Numerical Study of Aerodynamic Characteristics of Airflow Around NACA 0012 and NACA 4412 Airfoils at Re = 170000

In this paper, a numerical study of aerodynamic characteristics using CFD method of the two dimensional flow around NACA 0012 and NACA 4412 airfoils at various angles of attack from -10 0 to 25 0 , and operating at Re 170000 is presented. The formation of flow patterns, pressure distribution, pressure coefficient (Cp), lift coefficient (Cl), drag coefficient (Cd), moment coefficient (Cm), Skin friction coefficient distribution along the chord. The Y+ wall value curve as a validation that the mesh formation is in accordance with the requirements, i.e.: Y+ wall value below 10 for viscous dominate flows and for SST turbulent model viscous. The geometry of NACA 0012 and NACA 4412 airfoils are created using SolidWorks, CFD analysis is carried out using Ansys Workbench 2018. The results show that the stall condition at NACA 0012 is at an angle of attack of 15 0 , while at NACA 4412 it is at 16 0 . Under the stall conditions, Cl of NACA 4412 is approximately 13 percent higher than NACA 0012, while the value of Cd shows almost the same value for both NACA 0012 and NACA 4412.


INTRODUCTION
Due to its shape, the airfoil is designed to take advantage of the airflow that passes through it as it moves against the air to generate lift and reduce drag.However, there is no single optimal airfoil design because the design will always vary for each use case.For example, while a fixed-wing streamliner produces very little drag, it may not generate enough lifting power for an aircraft to take off.Going even further, airfoils designed for low-altitude flights will have different characteristics than high-altitude ones [1].
The National Advisory Committee for Aeronautics (NACA) developed and tested a series of airfoils and tested a series of airfoils, denoted as NACA Airfoils.The airfoil geometry equation is generic and can be parameterized with any four integers to visualize any member of the 4-Series NACA family [2], including here NACA 0012 and NACA 4412.
In aviation, NACA 0012 is generally utilized as a rudder, flap in an airplane, or helicopter rotor, and by its design, it can help the plane in stability, speed control, and landing.NACA4412 is used in many aircraft, including single-engine, two-seat light airplanes with high wings.Nowadays, where the world is looking for a substitute for using fossil energy, whose energy reserves are decreasing, both NACA 0012 and NACA 4412 airfoils are widely used along the entire blade span of a small horizontal axis wind turbine and also a vertical axis wind turbine as an alternative to environmentally friendly energy [3,4,5].
Based on the NACA 4-digit series classification, although both NACA 0012 and NACA 4412 have the same maximum airfoil thickness, they will produce different aerodynamic characteristics [6,7,8].This paper discusses the characteristics of airflow passing through the NACA 0012 and NACA 4412 aerofoil profiles to obtain aerodynamic parameters, such as streamlined patterns and pressure distribution that occur on the upper surface and on the lower surface of the airfoil, so as to generate aerodynamic forces.This paper also discusses the lift force coefficient and drag coefficient as functions of variation in angle of attack, the skin friction distribution along the chord length, and the Y plus value as a validation for the formation of a fine meshing.

LITERATURE REVIEW
The following are some reviews of previous studies related to this paper.
Eleni Douvi & Dionosios P. Margaris, 2012, [3]: Predicted the flow field and aerodynamic characteristics of a NACA 0012 airfoil in dry and wet (heavy rain) simulated condition by using a computational fluid dynamics, operating at Reynolds numbers Re=1×10^6 and Re=3×10^6, Lift and Drag coefficients are computed at various angles of attack, The result in dry condition, the lift coefficient curve is about 16 degree of angle of attack before stall, and the drag coefficient curve increase rapidly after stall angle of attack.
S. Y. Win and M. Thianwiboon, 2021, [4]: Investigated the changes of the lift-to-drag ratio of an airfoil NACA 4412 with the variation angles of attack with a fixed speed of 30 m/s by using Computational fluid dynamics (CFD) to calculate the aerodynamic coefficients acting on it.The result shown when the angle of attack decreases is that the lift coefficient increases and the drag coefficient decreases, resulting in a high lift-to-drag ratio.Sadikin, A. et al., 2018 [11]: presented a computational fluid dynamics (CFD) simulation of air flow past a 2D model NACA 0012 airfoil at a high Reynolds number (Re = 3.0 x 106) at various angles of attack (-10° to 15°) using three turbulence models: Spalart-Allmaras, Realizable k-Ɛ, and k-ω shear stress transport (SST).It is observed that the Realizable k-Ɛ eliminates the small separation bubble on the upper surface of the airfoil, delaying separation flow.
A. Mahato et al., 2023[12]: Comparing of performance between symmetrical (NACA 0012) and asymmetrical (NACA 4418) airfoils in the application of Wind Turbines by using CFD simulation using Ansys software and applying the one-equation Spalart Allmaras turbulence model with different angles of attack.This study finds the NACA 4418 asymmetrical airfoil is preferable to the symmetrical NACA0012 airfoil for Horizontal Axis Wind Turbines (HAWT), and the stall condition is up to 16 degrees of angle of attack.
Ibrahim Gov and M. H. Dogru, 2020 [13]: discussed the aerodynamic performance optimization of NACA 0012 studied by using CFD analysis.The results show that the performance achieved an approximately 5% improvement in angle of attack and stall condition at 20 degrees.
E. ÖZKÖKDEMİR, D. F. KURTULUŞ, 2016 [14]: Investigated the aerodynamic heating of the structural integrity of wings fins under thermal loading with the NACA 0012 airfoil profile by twodimensional analysis models.The results obtained from the heat transfer coefficient and the static temperature on the airfoil, in addition to the pressure distribution, result in deformation on the airfoil.
Hoffman J. et al., 2016 [15]: working on a 3D simulation of NACA 0012 at a high Reynolds number studied with various angles of attack, the result of the simulation captured the stall condition reduces drastistically lit per drag ratio.
S. Gowda, 2019 [16]: comparing the aerodynamic parameters for a range of angles of attack and the performance of NACA 4412 and Naca 2412 airfoils and discussing that at the same flow conditions, the stall angle for 4412 is 8 degrees and NACA 2412 is a 12 degrees angle of attack.Michna and K. Rogowski, 2022 [17]: discuss the effect of the Reynolds number and the turbulence intensity on the lift and drag coefficients of the NACA 0018 airfoil under a range of Reynolds numbers from 50,000 to 200,000 for the range of angles of attack from 0 to 10 degrees.The result was a drag increase from an angle of attack of 6 degrees at a Reynolds number of 50,000.For higher Reynolds numbers, the drag coefficient increases smoothly as a function of the angle of attack.

NACA Airfoils
An airfoil is a profile structure designed to manipulate the flow of a fluid to produce a reaction, which in an aircraft's case is an aerodynamic lift.The airfoil shape is a cross section of the wings of fixed wing aircraft and is designed to take advantage of the natural response of air flow when disrupted.When air passes through an airfoil (wing), there is a positive pressure lifting action from the air below the wing (lower surface) and a negative pressure lifting action from lowered pressure above the wing (upper surface).
Drag Coefficient: Where:  = Lift,  = Drag,  = dynamic pressure, and  = surface are.So, the purpose of an airfoil is to generate lift and reduce drag.There is no single optimal design of airfoil; the design characteristics must be carefully considered when constructing an aircraft, as optimal characteristics can vary depending on the aircraft's intended purpose.
NACA airfoils are airfoil shapes for aircraft wings developed by the National Advisory Committee for Aeronautics.The NACA four-digit wing sections, such as NACA 0012 and NACA 4412, define the profile by: one digit describing maximum camber as a percentage of the chord; one digit describing the distance of maximum camber from the airfoil leading edge in tens of percents of the chord; and the last two digits describing the maximum thickness of the airfoil as a percent of the chord.Based on this description, NACA 0012 means an airfoil profile that does not have camber, or is called a symmetrical airfoil, and has a maximum thickness of 12% chord.NACA 4412 means a cambered airfoil has a maximum camber of 4% chord at 0,4. chord and a maximum thickness of 12% chord [18].The nomenclature of an airfoil is shown in Figure 2.

Spalart Allmaras
The Spalart-Allmaras turbulence model is a one-equation alternative that can be solved in less time than the full Navier-Stokes equations, because there is no need to use the local thickness of the shear layer along a turbulent flow to calculate a length scale [19,20].
Formation Equation: With; The constants are:

Type of Research
Type of research conducted using simulation analysis using ANSYS WORKBENCH 2018 software related to aerodynamics, especially fluid flow around an airfoil, with the aim of finding the effect of angle of attack on the lift and drag coefficients and to find the critical angle of attack that will affect the stall phenomenon.

Simulation Flowchart
The data using software simulations are carried out in the following steps:  Meshing functions as a geometry control volume divider; smaller, more refined elements with a purpose get more convergent results in the analysis.Figure 5 shows quadrilateral cells for meshing NACA 0012 and NACA 4412 airfoils.Control volume and computational mesh around the airfoil: NACA 0012 and NACA 4412 can be seen in Figure 6.The mesh in the far field domain is coarsened on the approach part of the boundary but gets closer gradually when close to the surface airfoils.In order for simulation to approximate the actual state, the y+ value, which serves as the first layer parameter input indicating the first thickness of the boundary layer, must be less than 10

Determine the Simulation Parameters
The following are the parameters, viscous models, and boundaries used in this calculation, shown in Table 1 below.4. Check and Visualize the Result This is the last step of simulation, where the calculation of the expected aerodynamic parameters is obtained, like velocity distribution, as shown in Figure 7, and pressure distribution, as shown in Figure 8. Figure 9 shows the distribution of pressure that occurs around the NACA airfoil 0012 and NACA 4412 at angles of attack 10, 14, and 17 degrees expressed by color differences occurring, ranging from blue with a speed close to zero, to red, which is the maximum pressure, of 1,00087 atm or 101413,15 Pa. on both airfoils visible areas with high pressure are located on the lower surface, and areas with low pressure occur on the upper surface.There is also a movement in the area of low pressure (represented by yellow color) on the upper surface that is getting closer to the leading edge with increased attack angle.This phenomenon occurs both in NACA 0012 and NACA 4412.At AOA 17 degrees the movement has already reached the leading edge.The low-pressure area formed on the upper surface NACA 4412 is wider compared to NACA 0012, which means the value of different pressure between upper and lower surface on NACA4412 is greater than NACA 0012.

Velocity Distribution
On the left side Fig. 9 shows the velocity distribution that occurs on the NACA 0012 with an attack angle of 10 degrees, 14 degrees and 17 degrees.The color blue indicates the speed equal zero, then the velocity will increase, represented by the color blue, green, yellow, and finally red, which represents the maximum speed value of 35 m/s.As the angle of attack increases, the area with the blue color increases.It can be analyzed that the presence of a zero speed indicates the vortex that occurs, and with the increasing angle of attack the formation of vortex will be greater.
On the right side of Figure 10 the area with zero speed m/s on NACA 4412 is wider compared to NACA4412 at an attack angle of 14 degrees.This indicates that a faster flow separation occurs on a cambered airfoil than a symmetrical airfoil at the same attack angle.From Figure 12.(a) above it can be seen that the maximum lift coefficient of 1.65 occurs at an angle of attack of 15 degrees, which means it is a critical angle of attack for NACA 4412, while in Figure 12.(b) it can be seen that the maximum lift coefficient occurs at an angle of attack of 16 degrees, with a value of Cl = 1.49.

Lift Coefficient, Cl
From Figure 13 below it can be seen that the Cl max value at NACA 4412 is greater than that at NACA 0012, but the max value is obtained at an angle of attack of 15 degrees, which is lower than the angle of attack of 16 degrees at NACA 0012 airfoil, which means NACA 0012 airfoil can delay stall compared to NACA 4412.From Figure 15 it can be seen that the pressure coefficient obtained by NACA 4412 is greater than NACA 0012, so the lift that occurs at NACA 4412 will also be greater than NACA 0012.There is no change of value in the skin friction coefficient for both NACA 0012 and NACA 4412 with variations angle of attack, as shown in Figure 16.From Figure 17 and 18 it can be seen that both at NACA 0012 and NACA 4412, at each variation of the angle of attack a Y plus value below 3 is obtained, which means that the meshing formation of this paper is appropriate.

Figure 1 .
Figure 1.Open circle gas turbine power plant system

Figure 9
Figure 9 Pressure distribution around NACA 0012 and NACA 4412 at various angle of attack

Figure 10 VelocityFigure 11 .Figure 12 .
Figure 10 Velocity Distribution around NACA 4412 at various angle of attack and comparison between NACA 0012 and NACA 4412 at 14 degrees of AOA

Figure 16 .
Figure 16.Skin friction coefficient on NACA 0012 at 10 degrees and 14 degrees angle of attack

Figure 17 .
Figure 17. the Y+ wall value of NACA 0012 at 17 degrees agle of attack

Table 1 .
The Simulation Parameters