Group 5 Report Edit 2

Project Viento

A Report

submitted by

GROUP 5

as part of the course

AS 5220: STRUCTURAL DESIGN

DEPARTMENT OF AEROSPACE ENGINEERING

INDIAN INSTITUTE OF TECHNOLOGY MADRAS,

CHENNAI

27 AUGUST 2014

ACKNOWLEDGEMENTS

Firstly, we would like to thank our Professors Mr. Joel George M and Mr. G Rajesh

for giving us an opportunity to do the course AS5210: Aerodynamic Design. They

provided us with invaluable information about systematic way of designing an air-

craft, which was very helpful in preparing our reports. We also thank our TAs for

giving us valuable guidance on the reports we prepared. All thanks to the references

especially Aircraft Design by Raymer and internet where we got enormous historical

data relating to aircraft that were similar to ours.

i

ABSTRACT OF THE PROJECT

This can be in two paragraphs. First paragraph should start with an introduction

including the motivation behind selecting the mission requirements (as in why did

you choose the desired payload, endurance etc.) followed by a brief description of the

proposed aircraft.

Second paragraph can be on the work done so far in designing and fabrication of

the aircraft. Overall, abstract should be less than or equal to 2 pages.

iii

TABLE OF CONTENTS

ACKNOWLEDGEMENTS i

ABSTRACT OF THE PROJECT iii

LIST OF TABLES vii

LIST OF FIGURES ix

ABBREVIATIONS AND NOTATIONS xi

1 INTRODUCTION 1

1.1 MISSION REQUIREMENTS . . . . . . . . . . . . . . . . . . . . . 1

1.2 CONFIGURATION CHOICE . . . . . . . . . . . . . . . . . . . . . 2

1.3 SUMMARY OF WORK DONE AS PART OF THE AERODYNAMICDESIGN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1.3.1 Weight Estimate [Raymer (1990)] . . . . . . . . . . . . . . . 2

1.3.2 Airfoil and Wing Geometry [Raymer (1990)] . . . . . . . . . 3

1.3.3 Sizing of Fuselage, Tail and Control Surfaces [Raymer (1990)] 3

1.3.4 Three View of plane . . . . . . . . . . . . . . . . . . . . . . 4

1.3.5 Landing Gear Configuration . . . . . . . . . . . . . . . . . . 4

1.3.6 Propeller Selection . . . . . . . . . . . . . . . . . . . . . . . 5

1.3.7 Propeller Performance . . . . . . . . . . . . . . . . . . . . . 5

1.3.8 Center of Gravity . . . . . . . . . . . . . . . . . . . . . . . . 6

1.3.9 Improved Lift Characteristics . . . . . . . . . . . . . . . . . 7

1.3.10 Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.3.11 Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

1.3.12 V-n Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . 8

1.3.13 Performance Evaluation . . . . . . . . . . . . . . . . . . . . 9

1.4 Bill of Materials with suggested Vendors . . . . . . . . . . . . . . . 10

2 Wing Design 11

v

2.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.2 Lift Distribution - Schrenks Method . . . . . . . . . . . . . . . . . 11

2.2.1 Approximate Lift Distribution . . . . . . . . . . . . . . . . . 12

2.2.2 Shear Force Diagram . . . . . . . . . . . . . . . . . . . . . . 13

2.2.3 Bending Moment Diagram . . . . . . . . . . . . . . . . . . . 13

2.3 Preliminary Design . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

2.4 Matlab Code . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

LIST OF TABLES

1.1 Mission Requirements . . . . . . . . . . . . . . . . . . . . . . . . . 1

1.2 Weight of RC plane . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

1.3 wing geometry parameters . . . . . . . . . . . . . . . . . . . . . . . 3

1.4 Control surface parameters and tail parameters . . . . . . . . . . . 3

1.5 dimension and location of wheels . . . . . . . . . . . . . . . . . . . 5

1.6 Propeller parameters . . . . . . . . . . . . . . . . . . . . . . . . . . 6

1.7 CL for different flight segment . . . . . . . . . . . . . . . . . . . . . 7

1.8 CD at different flight stage . . . . . . . . . . . . . . . . . . . . . . . 7

1.9 stability parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

1.10 Performance parameters for different flight segments . . . . . . . . . 9

1.11 Bill of materials. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

vii

LIST OF FIGURES

1.1 Mission profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1.2 Three View of plane . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

1.3 MA 11 83 Blade . . . . . . . . . . . . . . . . . . . . . . . . . . 51.4 Location of center of gravity . . . . . . . . . . . . . . . . . . . . . . 6

1.5 Drag polar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

1.6 V-n diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.1 Lift Distribution on the aircraft wing . . . . . . . . . . . . . . . . . . 12

2.2 Lift Distribution on the aircraft wing . . . . . . . . . . . . . . . . . . 13

2.3 Shear Force Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

2.4 Bending Moment Diagram . . . . . . . . . . . . . . . . . . . . . . . . 14

ix

ABBREVIATIONS AND NOTATIONS

2D Two dimensional3D Three dimensional Angle of attackb Wing spanAR Aspect ratioJ Advance ratioCt Thrust coefficientCp Power coefficientn Propeller efficiencyCs Speed power coefficientTAF Total activity factorCG Center of gravityCL Lift coefficientCD Drag coefficientSM Stability marginhNP Neutral point locationCm Coefficient of momentCm slope of Cm vs curve

xi

CHAPTER 1

INTRODUCTION

Aim of this project is to design a portable UAV which can carry highest payload

possible while simultaneously pursuing the lowest empty weight possible. It can also

transform itself for different mission specification, with an effective modular design.

In aerodynamic design of RC plane we followed many methods to get an idea of

weight, size, dimension and performance etc. of RC plane. We have also done the

stability analysis and calculated the parameters affecting the performance of plane.

1.1 MISSION REQUIREMENTS

Motivation for making this plane was to compete in SAE(Society of Automotive

Engineers) competition, but it can also serve many other purposes. As we started

designing this plane we set some mission requirement for our plane which are given

below:

Table 1.1: Mission Requirements

Cruise altitude 30 mCruise velocity 35 kmphCeiling 50 mRange 120 mLanding Distance 50 m

And mission profile for our plane:

Figure 1.1: Mission profile

1.2 CONFIGURATION CHOICE

We have chosen a biplane configuration for wing design and T-tail configuration for

tail design.Following are the salient features of the configuration considered:

Biplane configuration was chosen because it produces more lift compare to

monoplane so we can reduce our wingspan.It is also provide a strong structure

and it is easy to manufacture.

T-tail configuration for tail design is chosen because it allows clean airflow for

better stability at low speed. It is also cause for high gliding performance.

1.3 SUMMARY OF WORK DONE AS PART OF THE AERODYNAMIC

DESIGN

1.3.1 Weight Estimate [Raymer (1990)]

In first weight estimation we used the historical data of RC plane with similar mission

requirement and profile to get weight of our RC plane. In second weight estimation

we used powerplant weight to calculate weight. Result of first and second weight

estimation are given below.

2

Table 1.2: Weight of RC plane

first weight estimation 1489 g

second weight estimation 1586.19 g

1.3.2 Airfoil and Wing Geometry [Raymer (1990)]

We selected S1223 airfoil based on our mission requirement after comparison between

many airfoils data. From the historical data of biplane we have chosen no stagger

(stagger matters for pilot visibility and ours is UAV)for our plane and we have fixed

the following parameters of wing geometry for our biplane from historical data:

Taper ratio 0.45Aspect ratio 6

Gap ratio 1Dihedral angle 2o

Table 1.3: wing geometry parameters

1.3.3 Sizing of Fuselage, Tail and Control Surfaces [Raymer (1990)]

In sizing process we have calculated the dimension of fuselage, tail and control surfaces

for our design and requirement. Which is given below:

Fuselage length: 659 mm

component Chord AreaAileron 28.6 mm 48.72 cm2

Elevator 20.67 mm 48.9 cm2

Rudder 19 mm 21.49 cm2

Control surface parameters

Horizontal tail Vertical tailLHT 39.58 cm LV T 39.58 cmSHT 173 cm

2 SV T 79 cm2

AR 4 AR 1.5CrHT 9.07 cm CrV T 8.67 cmCtHT 4.08 cm CtV T 3.9 cmbHT 26.3 cm hV T 12.57 cm

Tail parameters

Table 1.4: Control surface parameters and tail parameters

3

1.3.4 Three View of plane

Three view of plane has been drawn using autocad software after having complete

idea of dimension of our plane.

Figure 1.2: Three View of plane

1.3.5 Landing Gear Configuration

Tail dragger configuration has been selected for landing gear configuration because it

has many advantage and disadvantage doesnt matter for RC plane.

We have followed the conventional tail dagger configuration and have calculated

the diameter of wheels. We have also calculated the location of landing gear wheels

from nose and the strut length of landing gear which is given below:

strut length: 6.60 cm

4

wheels Diameter Width Location (from nose)Main wheel 4.56 cm 0.31 cm 21.55 cmTail wheel 3.8 cm 0.23 cm 60.90 cm

Table 1.5: dimension and location of wheels

1.3.6 Propeller Selection

We have calculated the diameter of three bladed propeller from empirical relation,

then we have calculated pitch and we have selected a 3-bladed 11 x 8 propeller which

has approximately same characteristics.

A fuselage-mounted tractor configuration is chosen, where the propeller is in front

of the attachment point. This configuration is allowed for smaller tail area and more

stability.

Figure 1.3: MA 11 83 Blade

1.3.7 Propeller Performance

Performance parameters has been calculated to find the efficient operating conditions

for propeller. We have used empirical relation and power efficiency vs advance ratio

graph to calculate these parameters which has been tabulated below:

5

Parameters symbols valuesAdvance ratio J 0.55 < J < 0.65

Thrust coefficient Ct 0.148Power coefficient Cp 0.1023

Propeller efficiency n 0.8Speed power coefficient Cs 0.868

Total activity factor TAF 468.51

Table 1.6: Propeller parameters

1.3.8 Center of Gravity

Center of gravity is very important for performance and stability analysis. We have

found it using component method. In component method we calculate CG of different

components of airplane before calculating the CG of plane.

The center of gravity of RC plane was found to be at 0.3015 meters from the nose

and at a height of 0.0014 meters from the nose.

Figure 1.4: Location of center of gravity

6

1.3.9 Improved Lift Characteristics

In this we have calculated the lift coefficient CL at different flight segment using basic

lift and weight relationship for different segment. We have also calculated CLmax and

Lmax using empirical relations. CL for different segment is given below:

Take off 1.706Climb 1.275Cruise 0.998

Landing 1.489

Table 1.7: CL for different flight segment

Also CLmax and Lmax are 1.845, 8.68o respectively.

1.3.10 Drag

Parasite drag, Lift-dependent drag, Skin friction drag, Flat-plate skin friction coeffi-

cient and component form factors, miscellaneous drag coefficient, leakage and protu-

berance drag were considered for the drag coefficient calculation.

There are two method to calculate parasite drag Cd0: Equivalent skin friction

method and Component build up method. By following these method we got Cd0

= 0.04311. Then we have found the lift dependent drag. From parasite and lift

dependent drag we have plotted the drag polar diagram and have calculated CD at

different flight segment.

Stage CDTake off 0.1790Climb 0.1198Cruise 0.0989

Landing 0.1071

Table 1.8: CD at different flight stage

7

Figure 1.5: Drag polar

1.3.11 Stability

Stability and performance analysis is very important to make a plane fly. In this

we have discuss the parameters which affect stability by affecting neutral point and

stability margin. We have also calculated those point location by using moment

equilibrium and aerodynamics relations.

hNP 2.9916hCG 2.7585SM 0.2331Cm -1.3892

Table 1.9: stability parameters

1.3.12 V-n Diagram

V-n diagram is a type of flight envelope which limits the manoeuvre boundaries for

a given aircraft. This envelope demonstrate the variations of airspeed versus load8

factor.

Figure 1.6: V-n diagram

1.3.13 Performance Evaluation

We have calculated the performance parameter for different flight segment which has

been given below:

Table 1.10: Performance parameters for different flight segments

CruiseVcruise 13.9 m/sT/W 0.1284W/S 118.085 N/m2

Velocity for minimum thrust 15.673 m/sMinimum thrust 1.939 N

Velocity for minimum power 11.91 m/sMinimum power 26.67 WMaximum range 29.45 km(if in automated UAV mode)Maximum power 35.7 min(depending on battery)

ClimbingBest climb angle 13.56o

Velocity 11.74 m/sDrag 2.1749 N

Rate of climb 2.754 m/s

GlidingVelocity for minimum sink rate 11.91 m/s

Minimum sink rate 3.845 m/s

9

LandingGround roll 60.989 m

Flare distance 27.741 mApproach distance 20.987 m

Total landing distance 109.717 m

Level turnMaximum turn rate 1.539 rad/sTurn rate for loiter 1.408 rad/s

1.4 BILL OF MATERIALS WITH SUGGESTED VENDORS

Table 1.11 gives the details of the materials required for fabrication and their price.

Table 1.11: Bill of materials.

Components Price(Rs.)Motor 1000

Battery 5000Balsa wood 5000Aluminium 1000

ESC(electronic speed controller) 1000Transmitter and receiver 16500

Servo motors 4800Propellers 500

Miscellaneous 2000Total 20,300(without transmitter and receiver)

10

CHAPTER 2

Wing Design

2.1 INTRODUCTION

In this chapter, we try to find an approximate solution to the load distribution on

the wing. Using the load distribution, we calculate the very important Shear Force

Diagram (SFD) and the Bending Moment Diagrams (BMD).

We end with a conceptual design of the wing structure based on historical choices

made in wing construction.

2.2 LIFT DISTRIBUTION - SCHRENKS METHOD

To design any structure, it is important to find the loads which act on the body. In

case of the wings of an aircraft, the important loads are the weight of the aircraft and

the lift distribution on the wings.

Unlike infinite wings, finite wings have an uneven lift distribution. Delving into

some basic aerodynamics of finite wings, we can find that the lift distribution on

an elliptical wing planform is also elliptical. But, as is our case, trapezoidal (and

rectangular) wings do not follow this method.

To overcome this problem, we use the method suggested by O.Schrenk in his

NACA paper, which provides a way to calculate the approximate Lift distribution on

the wings. The exact method can be found in the paper and various other resources

(which have been referenced in the references [Schrenk (1940)]).

This method is commonly used to determine overall span-wise lift distribution,

especially at the preliminary design stage for low sweep and moderate to high aspect

ratio wings on FW aircraft. The method states that the resultant load distribution is

an arithmetic mean of:

1. A load distribution representing the actual planform shape

2. An elliptical distribution of the same span and area

An elliptical distribution is presented in figure below, here the semi-span wing area

= area of Elliptic quadrant = S/2.

Area :S

2=

1

4[(pi

2)(2a)(b)]

a =4S

pib

For an ellipse:y2

( b2)2

+c2ya2

= 1

cy = (4S

b)

(1 (2yb

)2)

To convert into a load distribution, we put Wy (N/m) in place of Cy and put L (N)

in place of S.

wy = (4L

b)

(1 (2yb

)2)

Figure 2.1: Lift Distribution on the aircraft wing

2.2.1 Approximate Lift Distribution

The following diagram shows the approximate lift distribution on the aircraft wing

which was found using Schrenks method:

12

Figure 2.2: Lift Distribution on the aircraft wing

2.2.2 Shear Force Diagram

The following diagram shows the Shear Force Diagram found using lift distribution

obtained by Schrenks method:

Figure 2.3: Shear Force Diagram

2.2.3 Bending Moment Diagram

The following diagram shows the Bending Diagram found using lift distribution ob-

tained by Schrenks method:13

Figure 2.4: Bending Moment Diagram

2.3 PRELIMINARY DESIGN

The following characteristics of the wing have been provisionally decided as suggested:

1. Type of Wing - Our plane is a biplane and the wings will be supported using

struts between the wings. Therefore it is a kind of braced cantilever wing.

2. Covering - Since we are trying to build a light plane, we have decided on our

wings to be fabric covered.

3. Skin type - Again due to considerations of weight, it is better to choose the

Stressed Skin construction technique.

4. Assembly - We would need more information to decide on this. But considering

the size of the plane, it is better to have lesser sections. Hence we think that a

2-peice wing assembly would be better

5. Root joint - Both wings have a dihedral angle. This makes a continuous

attachment difficult. Hence a wing box will be necessary to attach the wings

to the fuselage.

14

2.4 MATLAB CODE

This is the Matlab code used to generate the results shown above:

clc;

clear all;

%Schrenk's Method

%Wing Dimensions (in mm)

c root = 143.5;

c tip = 64.5;

b = 628;

%Maximum Loading Case

n=2.16;

V=15.126;

W=1.6*9.8;

%Wing profile

x=0:0.0001:(b/2);

y=-((0.1257961783439490445859872611465)*x - (c root/2));

hold on

plot(x,y);

%Wing Area and Equivalent Ellipse

area=0.5*(c root+c tip)*(b/2);

S=2*area*((10(-3))2);

a=2*area/(pi*(b/2)); %a for ellipse

z=(2*a/b)*sqrt(((b/2)2)-(x.2));%ellipse

c ell = a*sqrt(1-((2*y/b).2));

mean = (y+z)./2; %schrenk's method

plot(x,z,'g');%plotting Ellipse

plot(x,mean,'r');%plotting Schrenk's representative lift distribution

hold off

C L = n*W/(0.5*1.225*V*V*S);

L=n*W;

w y = (2*L)/(pi*b)*sqrt(1-((2*x/b).2)); %Actual Lift Distribution

15

SF total = trapz(x,w y); %Reaction at Root

%Shear Force Diagram

i=1;

for span = 0.1:0.1:b/2

store=0:0.1:span;

w y2 = (2*L)/(pi*b)*sqrt(1-((2*store/b).2));

SF x(i)=SF total - trapz(store,w y2);

i=i+1;

end

span=0.1:0.1:b/2;

figure

plot(span,SF x);%Shear Force Diagram

%Bending Moment Diagram

i=1;

j=1;

dx=0.001;

for span = 0.0:0.1:b/2

b y=0;

for store=b/2:-0.001:span

b y = b y + (dx/1000)*(2*L)/(pi*b)*sqrt(1-((2*store/b).2))*(store-span);

end

BM x(i)=(b y);

i=i+1;

end

span=0.0:0.1:b/2;

figure

plot(span,BM x);%Bending Moment Diagram

16

REFERENCES

1. Raymer, X. Y., Design. ABC Press, 1990, 1 edition.

2. Schrenk, O., A simple approximation method to obtain the spanwise lift distribution.NACA, 1940.

17

Documents

Group 5 Report Edit 2