Experimentaland Numerical Analysis over the Truncated Airfoil with Slotted Flap Configuration

Abstract

Objectives: The objective of this work is to investigate the flow over the truncated NACA2414 airfoil with slotted flap
configuration. Investigation is carried using wind tunnel and commercial CFD tools with the consequences of the variation
of speed, angle of attack of control surface and effective angle of attack. Methods/Statistical Analysis: Wind
tunnel tests are performed on a reduced scale model (3:10) in wind tunnel at Manipal University. Experimentally determined
static pressure distribution data at equal interval stations, placed over the wing are used to generate pressure
data. The results obtained are used in turn as a benchmark to validate the CFD simulations. Findings: Findings clearly
states that the blunt trailing edge has high drag coefficient. Although off set cavity and cavity has proven to be the
ideal modification for trailing edge shape in a plane wing from the study. It can be concluded that the same is not true
when such modifications are applied to slotted flap configuration. Application/Improvements: Further studies can be
carried out with the alternative designs and a rigorous research can contribute in efficient designs in high lift devices.

Full text

Experimental and Numerical Analysis over
the Truncated Airfoil with Slotted Flap
Conguration
U. Shiva Prasad1*, S. Ajay Varma1, R. Suresh Kumar2, K. M. R. Sree Vaibhav1 and
C. H. Satya Sandeep1
1Institute of Aeronautical Engineering, Dundigal, Hyderabad – 500 043, Telangana, India;
shivaprasad047@gmail.com, sagiajayvarma@gmail.com, rsuresh64@gmail.com, chsatyasandeep@gmail.com
2Department of Aeronautical and Automobile Engineering, Manipal Institute of Technology, Manipal University,
Manipal – 576104, Karnataka, India;
sreevaibhav@outlook.com
Abstract
Objectives         
            -
              Methods/Statistical Analysis: 
               -
  
               Findings -
                    
        
   Application/Improvements: 
       
*Author for correspondence
1. Introduction
Optimization of the aircra performance has been the
motivation behind the ever-expanding aircra industry.
Slightest of the design improvements has the potential
to boost the performance of the aircras in the long run.
e Slotted aps design has been used to increase the li
coecient of the primary wing to a considerable extent.
Any geometrical variation in the trailing edge of an airfoil
implies a considerable change in li and drag characteris-
Indian Journal of Science and Technology, Vol 11(22), DOI: 10.17485/ijst/2018/v11i22/115627, June 2018
ISSN (Print) : 0974-6846
ISSN (Online) : 0974-5645
tics of the airfoil1. is change in the li characteristics are
the byproducts of eddies formed at the end of the trailing
edge of the airfoil. e shape and nature of these eddies
depends on the several factors like the shape of the airfoil,
Reynolds number and angle of the incidence of the airfoil.
Dierent shapes for the trailing edges are considered to
study the eectiveness of the designs.
Rapid increases in the computational power of the
modern-day computers have assisted researchers across
the world to resort to the ComputationalApproaches to
Keywords:

Experimental and Numerical Analysis over the Truncated Airfoil with Slotted Flap Conguration
Indian Journal of Science and Technology
Vol 11 (22) | June 2018 | www.indjst.org
2
simulate the required environment to study the problem.
Computational uid dynamics is extensively used in aero-
dynamics and it has been proven to be a reliable source
for predicting the ow intricacy. In the context of this
study, several other studies carried out by independent
researchers have proven that oset cavity exhibits 50%
lesser drag compared to splitter plate and cavity congu-
rations. Similarly,trailing edge modicationslike splitter
plate method, Trailing edge wedge method, ventilated
cavity method and M-shaped serration has a wing con-
guration has been to know the most favorable shape for
the drag reduction2.
2. Methodology
Due to the experimental constraints of the wind tunnel,
Experiments were carried out on a 3:10 scaled model
witha chord length of 30 cm. e Computational analyses
were carried out at the true lengths and the experimental
model was manufactured by the use of CNC machine.
e placement of ribs was taken by considering the
structural factors of wing as shown in Figure 1. e
experimental analysis was carried by pressure tapping
methods. A total of 14 pressure taps have been arranged
symmetrically in order to calculate pressure coecient at
a speed of 30m/s.
e airfoils have been modeled in ICEM. Sections
of the NACA 2414 airfoil has been selected as the base
wing. e model is scaled to 25 percent of the one meter
chord and 0.02 gap is given in between the main and ap
as shown in Figure 2. Domain has been modeled to be 10
times larger than the chord of the airfoil. Control surfaces
are primarily used to during the take-o and the land-
ing phases of the ight. A nal angle of 12 degrees has
been obtained through an incremental angle of attack of 4
degrees for the iteration.
Multi body computational analyses have been one
of the most dicult problems due to the requirement of
higher mesh resolution at the interface. e gap between
the main wing and the control surface tend to complicate
the ow structure. e rst element height was chosen to
be 1.2E-04. Analysis had been carried out at a Reynolds
number of 2.1E6. Spalart-Allmaras model is a one equa-
tion model specically designed for the application
involving wall bounded ows and boundary layers with
steep pressure gradients. Advantage of Spalart-Allmaras
not only increases the accuracy of the solution3, but also
has an advantage of y+ insensitivity towards wall treat-
ment. K-Epsilon and K-Omega have also been considered
for the study and a grid independence study has been car-
ried out and the results are as follows
Figure 1. Depicts the wing skeletal body with the
placement of the ribs along the span of the wing.
Figure 2. Meshed model.

U. Shiva Prasad, S. Ajay Varma, R. Suresh Kumar, K. M. R. Sree Vaibhav and C. H. Satya Sandeep
Indian Journal of Science and Technology 3
Vol 11 (22) | June 2018 | www.indjst.org
3. Results
In the current section results are presented for the dif-
ferent sections of trailing edge modications of NACA
2412. e test results are compared with experimental
tests which were carried at MIT-MU using the sub-sonic
research tunnel facility.
In the current paper two trailing edge modications
are made as shown in the Figure 3 and 4 length of the
chord trailing edge was selected based the theory devel-
oped1,2. To obtain the eective results the trailing edge
was varied dierent angle of attack in range between 0 to
12 degree experimentally and numerically using 43 lakh
nodes computationally eective results using ANSYS
Fluent which is detailed in Table 1.
3.1 Validation of Results
To validate the results eciently and present eectively
both experimental and numerical analysis were carried
for the eight cases of the trailing edge modications with
dierent angle of attack. Results presented in the Figure
5-14 shows the points are nearly matched with numerical
results on a scale of unit chord length.
Applying the methodology the results are presented in
gures which shows that the main wing and the attached
trailing edge wing with main wing or xed wing trialing
modication i.e., cavity and blunt. e discussion will
start with blunt edge and followed by the cavity with the
gap between main wing and secondary wing are xed.
Figure 3. NACA 2414 airfoil with blunt shaped trailing
edge.
Figure 4. NACA 2414 airfoil with cavity shaped trailing.
Nodes Co-ecient of Li Co-ecient of Drag Co-ecient of
Moment
1.1 Million 0.50860 0.02225 0.24147
2.9 Million 0.49437 0.022082 0.23619
4.3 Million 0.49498 0.022048 0.2349
Table 1. Grid analysis (dual blunt with 12-degree control surface deection)

Experimental and Numerical Analysis over the Truncated Airfoil with Slotted Flap Conguration
Indian Journal of Science and Technology
Vol 11 (22) | June 2018 | www.indjst.org
4
In case of cavity section the results of pressure coef-
cient at the leading edge of the airfoil reached maximum
pressure for all the cases which shows that the maximum
pressure or stagnation pressure is raised on the leading
edge. As the results with trailing edge deections the vari-
ation of pressure coecient in that is of primary interest,
in investigating the airfoil to reduce drag over the airfoil
section, this is of interest for many years4,5. e deection
Figure 5. Cavity control surface at 0 degree with free
stream 0 degree.
Figure 7. Cavity control surface at 8 degree with free
stream 0 degree.
Figure 6. Cavity control surface at 4 degree with free
stream 0 degree.
Figure 8. Cavity control surface at 12 degree with free
stream 0 degree.

U. Shiva Prasad, S. Ajay Varma, R. Suresh Kumar, K. M. R. Sree Vaibhav and C. H. Satya Sandeep
Indian Journal of Science and Technology 5
Vol 11 (22) | June 2018 | www.indjst.org
of trailing edge variation has increased with increase in
pressure over the secondary airfoil section from which it
clearly states the increment in pressure results in increased
li contribution which is discussed in preceding sections
regarding the li and drag variation. e pressure distri-
bution over the primary and secondary section closely
matches with experimental data, except one point where
the suction pressure has increased with secondary wing
angle of attack this is due to the reason that ow induced
Figure 9. Blunt control surface at 0 degree with free stream
0 degree.
Figure 10. Blunt control surface at 4 degree with free
stream 0 degree.
Figure 11. Blunt control surface at 8 degree with free
stream 0 degree.
Figure 12. Blunt control surfaceat 12 degree with free
stream 0 degree.

Experimental and Numerical Analysis over the Truncated Airfoil with Slotted Flap Conguration
Indian Journal of Science and Technology
Vol 11 (22) | June 2018 | www.indjst.org
6
vortices eect was not clearly captured in the experimen-
tal analysis.
Blunt trialing edge airfoils are also of interest in the
current study which has high drag compared to the other
case because the low pressure in the wake acting on the
trailing edge. Figure 9-12 it clearly shows that the numeri-
cal results matches closely with experimental results but
the pressure variation at the gap has induced vortex ows
which increased the drag force comparatively. Results
clearly states that there are two stagnation points one over
each section.
4. Conclusion
Results obtained from the wind tunnel results justies
the accuracy of the turbulence model used in the com-
putational analyses. From the above experimental and
numerical results, it can be understood that the slight-
est of the variation in the geometrical variation would
bring signicant changes. e trailing edge modication
of main wing is depicted. It shows that the blunt has low
li coecient compared to the cavity although initially at
zero angle of attack it increased. With increment in trail-
ing edge angle of attack the li has increased till 11 degree
with increment in further angle result has varied by the
ow induced vortices which has inuenced the cavity
section resulting in the low li coecient. It clearly states
that the blunt trailing edge has high drag at low angle of
attack, which has reached to minimal point at angle of
4 degree in further increment the coecient of drag has
increased. Comparatively the case of cavity of low drag
at even zero angle of attack the variation of drag has
increased with angle. is results we can state that clearly
states that the blunt trailing edge has high drag coecient.
Although o set cavity and cavity has proven to be the
ideal modication for trailing edge shape in a plane wing
from the study1. It can be concluded that the same is not
true when such modications are applied to slotted ap
conguration.
5. Acknowledgements
Authors express sincere gratitude to all those who have
directly or indirectly contributed to this research work.
Figure 14. Drag plots of blunt and cavity shaped trailing
edged aerofoil in slotted ap conguration.
Figure 13. Li plots of blunt and cavity shaped trailing
edged aerofoil in slotted ap conguration.

U. Shiva Prasad, S. Ajay Varma, R. Suresh Kumar, K. M. R. Sree Vaibhav and C. H. Satya Sandeep
Indian Journal of Science and Technology 7
Vol 11 (22) | June 2018 | www.indjst.org
Also thank MIT-Manipal University, Manipal, for the
computational and experimental support given.
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3. Prasad US, Ajay VS, Rajat RH, Samanyu S. Aerodynamic
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5. Lávička D, Matas R. Computation of drag and li coe-
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