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International Journal of Mechanical Engineering and Technology (IJMET) Volume 8, Issue 6, June 2017, pp. 471–479, Article ID: IJMET_08_06_049

Available online at http://www.iaeme.com/IJMET/issues.asp?JType=IJMET&VType=8&IType=6

ISSN Print: 0976-6340 and ISSN Online: 0976-6359

© IAEME Publication Scopus Indexed

CONCEPTUAL DESIGN OF FLYING VEHICLE

Alka Sawale

Assistant Professor, Department of Aeronautical Engineering,

MLR Institute of Technology, Hyderabad, India

Sreekanth Sura


MLR Institute of Technology, Hyderabad, India

Anitha D


Institute of Aeronautical Engineering, Hyderabad, India

B. Subbaratnam

Professor and Head, Department of Mechanical Engineering,

Vardhaman College of Engineering, Hyderabad, India

ABSTRACT

The primary objective of this report is to design a solar powered unmanned aerial

vehicle with less weight of around 3kg.Another objective of this report is to provide an

initial selection of the solar powered UAV. A comparative study of other solar

powered will be done to have configurations of this solar powered UAV. For initial

configuration there will be a weight estimation for this UAV, initial selection of weight

distribution will be discussed. Since Power is most crucial parameter for solar

powered UAV, so will be looking at fundamental equations of power. Finally a drag

polar estimation will be done.

Key words: Unmanned Aerial Vehicle (UAV), solar power, weight estimation, weight

distribution.

Cite this Article: Alka Sawale, Sreekanth Sura, Anitha D and B. Subbaratnam.

Conceptual Design of Flying Vehicle. International Journal of Mechanical

Engineering and Technology, 8(6), 2017, pp. 471–479.

http://www.iaeme.com/IJMET/issues.asp?JType=IJMET&VType=8&IType=6

1. INTRODUCTION

A flying wing is sometimes represented as theoretically the most aerodynamically efficient

(lowest drag) design configuration for a fixed wing aircraft. It also would offer high structural

efficiency for a given wing depth, leading to light weight and high fuel efficiency. Because it

lacks conventional stabilizing surfaces and the associated control surfaces, in its purest form

Conceptual Design of Flying Vehicle


the flying wing suffers from the inherent disadvantages of being unstable and difficult to

control.

These compromises are difficult to reconcile, and efforts to do so can reduce or even

negate the expected advantages of the flying wing design, such as reductions in weight and

drag. Moreover, solutions may produce a final design that is still too unsafe for certain uses,

such as commercial aviation. Further difficulties arise from the problem of fitting the pilot,

engines, flight equipment, and payload all within the depth of the wing section. Other known

problems with the flying wing design relate to pitch and yaw. Pitch issues are discussed in the

article on tailless aircraft. In some flying wing designs, any stabilizing fins and associated

control rudders would be too far forward to have much effect, thus alternative means for yaw

control are sometimes provided. One solution to the control problem is differential drag: the

drag near one wing tip is artificially increased, causing the aircraft to yaw in the direction of

that wing. Typical methods include. A consequence of the differential drag method is that if

the aircraft maneuvers frequently then it will frequently create drag. So flying wings are at

their best when cruising in still air: in turbulent air or when changing course, the aircraft may

be less efficient than a conventional design.

Figure 1 Bi-directional flying wing. Top-down view

The supersonic bi-directional flying wing design comprises a long-span low speed wing

and a short-span high speed wing joined in the form of an unequal cross. The proposed craft

would take off and land with the low-speed wing across the airflow, then rotate a quarter-turn

so that the high-speed wing faces the airflow for supersonic travel has funded a study of the

proposal The design is claimed to feature low wave drag, high subsonic efficiency and little or

no sonic boom. The proposed low-speed wing would have a thick, rounded airfoil able to

contain the payload and a long span for high efficiency, while the high-speed wing would

have a thin, sharp-edged airfoil and a shorter span for low drag at supersonic speed.

The Defense Advance Research Academy (DARA) has researched a solar powered HALE

UAV. The main idea behind the project called vulture is to combine the key benefits of both

an aircraft and a satellite into one system and to keep these systems in air continuously for 5

years. If the analysis is successful we will undergo the process of making prototype of this

UAV with some advanced features included in it.

Alka Sawale, Sreekanth Sura, Anitha D and B. Subbaratnam


2. MISSION SPECIFICATION

2.1. Mission Profile

The mission profile for this mission is shown as:-

Figure 2 Mission Profile

Table 1 Dimensions

WING SPAN 280

ROOT CHORD 80

TIP CHORD 50

SWEEP ANGLE 33.34 degrees

WING AREA 18200

2.2. Swept Wings

2.2.1. Neutral Point and Stability

We have already learned that the center of gravity must be located in front of the neutral

point. While the n.p. of an un swept, rectangular wing is approximately at thec/4point, the n.p.

of a swept, tapered wing must be calculated. The following procedure can be used for a

simple, tapered and, swept wing. First, we

Calculate the mean aerodynamic chord length of a tapered wing, which is independent

from the sweep angle:

With the root chord lr, the tip chord lt and the tap. We can also calculate the span wise

location of the mean Chord , using the span b,

The n.p. of our swept wing can be found by drawing a line, parallel to the fuselage center

line, at the spanwise station y. The chord at this station should be equal to . The n.p. is

approximately located at the c/4 point of this chord line (see the sketch below).



Figure 3 Geometric parameters of a tapered, swept wing

Instead of using the graphical approach, the location of the neutral point can also be

calculated by using one of the following formulas, depending on the taper ratio:

, if taper ratio> 0.375

, if taper ratio< 0.375

The c.g must be placed in front of this point, and the wing may need some twist (washout)

to get a sufficiently stable wing.

2.3. Finding the Required Twist ßreq

Using graph 1, we enter the graph with the aspect ratio AR on the horizontal axis, and draw a

vertical line upwards, until we intersect the curve, corresponding to the sweep angle of the c/4

line. Continuing to the axis on the left border, we find the standard value β*req for the

required twist angle. This standard value is valid for a wing, which is trimmed at = 1.0

and has a stability coefficient of β* =10% (see above), and Uses airfoils with a moment

coefficient of zero.

From the standard value we calculate the true, required twist angle, using the formula

inset into the graph. Therefore, we calculate the ratio of our target lift coefficient to the

standard lift coefficient (CL/ ) and the ratio of our desired stability coefficient to the

standard . We see, that a reduction of the lift coefficient to CL=0.5 also reduces the

required twist by 50%. Also, if we use a smaller stability margin β, we need a smaller amount

of twist.

Figure 4 Finding the Required Twist Graphs



2.3.1. Variation of zero lift angle

If we use different airfoils at root and tip, they may have different zero lift directions, which

influences the equilibrium state. The geometric twist has to be reduced by the difference of

the zero lift directions β0 of tip and root sections:

Using the same airfoil for both sections, we can set β0 to zero.

2.3.2. Influence of the Airfoil Moment coefficients

The moment coefficient of the airfoils contributes to the equilibrium, and has to be taken into

account for the calculation of the twist. Graph 2 can be used to find the equivalent twist due to

the contribution of Cm, which has to be subtracted from the required twist. If we use airfoils

with positive moment coefficients, the contribution will be positive, which results in a

reduction of the amount twist, highly cambered airfoils yield negative values βCm, which

force us to build more twist into the wing. Similar to the previous graph, We enter with the

aspect ratio, intersect with the sweep curve and read the value for βCm from the left hand

axis.

Figure 5 Finding the additional twist due to the airfoils moment coefficient.

Again, the graph has been plotted for a certain standard condition, which is a moment

coefficient of cm* = 0.05 (note: positive value). We apply the ratio of the moment

coefficients (cm/cm*) to find the contribution βCm of the moment coefficient to the

geometric twist. This contribution has to be subtracted from the required twist angle, too.

Using the usual, cambered airfoils with negative moment coefficients will change the sign

of the ratio cm/cm*, which results in negative β°Cm values. This means, that the subtraction

from βreq will actually be an addition, increasing the geometric twist angle. If we have

different airfoils at root and tip, we can use the mean moment coefficient (cm,tip+ cm,root)/2

to calculate the ratiocm/cm*.Finally, we can calculate the geometric twist angle βgeo, which

has to be built into the wing:

.

S = (l_r + l_t)/2 * b = 0.5085 m²

And the aspect ratio

AR = b²/S = 11.0

And the mean moment coefficient

cm = (cm,r + cm,t)/2 = 0.02 .



Using graph 1, we find β*req = 11.8°, which has to be corrected to match our design lift

coefficient and the desired stability margin:

11.8 * (0.5/1.0) * (0.05/0.1) = 2.95°

This means that our model would need a twist angle of 2.95° (wash out) from rot to tip, if

we would use a symmetrical airfoil section.

The difference of the zero lift angle of tip and root section is

Now we read the twist contribution of the moment coefficient from graph 2, which is

β*Cm = 5.8°, which has to be corrected for our smaller mean moment coefficient:

5.8 * (0.02/0.05) = 2.32°

Finally, we calculate the geometric twist from

2.95° - 0.8° - 2.32° = -0.17°

The negative value means, that we could use a small amount of wash-in! This is because

we have already enough stability due to the selection of airfoils with reflexes camber lines.

Since the calculated amount is very small, we can use the same angle of incidence for the root

and tip ribs. Since the presented method is not perfect, we can assume accuracy to 1 degree,

which is also a reasonable assumption for the average building skills.

3. OVERALL WING AND AIRFOIL CONFIGURATION

Based on Roskam, conventional configuration is used for the HALE-SUPAV. The UAV will

not have higher range and thus will preliminary fly on land and if necessary on water. Adding

another alternative fuel system would give more rang but will also increase complexity and

weight, which is not recommended.

The geometry of the wing should have negligible sweep because the aircraft will be

operating at low speeds. Sweep will also increase weight and reduce available solar cell area,

both of which will hinder the aircraft’s performance. An initial airfoil selection will be the

Selig 1223, as shown in Figure. This airfoil has12.14% maximum thickness-to-chord ratio at

roughly 20% from the leading edge.

The Selig 1223 airfoil was chosen as the initial configuration because it has all the

characteristics, which requires in solar power high altitude and long endurance airplanes. The

first important one is that it a low Reynolds number will be generated throughout the mission,

and therefore an airfoil that has ideal characteristics at low speed has been chosen. High lift to

drag ratio is also one of the important characteristics. Using XFLR5 software that analyses the

airfoils, a graph of L/D vs. angle of attack was created and is shown in Figure 8. A legend that

is used for the different Reynolds number used is shown in figure 5. As the figures show, not

only does this airfoil have a high lift-to-drag ratio, but also it has a fairly wide operating angle

of attack where the lift-to-drag ratio is optimum. Other airfoil characteristics are shown in

figures.



Figure 6 Selig 1223 Airfoil

4. ANALYSIS

Figure 7 Plot Points

Figure 8 CL/CD Vs ALPHA Figure 9 CL-ALPHA Curve

Figure 10 CM-ALPHA Curve Figure 11 CL/CD Curve



The drag polar will now be calculated for the airplane itself. Using the value of the

calculated zero-lift drag coefficient, the overall aircraft drag coefficient can be calculated

using. Assuming a lift coefficient of 1.2169 from Section 8.3, as well as an Oswald efficient

factor of 0.9, the total aircraft drag coefficient can be found as 0.0231.This will be used to

find lift to drag ratio as shown: XFLR5 program also used to find out lift to drag ratio for this

airplane and the data are shown in appendix F. During the analysis, three different values of

drag coefficient have come out hence the lift to drag ratios.

5. RESULT AND DISCUSSION

The drag polar explains in table 9 shows that lift to drag ratio ranges from 38to 52. The major

difference is in analysis because other two have less difference. The analysis does not include

fuselage drag coefficient as other so that analysis should be most inaccurate among others.

The average value between these two approaches can be used in further analysis if needed.

5.1. Flying wing CG calculator

Figure 12 CG Calculations

Figure 13 Solid works Model Figure 14 Side View with Winglet

6. CONCLUSIONS

The current desire for a greener society, an alternative source of energy for aircraft is needed.

There are many alternative energy solutions that are promising including bio-fuel and

hydrogen fuel cells, but nothing is as limitless as solar technology. As, mentioned throughout

the project, the application of high altitude long endurance UAVs can potentially be very

large, whether it is in weather surveillance, studying natural disaster, or fire direction. The

solar power UAV design discussed weight, has a large wingspan of 280mm, and hold upto

300grams of payload, which is more than enough for all the surveillance and autopilot

instruments. The advances in solar technology have made it so the concept of solar powered

UAVs and MAVs is not just a theory anymore. Solar power airplanes are necessary for

greener society and can be an important part of the future of aviation.

Hence we consider this UAV for future scope of making it a solar powered UAV.



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