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.

New combinational active control strategy for improving aerodynamic characteristics of airfoil and rotor

2019 ◽  
Vol 234 (4) ◽  
pp. 977-996
Author(s):  
Yi-yang Ma ◽  
Qi-jun Zhao ◽  
Guo-qing Zhao
Keyword(s):  
Flow Control ◽  
Control Strategy ◽  
Active Flow Control ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Synthetic Jet ◽  
Dynamic Stall ◽  
Trailing Edge ◽  
Trailing Edge Flap ◽  
Cfd Method

In order to improve the aerodynamic characteristics of rotor, a new active flow control strategy by combining a synthetic jet actuator and a variable droop leading-edge or a trailing-edge flap has been proposed. Their control effects are numerically investigated by computational fluid dynamics (CFD) method. The validated results indicate that variable droop leading-edge and synthetic jet can suppress the formation of dynamic stall vortex and delay flow separation over rotor airfoil. Compared with the baseline state, Cdmax and Cmmax are significantly reduced. Furthermore, parametric analyses on dynamic stall control of airfoil by the combinational method are conducted, and it indicates that the aerodynamic characteristics of the oscillating rotor airfoil can be significantly improved when the non-dimensional frequency ( k*) of variable droop leading-edge is about 1.0. At last, simulations are conducted for the flow control of rotor by the combinational method. The numerical results indicate that large droop angle of variable droop leading-edge can better reduce the torque coefficient of rotor and the trailing-edge flap has the capability of increasing the thrust of rotor. Also, the synthetic jet could further improve the aerodynamic characteristics of rotor.

Flow separation control of NACA-2412 airfoil with bio-inspired nose

2019 ◽  
Vol 91 (7) ◽  
pp. 1058-1066 ◽  
Author(s):  
Mohamed Arif Raj Mohamed ◽  
Ugur Guven ◽  
Rajesh Yadav
Keyword(s):  
Flow Control ◽  
Flow Separation ◽  
Control Method ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Separation Control ◽  
Cetacean Species ◽  
Aerodynamic Efficiency ◽  
Content Type ◽  
Flow Separation Control

Purpose The purpose of this paper is to achieve an optimum flow separation control over the airfoil using passive flow control method by introducing bio-inspired nose near the leading edge of the NACA 2412 airfoil. Design/methodology/approach Two distinguished methods have been implemented on the leading edge of the airfoil: forward facing step, which induces multiple accelerations at low angle of attack, and cavity/backward facing step, which creates recirculating region (axial vortices) at high angle of attack. Findings The porpoise airfoil (optimum bio-inspired nose airfoil) delays the flow separation and improves the aerodynamic efficiency by increasing the lift and decreasing the parasitic drag. The maximum increase in aerodynamic efficiency is 22.4 per cent, with an average increase of 8.6 per cent at all angles of attack. Research limitations/implications The computational analysis has been done for NACA 2412 airfoil at low subsonic speed. Practical implications This design improves the aerodynamic performance and increases structural strength of the aircraft wing compared to other conventional high-lift devices and flow-control devices. Originality/value Different bio-inspired nose designs which are inspired by the cetacean species have been analysed for NACA 2412 airfoil, and optimum nose design (porpoise airfoil) has been found.

Flow separation control using a bio-inspired nose for NACA 4 and 6 series airfoils

2021 ◽  
Vol 93 (2) ◽  
pp. 251-266
Author(s):  
Mohamed Arif Raj Mohamed ◽  
Rajesh Yadav ◽  
Ugur Guven
Keyword(s):  
Flow Control ◽  
Flow Separation ◽  
Control Method ◽  
Leading Edge ◽  
Aircraft Design ◽  
Separation Control ◽  
Cetacean Species ◽  
Aerodynamic Efficiency ◽  
Content Type ◽  
Flow Separation Control

Purpose This paper aims to achieve an optimum flow separation control over the airfoil using a passive flow control method by introducing a bio-inspired nose near the leading edge of the National Advisory Committee for Aeronautics (NACA) 4 and 6 series airfoil. In addition, to find the optimised leading edge nose design for NACA 4 and 6 series airfoils for flow separation control. Design/methodology/approach Different bio-inspired noses that are inspired by the cetacean species have been analysed for different NACA 4 and 6 series airfoils. Bio-inspired nose with different nose length, nose depth and nose circle diameter have been analysed on airfoils with different thicknesses, camber and camber locations to understand the aerodynamic flow properties such as vortex formation, flow separation, aerodynamic efficiency and moment. Findings The porpoise nose design that has a leading edge with depth = 2.25% of chord, length = 0.75% of chord and nose diameter = 2% of chord, delays the flow separation and improves the aerodynamic efficiency. Average increments of 5.5% to 6° in the lift values and decrements in parasitic drag (without affecting the pitching moment) for all the NACA 4 and 6 series airfoils were observed irrespective of airfoil geometry such as different thicknesses, camber and camber location. Research limitations/implications The two-dimensional computational analysis is done for different NACA 4 and 6 series airfoils at low subsonic speed. Practical implications This design improves aerodynamic performance and increases the structural strength of the aircraft wing compared to other conventional high lift devices and flow control devices. This universal leading edge flow control device can be adapted to aircraft wings incorporated with any NACA 4 and 6 series airfoil. Social implications The results would be of significant interest in the fields of aircraft design and wind turbine design, lowering the cost of energy and air travel for social benefits. Originality/value Different bio-inspired nose designs that are inspired by the cetacean species have been analysed for NACA 4 and 6 series airfoils and universal optimum nose design (porpoise airfoil) is found for NACA 4 and 6 series airfoils.

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.

scholarly journals The Suction Control Characteristics of Flow Separation on NACA 23012

2020 ◽  
Vol 2 (1) ◽  
pp. 1-15
Author(s):  
Moses Oluwatosin Julius ◽  
Saheed Adewale Adio ◽  
Adam Olatunji Muritala ◽  
Oluwasanmi Alonge
Keyword(s):  
Flow Separation ◽  
Aerodynamic Characteristics ◽  
Control Technique ◽  
Navier Stokes ◽  
Suction Control ◽  
Lift Enhancement ◽  
Stall Angle ◽  
The Impact ◽  
Drag Ratio ◽  
Lift To Drag Ratio

The enormous loss of momentum leads to stall and adversely affects the aerodynamic performance of aeroplane wings which may lead to a disaster, more importantly, risking the safety of the aeroplane by putting lives of passengers on it in danger. Therefore, this paper focuses on the enhancement of aerodynamic characteristics of NACA 23012 through the mitigation of flow separation and delay of the stall at higher angles of attack by using suction for Reynolds number (Re) = 3.4 x 106 . Considering the different suction features such as suction width, suction position, and suction coefficient, the separation delay capability of a suction control is studied. Also, the lift to drag ratio and the impact of energy consumption variation during the control technique are used for estimating the control effects. The Reynolds Average Navier-Stokes (RANS) equations are employed together with the Menter’s shear stress turbulent model. The result of this study revealed that the jet position just behind the separation point at 0.2 % of the chord length shows an outstanding control outcome on the separation and stall, thereby increasing the lift. The lift to drag ration increased proportionately when the suction jet coefficient was increased. At suction coefficient Cq = 0.00225, a 92.1% drag reduction and 72.7% lift enhancement is observed. Hence, the stall angle is moved beyond 21.5o from an initial angle of 16° and the more energy was saved at a high angle of attack.

Helicopter Fuselage Model Drag Reduction by Active Flow Control Systems

2019 ◽  
Vol 64 (2) ◽  
pp. 1-15 ◽  
Author(s):  
Fabrizio De Gregorio
Keyword(s):  
Experimental Investigation ◽  
Flow Control ◽  
Control Systems ◽  
Flow Separation ◽  
Laboratory Tests ◽  
Active Flow Control ◽  
Aerodynamic Characteristics ◽  
Pulsed Jets ◽  
Wide Range ◽  
Active Flow

A comprehensive experimental investigation of a helicopter blunt fuselage model was carried out to evaluate the effectiveness of active flow control (AFC) systems in reducing parasite fuselage drag. The main objective was to demonstrate the capability of different active technologies to decrease fuselage drag by alleviating the flow separation in the loading ramp region of large transport helicopters. The work was performed on a simplified blunt fuselage at model scale. Two different flow control actuators were considered for evaluation: steady blowing and unsteady blowing (i. e., pulsed jets). Laboratory tests of each individual actuator were performed to assess their performance and properties. The fuselage model was investigated with and without the AFC systems located along the loading ramp edges. Significant drag reductions were achieved for a wide range of fuselage angles of attack and sideslip angles without negatively affecting other aerodynamic characteristics.

scholarly journals Effects of leading-edge rod on dynamic stall performance of a wind turbine airfoil

2017 ◽  
Vol 231 (8) ◽  
pp. 753-769 ◽  
Author(s):  
Junwei Zhong ◽  
Jingyin Li ◽  
Penghua Guo
Keyword(s):  
Stokes Equations ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Dynamic Stall ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Moment Coefficient ◽  
Wind Turbine Airfoil ◽  
Stall Angle ◽  
The Moment

A cylindrical rod placed at the leading edge of the S809 airfoil is used as an alternative for the conventional vortex generators. In this paper, extensive numerical investigations have been conducted on the effects of the rod on the static and dynamic stall performance of the S809 airfoil. The flows around the stationary and sinusoidally oscillating S809 airfoils at Re = 106 are simulated by solving the unsteady two-dimensional Reynolds-averaged Navier–Stokes equations with the Shear Stress Transport k–ω model. For the stationary airfoil, the leading edge rod can effectively enhance the aerodynamic characteristics of the airfoil and delay the stall angle, with the maximum lift–drag ratio increased by 30.7%. For the airfoil undergoing deep dynamic stall, the rod shows the capacity of eliminating the dynamic stall vortex at the leading edge and suppressing the flow separation at the tailing edge. It also reduces the peak of the negative pitching moment and the hysteresis effects substantially, and eliminates the negative damping sub-loop of the moment coefficient. Moreover, the distance between the rod and the airfoil has a strong influence on the lift forces but little effect on drag and moment coefficients of airfoil under deep dynamic stall.

Numerical studies of active flow control on wing tip extension

2019 ◽  
Vol 91 (2) ◽  
pp. 346-352
Author(s):  
Petr Vrchota ◽  
Ales Prachar ◽  
Shia-Hui Peng ◽  
Magnus Tormalm ◽  
Peter Eliasson
Keyword(s):  
Flow Control ◽  
Mass Flow ◽  
Flow Separation ◽  
Physical Modeling ◽  
Active Flow Control ◽  
Leading Edge ◽  
Content Type ◽  
Tip Extension ◽  
Active Flow ◽  
Wing Tip

Purpose In the European project AFLoNext, active flow control (AFC) measures were adopted in the wing tip extension leading edge to suppress flow separation. It is expected that the designed wing tip extension may improve aerodynamic efficiency by about 2 per cent in terms of fuel consumption and emissions. As the leading edge of the wing tip is not protected with high-lift device, flow separation occurs earlier than over the inboard wing in the take-off/landing configuration. The aim of this study is the adoption of AFC to delay wing tip stall and to improve lift-to-drag ratio. Design/methodology/approach Several actuator locations and AFC strategies were tested with computational fluid dynamics. The first approach was “standard” one with physical modeling of the actuators, and the second one was focused on the volume forcing method. The actuators location and the forcing plane close to separation line of the reference configuration were chose to enhance the flow with steady and pulsed jet blowing. Dependence of the lift-to-drag benefit with respect to injected mass flow is investigated. Findings The mechanism of flow separation onset is identified as the interaction of slat-end and wing tip vortices. These vortices moving toward each other with increasing angle of attack (AoA) interact and cause the flow separation. AFC is applied to control the slat-end vortex and the inboard movement of the wing tip vortex to suppress their interaction. The separation onset has been postponed by about 2° of AoA; the value of ift-to-drag (L/D) was improved up to 22 per cent for the most beneficial cases. Practical implications The AFC using the steady or pulsed blowing (PB) was proved to be an effective tool for delaying the flow separation. Although better values of L/D have been reached using steady blowing, it is also shown that PB case with a duty cycle of 0.5 needs only one half of the mass flow. Originality/value Two approaches of different levels of complexity are studied and compared. The first is based on physical modeling of actuator cavities, while the second relies on volume forcing method which does not require detailed actuator modeling. Both approaches give consistent results.

Aerodynamic performance investigation of sweptback wings with bio-inspired leading-edge tubercles

2021 ◽  
Author(s):  
V. T. Gopinathan ◽  
J. Bruce Ralphin Rose
Keyword(s):  
Flow Control ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Aerodynamic Performance ◽  
Laminar Separation ◽  
Flow Separation Control ◽  
Pressure Distributions ◽  
Lifting Surfaces ◽  
Naca 0015

The aerodynamic behavior of sweptback wing configurations with bio-inspired humpback whale (HW) leading-edge (LE) tubercles has been investigated through computational and experimental techniques. Specifically, the aerodynamic performance of tubercled wings with symmetric (NACA 0015) and cambered (NACA 4415) airfoils is validated against the baseline model at various angles of attack ([Formula: see text]. The [Formula: see text]/[Formula: see text] ratio of the HW flipper is strategically reduced to 0.15 for ascertaining the flow control potential of the bio-inspired wings with sweptback configuration. It is a novel effort to quantify the effect of the leading-edge protuberances on stall delay, flow separation control and distribution of streamline vortices at unique [Formula: see text]/[Formula: see text] ratio outside the thickness range of HW flipper morphology. Four tapered sweptback wing models (Baseline A, Baseline B, HUMP 0015, HUMP 4415) are used with the amplitude-to-wavelength ([Formula: see text] ratio of 0.24 and Reynolds number about [Formula: see text]. The chordwise pressure distributions are recorded at the peak, mid and trough regions of the tubercled wings through a detailed wind tunnel testing and validated with numerical analysis. Additionally, the flow characteristics over the bio-inspired surfaces have been qualitatively analyzed through the laser flow visualization (LFV) technique to reveal the influence of laminar separation bubbles (LSBs). The essential aerodynamic characteristics such as boundary layer trip delay, vortex mixing, stall delay, and flow control at different AoA are addressed through consistent experimental data. As the sweptback configuration is a primary choice for airplane wings, the improved aerodynamic characteristics of the tubercled wings can be effectively utilized for the design of novel lifting surfaces, hydroplanes and wind turbines in the near future.

Experimental investigation of the effect of active flow control on the wake of a wind turbine blade

2021 ◽  
pp. 095440622110089
Author(s):  
GholamHossein Maleki ◽  
Ali Reza Davari ◽  
Mohammad Reza Soltani
Keyword(s):  
Experimental Investigation ◽  
Wind Turbine ◽  
Turbine Blade ◽  
Active Flow Control ◽  
Leading Edge ◽  
Wind Turbine Blade ◽  
Plasma Actuators ◽  
Plasma Actuation ◽  
Surface Pressure Distribution ◽  
Stall Angle

An extensive experimental investigation was conducted to study the effects of Dielectric Barrier Discharge (DBD), on the flow field of an airfoil at low Reynolds number. The DBD was mounted near the leading edge of a section of a wind turbine blade. It is believed that DBD can postpone the separation point on the airfoil by injecting momentum to the flow. The effects of steady actuations on the velocity profiles in the wake region have been investigated. The tests were performed at α = 4 to 36 degrees i.e. from low to deep stall angles of attack regions. Both surface pressure distribution and wake profile show remarkable improvement at high angles of attack, beyond the static stall angle of the airfoil when the plasma actuation was implemented. The drag calculated from the wake momentum deficit has further shown the favorable role of the plasma actuators to control the flow over the airfoil at incidences beyond the static stall angle of attack of this airfoil. The results demonstrated that DBD has been able to postpone the stall onset significantly. It has been observed that the best performance for the plasma actuation for this airfoil is in the deep stall angles of attack range. However, below and near the static stall angles of attack, plasma augmentation was pointed out to have a negligible improvement in the aerodynamic behavior.

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