design and aerodynamic simulation of blended wing airplane

Design and Aerodynamic simulation of

Blended wing airplane developed using

NASA SC(2)–0714 airfoil

R. Achuthan1

1Student,Department of Mechanical Engineering

Kongu Engineering College, Perundurai- 638060, Tamilnadu, India

[email protected]

S. Gowri shankar2

2Student,Department of Mechanical Engineering


Email- [email protected]

K.S.K. Sasikumar3

3Associate Professor, Department of Mechanical Engineering


Email- [email protected]

Abstract- Narrow body airliners consume lots of fuel and they experience higher drag force which affect their

performance. To overcome this issue, a new concept called blended wing body (BWB) is introduced. The BWB approach

reduces wetted area and form drag at wing-body junction. Supercritical airfoil is invented primarily to delay the onset of

wave drag in the transonic speed range. So, one of its type NASA SC (2)-0714 airfoil is chosen and from which a new

BWB 3-D model is developed using Solidworks 2017. By setting the inlet freestream velocity of 250 m/s, computational

fluid dynamics (CFD) analysis is carried out over the model at various angles of attack (AOA) such as 00,50,100 up to 350

using the software STAR CCM+13, R8 where revision of boundary condition is done at each AOA to speed up the process

and k-ω SST (Shear Stress Transport) turbulence model is chosen. Aerodynamic properties such as Lift, drag, lift

coefficient, drag coefficient and moments are calculated at each AOA and their respective velocity and pressure plots are

obtained. Results show that the lift-to-drag ratio which determines the performance of the plane is maximum at 00 AOA.

Its value is 15.116. Also, the critical AOA, an AOA up to which the plane can have maximum lift beyond which it will

reach stall where plane’s lift suddenly decreases, is also found to be 250 which is a greater than narrow body airliners.

Thus, the performance and properties of new BWB model are studied.

Keywords – Blended wing body; Supercritical airfoil; K-ω SST Turbulence; Lift-to-drag ratio; Performance; Critical

angle of attack; Stall (Fluid dynamics).

I. INTRODUCTION

The technology has been developing day by day as the research on new developing new concepts and ideas keeps

rising. The main reason for such a technological development is to comfort the humans and mainly all living beings

in the world. Various transportation systems have been developed such as roadways, railways, seaways and airways.

Among which one of the fastest as well as the safest way of transport is airways. Aviation industry is essential for

the transportation of both the people and cargo. So the number of aviation industries are on the race of developing

and manufacturing efficient airplanes keeping in mind the safety is the first priority.

The types of aircrafts are subsonic, supersonic and hypersonic planes. A subsonic aircraft is an aircraft which runs at

an max speed less than the speed of sound (Mach 1). It describes that the aircraft flies below its critical Mach

number, around 0.8. As of now the most commonly the commercially used planes are subsonic planes. Supersonic

and hypersonic are only used for research purposes because it is not suitable for passengers. So planes are efficient

there are some factors which affect their performance. One of the biggest airplanes, Airbus A380 which shows good

performance at higher altitude as said by Olejniczak et al. 2019 [1] but it consumes lot of fuel. The drawbacks of

subsonic planes are,

Journal of Xi'an University of Architecture & Technology

Volume XIII, Issue 3, 2021

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High fuel consumption during take-off.

Higher drag force acting over the body.

Large number of functional parts and high development costs.

In order to overcome these drawbacks, the research is going on to change the whole structure of the aircrafts and lots

of engineers and researchers are working on it. On doing the research, they first relied on Wind tunnel tests but it is

costly so they shifted to utilize the CFD packages like Ansys Fluent, Star CCM+, Openfoam to run the simulations

to find out the behaviour of aircrafts at various conditions. Finally a new approach called Blended Wing Body

(BWB) is introduced. The aircraft has distinct wing and body structures, which are smoothly blended together

having no clear dividing line. The development of BWB is not easy as Larkin et al. 2017 [2] developed with an

implementation of stabilizers andPanagiotou et al. 2018 [3] developed optimal airframe engine combination. This

type of model require lots of design optimization techniques right from the airfoil section, to get the better results.

II.STATE OF ART OF BLENDED WING AIRPLANE DEVELOPMENT

Olejniczak et al. 2019 [1] developed Airbus A380 3D model using solidworks and analyzed using Ansys fluent. Geometry setup, boundary conditions and meshing was done on the model. The altitude value is set to 13000m, atmospheric pressure is 16,000Pa, air density is 0.2556 kg/m

3, and air velocity is 250 m/s. Then the results were

plotted such as pressure and velocity distribution. The similar experiment was carried out at various angles of attack (AOA), tabulated the lift, drag coefficients and plotted in the graphs at different AOA. From the plot obtained, at AOA angle 17

0 the coefficient of lift suddenly drops and coefficient of drag suddenly rises. So, it was concluded that

its critical AOA was 170.

Larkin et al. 2017 [2] developed a new model of Blended wing airplane in which its design is optimized in such a way that the plane can have less drag. Two BWB models have been developed one with twin vertical stabilizers and one with twin inclined stabilizers. It is found that inclined stabilizers generate greater drag than vertical ones. But the inclined stabilizers generate good response curves than vertical ones.

Panagiotou et al 2018 [3] developed a new concept of BWB in a form of UAV design. a step-by-step layout design study was conducted to define the key characteristics and select the optimal airframe-engine combination. When compared with the conventional UAV model the difference in aerodynamic efficiency is at the order of magnitude of 30%.

A new model of BWB Airplane with NACA-0012 airfoil was developed by Dakka Dr et al. 2019 [4]. It was modelled with different percentage areas of airfoil around different spanwise locations. Then the CFD Simulations were carried over the model. The actual model was fabricated and the wind tunnel tests were also carried out for the validation. This BWB study proved to have an L/D increase of 9.4% at α = 5

0 than a conventional aircraft (Tube and wing)

comparison. Overall, the work has concluded that the BWB performs 9.4% better than the conventional comparison.

Flying V aircraft , a tailless, V-shaped flying wing aircraft was developed by Faggiano et al. 2017 [5]. The goal of the study is to estimate the lift-to-drag ratio of the configuration at the cruise condition: M = 0:85, h = 13; 000m, and CL = 0:26. Two design approaches are investigated: a dual-step optimization, where planform and airfoil variables are subsequently varied, and a single-step optimization, where planform and airfoil variables are varied simultaneously. Lift to drag ratio obtained was 23.7 which is 25% greater than the NASA benchmark value(18.9).

Sethunathan et al. 2014 [6] conducted CFD analysis for various supercritical airfoils such as 0406, 0412, 0706 and 1006. They performed the analysis of these airfoils individually using Ansys ICEM-CFD.The pressure distribution was one to verify global and local effects of airfoil structure.The test was conducted in subsonic velocity of 25m/s. The most important thing is that a cusp like structure at the trailing edge of an unsymmetrical airfoil produces a very high improvement in climbing performance of an aircraft (Sethunathan et al. 2014) The result shows that the supercritical airfoil (0406) configuration reduced the drag and improve coefficient of lift (CL) by 15-20% compared with the baseline model.

Khamedov et al. 2017 [7] conducted a computational study on NACA SC(2)-0714 airfoil was at mach number of 0.72 and Reynolds number of 35x10

6. The airfoil was analysed using ANSYS CFX at angles of attack of 2

0 and 10

0 K-ω

and Shear stress transport (SST) turbulence models were chosen. Two suction slots were introduced along the airfoil contour to determine their control effectiveness. They found that active control is ineffective at low angles of attack, but very effective to increase airfoil performance at high angles of attack. After doing the literature review of the appropriate articles we can say that airplane’s design and the selection of suitable airfoil plays a crucial role in developing the BWB model as both the performance and its behaviour at various angles of attack have to be studied. Thus, we can say that,



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• The blended wing body approach will enhance performance, fuel economy and reduction in drag. It enhances more stability.

• Supercritical airfoil is designed to delay the onset of wave drag in the transonic speed range. When air flows over its top surface it does not speed up as much as over curved upper surface.

• The presence of cusp like structure at the trailing edge of the supercritical airfoil will improve the climbing performance of the airplane.

• Angle of attack (AOA) is the angle between reference line of body and direction of wind. Critical AOA is the max angle above which the plane will not get required lift force that it will experience stall.

• Lift to drag ratio determines the performance of airplane as this ratio is dependent on shape of airplane not the size of airplane.

Figure 1. Conventional and supercritical airfoils

III. PROBLEM DEFINITION AND OBJECTIVE

3.1. Problem definition

The need for blended wing body approach is that the existing model (Tube and wing configuration) suffers somewhat high drag force and high fuel consumption. Various research has been done on design of body of the plane. A new concept called Blended Wing Body (BWB) is introduced which reduces wetted area and form drag at wing-body junction.Very limited studies are available for selecting the suitable airfoil. Supercritical airfoil [8] is designed primarily to delay the onset of wave drag in the transonic speed range. One of its type is NASA SC(2)-0714. It is a non-symmetrical airfoil. Even though these types of airfoils are used for supersonic airplanes, it can also be used for medium speed(subsonic) aircrafts [6]. So, using this supercritical airfoil the BWB model is developed and analyzed to study aerodynamic properties and to measure its performance and critical angles of attack.

3.2 Objective

Figure. 2. Flowchart of BWB development process



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IV. METHODOLOGY

A methodology is a representation, which project managers employ for the design, planning, execution and accomplishment of their project objectives. It incorporates the theoretical analysis of methods and principles associated with a branch of knowledge. In this work, we will start from the design part to analysis part in a step-wise manner.

4.1 Design process

First the nomenclature of the airfoil is clearly studied. The design of NASA SC(2)-0714 airfoil is generated from data file values available at website: (http://airfoiltools.com/airfoil/details?airfoil=nasasc2-0714-il). This is a modified type of airfoil as the data file values are corrected by Raymer w/one method [9].

Figure 3. Airfoil nomenclature

The data file is copied to Excel sheet and arranged in order of X, Y and Z. Values of Z are assigned 0. The data file values of airfoil are saved as type text(delimited). Now CAD package Solidworks 2017 is loaded. Go to Insert, select Curve and select Curve through XYZ points. The type .rtf is selected and the airfoil date file is loaded. Using this section as base the entire BWB model is developed.

(a) (b)

Figure 4. NASA SC(2) – 0714 airfoil. a) Outline of airfoil b) Generated airfoil

(a) (b)

Figure 5. a) Conceptual 3D model of BWB Airplane b) 2D Sketch of BWB Plane and its engines



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The conceptual model of airplane is developed by referring the previous BWB designs in the research articles and

the Raymer’s airplane design approach [10]. The 3D model is an optimized one which is done by repeating the CFD

process again and again for 00 AOA till we arrive desired velocity, pressure plots and performance which we will be

discussing further. The point we try to convey here is that the developed 3D model is actually done on trial and error

way in CFD process to get optimized design. The entire span of the BWB Plane consists of the same NASA SC(2)-

0714 Airfoil. It is developed using surface modeling approach and it is saved in Parasolid format to convert it into

solid model and for importing it to the Star CCM+ 13, R8.

4.2 CFD Analysis

Star CCM+ is one of the leading CFD softwares which have been used by many automobile and aerospace

companies. Computational fluid dynamics involves three major steps. They are Preprocessing, Meshing and Post

processing. In order to perform CFD analysis using this software, the processes [11] involved in order are

Preparation of geometry, Construction of simulation topology, Generation of mesh, Defining the Physics and

Boundary condition, preparation for analysis, running the simulation, and finally analyzing the results.

In geometry preparation, a sketch of rectangle is drawn to cover the region around the plane. Sketch is extruded to

create a 3D Domain around the BWB plane. The domain should be created in such a way that inlet surface is nearer

to3D model and outlet surface is far away from it in order to avoid reversedflow at the outlet during simulation.

Using Boolean subtract, the BWB plane body is subtracted from the domain.

Figure 6. Geometry setup and face names of enclosure and airplane

The next process is construction of simulation topology which is shown in Table 1.

Table -1 Topology table

S.no Boundaries Type

1 Top, Bottom, S1, S2 Symmetry plane

2 BWB, E1, E2 Wall

3 Inlet Velocity inlet

4 Outlet Pressure outlet



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The meshing process is very essential. Automated mesh option is selected under which the following meshers are

enabled. Surface remesher, trimmed cell mesher, prism layer mesher. For refining the mesh as done in [12] and [13],

various inputs have to be given. A section plane is created at the center of the plane to have an overall view of the

mesh around the domain. In Target surface size, size type is set to absolute and absolute size is set to 0.1 m. In

Minimum surface size, size type is set to absolute and absolute size is set to 0.01 m. In surface curvature, curvature

deviation distance is enabled, #points/circle is set to 60, #Max points/circle is set to 360, curvature deviation

distance is set to 0.01 m. In surface proximity search ceiling is enabled. Surface growth rate is set to 1.2. Number of

prism layers is set to 5. Prism layer stretching is set to 1.2. In prism layer total thickness, size type is is set to

absolute and absolute size is set to 0.1 m. In maximum cell size, size type is set to relative to base and percentage of

base is set to 100.

(a) (b)

Figure 7. a) Mesh of BWB plane and b) Section mesh around its enclosure

The next step is setting up the physics and boundary condition. Implicit unsteady flow type is chosen. Domain

is filled with gas. Segregated flow type is chosen. Air is set to have constant density. Flow is set to turbulent. K-ω

SST(shear stress transport) turbulence [7] is chosen. K-ω SST turbulence model is used assuming near wall

condition. It is a two-equation eddy viscosity model. It can be used for boundary layer problems, where the

formulation works from the inner part throughout the viscous sub-layer upto the walls. This model is used where

Reynolds number is low.ie, flow is more viscous. Governing equations as derived by Menter [14] and [15], are the

equations in which the solver tries to solve for each regions of the mesh. This process will be running in the

background of the software when the simulation is started. Next step is to set the boundary conditions. The initial

freestream velocity value of (250,0,0) is assigned as the initial condition under both the physics and boundaries

options.

The next step is preparation for analysis. We are going to extract lift, drag, moment and their coefficients at 00

angle of attack. Lift, drag, moment plots are assigned.Lift, lift coefficient and moment directions are assigned

(0,1,0).Drag and drag coefficient directions are assigned (1,0,0).The parts where we are going to calculate the lift,

drag, moment forces and their coefficients are BWB, E1, E2. Frontal area report [16] is plotted by assigning (1,0,0)

as view up position, (0,1,0) as normal position and is taken as reference area for calculating coefficient of lift and

drag. Reference area is found to be 3.157728m2. Reference air Density is assumed 1.225kg/m

3.Plot windows for lift,

moment, drag forces, their coefficients, velocity and pressure plot windows are opened.

Next process is running the simulation. First the process is initialized and the run button is clicked. When run

button is clicked, the solver tries to solve the governing equations at each region of the mesh surface by iterative

method and plot the results. Sometimes the solution will reach converged, the process stops itself. Sometimes it will

not stop, i.e. convergence will not be reached.In order to get the desired result, the process is manually stopped if

there is only little difference between the successive iteration plots as done in Figure9d, before which the it should

not be stopped.



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Final step is analyzing the results. The lift, drag, their coefficients and the moment values are plotted in the

Table.3. The pressure and velocity plots are displayed in Figure. 11 and Figure. 12. Lift coefficient is a

dimensionless coefficient that relates the lift force(L) generated by a lifting body to the fluid density(ρ) around the

body, the fluid velocity(V) and its frontal area(A).

Cl =L

(A × 0.5 × ρ × V ) (1)

Cd =D

(A × 0.5 × ρ × V ) (2)

Similarly, the same process is repeated for calculating forces, their coefficientsmoments at various angle of attack

and plots are obtained.For doing this there is no need to edit the geometry again, it may take muchprocessing time.

Instead of doing so we can change the boundary condition values and direction vectors of report values to calculate

results. Those values are given in the Table 2.According to boundary condition inputs as per NASA airplane body

forces in a climb [17], lift force will be always generated perpendicular to the free stream velocity, and drag force

acts towards the free stream irrespective of geometry of the body. The main advantage of revising boundary

condition is that we don’t have to repeat the steps again for successive AOA, thus reducing processing time.

(a) (b)

Figure 8. a) Inlet Boundary condition and forces acting at an AOA (actual) b) Revised inlet boundary conditions and forces acting without

changing the geometry

(a) (b)

(c) (d)

Figure 9. (a) Lift, (b) Drag, (c) Moment and (d) Residual monitor plots(iterations) at 00 AOA (Analysis done up to 300 iterations)



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Table

Negative sign in table indicates the direction is set to opposite as shown in Figure. 8b.

Figure 10. Velocity streamline around BWB (At 0

(a) (b)

AOA(α)

(Degree)

Direction Vectors of D, Cd

(cosα, sinα,0) (m)

0 (1,0,0)

5 (0.9962,0.08715,0)

10 (0.9848,0.17365,0)

15 (0.9659,0.2588,0)

20 (0.93969,0.3420,0)

25 (0.9063,0.42262,0)

30 (0.866025,0.5,0)

35 (0.81915,0.57357,0)

Table -2 Inlet boundary conditions and directions

Negative sign in table indicates the direction is set to opposite as shown in Figure. 8b.

Figure 10. Velocity streamline around BWB (At 00AOA)

Direction Vectors of D, Cd

Direction Vectors of L, Cl

and Moment

(-sinα, cosα,0) (m)

Inlet Velocity (V) Components

(Vcosα, Vsinα, 0) (m/s)

(0,1,0) (250,0,0)

(0.9962,0.08715,0) (-0.08715,0.9962,0) (249.04867,21.7889,0)

(0.9848,0.17365,0) (-0.17365,0.9848,0) (246.2019,43.41204,0)

(-0.2588,0.9659,0) (241.4814,64.70476,0)

(0.93969,0.3420,0) (-0.3420,0.93969,0) (234.923,85.5050,0)

(0.9063,0.42262,0) (-0.42262,0.9063,0) (226.5769,105.65456,0)

(-0.5,0.866025,0) (216.50635,125,0)

(0.81915,0.57357,0) (-0.57357,0.81915,0) (204.788,143.3941,0)

Inlet Velocity (V) Components

, 0) (m/s)

(249.04867,21.7889,0)

(246.2019,43.41204,0)

(241.4814,64.70476,0)

(234.923,85.5050,0)

(226.5769,105.65456,0)

(216.50635,125,0)

(204.788,143.3941,0)



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(c) (d)

Figure 11. At 00 AOA, Pressure distribution

Figure 12. At 250 AOA, Pressure distribution(a) at front and(b)

Table

S.no

1

2

3

4

5

6

7

8

From the Table. 4 it is clearly shown that the value of lift and lift

AOA, the lift and lift coefficient values start decreasing. The drag, drag coefficient and moments values increases as

AOA increases. So the critical angle of attack is found to be 25

performance can be found out using the formula,

AOA, Pressure distribution (a) at front and at (b) bottom, Velocity contour(c) at centre and(d)

(a) (b)

(c) (d)

AOA, Pressure distribution(a) at front and(b) at bottom, Velocity contour(c) at centre and(d) at engine region

Table -3 Simulations iterated till convergence for all AOA

AOA No. of iterations

0 300

5 500

10 900

15 1750

20 1800

25 2250

30 3675

35 4000

V. RESULTS AND GRAPHS

From the Table. 4 it is clearly shown that the value of lift and lift coefficient increases as AOA rises. But after 25


AOA increases. So the critical angle of attack is found to be 250 as we also see the variation in Fig

performance can be found out using the formula,

at engine region

bottom, Velocity contour(c) at centre and(d) at engine region

coefficient increases as AOA rises. But after 250


riation in Figure 13a. The



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k = tan α =L

D=Cl

Cd

(3)

Table 4 Values of lift, drag, their coefficients, moments and k

Figure. 13a shows that coefficient of lift is maximum at 250AOA and reduces as AOA rises. This value is greater

than the conventional airplanes as they have critical AOA in range 150 to 20

0 as it analyzed that Airbus A380 has

critical AOA of 170 [1]. From Figure. 13d Lift to drag ratio (k=L/D or CL/CD) is found to be maximum at 0

0 AOA.

Its value is 15.116 which is greater than as obtained in whose BWB model developed using NACA-0012 airfoil [4]

which has max k value is 11.08. Thus the performance of the airplane and critical angle of attack have been

improved when NASA SC-2(0714) airfoil is used.

(a) (b)

AOA

(α)

(Degree)

Lift (L)

(N)

Coefficient

of Lift

(Cl)

Drag (D)

(N)

Coefficient

of Drag

(Cd)

Moment(Nm) k=

(Cl/Cd)

or (L/D)

0 38835.179 0.33240 2569.2937 0.02199 7407.7588 15.116

5 48252.257 0.41290 6221.4140 0.05320 17436.579 7.7610

10 57878.220 0.49540 11162.8775 0.09554 32152.828 5.1850

15 71277.410 0.61006 18110.1720 0.15500 51840.369 3.9360

20 80006.650 0.68477 25898.3200 0.22166 74633.947 3.0890

25 92646.880 0.79296 36634.7420 0.31350 105592.505 2.5290

30 90003.970 0.77030 44517.7090 0.38102 128038.408 2.0217

35 84947.470 0.72706 51343.8330 0.43940 147869.540 1.6546



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(c) (d)

Figure 13. a) CL plot b) CD plot c) Moment plot d) k plot at successive AOA

VI. CONCLUSION

Using CFD approach the performance of the Blended wing airplane with supercritical airfoil is assessed at an

airflow velocity of 250 m/s and the Aerodynamic properties of the airplane are studied. From the results of the

analysis we can conclude that the,

Two-equation RANS (Reynolds averaged Navier stokes), k-ω SST turbulence model gives good results for

external aerodynamic analysis of BWB plane as it can be easily applied to viscous affected region without

further modification. Even though Spalart allmaras model can also be used, they have limitations under

predicting flow separation and formation of decaying turbulence.

The critical AOA, maximum AOA above which the airplane no longer experience sufficient lift force for the

developeD BWB model is 250 which is much greater than the Airbus A380(17

0).

The Cusp like structure at the trailing edge of the NASA SC(2)- 0714 airfoil has its effect on the both the

performance and the Lift enhancement.

The entire process of CFD analysis has been made simple, easier and faster as the revision of boundary

conditions (Figure. 8b) are made to perform analysis for each angle of attack instead of changing the geometry

(positioning the model at an angle) and repeating the process again. The latter process is more time consuming.

The performance is given by L/D ratio which reduces as AOA increases and its maximum value is 15.116,

showing 36.426% improvement than NACA-0012 airfoil type used in BWB [4].

(a) (b)

Figure 14. Comparison charts a) Max L/D value for airplanes b) Critical AOA for airplanes.

NOMENCLATURE

A Reference area or frontal area CD Coefficient of drag CL Coefficient of lift CP Centre of pressure D Drag k Lift to drag ratio L Lift RL Reference line or chord line V Freestream inlet velocity α Angle of attack ρ Air density



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ACKNOWLEDGEMENTS

We would like to thank our project supervisor Dr. K.S.K. Sasikumar for his valuable guidance to our project. We

would like to thank our Head of the department Dr. R. Rajasekar for providing all the facilities to carry out this

project work and whose encouraging part is indispensable for our success.

CONFLICT OF INTEREST

No potential competing interest was reported by the authors.

REFERENCES

[1]. Olejniczak, D. and Nowacki, M., 2019. Assessment of the selected parameters of aerodynamics for Airbus A380 aircraft on the basis of

CFD tests. Transportation Research Procedia, 40, pp.839-846.

[2]. Larkin, G. and Coates, G., 2017. A design analysis of vertical stabilisers for Blended Wing Body aircraft. Aerospace science and

technology, 64, pp.237-252.

[3]. Panagiotou, P., Fotiadis-Karras, S. and Yakinthos, K., 2018. Conceptual design of a blended wing body MALE UAV. Aerospace Science

and Technology, 73, pp.32-47.

[4]. Dakka Dr, S. and Johnson, O., 2019. Aerodynamic Design and Exploration of a Blended Wing Body Aircraft at Subsonic Speed.

International Journal of Aviation, Aeronautics, and Aerospace, 6(5), p.17.

[5]. Faggiano, F., Vos, R., Baan, M. and Van Dijk, R., 2017. Aerodynamic Design of a Flying V Aircraft. In 17th AIAA Aviation Technology,

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[6]. Sethunathan, P., Niventhran, M., Siva, V. and Kumar, R.S., 2014. Analysis of aerodynamic characteristics of a supercritical airfoil for low

speed aircraft. International Journal of Research in Engineering and Technology, 3(06), pp.179-183.

[7]. Khamedov, R., Baitlessov, R. and Rojas-Solórzano, L., 2017, November. CFD Study of Effects of Boundary Layer Suction on Transonic

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[8]. Supercritical airfoil, Aircraft wing design https://en.wikipedia.org/wiki/Supercritical_airfoil

[9]. SC(2)-0714 Supercritical airfoil (coordinates from Raymer w/one correction) http://airfoiltools.com/airfoil/details?airfoil=nasasc2-0714-il

[10]. Raymer, D., 2012. Aircraft design: a conceptual approach. American Institute of Aeronautics and Astronautics, Inc..

[11]. Flow Investigations with STAR CCM+ Tutorial: Fundamentals https://www.youtube.com/watch?v=Y-clr8RwnQw

[12]. Star-CCM+: Surface/Volume/Automated Mesh Basics and Initial Mesh https://www.youtube.com/watch?v=KAcfF3XXMSo

[13]. STAR CCM+ Helicopter Fuselage Aerodynamics https://www.youtube.com/watch?v=f2vByNyutw4

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[16]. Incompressible Flow over NACA Aerofoil - CFD using Star ccm https://www.youtube.com/watch?v=l6y80C2sVlE&t=4079s

[17]. NASA Glenn research center, Forces in a climb. https://www.grc.nasa.gov/www/k-12/airplane/climb.html



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