Aerodynamics Analysis Comparison between NACA 4412 and NREL
S823 Airfoils
SAYEL M. FAYYAD*, AIMAN AL ALAWIN, SULEIMAN ABU-EIN,
ZAID ABULGHANAM, ABDEL SALAM ALSABAG, MOHANNAD O. RAWASHDEH,
MUNTASER MOMANI, WALEED MOMANI
Department of Mechanical Engineering, Faculty of Engineering Technology,
Al-Balqa Applied University,
Amman,
JORDAN
*Corresponding Author
Abstract: - This paper presents a study of the aerodynamics of a wing or bluff bodies and compares different
wing types' behavior against aerodynamic forces. NACA 4412 and NERL S823 airfoils will be analyzed
numerically using the ANSYS simulation. The methodology used in this paper depends on collecting data from
the last studies, studying the analyzed airfoil models, and constructing an analytical model to show the
aerodynamic effects on NACA 4412 and NERL S823 airfoils, and find the total solution. A comparison
between NACA 4412 airfoil and NREL'S S823 is presented. It was found that the lift coefficient for NACA
4412 values is higher than that of NREL S823 airfoil but for NACA 4412 such values are decreasing as the
angle of attack (AoA) is increasing till 8ᵒ of AoA after that Cl values are increasing slightly. In contrast, for
NREL S823 airfoil the values of lift coefficient (Cl) are increasing with AoA till 8ᵒ after that they become
constant or slightly decreasing, while for drag coefficient, it can be noticed that values of drag coefficient (Cd)
for NACA 4412 are lower than that of NREL S823 airfoils and for all values of angle of attack, also values for
both airfoils are decreasing with AoA till 8° and then slightly increased.
Key-Words: - Aerodynamics, Airfoils, NACA 4412 airfoil, ANSYS, simulation, Drag Coefficient, Lift
Coefficient, NREL S823 airfoil.
Received: February 26, 2023. Revised: December 15, 2023. Accepted: February 19, 2024. Published: April 2, 2024.
1 Introduction
Aerodynamics is the science of how a body travels
through the air. As a result, it is a branch of
dynamics concerned with the motion of air and
other gases, as well as the forces acting on a moving
or stationary object in an air current. As a result,
there are three main components to flight
aerodynamics. Examples of these components are
airplanes, relative winds, and the atmosphere.
An airfoil is a surface that is designed to elicit a
certain reaction from the air it passes through. As a
result, an airfoil is any component of an aircraft that
transforms air resistance into a force useful for
flight. A propeller's blades are so engineered that as
they revolve, their form and location generate a
stronger pressure to build up behind them than in
front of them so that they pull the airplane forward.
The objective of this study is to numerically
evaluate several types of airfoils (wings or bluff
bodies) with varied parameters using ANSYS and
then compare the results to determine the ideal
conditions for airfoil designs, including geometry.
Two types of airfoils are being studied: NACA 4412
and NREL airfoils. The major goal of this work, as
described above, is to examine the NACA 4412 and
NREL's airfoils using the ANSYS simulation and
compare the findings with varied airfoil geometry
and aerodynamic circumstances. Any airfoil
contains top and lower surfaces. The essential point
is the higher density of streamlines above the wing,
even though the top surface of the average wing
profile is curvier than the lower surface. The larger
the density of streamlines, the faster the air flows.
According to Bernoulli's principle, a rise in fluid
speed happens at the same time as a decrease in
pressure or potential energy. This is identical to the
energy conservation principle. The total of all kinds
of mechanical energy in a fluid along a streamline is
the same at all places along that streamline in a
steady flow, [1].
Because of the effect of the wing planform,
airfoil section properties differ from wing or aircraft
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properties. From root to tip, a wing can have varied
airfoil sections with taper, twist, and sweepback.
The action of each part along the span determines
the wing's resulting aerodynamic qualities, [1]. The
lift over drag (L/D) ratio is often used to determine a
wing's efficiency. This ratio changes depending on
the angle of attack, but it always reaches a
maximum value for a specific angle of attack. The
wing has reached its optimum efficiency at this
angle. The shape of the airfoil is the factor that
defines the most efficient angle of attack for the
wing, as well as the degree of efficiency. The
maximum thickness of the most efficient airfoils for
common usage is found roughly one-third of the
way back from the leading edge of the wing,
according to research, [1]. High-lift wings and high-
lift devices for wings have been developed by
shaping the airfoils to produce the desired effect.
The amount of lift produced by an airfoil will
increase with an increase in the wing chamber. An
increase in the wing chamber will enhance the
amount of lift produced by an airfoil. The curvature
of an airfoil above and below the chord line surface
is referred to as a camber. The upper chamber
denotes the upper surface, the lower camber denotes
the lower surface, and the mean camber denotes the
section's mean line. Camber is positive when the
chord line departs inward, and negative when it
departs outward. As a result, the upper surface of
high-lift wings has a considerable positive camber
and the lower surface has a slight negative camber.
By enlarging the upper chamber and producing a
negative lower chamber, wing flaps allow a regular
wing to approximate this state [1].
It's also known that the bigger the wingspan is in
comparison to the chord, the more lift is obtained.
Aspect ratio is the term for this comparison. The
greater the lift, the higher the aspect ratio. Despite
the advantages of increasing the aspect ratio,
structural and drag factors were determined to be
significant limits. The total amount of drag on an
aircraft is made up of many drag forces with three
main: Parasite drag; Profile drag and Induced drag.
Parasite drag is the result of a complex
interaction of many drag forces. Any exposed thing
aboard an aircraft creates air resistance, and the
more objects in the airstream, the parasite drag will
be greater. While parasite drag can be decreased by
decreasing the number of exposed parts to a
minimum and simplifying their design, the sort of
parasite drag that is the most difficult to reduce is
skin friction. There is no such thing as a perfectly
smooth surface. When inspected under
magnification, even machined surfaces have a
ragged, uneven appearance. The air near the surface
is deflected by these jagged surfaces, creating
resistance to smooth circulation. By adopting glossy
flat finishes and removing protruding rivet heads,
roughness, and other abnormalities, skin friction can
be decreased.
NACA 4412 and NREL’s airfoils
The NACA four-digit wing sections define the
profile as follows:
1. One digit describing the maximum camber as
a percentage of the chord
2. One digit describing the distance of maximum
camber from the airfoil leading edge in tens of
percent of the chord
3. Two digits describing the maximum thickness
of the airfoil as a percent of the chord, [2].
From 1984 to 1993, the NREL S823 designed
and developed seven families of airfoils, each with
23 variants suitable for different rotor diameters.
The NREL S823 airfoil (Figure 1) was chosen from
among the 23 airfoil variants based on the
availability of experimental data, [3] and [4]. The
NERL S823 was compared to another designated
airfoil DU 06-W-200 which was considered to be
laminar and unsymmetrical and designed for vertical
axis wind turbine at Delft University of Technology
in the year 2006 (Figure 1), [5].
Fig. 1: NREL's S823- (ONERA OA213 AIRFOIL -
NERA/Aerospatiale OA213 rotorcraft)
Many studies have been conducted on this
critical issue. The application of Computational
Fluid Dynamics (CFD) in the simulation and design
of high subsonic transport aircraft wings.
RAMPANT, an unstructured, multigrid flow solver,
was used to perform the computation. CATIA was
used to create a 2-D and 3-D modeling of the wing.
The grid of the wing was created by using TGrid
and preBFC software. The paper describes the grid
creation technique as well as the application of CFD
to the wing design process. It then goes over the
advantages and disadvantages of using the
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aforementioned tools. The wing is then optimized
using the results of the aerodynamic analysis. It
concludes with a discussion of the findings and
suggestions future for research, [6]. The authors in
[7], addressed the issue of ionic flapping aircraft
gliding performance by numerical simulation
method implementing two-way fluid-structure
interaction (FSI), the investigation included the
angle of attack, a rigid and flexible wing, the elastic
model effects and velocity on the aerodynamic
features of a gliding aircraft, at an angle of attack of
10°, minor effect on the aerodynamic performance
of the aircraft was observed, holding maximum lift
to drag ratio for both the flexible and rigid wings]. It
was also found that with an increase in the gliding
speed, the lift force increased while the lift can’t
support the gliding movement at low speed. To
achieve gliding, the weight of the micro air vehicle
is kept under control at around 3 g with the gliding
speed assured to be more than 6.5 m/s. The findings
of this study have significant implications for the
design of bionic flapping aircraft. The authors in [8],
discussed the most prominent applications of
morphing concepts for both two and three-
dimensional wing models. Various methods and
tools usually used for the design and analysis of
these concepts, ranging from aerodynamic to
structural analyses, and from control to optimization
aspects, are discussed. During the review process, it
became clear that the acceptance of morphing
concepts for routine use on aerial vehicles is still
limited, and some reasons for this are given. Lastly,
promising future applications are identified.
Designing the blade for low wind power density
regions was discussed in [9]. Wind turbine blade
aerodynamic airfoils have a significant impact on
wind turbine aerodynamic efficiency. This entails
selecting an appropriate airfoil section for the
proposed wind turbine blade. In their study, NACA
4412 airfoil profile was used to analyze wind
turbine blades. GAMBIT 2.4.6 is used to create the
airfoil geometry. CFD analysis is performed using
FLUENT 6.3.26 at various angles of attack ranging
from 0° to 120°. The coefficients of lift and drag are
calculated for a Reynolds number of 1x105. A
comparative study of various airfoils from the
NACA and NREL Airfoil families is presented in
[10], with a focus on their suitability for small wind
turbines. Four comparison criteria have been
considered in this case. Maximum glide ratio at
lower and higher Reynolds numbers, angle of attack
difference between lower and higher Reynolds
numbers, and percentage deviation of maximum
glide ratio from stall point are the criteria. XFOIL
analysis using Q-blade software yields the data
required for comparing two families of airfoils,
revealing that NACA airfoils have better average
performance criteria while NREL airfoils have
better stability criteria. The authors in [11],
determined aerodynamic coefficients for different
wing spans with various ground clearances, it was
found that short-span wings have the tendency to
delay the beginning of separation and eventually
lose negative lift. Due to vortices, there wasn’t a
significant change in the strength or size at the wing
end plate, these vortices, at short-span wings,
affected a larger percentage of the wing encouraging
the flow to stay attached and mitigate the opposite
pressure gradient which will lead to separation at
longer spans, as a result, it was demonstrated that
shorter span wings have lower lift coefficient as
compared to larger span wings. A reviewing for
flapping wing aerodynamics modeling, including
wing kinematics and the Navier-Stokes equation is
presented in [12]. Also reviewed was the
mathematical formulation of normal forces, chord-
wise forces, total forces, lift, and thrust. It has
recently been demonstrated that a flexible wing is
far superior to a rigid wing. The authors in [13],
investigated many flapping wing aerodynamics
topics numerically and experimentally. These topics
cover some of the most recent advances in flapping
wing aerodynamics, such as wake structure analysis,
the effects of airfoil thickness and kinematics on
aerodynamic performance, vortex structure analysis
around 3D flapping wings, and kinematics
optimization. Both experimental and numerical
approaches are used to investigate the wake
structures behind a sinusoidal pitching NACA0012
airfoil. The experiments are carried out using
Particle Image Velocimetry (PIV), and two types of
wake transition processes are distinguished, namely
the transition from a drag-indicative wake to a
thrust-indicative wake and the transition from a
symmetric wake to an asymmetric wake. The
developed SD solver's numerical results agree well
with the experimental results. The initial conditions,
such as the initial phase angle, are found
numerically to determine the deflective direction of
the asymmetric wake. The [14] is focused on
estimating the performance of a small wind turbine
blade with a suitable dimple arrangement at 25%
and the middle of the chord length for NREL S228
and S238, the conducted CFD analysis used k-Ɛ
turbulence model by ANSYS Fluent software by
which the aerodynamic performance and the
moment equations are solved, a delay flow
separation was observed at the dimple entrance
leading to creation of vortices, the investigation also
includes a simulation for the blade with an adequate
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overall performance by the help of GH BLADED
and the obtained results are discussed. According to
[15], an effort was made by simulating the selected
airfoils using Q Blade open-source software at the
National Renewable Energy Laboratory (NREL),
namely S823 and DU 06-W-200. Q Blade software
employs a special algorithm called the double
multiple stream tube (DMS) for the assessment of
horizontal axis wind turbines (VAWTs) and the
blade element method (BEM) for assessing
horizontal axis wind turbines (HAWTs). The
graphical user interface (GUI) of Q Blade includes
the viscous-in-viscid coupled panel process code
XFOIL for calculating the lift and drag coefficients
of an airfoil at any angle of attack (AoA). The
simulation is performed and compared at various
Reynolds numbers ranging from 1x105 to 3x105 for
both selected airfoils. For each applied Reynolds
number, results show that the S823 airfoil with a
higher lift coefficient up to 10° AoA, than the DU
06-W-200 airfoil has higher values, the pattern is
true for the lift-to-drag ratio. Lastly, the simulation
results are validated by comparing them to the
obtained experimental data, which shows good
agreement between the Q Blade simulation result
and those for experimental data. According to the
NACA four-digit wing sections, in which the profile
is defined as follows: one digit describes maximum
camber distance from the airfoil leading edge in tens
of percent of the chord; one digit describes
maximum camber as a percentage of the chord; and
then two digits describing the maximum thickness
of each airfoil as a percentage of the chord, [16].
When a stream of air flows over and under an
airfoil in motion, it produces a total aerodynamic
force. The point of impact is when the air splits and
flows around the airfoil. The point of collision
creates a high-pressure region or stagnation point.
The high-pressure region is often positioned near
the lower section of the leading edge, depending on
the angle of attack. This high-pressure region adds
to the overall force produced by the blade. The
entire aerodynamic force, also known as the
resultant force, may be separated into two
components: lift and drag. Lift acts on the airfoil
perpendicular to the relative wind. Drag is the
resistance or force that resists the airfoil's motion
through the air. It operates on the air.foil in a
direction that is parallel to the relative wind. Many
factors influence the overall lift produced by an
airfoil. Increased speed creates lift by creating a
bigger pressure difference between the top and
lower surfaces. Lift fluctuates with the square of the
speed, rather than increasing in direct proportion to
it, [17].
2 Problem Formulation
2.1 CFD Analysis of NACA 4412 and
NREL's S823 Airfoils
First of all, NACA 4412 airfoil coordinates file was
imported into ANSYS Design Modeler, and then a
C-type boundary was created around it for meshing,
CFD analysis, and post-processing results. A
pressure-based solver with a steady-state solution
was used in conjunction with the Spalart-Allmaras
viscous model. The fluid is air entering the domain
at a rate of 18 m/s, and the outlet boundary
condition is pressure-based. FLUENT generates a
residual for each governing equation that is solved,
and the residual indicates how well the present
solution meets the governing equation's discrete
form. The solution is iterated in this instance until
the residual for each equation is less than 1e-6. The
working pressure is 101.325 kPa, the turbulent
viscosity ratio is 10%, the airfoil chord length is 1m,
the pressure-velocity coupling scheme is SIMPLE,
and Second-Order Upwind is utilized to calculate
pressure and momentum. Figure 2 shows the NACA
4412 airfoil and its mesh, [18], [19], [20].
Fig. 2: The meshing used for NACA 4412
The study employed a steady-state, pressure-
based solver, and finite volume discretization to
solve the k-epsilon model’s governing equations.
Designers monitored the numerical solution error to
make sure it was converging properly. A 5 m radial
and 10 m long C-type computational domain was
selected. During simulations, air (density = 1.225
kg/m3 and dynamic viscosity = 1.7894 e-05 Pa s) is
employed as a fluid flow medium. This was
accomplished by using grids of varying sizes to
establish a mesh independence study. This was done
by raising the number of grid elements until the
solution demonstrated little change with additional
increases in the mesh density. The residuals of the
governing differential equation’s outcome variable
are used to assess the convergence speed throughout
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the iteration phase. Additionally, for each of the
incorporated force coefficients, the relative
differences between two sequential iterations are
used to verify convergence. Double precision, 2D
analysis was conducted using normal k-epsilon flow
equations, and a continuous flow of air was seen
around the plane’s perimeter. The lift force is
determined by the spacecraft’s weight, whereas the
drag force is determined by the airplane’s
aerodynamic efficiency and its wingspan. Table 1
shows boundary conditions applied on these airfoils.
Table 1. Boundary conditions applied
Air density
1.225 kg/m3
Viscosity
1.7894e-05 kg/m-s
Inlet velocity
18 m/s
Wall Motion
Stationary Wall
Shear Condition
No Slip
Outlet Gauge Pressure
0 Pa
Pressure-Velocity Scheme
Coupled
Pressure, Momentum, Turbulent
kinetic energy & dissipation rate
Second-order upwind
Gradient
Least Square Cell-Based
3 Results and Discussion
3.1 NACA 4412 Airfoil Results
The findings indicate an area of high pressure at the
leading edge (stagnation point) of the airfoil and a
low-pressure zone on the top airfoil surface. The
Bernoulli equation states that pressure and velocity
are inversely linked; thus, velocity will be lower in
high-pressure areas. The pressure applied to the
bottom surface of the airfoil was higher than the
pressure applied to the entering flow stream, and
therefore, the airfoil was simply forced upward,
perpendicular to the arriving flowing fluid. Figure 3
shows the Drag coefficient of NACA 4412 Airfoil
results.
Fig. 3: Drag coefficient graph
Figure 4 shows the lift coefficient as a function
with a number of iterations of analysis.
Fig. 4: Lift coefficient graph
Figure 5 shows the pressure coefficient as a
function of position (m) on the airfoil.
Fig. 5: Pressure coefficient chart
Figure 5 depicts a pressure coefficient chart; the
pressure distribution of NACA airfoil profiles is
estimated using the numerical panel technique for
2D lifting air flow circumstances. The fluctuation in
pressure coefficients along the chord is seen by
analyzing the airfoil shape exposed to various AOA
(angle of attack) conditions, including stalling
angles. The zero lift AOA of the profiles is also
examined to determine the impact of thickness-to-
chord ratios on airfoil properties. Figure 6 shows the
velocity contours of the NACA 4412 airfoil, it can
be noticed that red areas have maximum velocity
values while blue-colored areas have the minimum
values of velocity.
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Fig. 6: Velocity contours
Figure 7 shows the pressure contours of the
NACA 4412 airfoil, it can be noticed that red areas
have maximum pressure values while blue-colored
areas have the minimum values of pressure.
Fig. 7: Pressure contours
Figure 8 shows the pressure values at 0° attack
angle the maximum value of pressure at the tip of
the airfoil with 3141.270 Pa, while the minimum
pressure is 1570 Pa at the top of the airfoil.
Fig. 8: Pressure contours for 0° AoA
Figure 9 shows the velocity values at 0° attack
angle it is clear that the maximum value of velocity
at the top (upper surface) of the airfoil with 55 m/s,
while the minimum velocity is from 14-21 m/s at
the front and lower surface of the airfoil.
Fig. 9: Velocity contours for 0 deg. AoA
Figure 10 shows the pressure values at a 2°
attack angle it is clear that the maximum values of
pressure at the tip (front) of the airfoil with 155.252-
191.962 Pa, while the minimum pressure is -
138.428 Pa at the top (upper surface) of the airfoil
Fig. 10: Pressure contours for 2 deg. AoA
Figure 11 shows the velocity values at 2ᵒ attack
angle it is clear that the maximum values of velocity
at the top (upper surface) of the airfoil with 22.013-
24.459 m/s, while the minimum velocity is from
2.446-7.338 m/s at the front and lower surface of the
airfoil.
Fig. 11: Velocity contours for 2° AoA
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Figure 12, Figure 13, Figure 14, Figure 15,
Figure 16 and Figure 17 show values of pressure
and velocity at different values of attack angle (6, 8,
and 12 degrees) respectively.
Fig. 12: Velocity contours for 6° AoA
Fig.13: Pressure contours for 6° AoA
Fig. 14: Pressure contours for 8° AoA
Fig. 15: Velocity contours for 8° AoA
Fig. 16: Velocity contours for 12° AoA
Fig. 17: Pressure contours for 12° AoA
Table 2 shows the Final results of Cd and Cl for
NACA 4412.
Table 2. Results of Cd and Cl for NACA 4412
Angle of Attack
Drag coefficient
(Cd)
0
0.11
2
0.0137
6
0.0175
8
0.021
12
0.04
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A linear relation was observed between the
angle of attack (AoA) and Lift/Drag ratio up to 8°
AoA which means the Cd/Cl ratio increases with
increasing angle of attack up to 8°. On the contrary,
after 80 it showed an inverse relation, and the
Lift/Drag ratio started decreasing with increasing
AoA value.
3.2 NREL's S823 Airfoil
The airfoil families developed by the National
Renewable Energy Laboratory (NREL) are
generally resistant to relative roughness effects,
resulting in somewhat reduced yearly energy losses.
Additionally, the airfoils are usually modified to
have a thicker body, resulting in unexpected
performance characteristics. The use of blade tip
airfoils with a low Glade ratio and a scoop that
correlates with the control of the maximum power
may result in further performance improvement
while operating a stall-regulated turbine. This
allows 100 percent to fifteen tons of sweptback rotor
area for a given generator size, depending on the
design. The S-Series airfoils from NREL are
available in both thin and thick families. The thin
airfoil families are well suited for stalling controlled
wind turbines in situations where performance
losses due to airfoil change of state are critical
considerations. The change in the state of the airfoil
is not a significant disadvantage for variable pitch
and variable speed turbines. In most cases, the main
airfoil is used in conjunction with root and tip
airfoils. Most turbine blades are made up of a
circular portion that connects to the hub. NREL
airfoil curves are somewhat smoother and have a
distinctive form, even if the flow conditions change
while concerns with noise and discontinuity in
power production for stall-controlled wind turbines
may arise on airfoils with camber ridges on specific
NACA airfoils. Figure 18 shows the lift coefficient
of NREL’Ss823 Airfoil results with the number of
iterations.
Fig. 18: Lift coefficient graph for NREL's S823
Figure 19 shows the drag coefficient as a
function with the number of iterations of the
analysis.
Fig. 19: Drag coefficient graph for NREL's S823
Figure 20 shows the pressure coefficient as a
function of position (m) on the airfoil
Fig. 20: Pressure coefficient chart
Figure 21 shows the pressure contours of
NREL's S823 airfoil, it can be noticed that red areas
have maximum pressure values (3044.380 Pa) while
blue-colored areas has the minimum values of
pressure (608.876).
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Fig. 21: Pressure contours for NREL's S823 (AoA =
0 deg.)
Figure 22 shows the velocity contours of
NREL's S823 airfoil, it can be noticed that red areas
have maximum velocity values (71.782 m/s) while
blue-colored areas have the minimum values of
velocity (7.178 m/s).
Fig. 22: Velocity contours for NREL’s S823 (AoA
= 0 deg.)
Figure 23, Figure 24, Figure 25 and Figure 26
show the velocity and pressure distribution values at
8 and 12 attack angles respectively for NREL's S823
airfoil.
Fig. 23: Velocity contours for NREL's S823 (AoA =
8 deg.)
Fig. 24: Pressure contours for NREL's S823 (AoA =
8 deg)
Fig. 25: Pressure contours for NREL's S823 (AoA =
12 deg)
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Fig. 26: Velocity contours for NREL's S823 (AoA =
12 deg)
It can be noticed that for NREL's S823 airfoil,
there are no values determined at attack angles of 2
and 6 degrees. To compare the results of the
targeted airfoils at 1232877 Reynolds number and
the vast range of blade angles of attack from 0 to 12
degrees, the airflow simulations provided lift
coefficient and drag coefficients. The governing
equation was solved, and the flow issue was
addressed using the usual k-epsilon model, coupled
algorithm, and second-order upwind approach
provided in this study. The k-epsilon model with
better wall treatment is the most suited CFD model
due to its low error. The k-epsilon (k−ϵ) model for
turbulence is commonly used to simulate mean flow
characteristics for turbulent flow situations. It is an
Eddy viscosity model, which is a type of turbulence
model used to determine Reynolds stress. This is a
two-equation model. That is, in addition to the
conservation equations, it solves two transport
equations (PDEs) to account for historical effects
such as convection and turbulent energy diffusion.
Two variables are transported: turbulent kinetic
energy (k), which determines the energy in
turbulence, and turbulent dissipation rate (ϵ), which
defines the rate of dissipation of turbulent kinetic
energy, [21].
3.3 Comparison of the Two Airfoils
One objective of this study is to compare the values
of both lift and drag coefficient values for both
airfoils under study: NACA 4412 and NERL S823
airfoils at different attack angles, Table 3 shows a
comparison value of Drag coefficient and lift
coefficients at 0, 8, and 12 degrees of attack angles.
Table 3. CFD results comparison
NACA 4412
NREL’s S823
Angle
of
Attack
Lift
coefficient
(Cl)
Drag
coefficient
(Cd)
Lift
coefficient
(Cl)
Drag
coefficient
(Cd)
0
1.683
0.11
0.273
0.1355
8
1.113
0.021
0.9258
0.041
12
1.3376
0.04
0.983
0.094
Figure 27 shows a comparison between the
Drag coefficient values of the two airfoils NACA
4412 and NREL’s S823 airfoils.
From Table 3 and Figure 27, it can be noticed
that NACA 4412 airfoil has a more lift coefficient
values than that of NERL S823, while the NERL
S823 airfoil has more Drag coefficient values than
that of NACA 4412 airfoil at all attack angles this is
because of the airfoil shape and dimensions.
Fig. 27: Drag Coefficient of the two airfoils NACA
4412 and NREL S823
It can be noticed that values of Cd of NACA
4412 are lower than that of NREL S823 airfoils for
all values of angle of attack, also values for the two
airfoils are decreasing with AoA till 8 degrees and
then increase slightly. Figure 28 shows a
comparison between Lift coefficient values of the
two airfoils NACA 4412 and NREL’s S823 airfoils.
0
0,05
0,1
0,15
0 5 10 15
Cd
Attack Angle
Drag Coefficient of NACA 4412
and NREL S823 airfoils
Cd 4412 Cd NREL S823
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DOI: 10.37394/232013.2024.19.13
Sayel M. Fayyad, Aiman Al Alawin,
Suleiman Abu-Ein, Zaid Abulghanam,
Abdel Salam Alsabag, Mohannad O. Rawashdeh,
Muntaser Momani, Waleed Momani
E-ISSN: 2224-347X
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Volume 19, 2024
Fig. 28: Lift Coefficient of NACA 4412 and NREL
S823 airfoils
It can be noticed that the lift coefficient for
NACA 4412 values is higher than that of NREL
S823 airfoil but for NACA 4412 such values are
decreasing as AoA is increasing till 8 degrees of
AoA after that Cl values are increasing slightly. In
contrast, for NREL S823 airfoil the values of Cl are
increasing with AoA till 8 ᵒ after that they become
constant or decrease slightly.
From Table 4 the momentum exerted on the two
types of airfoils can be calculated using the
following equations
-For drag force (D)
D= (0.5)*Cd*A*ρ*V^2 (1)
Where Cd is the drag coefficient, A: is the reference
area, and ρ: is air density. And so, the momentum is
MD=D*S (2)
Where S is the distance (position from the tip of the
airfoil), the Lift force is given as:
L= Cl(A*0.5*ρ*V^2) (3)
So, the momentum resulting from lift force is given
as:
ML=L*S (4)
Using data from Table 3, the momentum
resulted from both drag and lift of the two wings (at
0, 8, and 12 AoA) as follows (assume ρ=1.4 kg/m3,
v is taken at average value, approximate area A=0.1
m2 (24 inches x 6 inches) for both). See Table 4 and
Table 5 for results.
Table 4. Drag and lift forces and momentum
calculations for NACA 4412 airfoil
NACA 4412
Angl
e of
Attac
k
Lift
coeffici
ent (Cl)
Drag
coeffici
ent (Cd)
Lift
force
N.
ML
N.m
(at
mid-
span
x=0.
3 m)
Drag
force(
N)
MD
0
1.683
0.11
152.6
82
45.8
0
9.98
3.00
8
1.113
0.021
19.95
5.98
0.376
0.11
12
1.3376
0.04
33.80
10.1
40
1.011
0.30
3
Table 5. Drag and lift forces and momentum
calculations for NREL S823 airfoil
NREL’s S823
Angl
e of
Attac
k
Lift
coefficie
nt (Cl)
Drag
coefficie
nt (Cd)
Drag
force(
N)
MD
(N.
m)
Lift
force
(N)
ML
(N.
m)
0
0.273
0.1355
12.22
3.66
24.77
7.43
8
0.9258
0.041
0.646
0.20
14.57
82
4.37
3
12
0.983
0.094
1.685
0.50
5
17.61
5
5.28
4
3.4 Discussion
It can be noticed that the drag coefficient for both
airfoils decrease as AoA increases to 9 degrees,
after this angle and for both airfoils it starts to
increase. Also, the CD for NACA 4412 is lower
than NERL S823 for 12 degrees (AoA). The lift
coefficient of NACA 4412 is higher than that of
NERL S823 for all angles of attack, CD for both
airfoils are decreasing as AoA increases to 9
degrees, then it increases slightly till 12 degrees for
NACA 4412 and slightly constant to decrease for
NERL S823 airfoil.
4 Conclusion
In the performance criterion, NACA airfoils have
shown superior results, while in the stability criteria,
things are opposite. Generally, ANSYS Fluent is
suitable for aerodynamic analysis of wind turbine
blades, results show that a turbulent layer produces
more significant drag at lower airfoil angles of
attack. In designing wind turbine blades, it is critical
to ensure that the airfoil utilized does not develop
any instabilities in operation. Special care is
required if using a pitch-controlled wind turbine.
Stabilizing the angle of attack is very important to
maintain proper operability. When looking at many
technical requirements that wind turbines must
satisfy, it is clear that one will have to choose
0
0,5
1
1,5
2
0 5 10 15
Cl
AoA (degrees)
Lift Coefficient of NACA 4412
and NREL S823 airfoils
NACA4412 NREL S823
WSEAS TRANSACTIONS on FLUID MECHANICS
DOI: 10.37394/232013.2024.19.13
Sayel M. Fayyad, Aiman Al Alawin,
Suleiman Abu-Ein, Zaid Abulghanam,
Abdel Salam Alsabag, Mohannad O. Rawashdeh,
Muntaser Momani, Waleed Momani
E-ISSN: 2224-347X
139
Volume 19, 2024
between several airfoils. These provide the
functions of crucial aerodynamic, mechanical,
reusability, and supportability requirements. Aspects
such as electromagnetic interference, acoustic noise
production, and aesthetic appearance are generally
expected to be less critical for alternate rotor
features. Traditionally, in aircraft lifting surface
theory, people assume a positive relationship
between high lift and low drag and that the lift-to-
drag ratio may be a significant concept. This art of
the debate is different from aircraft wing airfoils.
For the first issue, rotor performance reveals that the
product of the chord and the lift coefficient must be
greater than one. Operational at a better lift constant
will enable the use of smaller blades. When it comes
to vicious power losses, the overall viscous torsion
is determined by the L/D ratio of the airfoil, which
limits the vicious power losses, but the specific
amount of lift does not dictate viscous torsion itself.
The drag coefficient for both airfoils decreases as
AoA increases until 9 degrees, at which point it
begins to increase for both airfoils. Furthermore, for
12 degrees, NACA 4412 has a lower CD than
NERL S823 (AoA). For all angles of attack, the lift
coefficient of NACA 4412 is greater than that of
NERL S823. CD for both airfoils decreases as AoA
increases until 9 degrees, then increases slightly
until 12 degrees for NACA 4412 and remains
slightly constant to decrease for NERL S823 airfoil
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DOI: 10.37394/232013.2024.19.13
Sayel M. Fayyad, Aiman Al Alawin,
Suleiman Abu-Ein, Zaid Abulghanam,
Abdel Salam Alsabag, Mohannad O. Rawashdeh,
Muntaser Momani, Waleed Momani
E-ISSN: 2224-347X
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Volume 19, 2024
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Contribution of Individual Authors to the
Creation of a Scientific Article (Ghostwriting
Policy)
- Sayel M. Fayyad carried out the simulation
- Aiman Al Alawin: editing and writing
- Suleiman Abu-Ein: literature review
- Zaid Abulghanam: writing and figures
- Abdel Salam Alsabag: writing discussion
- Mohannad O. Rawashdeh: writing conclusion
- Muntaser Momani: Literature review.
- Waleed Momani: Methodology
Sources of Funding for Research Presented in a
Scientific Article or Scientific Article Itself
No funding was received for conducting this study.
Conflict of Interest
The authors have no conflicts of interest to declare.
Creative Commons Attribution License 4.0
(Attribution 4.0 International, CC BY 4.0)
This article is published under the terms of the
Creative Commons Attribution License 4.0
https://creativecommons.org/licenses/by/4.0/deed.en
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WSEAS TRANSACTIONS on FLUID MECHANICS
DOI: 10.37394/232013.2024.19.13
Sayel M. Fayyad, Aiman Al Alawin,
Suleiman Abu-Ein, Zaid Abulghanam,
Abdel Salam Alsabag, Mohannad O. Rawashdeh,
Muntaser Momani, Waleed Momani
E-ISSN: 2224-347X
141
Volume 19, 2024