Lift and Drag Forces of a High Efficiency Airfoil With an Embedded Rotating Cylinder

2016 ◽  
Author(s):  
Komal Gada ◽  
Hamid Rahai
Keyword(s):  
Flow Separation ◽  
High Efficiency ◽  
Active Flow Control ◽  
Tangential Velocity ◽  
Axial Flow ◽  
Flow Simulation ◽  
Leading Edge ◽  
Rotating Cylinder ◽  
Chord Length ◽  
Mean Velocity

Numerical investigations of an optimized thin airfoil with an active flow control device (rotating cylinder) embedded into the airfoil have been performed. The objective of the study was to investigate the possibility of using a rotating cylinder to maintain performance of micro aerial vehicles, MAVs, when significant and sudden variation in wind speed (example: gust) is present. The airfoil has a chord length of 19.66 cm and a span of 25 cm. The free stream mean velocity was 20 m/s which corresponds to a chord length Reynolds number of 2.54×105. Simulations were performed at 17 degrees angle of attack which include the initial angle that the cambered leading edge makes with the incoming axial flow. Simulation results for the airfoil without the embedded cylinder have shown flow separation at approximately 85% chord length. Then, a rotating cylinder with a 0.51 cm diameter was embedded into the airfoil, spanning the width of the airfoil at slightly downstream of the location of flow separation, i.e. at x/c = 0.848. There was 1 mm spacing between the cylinder and the airfoil, to allow cylinder’s rotation. Investigations were performed at different rotation speeds, corresponding to corresponding tangential velocities being higher than, equal to and less than the local axial freestream mean velocity. Results showed approximately 10% improvement in lift to drag ratio (L/D), when the tangential velocity was the same or higher than the local axial mean velocity.

scholarly journals Numerical Analysis of Effect of Leading-Edge Rotating Cylinder on NACA0021 Symmetric Airfoil

2019 ◽  
Vol 4 (7) ◽  
pp. 11-17
Author(s):  
Md. Abdus Salam ◽  
Vikram Deshpande ◽  
Nafiz Ahmed Khan ◽  
M. A. Taher Ali
Keyword(s):  
Free Stream ◽  
Numerical Study ◽  
Tangential Velocity ◽  
Leading Edge ◽  
Rotating Cylinder ◽  
Aerodynamic Performance ◽  
Stream Velocity ◽  
Surface Boundary ◽  
Lower Velocity ◽  
Numerical Studies

The moving surface boundary control (MSBC) has been a Centre stage study for last 2-3 decades. The preliminary aim of the study was to ascertain whether the concept can improve the airfoil characteristics. Number of experimental and numerical studies pointed out that the MSBC can superiorly enhance the airfoil performance albeit for higher velocity ratios (i.e. cylinder tangential velocity to free stream velocity). Although abundant research has been undertaken in this area on different airfoil performances but no attempt was seen to study effect of MSBC on NACA0021 airfoil for and also effects of lower velocity ratios. Thus, present paper focusses on numerical study of modified NACA 0021 airfoil with leading edge rotating cylinder for velocity ratios (i.e.) between 1 to 1.78 at different angles of attack. The numerical study indicates that the modified airfoil possess better aerodynamic performance than the base airfoil even at lower velocity ratios (i.e. for velocity ratios 0.356 and beyond). The study also focusses on reason for improvement in aerodynamic performance by close look at various parameters.

scholarly journals Influence of Chord Lengths of Splitter Blades on Performance of Small Axial Flow Fan

2015 ◽  
Vol 9 (1) ◽  
pp. 361-370
Author(s):  
Guoqi Li ◽  
Lifu Zhu ◽  
Yongjun Hu ◽  
Yingzi Jin ◽  
Toshiaki Setoguchi ◽  
...  
Keyword(s):  
Flow Characteristics ◽  
Axial Flow ◽  
Internal Flow ◽  
Leading Edge ◽  
Chord Length ◽  
Aerodynamic Noise ◽  
Axial Fan ◽  
Axial Flow Fan ◽  
Suction Surface ◽  
Small Axial Flow Fan

On the basis of small axial fan with five blades, 6 types of small axial flow fans with different chord lengths splitter blades were designed. Numerical simulation of 6 fan models with splitter blades and prototype fan were done by using Fluent. Based on the obtained simulation results, internal flow characteristics and aerodynamic noise were analyzed and compared. It indicates that: splitter blades with suitable chord length have improved significantly on internal flow characteristics, which inhibits backflow from pressure surface to the suction surface at blade tip and leading edge and restrains flow separation. The 6 model fans are better than prototype fan on aerodynamic noise improvement, but too long or too short chord lengths are both disadvantage to improve aerodynamic noise. The results reveal that 2/6, 3/6 and 4/6 chord length model have relatively better acoustic characteristics and internal flow characteristics. The research program will offer a reference for structural improvements and noise reduction on small axial flow fan.

Aerodynamic Characteristics of a Blended-Wing-Body Aircraft With A Serpentine Inlet Using Flow Control Techniques

2021 ◽  
Author(s):  
Min-Sik Youn ◽  
Youn-Jea Kim
Keyword(s):  
Flow Control ◽  
Flow Separation ◽  
High Efficiency ◽  
Active Flow Control ◽  
Quantitative Measure ◽  
Aerodynamic Characteristics ◽  
Interface Plane ◽  
Flow Distortion ◽  
Blended Wing Body ◽  
Active Flow

Abstract Demands of a modern aircraft regarding its aerodynamic performance and high efficiency are ever-growing. An S-shaped inlet, as known as a serpentine duct, plays a significant role in increasing fuel efficiency. Recently, the serpentine duct is commonly employed for military aircraft to block the front of the jet engine from radar. However, delivering a non-uniformly distorted flow to the engine face (aerodynamic interface plane, AIP) though a serpentine duct is inevitable due to the existence of flow separation and swirl flow in the duct. The effect of distortion is to cause the engine compressor to surge; thus, it may impact on the life-cycle of aircraft engine. In this study, aerodynamic characteristics of a serpentine duct mounted on a blended-wing-body (BWB) aircraft was thoroughly investigated to determine where and how the vortex flow was generated. In particular, both passive and active flow control were implemented at a place where the flow separation was occurred to minimize the flow distortion rate in the duct. The passive and active flow control systems were used with vortex generator (VG) vanes and air suctions, respectively. A pair of VG s have been made as a set, and 6 sets of VG in the serpentine duct. For the active flow control, 19 air suctions have been implemented. Both flow control devices have been placed in three different locations. To evaluate the performance of flow control system, it is necessary to quantify the flow uniformity at the AIP. Therefore, coefficient of distortion, DC(60) was used as the quantitative measure of distortion. Also, change in DC(60) value while the BWB aircraft is maneuvering phase was analyzed.

Numerical Investigation on Effect of Compressor Performance in Single Rotor With Micro-Vortex Generator

2017 ◽  
Author(s):  
Shan Ma ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Jinhua Lang ◽  
Haiyang Kuang
Keyword(s):  
Secondary Flow ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Axial Flow ◽  
Flow Simulation ◽  
Vortex Generator ◽  
Leading Edge ◽  
Future Research ◽  
Characteristic Line ◽  
Design Speed

The performance of axial compressor is considerably influenced by secondary flow, like corner separation between wall and blade in a compressor stage. An extensive experimental study of vortex generator (VG) applied on axial compressor was conducted by many scholars, in order to control these effects and improve the aerodynamic performance. According to their size, they are classified as traditional VGs (h/δ>0.5) and Micro-vortex generators (MVGs, h/δ = 0.1∼0.5).MVGs is one of the hot spots of present research to restrain the secondary flow. In order to investigate the effect of MVGs used in rotor, this study was carried out on Northwestern Polytecnical University rotor (NPU rotor), which is a subsonic axial flow compressor rotor. The Vane-MVGs were placed at a distance of 11% chord length ahead of the leading edge on the end-wall. The characteristic line of 54% (8130RPM), 71% (10792RPM) and 84% (12768RPM) design speed were calculated by steady 3D RANS simulations with Spalart-Allmar turbulence model and compared with the corresponding MVGs cases, respectively. Results show that the stall margins of the 3 speeds with MVGs were larger than baseline, but the efficiency and pressure ratio were reduced in different degrees. In this paper, the flow characteristics at 54% (8130RPM) design speed and the development process of vortex generated by MVGs are analyzed in detail. The influence of MVGs height and stagger angle on rotor performance is also discussed. Moreover, flow simulation of MVGs used on axial compressor single rotor’s hub offered a guideline to future research.

scholarly journals Active Flow Control over a NACA23012 Airfoil using Hybrid Jet

2021 ◽  
Vol 71 (6) ◽  
pp. 721-729
Author(s):  
Deepak Kumar Singh ◽  
Anuj Jain ◽  
Akshoy Ranjan Paul
Keyword(s):  
Flow Control ◽  
Flow Separation ◽  
Active Flow Control ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Navier Stokes ◽  
Flow Separation Control ◽  
Blowing Ratio ◽  
Maximum Improvement ◽  
Stall Angle

A time-dependent numerical simulation is performed to examine the flow separation control with the action of a hybrid jet (the combination of synthetic and continuous jets) over a NACA23012 airfoil. The unsteady Reynolds-averaged Navier–Stokes (URANS) simulation is performed with Spalart-Allmaras (SA) turbulence model to simulate the flow field around the airfoil to analyse the effect of the hybrid jet. A combined jet is placed at the point of flow separation on the upper surface of the airfoil which is located at the 12% of the chord length from the leading edge of the airfoil for a given flow configuration. Flow simulations are performed at a chord-based Reynolds number of 2.19 × 106 for the hybrid jet oscillating frequency of 0.159 at a blowing ratio of 3.0. The contribution of the continuous jet in the hybrid jet is evident by the flow control. Variation in the continuous jet velocity is studied, which improved the aerodynamic characteristics of the airfoil. The maximum improvement in lift to drag ratio is observed from 11.19 to 22.14 at an angle of attack of 22 degree. The stall angle also shows an enhancement from 18 degree to 20 degree.

Effects of roughness on the performance of axial flow cyclone separators using numerical simulation method

2019 ◽  
Vol 233 (7) ◽  
pp. 914-927
Author(s):  
Li Jun ◽  
Chunyuan Ma ◽  
Wang Tao ◽  
Jingcai Chang ◽  
Xiqiang Zhao
Keyword(s):  
Numerical Simulation ◽  
High Efficiency ◽  
Separation Efficiency ◽  
Tangential Velocity ◽  
Axial Flow ◽  
Simulation Method ◽  
Roughness Height ◽  
Wall Roughness ◽  
Numerical Simulation Method ◽  
Axial Flow Cyclone

An axial flow cyclone is a separator with high efficiency and low resistance. Researchers have extensively studied the structure and parameters that have the greatest influence on its performance. However, the influence of wall roughness on the performance of axial flow cyclones has been neglected for a long time. The wall roughness height can be changed by the manufacturing process and the effect of particles on the wall. Thus, in this study, the effects of roughness on an axial flow cyclone are investigated using a numerical simulation method. The Reynolds stress model and discrete phase model are used for gas and particle prediction and the simulation result were verified through experimentation. The results of the numerical simulation show that the roughness height has big influence on axial flow cyclones. The separation efficiency decreases and static pressure drop increases with increasing roughness height. This happens especially at high inlet velocity. The tangential velocity decreases, particularly near the inner surface of the cyclone, and axial velocity increases in the center of the pipe. The trends show that the degree of change reduced for all parameters with increasing roughness height.

Dynamic Characteristics of Flow Separation From a Low Reynolds Number Airfoil

2007 ◽  
Author(s):  
Daniel R. Morse ◽  
James A. Liburdy
Keyword(s):  
Reynolds Number ◽  
Flow Separation ◽  
Large Scale ◽  
Region Of Interest ◽  
Leading Edge ◽  
Chord Length ◽  
Vortical Flow ◽  
Local Circulation ◽  
Time Resolved ◽  
Freestream Velocity

This study examines the generation of large scale vortices caused by flow separation from a flat wing at various angles of attack. Time-resolved particle image velocimetry is used to determine the evolution and convective characteristics of the large scale structures. A rectangular airfoil with aspect ratio of 0.5 is used and data are collected at a Reynolds number of 23,500, for angles of attack from 0° to 20°. Data consists of two dimensional velocity fields obtained at 500 Hz located at the airfoil centerline. The region of interest is near the separation point but fields of view extend over approximately one half of the chord length from the leading edge to document the downstream progression of the large scale vortical flow elements. The velocity data were processed to identify the vorticity field dynamics in terms of the Kelvin-Helmholtz instability occurring near the leading edge. The vortical structures are identified using vortex detection based on local circulation. The convective nature of the vortex elements are shown to consist of merging, stalling and convecting, with convective velocities on the order of 20% of the freestream velocity with an associated Stouhal number based on chord length and freestream velocity of approximately 1.0.

Oscillating behaviour of laminar separation bubble formed on an aerofoil near stall

2004 ◽  
Vol 108 (1081) ◽  
pp. 153-163 ◽  
Author(s):  
K. Rinoie ◽  
N. Takemura
Keyword(s):  
Separation Bubble ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Chord Length ◽  
Flow Visualisation ◽  
Laminar Separation Bubble ◽  
Mean Velocity ◽  
Laminar Separation ◽  
Naca 0012 ◽  
Small Separation

Abstract Laminar separation bubbles formed on NACA 0012 aerofoil near the onset of a stall were investigated to clarify the behaviour of the laminar separation bubble. Measurements were done at a chord Reynolds number of 1·3 × 105. Mean velocity measurements indicate that the long bubble of about 35% chord length is formed at α = 11·5° after the short bubble burst occurred. However, the instantaneous flow visualisation picture indicates that the flow is strongly oscillating at this angle of attack. The phase averaging technique has been applied to analyse this oscillating behaviour. The results indicate that the flow is oscillating between a small separation-reattachment bubble formed near the leading-edge at about a 10% chord length and a large separated region extending over the aerofoil surface. It is suggested that this small separation-reattachment bubble has a similar flow structure to that of the short bubble formed at a lower angle of attack.

Fouling Influence Factor Analysis of Axial Flow Compressor

2012 ◽  
Vol 516-517 ◽  
pp. 619-622
Author(s):  
Hua Dong Yang ◽  
Hong Xu
Keyword(s):  
Influence Factor ◽  
Axial Flow ◽  
Flow Simulation ◽  
Leading Edge ◽  
Performance Parameters ◽  
Suction Surface ◽  
Axial Flow Compressor ◽  
Degradation Factor ◽  
Main Factors ◽  
Flow Compressor

Fouling is an important performance degradation factor of axial flow compressor. In order to reveal fouling mechanism, flow simulation of four cases of NASA rotor37 has been performed, such as clean compressor, compressor with roughness of 50μm, 100μm and 150μm. Thermodynamics performance parameters of compressor at different rotational speed with different roughness are discussed. And then fouling sources and its influence factor are analyzed. Research finds that gas contaminants can deposit in blade surface existed with internal oil and water. Simultaneously, humidity, temperature, flow velocity and contact area are the main factors for the formation of compressor fouling. Finally, research finds that fouling can easily formed in suction surface and fouling level of leading edge is more critical than other locations.

Fan tonal noise from aircraft aeroengines with short intake: A study at approach

2018 ◽  
Vol 17 (6-8) ◽  
pp. 600-623 ◽  
Author(s):  
S Guérin ◽  
A Holewa
Keyword(s):  
Flow Separation ◽  
Power Level ◽  
Leading Edge ◽  
Radial Component ◽  
Moderate Increase ◽  
Sound Waves ◽  
Mean Flow ◽  
Mean Velocity ◽  
Rotor Stator Interaction ◽  
The Mean

This work assesses the risks of increased fan noise for high bypass ratio aeroengines with short intakes. The close proximity between the fan and inlet contributes to the increase in radiation of the rotor-alone tones and reinforces the interaction of the inflow distortion with the fan. Thus, the closer the fan is to the inlet, the higher the risk for noise generation. This article discusses the results of Harmonic Balance simulations performed on a conceptual turbofan operated at the approach condition. The inflow distortion created by the nacelle incidence is dominated by the circumferential component [Formula: see text] = 1. Its presence is visible throughout the nacelle. A thorough analysis of the unsteady pressure and velocity fields shows that the new acoustic source created by the periodic unsteady loading of the rotor cutting the inflow distortion is negligible compared to the rotor–stator interaction. But the amplitude of the rotor–stator interaction tones is affected by the unsteadiness of the rotor wake shape, particularly in the tip region where a pronounced flow separation on the rotor blade is created at a certain range of azimuthal position. The variations of the flow incidence at the rotor leading edge, due to the axial and tangential components of the mean velocity, cannot explain that flow separation. Instead, the origin is attributed to the azimuthal variations of the radial component of the mean flow velocity near the casing which slightly points inward to the spinner, i.e. in the opposite direction of the casing contour line. The flow separation induces a pronounced scattering of the wake azimuthal components mw =  hB into [Formula: see text] in the tip region, whereas the same effect is rather limited on the rest of the blade height. This leads to a moderate increase of the tonal sound power level compared to the case with clean inflow. The azimuthal scattering due to the propagation of the sound waves through the distortion is found to be weak in the bypass duct. However, this effect is very important in the inlet lip vicinity, where the strong asymmetry of the flow modifies the path of the sound waves up to the far field.

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