Experimental Study of the Influence of Flap Angle Configuration on the Drag Coefficient of Airfoil Model of Homebuilt Aircraft

Homebuilt aircraft are experimental aircraft, of which at least 51% of the aircraft parts are self-supporting and not factory-made. The purpose of this study was to describe the characteristics of the flap angle configuration of a homebuilt aircraft airfoil and determine the drag coefficient (CD) of the airfoil on the flap angle configuration. This research method uses an experimental approach, which is carried out in a wind tunnel at the Fluid Mechanics Laboratory, Faculty of Engineering, Hasanuddin University Gowa. The homebuilt aircraft wing model uses the NACA 23012 airfoil model, by treating the main airfoil flap angle configuration (Fu) with the same control airfoil or aileron (Fk) flap angle at 5 (five) levels of flap angle variations, namely -150, 00, 150, 300, and 450. Furthermore, each configuration of the airfoil flap is flowed with air at the same speed of 22 m/s or at Reynolds number Re = 270,886 at 6 (six) levels of angle of attack (α), namely at angles of -150, -100, 00, 50, 100 and 150. The results of this study indicate that the maximum value of CD always occurs at 150 and angle Fk = 450 for all variations of the Fu angle. The maximum value of CD = 2.8656 occurs at the same angle Fu and angle Fk, namely at angles of 450 and α = 150, while the minimum value of CD = 0.3144 occurs at the same angle Fu and Fk, namely at angles 00 and α = -100.


Introduction
Airplanes flying at low-speed performance, such as landing or taking off, really need a high lift force to balance the weight of the aircraft.Conventional low-speed aircraft have a maximum lift coefficient (CL max) of about 1.4 or 1.5 [1].If a low stall speed is needed, a higher CL max must be obtained.One way to increase is to increase the surface of the wing, but this method will increase the value of drag.To overcome this problem is to use a tool called high lift devices.These high lift devices can change the characteristics of the airfoil, by increasing the CL max when needed, especially when operating at low speeds.
One type of high lift device is a flap.Mounting the flap on the wing is mounted on an airfoil with a hinged surface on the rear edge of the wing [2].If the flap is lowered, the airplane's stall speed will decrease.The flap reduces stall speed by increasing the wing surface area and thereby increasing the maximum lift coefficient [3].The flap can be seen when the plane is about to take off or land, because, in these two conditions, the plane is at low speed.So, increasing the lift required additional lift by expanding the surface of the wing.Self-propelled aircraft are aircraft with low speed, so the use of flaps is the main thing.
A lot of research on airfoils has been done to increase their performance.One of them is by adding across-section behind the airfoil which is usually called a flap.The flap itself is divided into several types including plain, split, fowler, and slotted.Airfoil research that was previously carried out experimentally has now begun to shift toward computing as technology develops.This is because experimental research costs money in making the prototype, so to overcome this, the research is carried out with dynamic fluid computational simulations.
Singh, 2017 [4] investigated the effect of plain flaps on the aerodynamic characteristics of the NACA 66-01 airfoil.The results of this study indicate that the stall angle increases and the wing performance is further improved when the angle of attack increases.The greater the deflection in the flap, the more flow is trapped under the airfoil, leading to a decrease in flow velocity and an increase in pressure under the airfoil.Simultaneously, an increase in the adverse pressure gradient with flap deflection results in greater flow separation behind the flap.FLUENT ANSYS analysis was used to validate the experimental results.
Katz & Largman, 2016 [5] investigated the aerodynamic performance of a two-element airfoil, with 900 trailing edge flaps experimentally.The flap is 5% longer than the chord, significantly increasing the lift of the baseline airfoil, over a wide range of angles of attack.The maximum lift coefficient on the wing flaps is also increased, while the lift-to-drag ratio decreases.
Mahmood et al. 1995 [6] conducted a comparative experimental and numerical investigation of airflow over a high-lift NACA 4412 airfoil with and without a flap section.The experimental studies focused on the flow over the top of the airfoil with particular interest in the separation region near the back edge.
Abdelrahman et al. 2020 [7] conducted a study aimed at reducing the negative effect, by adding more than one flap at different distances in the circulation area.Attaching the flap to the trailing edge of the airfoil increases the lift coefficient, with a negative effect represented by adverse flow circulation and pressure gradient around the flap.Saha et.al. 2017 [8] investigated the aerodynamic effect of the Gurney flap on the NACA 2412 airfoil for subsonic conditions.The flap height ranges from 2% to 5% of the airfoil chord length.The study was evaluated at an angle of attack of 00 to 320 below the subsonic Mach number.Numerical analysis was performed using the computational fluid dynamics program to predict the flow field.The analysis shows that the optimal size of the device is always below the boundary layer thickness at the trailing edge and tends to increase the llift-to-dragratio significantly.Gurney flaps with 2% chord height provide the best performance than 3%, 4% , and 5% chord Gurney flaps.This study concludes with the suggestion that the Gurney flap can cause a reduction in drag in the high lift area, thereby increasing the lift-to-drag ratio before stalling.Salam, et al 2021 [9] conducted a study using a Computational Fluid Dynamics (CFD) approach and an experimental program.The experimental approach was carried out in a wind tunnel at the Fluid Mechanics Laboratory, Faculty of Engineering, Hasanuddin University, Gowa.The homebuilt aircraft wing model is the NACA 23012 airfoil model, by modifying the thickness to chord ratio (t/c) at t/c = 9%, t/c = 12%, and t/c = 15%.Furthermore, each model was treated with a freestream velocity (U) of 40 m/s, with variations in the angle of attack (α) -200, -150, -100, -50, 00, 50, 100, 150, and 200.The results showed that the addition of the t/c ratio increased the maximum CL value.For the maximum value of CL obtained at = 150, namely at t/c = 9%, CL = 1.4299, at t/c = 12%, CL = 1.4466, and at t/c = 15 %, CL = 1, 4979.The maximum CL/CD of 1.4999 was obtained at t/c = 15 % and = 5˚, thus the most suitable homebuilt aircraft wing model is the NACA 23012 airfoil model with t/c = 15 %.

Research Method
This research was conducted using an experimental approach.The model of the flap configuration of the flap airfoil of the homebuilt aircraft wing or the test object was investigated at the sub-sonic wind tunnel facility available at the Fluid Mechanics Laboratory, Department of Mechanical Engineering, Faculty of Engineering Unhas in Gowa Regency. 1 (one) specimen model was made with a configuration of 5 (five) variations in the angle of the airfoil flap, namely at -150, 00, 150, 300, and 450 angles for the main airfoil flap and the control airfoil flap (aileron).The treatment of the airfoil flap angle configuration in question is that at each level of the main airfoil flap angle, 5 (five) variations of the aileron flap angle are given, namely at -150, 00, 150, 300, and 450 angles, then each airfoil flap configuration is assigned an airflow velocity, the same as 22 m/s, at 6 (six) levels of angle of attack, namely at -150, -100, 00, 50, 100 and 150 angles.
The test object is adjusted to the available wind tunnel test section because to get good test results data, the ratio of the cross-sectional area of the test object to the cross-sectional area of the channel or wind tunnel test section should not be greater than 1:3.For this reason, the airfoil size for the wing model of a homebuilt aircraft is on average 150 mm in length (chord), 200 mm in width (span) and 20 mm in thickness.The material used in the manufacture of the test object is wood.The equipment used is international standard testing equipment at the Fluid Mechanics Laboratory, Department of Mechanical Engineering, Faculty of Engineering, Hasanuddin University.This low-speed wind tunnel is made by Plint & Partners LTD Engineers, where the velocity of the air flowing through the test section (300 mm x 300 mm) is a maximum of 25 m/s.Fig. 2 below, shows the wind tunnel installation that will be used, while Fig. 3 shows the position of the test object in the wind tunnel and load balancing equipment to measure the resulting drag force.According to Cengel & Cimbala [10] the value of the drag coefficient (CD), can be determined based on equation (1) and equation ( 2) below.
Where CD = drag coefficient,  = air density (kg/m 3 ), U = freestream velocity (m/s), A = wing planform area (m 2 ), and FD = drag (N).To describe the characteristics of the drag coefficient, the Reynolds number (Re) must be determined from the following equation.
Table 1 below, shows the results of the drag coefficient (CD) on variations in the angle of attack (α) and control flap angle (Fk) at the main flap angle (Fu) -15 0 and Re = 270,886.From Table 1 it is found that at Fu = -15 0 , the value of the drag coefficient decreases from the angle of attack -15 0 to the angle of attack 0 0 , then increases in size from the angle of attack 0 0 to the angle of attack 15 0 , while the control flap angle also decreases from the angle of -15 0 to the angle of 0 0 , then The CD increases from the control flap angle 0 0 to the control flap angle 45 0 .The maximum drag coefficient of 2.6141 occurs at the angle of attack 15 0 and the angle of the control flap 45 0 , while the minimum drag coefficient of 0.8354 occurs at the angle of attack 0 0 and the angle of the control flap 0 0 .Based on these results, changes in the angle of attack of the airfoil and the angle of the control flap from negative to positive, the drag coefficient decreases, while changes in the angle of attack of the airfoil and the angle of the flap are positive, resulting in a decrease in the drag coefficient, produce a larger drag coefficient when the angle is enlarged.From the graph, it can be seen that the CD value increases with the increase in the control flap angle from 00 to 450 angle.In Fig. 4, information is also obtained that, at the same control flap angle, different CD values can be obtained when the angle of attack changes.The greater the value of the angle of attack, the greater the CD value.

CD
For control flap angles of -150 to 150, the smallest CD value is obtained at an angle of attack of 00, while for control flap angles of 300 and 450, the minimum CD value is obtained at an angle of attack of -100.The characteristics of CD show a pattern that tends to be the same for all levels of control flap angle.Table 2 below, shows the results of the drag coefficient (CD) on variations in the angle of attack (α) and control flap angle (Fk) at the main flap angle (Fu) 00 and Re = 270,886.From Table 2 it is obtained that at Fu = 00, the value of the drag coefficient decreases from the angle of attack -150 to the angle of attack -100, then increases from the angle of attack -100 to the angle of attack 150, while the control flap angle also decreases from the angle of -150 to an angle of 00, then increases from the control flap angle of 00 to the control flap angle of 450.The maximum drag coefficient of 2.5782 occurs at the angle of attack of 150 and the angle of the control flap 450, while the minimum drag coefficient of 0.3144 occurs at the angle of attack -100 and the angle of the control flap 00.Based on these results, changes in the angle of attack of the airfoil and the angle of the control flap from negative to positive, then the drag coefficient decreases, while the change in the angle of attack of the airfoil and the angle of the control flap is positive, a will produce a larger drag coefficient when the angle is increased.

CD
From the graph, it can be seen that the CD value increases with the increase in the angle of attack of the airfoil from -10 0 to the angle of attack of 15 0 for all levels of the control flap angle.In Figure 5, information is also obtained that, at the same control flap angle, different CD values can be obtained, when the angle of attack of the airfoil changes.The greater the value of the angle of attack of the airfoil, the greater the CD value.For control flap angle 0 0 , the smallest CD value is obtained at -15 0 airfoil angle of attack to 15 0 airfoil attack angle.Characteristics of CD show the same pattern for all control flap angle levels.The minimum CD value for all control flap angle levels was obtained at -10 0 airfoil angle of attack.
Table 3 shows the results of the drag coefficient (CD) on variations in the angle of attack of the airfoil (α) and the control flap angle (Fk) at the main flap angle (Fu) 150 and Re = 270,886.From table 3 it is found that at Fu = 150, the value of the drag coefficient decreases from the angle of attack of the airfoil -150 to the angle of attack of the airfoil -100, then increases from the angle of attack of the airfoil -100 to the angle of attack of the airfoil of 150, while the control flap angle also decreases from the angle of -150 to with an angle of 00, then it increases from the control flap angle 00 to the control flap angle 450.
The maximum drag coefficient of 2.6860 occurs at the angle of attack 150 and the control flap angle 450, while the minimum drag coefficient of 0.4042 occurs at the angle of attack -100 and the control flap angle 00.Based on these results, the change in the angle of attack of the airfoil and the angle of the control flap from negative to positive, the drag coefficient decreases, while the change in angle s Airfoil strike and control flap angle are positive, will produce a larger drag coefficient when the airfoil angle of attack and control flap angle is increased.Fig. 6 shows a graph of the experimental results of the drag coefficient on the airfoil angle of attack (α) for each level of the control flap angle (Fk), at freestream velocity (U) 22 m/s or Re = 270,886 and main flap angle (Fu) = 150.From the graph, it can be seen that the CD value increases with the increase in the angle of attack of the airfoil from -100 to the angle of attack of 150 for all levels of the control flap angle.In Figure 6, information is also obtained that, at the same angle of the control flap, it can produce different CD values, when the angle of attack of the airfoil changes.The greater the value of the angle of attack of the airfoil, the greater the CD value.For control flap angle 00, the smallest CD value is obtained at -150 airfoil angle of attack to 150 airfoil attack angle.Characteristics of CD show the same pattern for all control flap angle levels.The minimum CD value for all control flap angle levels was obtained at -100 airfoil angle of attack.Table 4 shows the results of the drag coefficient (CD) on variations in the angle of attack of the airfoil (α) and the angle of the control flap (Fk) at the main flap angle (Fu) 300 and Re = 270,886.From Table 4, it is found that at Fu = 300, the value of the drag coefficient decreases from the angle of attack of the airfoil -150 to the angle of attack of the airfoil -100, then increases from the angle of attack of the airfoil -100 to the angle of attack of the airfoil of 150, while the control flap angle also decreases from the angle of -150 to with an angle of 00, then it increases from the control flap angle 00 to the control flap angle 450.The maximum drag coefficient of 2.7488 occurs at the angle of attack of 150 and the control flap angle of 450, while the minimum drag coefficient of 0.4761 occurs at the angle of attack of -100 and the angle of the control flap of 00.Based on these results, the change in the angle of attack of the airfoil and the angle of the control flap from negative to positive, the drag coefficient decreases, while the change in the angle of Airfoil strike and control flap angle positive, will produce a larger drag coefficient when the airfoil angle of attack and control flap angle is Fig. 7 shows a graph of the experimental results of the drag coefficient on the airfoil angle of attack (α) for each level of the control flap angle (Fk), at freestream velocity (U) 22 m/s or Re = 270,886 and main flap angle (Fu) = 300.From the graph, it can be seen that the CD value increases with the increase in the angle of attack of the airfoil from -100 to the angle of attack of 150 for all levels of the control flap angle.In Figure 7, information is also obtained that, at the same angle of the control flap, it can produce different CD values, when the angle of attack of the airfoil changes.The greater the value of the angle of attack of the airfoil, the greater the CD value.For control flap angle 00, the minimum CD value is obtained at -150 airfoil angle of attack to 150 airfoil angle of attack.Characteristics of CD show the same pattern for all control flap angle levels.The minimum CD value for all control flap angle levels was obtained at -100 airfoil angle of attack.Table 5 shows the results of the drag coefficient (CD) on variations in the angle of attack of the airfoil (α) and the angle of the control flap (Fk) at the main flap angle (Fu) 450 and Re = 270,886.From table 5 it is found that at Fu = 450, the value of the drag coefficient decreases from the angle of attack of the airfoil -150 to the angle of attack of the airfoil -100, then increases from the angle of attack of the airfoil -100 to the angle of attack of the airfoil of 150, while the control flap angle also decreases from the angle of -150 to with an angle of 00, then it increases from the control flap angle 00 to the control flap angle 450.The maximum drag coefficient of 2.8656 occurs at the angle of attack of 150 and the angle of the control flap 450, while the minimum drag coefficient of 0.6019 occurs at the angle of attack -100 and the angle of the control flap 00.Based on these results, changes in the angle of attack of the airfoil and the angle of the control flap from negative to positive, the drag coefficient decreases, while changes in the angle of attack of the airfoil and the angle of the control flap are positive, which results in a larger drag coefficient when the angle of attack of the airfoil and the angle of the control flap is increased.Fig. 8 shows a graph of the experimental results of the drag coefficient on the airfoil angle of attack (α) for each level of the control flap angle (Fk), at freestream velocity (U) 22 m/s or Re = 270,886 and main flap angle (Fu) = 450.From the graph, it can be seen that the CD value increases with the increase in the angle of attack of the airfoil from -100 to the angle of attack of 150 for all levels of the control flap angle.In Figure 8, information is also obtained that, at the same control flap angle, different CD values can be obtained, when the angle of attack of the airfoil changes.The greater the value of the angle of attack of the airfoil, the greater the CD value.For control flap angle 0 0 , the minimum CD value is obtained at -15 0 airfoil angle of attack to 15 0 airfoil angle of attack.Characteristics of CD to show the same pattern for all control flap angle levels.The minimum CD value for all control flap angle levels was obtained.

CD
Based on Figs. 4 to 8 above, it shows the same characteristic pattern, and the minimum drag coefficient value is obtained at the angle of attack -10 0 and the control flap angle 0 0 for all variations of the main flap angle, except for the main flap -15 0 the minimum drag coefficient value is obtained, the angle of attack is 0 0 .For the maximum drag coefficient value is obtained at the angle of attack of 15 0 and the control flap angle is 0 0 for all main flaps.The value of the minimum drag coefficient for each level of the main flap angle is shown in Table 6 below, while the value of the maximum drag coefficient for each level of the main flap angle is shown in Table 7.

Fig. 1
below shows the model of the airfoil flap.

FkFuFigure . 2 3 .
Figure .2 Research Installation Figure 3. Position of the Test Object on the Research Installation

Figure 4 .
Figure 4. Graph of Drag Coefficient (CD) with Angle of Attack (α) to 5 (five) Levels of Control Flap Angle (Fk) at Re = 270,886 and Main Flap Angle (Fu) = -150 Fig. 4 shows a graph of the experimental results of the drag coefficient on the ang le of attack (α) for each level of the control flap angle (Fk), at freestream velocity (U) 22 m/s or Re = 270,886 and main flap angle (Fu) = -150.From the graph, it can be seen that the CD value increases with the increase in the control flap angle from 00 to 450 angle.In Fig.4, information is also obtained that, at the same control flap angle, different CD values can be obtained when the angle of attack changes.The greater the value of the angle of attack, the greater the CD value.

Figure 5 Figure 5 .
Figure5shows a graph of the experimental results of the drag coefficient on the airfoil angle of attack (α) for each level of the control flap angle (Fk), at freestream velocity (U) 22 m/s or Re = 270,886 and main flap angle (Fu) = 0 0 .

Table 1 .
Coefficient of Drag (CD) on Variations in Angle of Attack (α) and Control Flap Angle (Fk) at Re = 270,886 and Main Flap Angle (Fu) = -15 0

Table 2 .
Coefficient of Drag (CD) on Variations in Angle of Attack (α) and Control Flap Angle (Fk) at Re = 270,886 and Main Flap Angle (Fu) = 0 0

Table 3 .
Coefficient of Drag (CD) on Variations in Angle of Attack (α) and Control Flap Angle (Fk) at Re = 270,886 and Main Flap Angle (Fu) = 15 0

Table 4 .
Coefficient of Drag (C D ) on Variations in Angle of Attack (α) and Control Flap Angle (Fk) at Re = 270,886 and Main Flap Angle (Fu) = 30 0

Table 5 .
Coefficient of Drag (CD) on Variations in Angle of Attack (α) and Control Flap Angle (Fk) at Re = 270,886 and Main Flap Angle (Fu) = 45 0