Exploring the Influence of NACA0018 Airfoil Attack Angle on the

Airflow Characteristics Based on CFD

Hechong Chen

Hao Mei

and Guyu Yi

School of Mechanical Science and Engineering, Huazhong University of Science and Technology, Wuhan, 430074, China

International Business School, Henan University, Kaifeng, 475004, China

Air Traffic Management, Civil Aviation University of China, Tianjin, 300300, China

Keywords: NACA0018, CFD, Turbulence Model, Attack Angle, Aerodynamic Characteristics.

Abstract: In order to explore the influence of aerodynamic characteristics on aircraft flight performance and analyze the

effect of attack angle on the aerodynamic performance of NACA0018 airfoil, numerical simulation of

NACA0018 airfoil has been established by using FLUENT software. Firstly, by comparing the results of

computational models and experimental values, the numerical boundary conditions and turbulence models

applicable to airfoil calculations were determined. Secondly, combined with the principle of lift improvement,

the pressure and velocity field of NACA0018 at 0 ° and 4 ° two attack angles were analyzed. The

experimental results show that increasing the attack angle can significantly improve the lift, that is, the lift of

the airfoil at a 4 ° attack angle is greater than that 0 °. This study provides a reference for wing design and

aerodynamic performance analysis, which can optimize wing shape and structure based on changes in attack

angles under different flying conditions. In the future, further research will be conducted on the matching

between wing attack angles and wing types, flap lengths, flap deflection angles during different flight stages,

in order to optimize the aircraft performance comprehensively.

1 INTRODUCTION

In recent years, with the advancement of high-

performance computers, computational fluid

dynamics (CFD) has become a core technology in the

field of aerospace research. Researchers have studied

the impact of aircraft lift on the aerodynamic

performance of wing profiles, as well as the design

and failure issues of aircraft lift devices, through

numerical simulation methods. The NACA0018

airfoil, as a classic airfoil, has become a standard

model for wing aerodynamics research due to its

simple geometric shape and wide application. Attack

angle(AoA) can improve wing lift within a certain

range to meet various flight requirements by

increasing lift to varying degrees through attack

angle, especially during take-off and climb phases

(Fan, Pang, and Liu, 2004). Therefore, it is very

significant to understand the NACA0018 airfoil

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aerodynamic characteristics on aircraft controling

and wing designing.

As the attack angle increases, the air velocity

flowing through the wing will increase, the air

pressure will decrease, and the lift coefficient (CL)

will increase. When flying at a constant altitude, the

increase on CL indicates the decrease in required

ground speed. During take-off, the aircraft must

reach sufficient speed and attack angle conditions to

balance its lift and gravity. At the end of the ground

acceleration phase, the aircraft begins to lift its front

wheels. During this phase, it is necessary to maintain

acceleration and increase the attack angle to obtain

greater lift. The ground effect gradually decreases

until it leaves the ground (Airbus, 2002).

During level flight, lift and drag are balanced, and

the lift limit is reached when CL equals CLmax. At

this point, if the attack angle increases, stall will

occur. At high attack angles, the airflow separates

from the upper surface of the wing. If the attack angle

Chen, H., Mei, H. and Yi, G.

Exploring the Inﬂuence of NACA0018 Airfoil Attack Angle on the Airﬂow Characteristics Based on CFD.

DOI: 10.5220/0013444600004558

Paper published under CC license (CC BY-NC-ND 4.0)

In Proceedings of the 1st International Conference on Modern Logistics and Supply Chain Management (MLSCM 2024), pages 497-504

ISBN: 978-989-758-738-2

497

continues to increase, the separation point of the

airflow becomes unstable and rapidly fluctuates back

and forth. As a result, the pressure distribution

constantly changes and alters the position and

magnitude of lift. This effect is called chattering,

which manifests as intense vibrations. When AoA

reaches its maximum value, the separation point

moves further forward, and the airflow on the upper

surface achieves total separation. This phenomenon

leads to a significant loss of lift, known as stall

(Airbus, 2002).

Although attack angle technology has been

widely applied in aviation, simulation research on the

influence of different geometric parameters,

especially attack angle, on the aerodynamic

performance of airfoils is still relatively scarce. The

daily operation of large aircraft is related to the safety

of personnel and property, as well as the efficiency of

airlines. The refined design of aircraft components is

of great significance in ensuring that aircraft

manufacturing meets design and airworthiness

requirements. With the continuous development of

aviation technology, higher requirements have been

put forward for wing design (Wang Chunyan, 2023).

This study used the NACA0018 airfoil as the

research object. In the simulation environment, visual

flow fields were obtained under different

configurations by changing the attack angle. The

changes in lift, drag, and speed coefficients were

analyzed based on the exported data. This method

allows for the evaluation of the overall aerodynamic

performance of the wing under dynamic conditions,

providing a reference for aircraft design.

2 METHOD

2.1 Establishment of Geometric

Structure of Flow Field

The basic airfoil studied in this paper is NACA 0018,

its chord length is set to 1000mm, and it is modeled

in ug, as shown in Figure 1.

To create the flow field model, use a benchmark

equal to 15 times the airfoil's chord length, that is, the

diameter of the C-type area and the side length of the

square area are 15000mm. From the front center of

the airfoil section, the center line extends 100mm

inward as the circle center of the C-type flow field,

and the line extends 15000mm towards the rear edge

as the axis to draw the square flow field area,as shown

in the Figure 2.

According to Figure 3, the hollow airfoil section

Figure 1: NACA 0018 3D model.

Figure 2: Flow field model.

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Figure 3: Flow field model details.

Figure 4: The first set of edge size adjustment.

Figure 5: Second set of edge sizing adjustments.

area serves as the boundary between the airfoil and

the flow field area. The whole flow field is divided

into six regions with the center of the C-shaped flow

field and the apex of the tail edge of the airfoil as the

boundary (Apostol E I, Țîțu A M, 2023).

2.2 Grid Division

The boundary of each region of the flow field is

dimensioned in order to achieve a finer meshing of the

area near the airfoil surface. The boundary of the

region is named after the overall direction of the

diverting field.

The six vertical lines and the horizontal lines that

meet the airfoil's leading edge are taken as the first

group for edge size adjustment,as shown in Figure 4.

As seen in Figure 5, there are 200 partitions, a 40000

bias factor, and bias in the vicinity of the airfoil's

Exploring the Inﬂuence of NACA0018 Airfoil Attack Angle on the Airﬂow Characteristics Based on CFD

499

surface. The second set of lines for adjusting edge

sizing is the three long horizontal lines that follow the

airfoil's rear edge, with the number of partitions of

200 and the bias factor of 40000, also biased towards

the area close to the airfoil surface,as shown in the

Figure 6. According to the Figure 7, the two trailing

edge curves of the airfoil along with the short

horizontal lines on either side are used as the third set

for edge sizing, with a number of partitions of 300 and

no bias. The fourth group for edge dimensional

adjustment consists of the two airfoil leading edge

curves and the two quarter arcs, with the number of

partitions being 300 and without bias. Finally, six

areas are selected for face grid division, and

quadrilateral grid structure is adopted (Kaya M N,

Kok A R, Kurt H, 2021).

Generate the final mesh division and specfic

details shown in the Figure 8 and Figure 9.

Figure 6: Third set of edge size adjustments.

Figure 7: Fourth set of edge size adjustments.

Figure 8: Grid division results.

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Figure 9: Grid division details.

Figure 10: The relationship between the Angle of attack and the lift and drag coefficients.

2.3 Experimental Setup

The solver used in the experiment adopts the pressure

base type, the velocity format is absolute velocity, and

the plane transient measurement does not consider the

influence of gravity. The viscous SST k-omega model

is used for the flow field model (Molaa A A,

Abdulwahid M A, 2024). Second-order upwind

format was used for the discrete equation format, and

the COUPLED method was chosen for the pressure-

velocity coupling. The fluid used to fill the flow field

is air with a density of 1.225kg/m

and a viscosity of

1.7894*10-5kg/(m · s). Working conditions

Atmospheric pressure is 101325Pa. Set the residual to

10-6 and the number of iterations to 1000.

Experiment with 0° Angle of attack first. In the

velocity inlet, choose the velocity definition method

of the Components, because the Angle of attack is 0,

so do not set the speed in the Y direction, set the speed

in the X direction to 33m/s, the concentration of

turbulence is 5% and the ratio of turbulence viscosity

is 10. The gauge pressure of the pressure outlet is 0.

Set the force vector directions for drag to 1 and 0

(corresponding vectors for X and Y, same below), and

the force vector directions for lift to 0 and 1,

respectively. Initialize and iterate the calculation to

get the result of the response.

Then experiment with 4ºangle of attack. Similar

setup, but due to the change of the Angle of attack,

the direction of the velocity and the direction of the

force vector of the lift resistance must be changed in

response. According to the relevant theoretical

knowledge, the relationship between the Angle of

attack and the lift force and the drag coefficient can

be obtained as seen in the Figure 10.

First of all, for the inlet velocity is changed.,as

shown in the figure 10, the velocity in X direction is

and the velocity in Y direction is (Li S, Li Y, Yang C,

et al, 2018).Therefore, the inlet velocity is changed to

32.93m/s in X direction and 2.31m/s in Y direction in

the setting of inlet velocity. Accordingly, according to

the results in the Figure 11, the direction of the force

vector of the resistance is set to 0.998 and 0.07, and

the direction of the force vector of the lift is set to -

0.07 and 0.998. The final result is obtained by

iterative calculation under such conditions.

3 EXPERIMENTAL RESULTS

The lift force, drag coefficient and corresponding

pressure and velocity cloud maps at 0° and 4°

attack angles were obtained through the experiment.

The first is the result of 0° angle of attack. As

shown in Figure 11, in the static pressure diagram, the

pressure distribution around the airfoil is basically

gentle and uniform. Two blue areas on the upper and

lower parts of the airfoil represent two low pressure

areas. Since they are symmetrical airfoil, the pressure

in the both parts of the airfoil is basically the same.

Exploring the Inﬂuence of NACA0018 Airfoil Attack Angle on the Airﬂow Characteristics Based on CFD

501

The leading edge's stagnation point has a high-

pressure area, and a slightly pressurized area is

formed in the back edge of the wing. In the velocity

distribution cloud diagram, as shown in Figure 12, the

velocity at the stagnation point of the leading edge is

very low, and a long wake is generated at the trailing

edge, while the velocity above and below the airfoil is

high (Sun X, Zhou D, 2022). At the position close to

the wall, the velocity is basically zero due to the

absence of external environmental conditions. The

distribution of the two images at 0ºangle of attack

corresponds closely to the theoretical distribution, and

the results of simulation are satisfactory.

The next is the simulation results at 4ºangle of

attack. According to Figure 13, the static pressure

diagram shows that the high pressure stagnation point

area moves down due to the presence of the attack

Angle.what’s more, an obvious low pressure area is

generated above the airfoil. Such results meet our

experimental expectations, that is to say, the lift force

of the airfoil is significantly improved under this

condition (compared with the condition of 0° attack

angle). Similarly, in the velocity distribution diagram,

as shown in Figure 14, it can be seen that the velocity

of the area above the wing is larger, while the

stagnation point area moves down, and the wake rises

slightly. The simulation results at an Angle of attack

of four degree show that the airfoil's lift increases,

which is in line with the experimental expectation,

and verifies the important influence of the attack

angle on the aerodynamic characteristics of the

airfoil. The experimental results show that different

angles of attack have significant effects on the

aerodynamic performance of airfoil, especially the

generation and distribution of lift. Designers can

optimize the wing shape and structure according to

the change of the angle of attack under different flight

conditions to improve aircraft performance in

different flight stages (Liu et al, 2022). For example,

a wing with a larger Angle of attack can be designed

during takeoff and landing phases to enhance lift,

while it can be optimized for a smaller angle of attack

to reduce drag during cruise phases. Studying the

pressure distribution and velocity field at different

attack angles can help design and optimize the shape

of the wing and improve the lift and aerodynamic

efficiency of the aircraft (Huang S, Hu Y, Wang Y,

2021). Through simulation, the performance of the

wing in different flight states can be predicted, so as

to optimize its shape and structural design.

Figure 11: 0ºpressure cloud image.

Figure 12: 0ºvelocity cloud image.

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Figure 13: 0ºpressure cloud image.

Figure 14: 0ºvelocity cloud image.

4 CONCLUSION

In this paper,The NACA0018 airfoil's aerodynamic

performance at different angles of attack is analyzed

by means of computational fluid dynamics (CFD).

The experimental results show that, compared with 0

° attack angle, the lift of the wing at 4° attack angle

is improved, the velocity in the lower part of the wing

is reduced, and the pressure is increased. Such

analysis is not only a key theoretical understanding of

airfoil behavior, but also a practical design that helps

to optimize the performance of the wing in different

stages of flight, such as take-off, cruise and landing.

The results of this study will contribute to the wider

field of fluid dynamics and aerodynamics by

validating theoretical predictions and providing data

that can be used to improve computational models. In

addition, the findings will be directly applied to

improving aircraft performance through better wing

design, ultimately enabling more efficient and safer

flight operations.It is hoped that in the future, more

accurate CFD simulations will help designers better

understand the attack angle ’ s effect on lift and

optimize the design of the wing. Furthermore, It is

considered that Lift force and attack angle are non -

linear. In the future, the research on nonlinear

dynamics will be strengthened to improve wing

performance and achieve safer and more efficient

flight operations.

AUTHORS CONTRIBUTION

All the authors contributed equally and their names

were listed in alphabetical order.

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