stator vane in turbine design.

Control of the leakage flow and heat transfer characteristics at the blade tip by cavity filling

In order to reduce leakage flow, cavity tips are often used in high-pressure turbines, at the cost of higher thermal load compared to flat tips. To better control the leakage flow and heat transfer at the blade tip, a cavity-filling geometry is proposed in this paper, and the effects of the filling area are also investigated by numerical simulation. The results indicate that the new geometries maintain the advantages of cavity tips in terms of aerodynamic performance. With the cavity filling area increasing, the heat transfer of the new geometries shows a trend of initially increasing and then decreasing. When the front 15% of the blade tip chord length is filled, the average heat transfer of this configuration is reduced by 3.7% compared to the cavity blade tip, and the high heat transfer area is reduced by 23.3%, at the cost of more intensive hot spots. As the cavity-filling area increases from 5% to 15%, the size of the vortexes at the cavity reduces, resulting in a rise in the leakage. Similar behaviors of the tip heat transfer and leakage flow were also found when the turbine pressure ratios varied from 3 to 7, with the cavity-filling geometry.

Multidisciplinary robust optimization approach of fan rotors under structural constraints with blade curvature

The robust optimization approach is an important method to solve the problem of fans/compressors performance disparities caused by geometric tolerances. Nonetheless, since the optimum design of turbomachinery blades is multidisciplinary in nature, traditional robust optimization approaches that only consider aerodynamic performance may result in the blade strength exceeding their material limit. Therefore, it is necessary to investigate the multidisciplinary robust optimization approach (MROA) considering both aerodynamic and structural performance. For MROA, there is a big challenge of huge computation cost and runtime for the time-consuming high-fidelity CFD and FEM. In this paper, a new method to reduce the computation cost and runtime of MROA is proposed by using the blade curvature to approximate the evaluation of blade stress. Firstly, data mining by self-organizing mapping method is used to analyze and extract the constraint value of the blade curvature. And the blade curvature constraint value is used as a penalty function instead of the time-consuming high-fidelity FEM, which greatly reduces the computational cost and runtime of the multidisciplinary optimization. Then, Polynomial chaos Kriging (PCK) is then used as a surrogate model for uncertainty quantification (UQ) of aerodynamic performances, and one MROA based on the PCK model-based UQ method and the curvature constraints is developed. The proposed method is verified using a fan rotor as a case study. The results show that the maximum stress of the traditional optimization method was 453 MPa, which exceeded the yield limit of aluminum alloy (420 MPa). In contrast, the maximum stress of the method proposed in this paper was 343 MPa. In terms of aerodynamic performance, compared with the baseline, the mean efficiency of the traditional and the proposed method increased by 2.6% and 2.1%, with a variance reduced by 45.5% and 48.5%, respectively. Therefore, compared with the traditional robust optimization method, the proposed method reduces the maximum stress of the blade by 24.3% with almost the same improvement for aerodynamic performance means and disparities.

Multidisciplinary Design of an Electrically Powered High-Lift System

Abstract To date, design processes for electrically powered compressor are mainly based on separate processes for each individual component. Whereas the blading is often designed by an integrated aerodynamic and mechanical design optimization, additional components such as the electrical machine are usually not included. These approaches neglect the interactions of the individual components, which can influence the system performance. This paper demonstrates a multidisciplinary design approach, combining an optimization approach for a compressor stage and an electrical machine. The automated optimization process is based on an evolutionary algorithm, evaluating each individual of a population in terms of aerodynamic performance, structural integrity and performance of the electrical machine. This approach is applied to the design of a mixed-flow compressor for active high-lift applications in aircraft. The results suggest that the overall system efficiency is mainly influenced by the compressor stage, whereas the system mass is dominated by the electrical components which highlights the need to combine both optimization approaches. Key design parameters of high power-density electrical-machine designs are identified. A comparison between a previous compressor-only optimization and a new design based on the new multidisciplinary optimization confirms the improvements the latter optimization approach yields.

Numerical Study of Contra-Rotating Vertical Axis Wind Turbine H-Rotor Darrieus Type

The new Contra-rotating Darrieus turbine configuration has been invented to enhance the V ertical Axis Wind Turbine (VAWT) performance. This configuration increases the relative rotational speed of the generator, resulting in higher output power. It is well known that the increase can reach four times the output power. However, how the Darrieus turbine VAWT contra- rotating configuration influences its aerodynamic performance still needs to be discovered. This study investigates the aerodynamic performance of the contra-rotating configuration by comparing it to the single-rotating Darrieus turbine VAWT under the same conditions. The freestream speed is 5 m/s, with TSR varying from one to two intervals of 0.2. This research is being completed using Computational Fluid Dynamics (CFD) 3D cases with an Unsteady Reynold Average Navier-Stokes (URANS) equation as the turbulent model equation. The results of this study show that in terms of output power or Power coefficient (Cp), the contra-rotating has a greater value than the single- rotating configuration. However, in all TSR variations, contra-rotating outperforms single-rotating in terms of aerodynamic performance or moment coefficient (Cm). This is due to the fact that the aspect ratio of stage 1 contra-rotating rotor is lower than the single-rotating rotor, resulting in more significant blade tip losses in contra-rotating. Theflow was discovered through the gap between stages 1 and 2 contra-rotating, providing additional momentum. This phenomenon increases Cm at an azimuth angle of 200°-255°.

Wind Turbine Blade-Tip Optimization: A Systemic Computational Approach

Curved bladelets on wind turbine blades play an important role in improving the performance and efficiency of wind turbines. Implementing such features on the tip of wind turbine blades can improve their overall aerodynamic characteristics by reducing turbulence and loading without hindering lift generation and overall efficiency, thus leading to increased energy capture and reduced costs over the life of the turbine. Subjecting the integrated blade tip to optimization procedures can maximize its beneficial contribution to the assembly in general. Within this context, a systemic workflow is proposed for the optimization of a curved bladelet implemented on a wind turbine blade. The approach receives input in the form of an initial tip geometry and performs improvements in two distinct stages. Firstly, shape optimization is performed directly on the outer shape to enhance its aerodynamic properties. Subsequently, the topology of its interior structure is refined to decrease its mass while retaining its improved airflow characteristics. The proposed workflow aims to approach blade tip optimization holistically, both in terms of aerodynamic performance and structural capabilities; is computationally validated via fluid dynamics studies and finite element analysis to evaluate the performance augmentation achieved through it; and is further coupled with additive manufacturing for the production of prototype parts, benefiting from the manufacturing flexibility offered by digital fabrication technologies. The optimized bladelet model presented an approximate 30% improvement in the torque generated on it, while maintaining only 70% of its original mass, effectively contributing to a 0.81% increase to the total torque generated by the blade, consequently confirming the effectiveness of the proposed methodology.

Aerodynamic and structural multidisciplinary optimization design method of fan rotors based on blade curvature constraints

The optimization design of the actual fan/compressor rotors is a multidisciplinary problem. The aerodynamics and strength multidisciplinary optimization design of the fan/compressor rotors can improve the aerodynamic performance of rotors as much as possible under the condition of satisfying the structural strength requirements. However, the conventional multidisciplinary optimization of aerodynamics and structures will significantly increase the amount of computation cost and time period. Thus, it is difficult for conventional multidisciplinary optimization methods to be used in real-world engineering practices. In order to reduce the computational burden and time cost of the multidisciplinary optimization design, this paper firstly proposes a multidisciplinary optimization design method for rotors based on blade curvature constraints. In the optimization, the self-organizing map (SOM) method is applied to analyze and extract the constraint value of the blade curvature. And the blade curvature constraint value is used as a penalty function instead of the time-consuming high-fidelity FEM, which greatly reduces the computational cost and runtime of the multidisciplinary optimization. Besides, the Free-form deformation (FFD) method is adopted to realize the flexible deformation of the three-dimensional blade with the smallest number of design variables and polynomial chaos Kriging is used as a surrogate model for aerodynamic performance prediction in the optimization process. The method verification and mechanism analysis are carried out by studying a fan rotor. Data mining of FEM samples shows that, in the optimization design space, the maximum spanwise curvature of the blade is monotonically related to the maximum stress of the blade. And the maximum streamwise curvature, maximum spanwise slope, maximum streamwise slope, and maximum thickness of the blade have no related relationship with the maximum stress of the blade. Compared with the conventional multidisciplinary optimization design method based on the surrogate models of high-fidelity computational fluid dynamics (CFD) and finite element method (FEM), the time cost of the multidisciplinary optimization design method based on curvature constraints is reduced by 10 times, and the time cost of the multidisciplinary optimization is significantly reduced. The results show that, in terms of structural performance, the maximum stress of the aerodynamic optimization method without blade curvature constraints is 426 MPa, which exceeds the yield limit of aluminum alloy material (420 MPa), and the maximum stress of the multidisciplinary optimization design with blade curvature constraints proposed in this paper is 344 MPa. In terms of aerodynamic performance, the isentropic efficiency of the aerodynamic optimization method without blade curvature constraints and the method proposed in this paper are increased by 2.1% and 1.8%, respectively. Therefore, the proposed method reduces the maximum stress of the blade by 19.5% with little influence on the aerodynamic performance.

Implementation of various-fidelity methods for viscous effects modeling on the design of a waverider

In the present study, the implementation of three viscous effects modeling methods at hypersonic flow conditions is investigated. The design of a hypersonic passenger transport waverider configuration is employed as a case study. The methods, each of which has a different fidelity level, are evaluated in terms of their capability to predict aerodynamic performance, as well as their effect on the final geometric design of the waverider. More specifically, for the same mission profile and design constraints, three different waverider configurations (Cases A, B and C) are designed with the osculating cone methodology, and each employing a different viscous effect modeling method. These three configurations are designed (a) assuming a purely inviscid flowfield, (b) using Eckert's Reference Temperature and (c) Van Driest's viscous effects modeling semi-empirical methods respectively, each subsequent one of higher complexity and fidelity than the previous. An optimization loop, employing the Nelder-Mead simplex algorithm, is also incorporated into the overall design process. Each configuration is evaluated with all three viscous effects modeling methods and the analytical results are compared against those of CFD modeling. The CFD analyses are performed by solving the steady-state three-dimensional Euler and Reynolds Averaged Navier Stokes (RANS) equations. Overall, the three configurations are evaluated in terms of aerodynamic performance, design/optimization computational cost, and accuracy between the analytical and CFD results. All inviscid analytical calculations demonstrate very good agreement with the Euler CFD results. Indicatively, the lift-to-drag predictions deviate by as much as 2% between the two approached. The Van Driest's method predictions result to 5%-6.5% lower lift-to-drag ratios compared to the RANS CFD, and to higher wall shear stress values. Contrarily, the Ref. Temperature method predictions are in a much closer agreement to the RANS CFD results, leading to a slight overprediction (0%-2.4%) of lift-to-drag ratios. The incorporation of viscous effects in the design process reduces the waverider's aerodynamic efficiency by about 8.5%-11% and results in a 46%-98% increase in computational time, compared to the inviscid analysis.

Effects of Asymmetric Vane Pitch on Reducing Low-Engine-Order Forced Response of a Turbine Stage

Asymmetric vane pitch is a key technique to suppress the forced response of downstream rotor blades. To address the problem of low-engine-order (LEO) excitation with high amplitude under an asymmetric configuration (half-and-half layout) widely recognized in the previous literature, we first apply the in-house computational fluid dynamics code Hybrid Grid Aeroelasticity Environment to perform full-annulus unsteady aeroelasticity simulations of the turbine stage, comparing the resonance response of rotor blades on different asymmetric configurations and analyzing the flow field at the vane exit, as well as the excitation force, modal force, and maximum vibrational amplitude on the rotor blades. Second, we reveal that the potential field of the vane row is the main source of the LEO excitation caused by asymmetric configuration on rotor blades, the vane wake and potential field jointly determine the LEO excitation strength of rotor blades, and the vane pitch difference ΔS can be used to regulate the strength of the LEO excitation. Finally, based on an in-depth understanding of flow physics under an asymmetric configuration, a more preferable and effective asymmetric configuration (non-half two-segment layout) is proposed. Our findings demonstrate that, with the proposed asymmetric configuration, the amplitude of the vane passing frequency was reduced by 48.32% compared to the uniform configuration; furthermore, the maximum vibrational amplitude of the three-nodal-diameter response of the rotor blade at the three-engine-order crossing decreased by 45.49% compared to the half-and-half layout. The non-half two-segment layout also significantly improves upon the half-and-half layout in terms of aerodynamic performance. The results presented in this paper provide a good theoretical basis for reducing blade vibration by applying asymmetric vane pitch in engineering practice.

Unsteady Investigation of the Effect of Purge Air on the Flow Field Inside a Turbine Vane Frame Using Particle Image Velocimetry

Abstract Reducing greenhouse gas emissions and high fuel expenses motivates manufacturers to design more efficient aircraft engines. For directly driven jet engines, this is possible through increased by-pass ratios for high propulsive efficiencies. This implies a change in the operating condition of the low-pressure turbine (LPT) toward lower rotational speeds at larger diameters. Turbine vane frames (TVFs) guide the airflow from the high-pressure turbine (HPT) to the LPT in radial and circumferential direction. The TVF setup integrates turning vanes and thus removes the need for a vane blade row in the first LPT stage. Consequently, the TVF yields a benefit for overall engine weight and length, resulting in overall efficiency gains. Experimental measurements have been conducted at the two-spool test rig at the Graz University of Technology, consisting of a single-stage HPT, the TVF, and the first LPT rotor. Engine-relevant flow conditions are achieved at the TVF inlet, including HPT tip clearance and purge air effects. Particle image velocimetry (PIV) was used to capture the flow field in between two struts of the TVF upstream of the splitter vanes. Flow data in the area of strong transient interactions between the HPT and the TVF are recorded and discussed in terms of aerodynamic performance. To reveal the unsteady behavior of the fluid, the flow field has been recorded at six serial stator–rotor positions. Two data sets of varying HPT purge flows were obtained to characterize the effect of purge air inside the measurement domain.

Aerodynamic characteristics of co-flow jet wing with simple high-lift devices

Numerical investigations are conducted to explore the aerodynamic characteristics of three-dimensional Co-Flow Jet (CFJ) wing with simple high-lift devices during low-speed takeoff and landing. Effects of three crucial parameters of CFJ wing, i.e., angle of attack, jet momentum and swept angle, are comprehensively examined. Additionally, the aerodynamic characteristics of two CFJ configurations, i.e., using open and discrete slots for injection, are compared. The results show that applying CFJ technique to a wing with simple high-lift device is able to generate more lift, reduce drag and enlarge stall margin with lower energy expenditure due to the super-circulation effect. Increasing the jet intensity can reduce the drag significantly, which is mainly contributed by the reaction jet force. The Oswald efficiency factor is, in some circumstances, larger than one, which indicates the potential of CFJ in reducing induced drag. Compared with clean wing configuration, using CFJ technique allows the aerodynamic force variation less sensitive to the swept angle, and such phenomenon is better observed for small swept angle region. Eventually, it is interesting to know that the discrete slotted CFJ configuration demonstrates a promising enhancement in aerodynamic performance in terms of high lift, low drag and efficiency.

Aerodynamic Shape Optimisation of a Camber Morphing Airfoil and Noise Estimation

In order to decrease the emitted airframe noise by a two-dimensional high-lift configuration during take-off and landing performance, a morphing airfoil has been designed through a shape design optimisation procedure starting from a baseline airfoil (NLR 7301), with the aim of emulating a high-lift configuration in terms of aerodynamic performance. A methodology has been implemented to accomplish such aerodynamic improvements by means of the compressible steady RANS equations at a certain angle of attack, with the objective of maximising its lift coefficient up to equivalent values regarding the high-lift configuration, whilst respecting the imposed structural constraints to guarantee a realistic optimised design. For such purposes, a gradient-based optimisation through the discrete adjoint method has been undertaken. Once the optimised airfoil is achieved, unsteady simulations have been carried out to obtain surface pressure distributions along a certain time-span to later serve as the input data for the aeroacoustic prediction framework, based on the Farassat 1A formulation, where the subsequent results for both configurations are post-processed to allow for a comparative analysis. Conclusively, the morphing airfoil has proven to be advantageous in terms of aeroacoustics, in which the noise has been reduced with respect to the conventional high-lift configuration for a comparable lift coefficient, despite being hampered by a significant drag coefficient increase due to stall on the morphing airfoil’s trailing edge.

Effect of the wing airfoil shape on the aerodynamics and performance of a jet-trainer aircraft – An optimization approach

Among the key factors in developing the performance of military aircraft are its aerodynamic characteristics and performance. This research presents the effect of shape of the wing airfoil on the aerodynamic characteristics and performance of the popular jet trainer aircraft L-39C. The aerodynamic data of different airfoil shapes were used to determine the aerodynamic characteristics and performance of the L-39C for different airfoil shapes in an effort to optimize the aircrafts aerodynamic and performance. NACA 64A012 airfoil is currently used on the L-39C, however, there may exist many airfoils that may have potential to improve the aerodynamic characteristics and hence the aircraft performance. For this purpose, a group of NACA airfoils are selected, namely, NACA 4412, NACA 2415, NACA 64212 and NACA 64215. Each of these airfoil influences the aerodynamic characteristics of the aircraft and hence its performance. For each airfoil, the aircraft performance in terms of thrust required, power required, and rate of climb at different altitudes and airspeeds are calculated and analysed. The airfoil data were calculated at cruising flight at zero angle of attack to reduce the variables that can affect the calculations. The results of the calculation and analysis showed that NACA 4412 has significant results in terms of aerodynamic characteristics although in terms of aerodynamic performance, the NACA 64A012 and NACA 4412 showed lower thrust required and power required. NACA 4412 has a (C L )max of 1.67, whereas NACA 64A012 has a (C L )max of 1.336, indicating that the airfoil stalled early at higher altitudes. NACA 4412 also showed better results in terms of aerodynamic characteristics compared to the other selected airfoils although NACA 64A012 shows some variance in the results. NACA 4412 and NACA 64A212 have shown to be promising in aerodynamic characteristics and performance where one has its own benefits over the other. Although in the end, NACA 4412 may be recommended for its aerodynamic characteristics and performance.

CFD analysis of variable geometric angle winglets

PurposeWinglets play a major role in saving fuel costs because they reduce the lift-induced drag formed at the wingtips. The purpose of this paper is to obtain the best orientation of the winglet for the Office National d’Etudes et de Recherches Aérospatiales (ONERA) M6 wing at Mach number 0.84 in terms of lift to drag ratio.Design/methodology/approachA computational fluid dynamics analysis of the wing-winglet configuration based on the ONERA M6 airfoil on drag reduction for different attack angles at Mach 0.84 was performed using analysis of systems Fluent. First, the best values of cant and sweep angles in terms of aerodynamic performance were selected by performing simulations. The analysis included cant angle values of 30°, 40°, 45°, 55°, 60°, 70° and 75°, while for the sweep angles 35°, 45°, 55°, 65° and 75° angles were used. The aerodynamic performance was measured in terms of the obtained lift to drag ratios.FindingsThe results showed that slight alternations in the winglet configuration can improve aerodynamic performance for various attack angles. The best lift to drag ratio for the winglet was achieved at a cant angle of 30° and a sweep angle of 65°, which caused a 5.33% increase in the lift to drag ratio. The toe-out angle winglets as compared to the toe-in angles caused the lift to drag ratio to increase because of more attached flow at its surface. The maximum value of the lift to drag ratio was obtained with a toe-out angle (−5°) at an angle of attack 3° which was 2.53% greater than the zero-toed angle winglet.Originality/valueThis work is relatively unique because the cant, sweep and toe angles were analyzed altogether and led to a significant reduction in drag as compared to wing without winglet. The wing model was compared with the results provided by National Aeronautics and Space Administration so this validated the simulation for different wing-winglet configurations.

Effect of Ducted Multi-Propeller Configuration on Aerodynamic Performance in Quadrotor Drone

Motivated by a bioinspired optimal aerodynamic design of a multi-propeller configuration, here we propose a ducted multi-propeller design to explore the improvement of lift force production and FM efficiency in quadrotor drones through optimizing the ducted multi-propeller configuration. We first conducted a CFD-based study to explore a high-performance duct morphology in a ducted single-propeller model in terms of aerodynamic performance and duct volume. The effect of a ducted multi-propeller configuration on aerodynamic performance is then investigated in terms of the tip distance and the height difference of propellers under a hovering state. Our results indicate that the tip distance-induced interactions have a noticeable effect in impairing the lift force production and FM efficiency but are limited to small tip distances, whereas the height difference-induced interactions have an impact on enhancing the aerodynamic performance over a certain range. An optimal ducted multi-propeller configuration with a minimal tip distance and an appropriate height difference was further examined through a combination of CFD simulations and a surrogate model in a broad-parameter space, which enables a significant improvement in both lift force production and FM efficiency for the multirotor, and thus provides a potential optimal design for ducted multirotor UAVs.

Numerical and Experimental Study of Aerodynamic Performances of a Morphing Micro Air Vehicle

The present work is focused on the investigation of the aerodynamic performances of a novedous bioinspired morphing Micro Air Vehicle (MAV) with an adaptive wing structure geometry. For this purpose, a numerical study of Computational Fluid Dynamics (CFD) implemented by Ansys Fluent 15.0 was performed in order to obtain insight about the aerodynamic effect of wing structure deformation when morphing devices are used, and its influence on the global aerodynamic parameters related with aircraft performances. On the other hand, an experimental study using the Particle Image Velocimetry technique and balance measurements in a Low-Speed Wind Tunnel was conducted to obtain experimental information about performances measured to establish a comparison between both, experimental and numerical results. The Micro Air Vehicle (MAV) presents a Zimmerman wing with an Eppler 61 airfoil. Three different wing configurations according to curvature and thickness variations and for all angles of attack have been studied. A comparative analysis based on aerodynamic features is performed by an assessment of lift coefficient (CL), total aerodynamic drag coefficient (CD) and aerodynamic efficiency as lift/drag ratio (CL/CD) in order to conclude the best wing configuration in terms of aerodynamic performance.

A novel linearization approach of chord and twist angle distribution for 10 kW horizontal axis wind turbine

In the present work, the aerodynamic design of a 10 kW horizontal axis wind turbine rotor is performed using both ideal and actual rotor theories. The obtained nonlinear blade profile is optimized in order to enhance the aerodynamic performance and to ease the fabricating complexity. A unique linearization approach is employed to linearize the chord and twist distribution by dividing the congruent line of both ideal and actual models into equal divisions. The points along the identical tangent line are considered as floating new blade roots, whereas the blade tip was kept fixed based on the primary design. The linear profile based on the new value of blade root is described using algebraic equations. The local element torque, capacity factor, and the annual energy production based on the Weibull distribution are adopted to determine and assess the optimal blade profile. Both CFD investigation and the FEA analysis are performed in order to evaluate both primary and optimized blades. The aerodynamic and aeroelastic comparison in terms of power output, thrust force, blade tip deflection, and the equivalent stress distribution of both blade profiles has been done over a wide range of incoming wind speeds. Results show an enhancement in the aerodynamic performance in terms of power coefficient up to (5.9%) compared to the primary blade design. Moreover, the optimized blade has shown less tip deflection by 27.92% than the primary blade at low wind speed.

Numerical investigation of effect of crossflow transition on rotor blade performance in hover

In the present study, the effect of crossflow transition on the rotor blade performance in hover was investigated numerically. The simulations were conducted for the Pressure Sensitive Paint (PSP) rotor in hover by using a Reynolds-averaged Navier-Stokes (RANS) solver based on unstructured meshes. The k−ωSST model was used for simulating the turbulent flow fields and for calculating the turbulent eddy viscosity. For predicting the laminar-turbulent onset phenomena involving crossflow-induced transition, the γ−Reθt−CF+ transition model was adopted. For the comparison of the transition locations without considering the crossflow transition effect, simulations were also carried out using the γ−Reθt transition model. The calculations were made for the PSP rotor at collective pitch angles from four to 12 degrees with an interval of two degrees. To investigate the effect of blade tip Mach number, two blade tip Mach numbers of 0.585 and 0.65 were tested. The predicted results such as the transition onset location and the rotor aerodynamic performance in terms of thrust coefficient, torque coefficient and figure of merit were compared with experimental data. It was found that proper consideration of the effect of crossflow transition is critical for the accurate prediction of the laminar-turbulent transition onset location on rotor blades. The figure of merit also compares better with the experimental data when the effect of laminar-turbulent transition is included. With the addition of the crossflow transition, the figure of merit is slightly decreased, particularly at low thrust levels.

An Additively Manufactured Four-Sensor Fast Response Aerodynamic Probe

Abstract This study describes the design, development, and testing of a miniature fast response aerodynamic probe (FRAP) with four sensors (4S), which are able to perform measurements in the unsteady three-dimensional flow field. Moreover, the calibration and first results with the newly developed probe are provided. The miniature FRAP-4S demonstrates a 3 mm tip diameter, offering a 25% reduction in diameter size, in comparison to a first-generation FRAP-4S, without any loss in terms of measurement bandwidth. The 3 mm outer casing of the probe is additively manufactured with a high-precision binder jetting technique. In terms of aerodynamic performance, the probe demonstrates high angular sensitivity up to ± 18 deg incidence angle in both directions. To evaluate the measurement accuracy of the newly developed FRAP-4S, measurements are performed at the Laboratory for Energy Conversion (LEC) in both a round axisymmetric jet and an one-and-a-half stage, unshrouded and highly loaded axial turbine configuration. Turbulence measurements performed with the miniature FRAP-4S are compared against hot-wire studies in round free-jets found in the literature. Good agreement in both trends but also absolute values is demonstrated. Moreover, the performance of the probe is compared against traditional instrumentation developed at LEC, namely, miniature pneumatic and FRAP-2S probes. The results indicate that the FRAP-4S, despite its larger size in comparison to the other probes tested, can resolve the main flow patterns, with the highest deviations occuring in the presence of highly skewed and sheared flow. Furthermore, the additively manufactured probe was proven to be robust after more than 50 hours of testing in the representative turbine environment configuration. Finally, it should be highlighted that the newly developed FRAP reduces measurement time by a factor of three in comparison to FRAP-2S, which directly translates to reduced development time and thus cost during the turbomachinery development phase.