scholarly journals Numerical Simulation of the Anti-Icing Performance of Electric Heaters for Icing on the NACA 0012 Airfoil

Aerospace ◽  
2020 ◽  
Vol 7 (9) ◽  
pp. 123
Author(s):  
Sho Uranai ◽  
Koji Fukudome ◽  
Hiroya Mamori ◽  
Naoya Fukushima ◽  
Makoto Yamamoto
Keyword(s):  
Numerical Simulation ◽  
Lift Coefficient ◽  
Heating Surface ◽  
Leading Edge ◽  
Simulation Method ◽  
Chord Length ◽  
Ice Accretion ◽  
Suction Surface ◽  
Edge Surface ◽  
Naca 0012

Ice accretion is a phenomenon whereby super-cooled water droplets impinge and accrete on wall surfaces. It is well known that the icing may cause severe accidents via the deformation of airfoil shape and the shedding of the growing adhered ice. To prevent ice accretion, electro-thermal heaters have recently been implemented as a de- and anti-icing device for aircraft wings. In this study, an icing simulation method for a two-dimensional airfoil with a heating surface was developed by modifying the extended Messinger model. The main modification is the computation of heat transfer from the airfoil wall and the run-back water temperature achieved by the heater. A numerical simulation is conducted based on an Euler–Lagrange method: a flow field around the airfoil is computed by an Eulerian method and droplet trajectories are computed by a Lagrangian method. The wall temperature distribution was validated by experiment. The results of the numerical and practical experiments were in reasonable agreement. The ice shape and aerodynamic performance of a NACA 0012 airfoil with a heater on the leading-edge surface were computed. The heating area changed from 1% to 10% of the chord length with a four-degree angle of attack. The simulation results reveal that the lift coefficient varies significantly with the heating area: when the heating area was 1.0% of the chord length, the lift coefficient was improved by up to 15%, owing to the flow separation instigated by the ice edge; increasing the heating area, the lift coefficient deteriorated, because the suction peak on the suction surface was attenuated by the ice formed. When the heating area exceeded 4.0% of the chord length, the lift coefficient recovered by up to 4%, because the large ice near the heater vanished. In contrast, the drag coefficient gradually decreased as the heating area increased. The present simulation method using the modified extended Messinger model is more suitable for de-icing simulations of both rime and glaze ice conditions, because it reproduces the thin ice layer formed behind the heater due to the runback phenomenon.

scholarly journals Active Control of a Stalled Airfoil Through Steady or Unsteady Actuation Jets

2015 ◽  
Vol 137 (9) ◽  
Author(s):  
V. G. Chapin ◽  
E. Benard
Keyword(s):  
Active Control ◽  
Computational Study ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Position Angle ◽  
Navier Stokes ◽  
Suction Surface ◽  
Parametric Studies ◽  
Naca 0012

The active control of the leading-edge (LE) separation on the suction surface of a stalled airfoil (NACA 0012) at a Reynolds number of 106 based on the chord length is investigated through a computational study. The actuator is a steady or unsteady jet located on the suction surface of the airfoil. Unsteady Reynolds-Averaged Navier–Stokes (URANS) equations are solved on hybrid meshes with the Spalart–Allmaras turbulence model. Simulations are used to characterize the effects of the steady and unsteady actuation on the separated flows for a large range of angle of attack (0 < α < 28 deg). Parametric studies are carried out in the actuator design-space to investigate the control effectiveness and robustness. An optimal actuator position, angle, and frequency for the stalled angle of attack α = 19 deg are found. A significant increase of the lift coefficient is obtained (+ 84% with respect to the uncontrolled reference flow), and the stall is delayed from angle of attack of 18 deg to more than 25 deg. The physical nonlinear coupling between the actuator position, velocity angle, and frequency is investigated. The critical influence of the actuator location relative to the separation location is emphasized.

Numerical investigation of the unsteady aerodynamics of NACA 0012 with suction surface protrusion

2019 ◽  
Vol 92 (2) ◽  
pp. 186-200
Author(s):  
Aslesha Bodavula ◽  
Rajesh Yadav ◽  
Ugur Guven
Keyword(s):  
Vortex Shedding ◽  
Stokes Equations ◽  
Lift Coefficient ◽  
Unsteady Aerodynamics ◽  
Leading Edge ◽  
Navier Stokes ◽  
Suction Surface ◽  
Content Type ◽  
Energy Harvesters ◽  
Naca 0012

Purpose The purpose of this paper is to investigate the effect of surface protrusions on the flow unsteadiness of NACA 0012 at a Reynolds number of 100,000. Design/methodology/approach Effect of protrusions is investigated through numerical simulation of two-dimensional Navier–Stokes equations using a finite volume solver. Turbulent stresses are resolved through the transition Shear stress transport (four-equation) turbulence model. Findings The small protrusion located at 0.05c and 0.1c significantly improve the lift coefficient by up to 36% in the post-stall regime. It also alleviates the leading edge stall. The larger protrusions increase the drag significantly along with significant degradation of lift characteristics in the pre-stall regime as well. The smaller protrusions also increase the frequency of the vortex shedding. Originality/value The effect of macroscopic protrusions or deposits in rarely investigated. The delay in stall shown by smaller protrusions can be beneficial to micro aerial vehicles. The smaller protrusions increase the frequency of the vortex shedding, and hence, can be used as a tool to enhance energy production for energy harvesters based on vortex-induced vibrations and oscillating wing philosophy.

Unsteady Numerical Simulation of Shock Systems in Vaneless Counter-Rotating Turbine

2005 ◽  
Author(s):  
Huishe Wang ◽  
Qingjun Zhao ◽  
Xiaolu Zhao ◽  
Jianzhong Xu
Keyword(s):  
Numerical Simulation ◽  
Mach Number ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Nozzle Design ◽  
Suction Surface ◽  
Rotor Wake ◽  
Wake Interaction ◽  
Unsteady Numerical Simulation ◽  
Almost All

A detailed unsteady numerical simulation has been carried out to investigate the shock systems in the high pressure (HP) turbine rotor and unsteady shock-wake interaction between coupled blade rows in a 1+1/2 counter-rotating turbine (VCRT). For the VCRT HP rotor, due to the convergent-divergent nozzle design, along almost all the span, fishtail shock systems appear after the trailing edge, where the pitch averaged relative Mach number is exceeding the value of 1.4 and up to 1.5 approximately (except the both endwalls). A group of pressure waves create from the suction surface after about 60% axial chord in the VCRT HP rotor, and those waves interact with the inner-extending shock (IES). IES first impinges on the next HP rotor suction surface and its echo wave is strong enough and cannot be neglected, then the echo wave interacts with the HP rotor wake. Strongly influenced by the HP rotor wake and LP rotor, the HP rotor outer-extending shock (OES) varies periodically when moving from one LP rotor leading edge to the next. In VCRT, the relative Mach numbers in front of IES and OES are not equal, and in front of IES, the maximum relative Mach number is more than 2.0, but in front of OES, the maximum relative Mach number is less than 1.9. Moreover, behind IES and OES, the flow is supersonic. Though the shocks are intensified in VCRT, the loss resulted in by the shocks is acceptable, and the HP rotor using convergent-divergent nozzle design can obtain major benefits.

scholarly journals Numerical Simulation and Hydrodynamic Performance Predicting of 2 Two-Dimensional Hydrofoils in Tandem Configuration

2021 ◽  
Vol 9 (5) ◽  
pp. 462
Author(s):  
Yuchen Shang ◽  
Juan J. Horrillo
Keyword(s):  
Numerical Simulation ◽  
Numerical Model ◽  
Resource Constraints ◽  
Lift Coefficient ◽  
Parametric Method ◽  
Experimental Results ◽  
Hydrodynamic Performance ◽  
Tandem Configuration ◽  
Artificial Neural Network Ann ◽  
Naca 0012

In this study we investigated the performance of NACA 0012 hydrofoils aligned in tandem using parametric method and Neural Networks. We use the 2D viscous numerical model (STAR-CCM+) to simulate the hydrofoil system. To validate the numerical model, we modeled a single NACA 0012 configuration and compared it to experimental results. Results are found in concordance with the published experimental results. Then two NACA 0012 hydrofoils in tandem configuration were studied in relation to 788 combinations of the following parameters: spacing between two hydrofoils, angle of attack (AOA) of upstream hydrofoil and AOA of downstream hydrofoil. The effects exerted by these three parameters on the hydrodynamic coefficients Lift coefficient (CL), Drag Coefficient (CD) and Lift-Drag Ratio (LDR), are consistent with the behavior of the system. To establish a control system for the hydrofoil craft, a timely analysis of the hydrodynamic system is needed due to the computational resource constraints, analysis of a large combination and time consuming of the three parameters established. To provide a broader and faster way to predict the hydrodynamic performance of two hydrofoils in tandem configuration, an optimal artificial neural network (ANN) was trained using the large combination of three parameters generated from the numerical simulations. Regression analysis of the output of ANN was performed, and the results are consistent with numerical simulation with a correlation coefficient greater than 99.99%. The optimized spacing of 6.6c are suggested where the system has the lowest CD while obtaining the highest CL and LDR. The formula of the ANN was then presented, providing a reliable predicting method of hydrofoils in tandem configuration.

An Investigation on Numerical Simulation Method for Aero-Acoustics Based on Acoustics Analogy

2013 ◽  
Vol 444-445 ◽  
pp. 462-467
Author(s):  
Dang Guo Yang ◽  
Yong Hang Wu ◽  
Jin Min Liang ◽  
Jun Liu
Keyword(s):  
Numerical Simulation ◽  
Near Field ◽  
Time Lag ◽  
Sound Field ◽  
Leading Edge ◽  
Simulation Method ◽  
Three Dimension ◽  
Sound Waves ◽  
Acoustic Analogy ◽  
Numerical Simulation Method

A numerical simulation method on noise prediction, which incorporates aerodynamics and sound wave equations based on acoustic analogy, is presented in the paper. Near-field unsteady aerodynamic characteristic can be obtain by large eddy simulation (LES), and far-field propagation of sound waves and spatial sound-field can be obtain by solving the time-domain integral equations of Ffowcs Williams and Hawings (FW-H). Based on the method, a numerical simulation was done on a two-dimension cylinder and a three-dimension flat plate with blunt leading edge. The agreement of numerical results with experiment data validated the Feasibility of the method. The results also indicate that LES can describe vortex generation and shedding in the flow-fields, and FW-H formulation, which has taken time-lag between sound emission and reception times into account, can simulate time-effect of sound propagation toward far-fields.

Ice Accretion Prediction for Single Airfoil and Multi-Element Airfoil

2011 ◽  
Vol 138-139 ◽  
pp. 140-145
Author(s):  
Zhi Guo Sun ◽  
Cheng Xiang Zhu ◽  
Chun Ling Zhu
Keyword(s):  
Leading Edge ◽  
Ice Accretion ◽  
Runge Kutta Method ◽  
Wide Range ◽  
Ice Accretions ◽  
Airflow Field ◽  
Grid Technique ◽  
Droplet Trajectory ◽  
Naca 0012 ◽  
Methods Numerical

Ice accretion on aircraft components is an enormous threat to flight safety. In this paper, ice accretions on the leading edge of the NACA 0012 airfoil and the NLR 7301 multi-element airfoil with flap are predicted using the icing code developed by us. This code mainly contains five modules which are grid module, airflow module, droplet module, heat module, and boundary reconstruction module. The effectiveness and robustness of this code are tested by executing the five modules orderly and repeatedly. The Spalart-Allmaras one-equation turbulence model is adopt to calculate the viscous airflow field and the four-order Runge-Kutta method is used to solve the droplet trajectory equations. In order to enhance the efficiency of the icing calculations, the multi-block grid technique is integrated into the grid module. Based on the above methods, numerical results in both two cases are presented and the necessary comparisons with the experimental data are given in corresponding chapters. The computational results show that performance of the icing code is very good for the wide range of icing conditions.

scholarly journals Effect of Leading-Edge Bulges on Aerodynamic Characteristics of Bionic Wingsail

2021 ◽  
Author(s):  
Chen Li ◽  
Peiting Sun ◽  
Hongming Wang
Keyword(s):  
Stokes Equations ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Navier Stokes ◽  
Suction Surface ◽  
Ansys Fluent ◽  
Navier Stokes Equations ◽  
Extension Direction ◽  
Lift And Drag

The leading-edge bulges along the extension direction are designed on the marine wingsail. The height and the spanwise wavelength of the protuberances are 0.1c and 0.25c, respectively. At Reynolds number Re=5×105, the Reynolds Averaged Navier-Stokes equations are applied to the simulation of the wingsail with the bulges thanks to ANSYS Fluent finite-volume solver based on the SST K-ω models. The grid independence analysis is carried out with the lift and drag coefficients of the wingsail at AOA = 8° and AOA=20°. The results show that while the efficiency of the wingsail is reduced by devising the leading-edge bulges before stall, the bulges help to improve the lift coefficient of the wingsail when stalling. At AOA=22° under the action of the leading-edge tubercles, a convective vortex is formed on the suction surface of the modified wingsail, which reduces the flow loss. So the bulges of the wingsail can delay the stall.

scholarly journals Research on the forward flight performance of rotor based on variable-droop leading edge

2021 ◽  
Vol 39 (3) ◽  
pp. 668-674
Author(s):  
Congcong Li ◽  
Yongjie SHI ◽  
Guohua Xu ◽  
Xingliang Liu
Keyword(s):  
Flow Control ◽  
Control Method ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Integral Form ◽  
Simulation Method ◽  
Aerodynamic Characteristics ◽  
Control Effect ◽  
Forward Flight ◽  
Force Coefficient

Aiming at the dynamic stall phenomenon of the retreating side of the rotor in forward flight, the existing flow control method of dynamic leading edge droop was applied to the flow control of forward-flying rotor at three-dimensional scale. A numerical simulation method based on variable droop leading edge is established in this paper. The seesaw rotor is taken as the research object, the moving overset mesh method and RBF grid deformation technology are used, the integral form of Reynolds average N-S equation is the main control equation. The influence of the dynamic leading edge at r/R=0.75~1 on the aerodynamic characteristics of the rotor when the forward ratio is 0.3 is investigated. It is found that variable droop leading edge on the retreating side can effectively inhibit the generation and development of separation vortices near the trailing edge, and has a significant effect on lifting lift coefficient and section normal force coefficient, reducing torque coefficient, and thus improving the equivalent lift-drag ratio of the rotor. In a certain range, the control effect is better with the increase of the droop amplitude under the leading edge.

Numerical Investigation of Bionic Rudder With Leading-Edge Protuberances

2019 ◽  
Vol 142 (1) ◽  
Author(s):  
Hongtao Gao ◽  
Wencai Zhu
Keyword(s):  
Electron Microscopy ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Hydrodynamic Performance ◽  
Spanwise Direction ◽  
Flow Mechanism ◽  
Suction Surface ◽  
Edge Profile ◽  
Drag Ratio ◽  
Lift To Drag Ratio

The duck's webbed feet are observed by using electron microscopy, and observations indicate that the edges of the webbed feet are the shape of protuberances. Therefore, the rudder with leading-edge protuberances is numerically studied in the present investigation. The rudder has a sinusoidal leading-edge profile along the spanwise direction. The hydrodynamic performance of rudder is analyzed under the influence of leading-edge protuberances. The present investigations are carried out at Re = 3.2 × 105 and 8 × 105. In the case of Re = 3.2 × 105, the curves of lift coefficient illustrate that the protuberant leading-edge scarcely affects the lift coefficient of bionic rudder. However, the drag coefficient of the bionic rudder is markedly lower than that of the unmodified rudder. Therefore, the lift-to-drag ratio of the bionic rudder is obviously higher than the unmodified rudder. In another case of Re = 8 × 105, the advantageous behavior of the bionic rudder with leading-edge protuberances is mainly performed in the post-stall regime. The flow mechanism of the significantly increased efficiency by the protuberant leading-edge is explored. It is obvious that the pairs of counter-rotating vortices are presented over the suction surface of bionic rudder, and therefore, the flow is more likely to adhere to the suction surface of bionic rudder.

Experimental Study of Self- Adjustable Flap for Fixed Wing MAV

2019 ◽  
Author(s):  
Anand Verma ◽  
Vinayak Kulkarni
Keyword(s):  
Reynolds Number ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Open Circuit ◽  
Suction Surface ◽  
External Energy ◽  
Subsonic Wind Tunnel ◽  
Aerodynamic Performances ◽  
Lift Curve ◽  
Lift Enhancement

Abstract Current investigation is focused on experimental analysis for the improvement of aerodynamic performances of low aspect ratio wings (S5010 profile, AR = 1,2) at low Reynolds number. In this study self-adjustable flap is used as a lift enhancement device, is mounted on the suction surface of the wing. Present mounting allows the flap to rotate freely about its leading-edge without supplying any external energy. All the experiment performed in open circuit subsonic wind tunnel for the chord-based Reynolds number is 105. It has been observed, the flap remains attached to the wing surface at low angles of attack. But, as the angle of attack increases, flow separation initiates to occur along the trailing edge which lift ups the flap and adjusts itself in a situation so that the flow reattaches. Hence, the lift curve, for pre-stall angles is almost identical in both clean and flapped configurations. Further, both the flap locations are seen to improve the lift coefficient than the clean airfoil for post-stall angles. The flap is more effective for both the models when it is located at a position of 70% of chord length.

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