Passenger Aircraft Design

1

Chapter 1

INTRODUCTION

This chapter provides an overview of the low cost town to town passenger aircraft

and its design techniques. The major objectives of the whole project and the project

schedule are also being described in this chapter.

1.1 INTRODUCTION:

This report will detail the conceptual design and sizing of a low cost town to town

aircraft to fulfill the civil aviation requirements of India and recommended ways and

means of establishing a viable civil aviation industry, based on the feasibility study

report submitted by NAL to Research council in November 1990 which is eventually

lead to the design and development of NAL Saras. The primary goal of this aircraft is

to replace the NAL Saras which is under development now, with a cheaper, higher

efficient aircraft. In addition to replacement of the Saras, the role of this new air

vehicle will be to bridge the gap between small business jets and large commercial

aircrafts, hence the name LTA.

To facilitate the conceptual design and analysis of this aircraft a spreadsheet MIT

Aero Design tool was created using an energy based approach for aircraft design.

This method outlined in AD class. Mattingly and Roskam are the basis of this report.

The reader of this report, armed with the MITAD tool, Mattingly, and Roskam,

should be capable of repeating the conceptual design of this LTA Spatz vehicle

outlined in this report or any other fixed wing aircraft.

The initially found design solution LTA named “Spatz” is compared to the current

NAL Saras.

2

Figure 1: NAL Saras 2-View( Google images)

Figure 2: LTA“Spatz” 3-view Sketch Using Catia v5r20

*Note about LTA“Spatz” drawings: All “Spatz” drawings were made using CATIA

V5R20.

3

1.2 MOTIVATION:

India aims to start making its own commercial aircraft in a bid to cash on a boom

in the domestic civil aviation sector.

There is potential for short-distance, low-cost carriers as operators are looking for

cost-effective and right-sized aircraft, while passengers want lower air fares.

Europe's Airbus expects India will need more than 1,000 aircraft over the next 20

years at a cost of $138 billion.

India's passenger numbers will expand by 15 percent annually over the next five

years, making it the fastest-growing market in the world.

The expansion of India's middle class, spurred by the country's growing economy,

has fuelled air travel.

1.3. OBJECTIVES:

The objectives of the whole project are:

1. Conceptual design and sizing of a low cost town to town aircraft to fulfill the

civil aviation requirements of India and recommended ways and means for

establishing a viable civil aviation industry.

2. The primary goal of this aircraft is to replace the NAL Saras which is under

development now, with a cheaper, higher efficient aircraft

3. To facilitate the conceptual design and analysis of this aircraft a spreadsheet

tool is to be created using Roskam approach for aircraft design.

1.4. TARGET SPECIFICATIONS:

The main target of the project is to design a low cost town to town passenger aircraft

which can replace the NAL Saras which is under development now, with a cheaper,

higher efficient aircraft.

1.5. PROJECT WORK SCHEDULE:

Month Activities

February 2012 Preparing synopsis and Literature review

February and March,

2012

Conceptual Design of the aircraft

March,2012 Midterm project evaluation and presentation.

March to April,2012 Preliminary design of the aircraft.

May,2012 Thesis writing and submission

Table: 1.5 project schedule

4

1.6. ORGANISATION OF PROJECT WORK:

The whole thesis is divided into five chapters. The first chapter gives

Introduction to an overview of the low cost town to town passenger aircraft and its

design techniques. The major objectives of the whole project and the project schedule

are also being described in this chapter. It is followed by the important objectives of

the project followed by the target specifications of this thesis work.

The second chapter consists of the spreadsheet MIT-AD tool, in which the

basis of the report is written on, is meant to be a stand-alone tool for any initial fixed

wing design application. For this reason all references, tables, and figures used in

sizing the LTA Spatz is located in the MIT-AD tool.

The third chapter consists of the Methodology of the thesis work. We design

low cost town to town passenger aircraft. The fourth chapter consists of the important

results associated with the validation.

The last chapter consists of the conclusion and future scope of the work.

5

Chapter 2

BACKGROUND THEORY

2.1 MIT Aero Design (MIT-AD) Tool Description

The accompanying spreadsheet MIT-AD tool, in which the basis of the report is

written on, is meant to be a stand-alone tool for any initial fixed wing design

application. For this reason all references, tables, and figures used in sizing the LTA

Spatz is located in the MIT-AD tool.

2.2 Layout and Features

2.2.1 Layout

The MIT-AD tool used for this project was created in Microsoft Excel. Key functions

are split into different „tabbed‟ worksheets and labeled by function. Tabs include:

1) Weight Estimation – in this tab following calculations are included

a. Fuel Fractions – Fuel fractions for various „types‟ of aircraft are listed

here from Roskam text.

b. Weight Empty Data – Historical data used for initial TOGW and WE

for a variety of aircraft „types‟ can be found in this tab.

c. Mission Profile – This tab is just a place holder for inserting the

mission profile for reference. For the LTA“Spatz”, the mission profile

chosen is shown in this tab

d. Mission Analysis – Based on user inputs, this tab is where the

preliminary sizing of the aircraft‟s take-off weight and empty weight

are calculated. Figure3 gives a screen shot of the mission analysis tab.

After the initial inputs are entered, the user simply clicks the „Final

Weight Empty‟ button, and an Excel macro will vary take-off weight

until empty weight calculated equals empty weight historical (or

geometrically based).

e. Mission Summary – The mission summary collects all the input and

output in one clear concise table, giving the user the ability to compare

all segments of the mission.

2) Sizing Analysis – Aircraft performance requirements are calculated in this tab

using the Mattingly “Master Equation” Examples of performance

6

requirements calculated for the LTA “Spatz”: max speed, service ceiling, and

take-off distance.

2.2.2 Inputs

All MIT-AD tool inputs are labeled with references on where the current example

Spatz variables can be found. Some of these reference tables and figures are shown in

this report or in the appropriately labeled tabs in the MIT-AD tool.

2.2.3 Outputs

Primary outputs from the MIT-AD tool are shown in green colour. The calculated

empty weight is shown for the given mission requirements. The user pushes the „Find

Empty Weight‟ Button to calculate the empty weight. In addition to mission analysis,

wing and engine sizing are doing with constraint analysis found in the „constraint

analysis‟ tab. The user can input performance requirements.

7

Chapter 3

METHODOLOGY

3.1. INTRODUCTION:

The MIT-AD tool uses an energy balance approach of kinetic and potential energy

to size aircraft, which is thoroughly described in Mattingly, Raymer, and Roskam.

This report will only review the basics of the method, and the remainder of the report

will detail how the MIT-AD tool was used to design the LTA “Spatz” vehicle.

Roskam outlines seven steps to sizing a vehicle. These steps are labeled in the

MIT-AD tool under the „mission analysis‟ tab shown in figure 3:

1) Determine payload based on mission requirements.

2) Guess initial take-off weight from historical data trends.

3) Use fuel fractions to estimate weight of fuel burned during the mission.

4) Calculate operational weight empty.

5) Calculate weight empty.

6) Use historical trends to determine the weight empty for the initial take-off

gross weight guess, figure 5. Note: That this step can be replaced by using

actual weight estimates based on the geometry or individual weight trends for

components. This will tend to give a more representative answer.

7) Change initial take-off weight guess and repeat steps 3 thru 6 until the weight

empty calculated matches the weight empty from historical data

The purpose of this chapter is to illustrate the process of design of a low cost town

to town passenger aircraft. The design process is started after literature survey. I

divided the process into following pivot points.

Requirements

Weight Estimation

Sizing

Airfoil selection

Wing design

Fuselage design

CG Location calculations

Tail design

Landing gear design

8

Table 1: In the MIT-AD tool the „weight estimation‟ Tab is laid out in order of the

seven steps to Aircraft Sizing from Roskam

3.2 Requirements

After literature survey we fixed the requirements as given below

Fixed Configuration:

19 Seater

Wing mounted Twin propeller

Conventional tail configuration

9

Tricycle Landing gear

High , Tapered wing

Crew = 2 Pilots

2 abreast Cabin

Full composite Fuselage

Short haul flight

3.3 Weight Estimation

The second pivot point in our conceptual design analysis is the preliminary estimation

of the gross weight of the airplane. From this step onwards every designing process

are done in MIT AD tool,

There are various ways to subdivide and categorize the weight components of an

airplane. The following was our choice.

Crew weight: The crew comprises the people necessary to operate the airplane

in flight. For our airplane, the crew is simply the 2 pilots.

Payload weight: The payload weight is what the airplane is intended to

transport, in our project its 19 passengers & their baggage.

Fuel weight: This is the weight of fuel in the fuel tank. Since fuel is consumed

during flight, this is a variable, decreasing with time during flight.

Empty weight: This weight of everything else- the structure, engines,

electronics equipment, landing gear, fixed equipment and anything else that is

not crew, payload, or fuel

The whole weight estimation includes

Mission profile & Payload

Guessing an Initial Take-Off Gross Weight from Historical Data

Mission Analysis using Fuel Weight Fractions & Empty weight calculation

3.4 Mission Profile and Payload

First a mission profile is defined by some set of requirements; in this case it is a LTA

Spatz to replace NAL Saras aircraft. This includes Engine start and warm up, taxi,

take-off, climb, cruise, decent and landing. Figure shows the mission profile used in

this design study. For Payload calculations we had to input the number of passengers,

10

number of crew, allowable baggage weight and average weight of a passenger. From

these values the total payload is calculated using the equation given below.

( ) ( ) (1)

Figure 3: LTA Spatz Mission Profile to be Used for Mission Analysis

Payload is usually accompanied with a mission profile such as the one given for the

LTA mission. Payloads for this mission range are passengers and cargo of 4200 lbs;

therefore, the most limiting design constraint of 4200 lbs is used. In addition to the

weight of the payload the drag of the aircraft will also go up due to the additional

excrescences. Table 2 outlines the Passenger weight, baggage and crew weight.

Table 2: Payload/Crew for Lta Mission

11

3.5 Guessing an Initial Take-Off Gross Weight from Historical Data

To get an order of magnitude for what a new aircraft should look like and weight, it is

common to begin with historical data. First, a class of aircraft with similar mission

requirements and/or capabilities is selected. For purpose of this study, the Let L 410,

NAL Saras, CASA Aviocar all fall into the „LTA‟ class of aircraft. Figure 2.8 [1]

.

Shows an historical regression of takeoff gross weight (TOGW) to weight empty

(WE). This plot will later be used to correct the initial TOGW guess with the WE that

is acquired from the mission analysis portion of the MIT-AD tool.

Table 3: In MIT AD TOOL Weight take off guess based on historical data

3.6 Mission Analysis using Fuel Weight Fractions

The end goal for running mission analysis is simply to find the total fuel

burned during the mission. To accomplish this, the mission is broken into small

manageable segments, in which fuel fractions can be calculated or looked up. For

short relatively simple segments such as taxi, take-off, climb, descent and landing,

table 3, found in Roskam can be used for a variety of aircraft.

The LTA “Spatz” is classified in the Regional TBP (Turbo Prop.) airplane

type in table3. The fuel fractions used for “Spatz” are highlighted by the yellow box.

These fuel fractions are multiplied together and then multiplied by the TOGW

to determine the total fuel burned for the given mission.

For cruise mission, the range needs to be factored into the fuel fraction. To

determine the fuel burned during a cruise, the Breguet Range equation is re-arranged

into the form shown in equation 1from Roskam. Values for Cp can be found from

table 2.8 [1]

. Table 4 can also be used to find typical values of L/D; however, if a drag

polar is assumed the user can calculate L/D based on instantaneous mission weight.

( ⁄ ) ( ⁄ ) ( ⁄ ) (2)

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3.7 Empty Weight Calculation

In this step we are calculating operational and empty weight. Table 6 shows the

equations used and the values for the LTA “Spatz”. The difference between

operational and empty weight are the mission specific equipment that goes into the

aircraft. For the LTA “Spatz” all mission equipment is assumed to be in the weight

empty. Final weight empty for the “LTA Spatz” is calculated to be 7500 lbs

Table 4: Operational and Empty Weight for the LTA “Spatz”

3.8 Allowable Empty Weight from Historical Data

The allowable empty weight based on historical data can be found one of two ways:

1) For a more representative weight that captures the individual weight break

down for all major aircraft components, the user can choose to gather

historical TOGW trends by component. Many of these trend equations can be

found in Raymer , or Roskam

2) For the most accurate method, the user can have a preliminary designer mock

up each component, and have a weights engineer give estimates based on

wetted area, structural thickness, and volumes, for each component. This

however, is time consuming and manpower intensive; therefore, is usually

saved for the preliminary design phase, as opposed to the conceptual design

phase.

13

The initial take off gross weight guess is adjusted for several iterations before the

empty weight shown in tables were found. Based on a TOGW of 14111 lbs the “LTA

Spatz” weight empty is 7500 lbs

Table 5: Final Empty weight for the LTA “Spatz”

3.9 Sizing

3.9.1 Stall Speed Sizing

A propeller driven airplane must have a power-off stall speed of no more than 50 kts

at sea level with flaps full down (i.e. landing flaps). With flaps up the stall speed is to

be less than 60 kts. Both requirements are to be met at take-off gross weight, Wto.

From Table 3.1[1]

it is seen that the following maximum lift coefficient value are

written within the state of art

CLmax= 1.60 and CLmax L= 2.00 (3)

VS = (2(W/S) / ΡCLMAX).5

(4)

With the help of equation 3.1 it now follows that:

To meet the flap down requirements:

(W/S)TO < 17.0psf

To meet the flap up requirements:

(W/S)TO < 19.5psf

Therefore to meet both the requirements (W/S)TO < 17.0

Figure3.1 from Reference [1]

illustrates this. Because the stall speed requirements were

formulated as a power-off requirement, neither power loading nor thrust loading are

important in this.

14

Table 6: IN MIT AD TOOL Sizing to stall speed requirements data

3.9.2 Sizing to Take-off distance

Take-off distances of aircrafts are determined by following factors

• Take-off weight

• Take-off speed

• Weight to power ratio

• Propeller characteristics

• Aerodynamic drag coefficient

• Ground friction coefficient

• Pilot technique

Value of air density ratio σ = 0.8616 at 5000ft

For civil aircrafts requirements of FAR 23 must be satisfied

– Take-off distance must be less than 1500ft

– Take-off distance = 1.66 x Take-off ground run

– STO = 1.66 x STOG

– Take-off Parameter for FAR 23 , TOP23

– Its dimension is lbs2/ft

2hp

STO = (8.134 x TOP23) + 0.01494 x (TOP23)2 (5)

15

Input the value of air density ratio and value of take-off ground run; value of STOG

should satisfy FAR 23 conditions. Then solve the quadratic equation manually and

input in to the MIT AD tool after that get the value for (W/S)*(W/P)/ (CLMAX) and

compare it with the take-off facility of the airport, Hence make sure that the take-off

distance requirements are met. To calculate weight to power ratio, we use the

following equation

⁄ (6)

To calculate total takeoff power PTO we use following equation

⁄ (7)

Table 7: IN MIT AD TOOL Sizing to take off requirements data

3.9.3Sizing to Landing distance

Landing distance of an airplane is determined by four factors:

Landing weight, WL

Approach speed, VA

Deceleration method used

Flying qualities of the airplane

Pilot technique

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Landing distance requirements are nearly always formulated at the design landing

weight, WL of an airplane. Table shows how WL is related to WTO for different types

of airplanes.

Kinetic energy considerations suggest that the approach speed should have a

„square‟ effect on the total landing distance. After an airplane has touched down, the

following deceleration methods can be used. Brakes, Thrust reversers, Parachutes,

Arresting systems.

For landing sizing we do following calculations for

Approach speed, VA which is equals to 1.3 times stall speed,

(8)

Landing ground run, SLG which is equals to 0.265 Vstall 2

(9)

Total Landing distance, SL which is equals to 1.938 times SLG

Figure 4: Definition of FAR 23 Landing Distance (from Airplane Design by Roskam

page 109)

User interface of MIT AD tool for the above calculations shown below

Table 8: IN MIT AD TOOL Sizing to landing distance requirements data

17

3.9.4 Sizing to climb requirements

For climb requirements we went through table 3.5 [1]

which describes

regression line coefficients C & D for different types of aircrafts. From these

coefficients we calculated the value of wetted wing area, using following equation.

(10)

The value of skin friction coefficient is chosen as 0.003 which is reasonable

for new generation aircrafts. From this value and using table 3.4 [1]

I found the values

of coefficients A & B. from these we calculated the value of equivalent parasite area,

using following equation

(11)

Then for climb requirements we need to input the value of Aspect ratio,

⁄ (12)

From this we calculated the value of Oswald efficiency factor, which is given by the

equation ( ) (13)

Next we found the value of maximum wing loading using the following equation

⁄

( )

(14)

From the values of wing loading and total takeoff weight we calculated the value of

wing area, S using

⁄ (15)

From the values of equivalent parasite area and wing area we calculated the value of

zero lift drag coefficient, using

(16)

From the values of Oswald efficiency factor, maximum lift coefficient, Aspect ratio

and CDO we calculated the value of total drag coefficient CD using following equation

(17)

18


Table 9: IN MIT AD TOOL Sizing to climb requirements data

3.10 Airfoil selection

In this step we select the airfoil for the main wing. The process starts with the

values of cruise speed and coefficient of lift provided by high lift device, which are

chosen as 189kts and 0.8 respectively.

Next step is the calculation of average weight during flight, which id given by the

equation given below

[ ( )] (18)

From the calculated value of average weight during cruise we calculated the value of

cruise lift coefficient using the following equation

(19)

Then we calculated the value of wing cruise lift coefficient using the following

equation

(20)

From the calculated value of wing cruise lift coefficient we calculated the value of

airfoil ideal lift coefficient using the following equation

19

⁄ (21)

Next is the calculation of maximum lift coefficient, which is given by the equation

(22)

Then we calculated the value of wing max. lift coefficient using the following

equation

(23)

From the calculated value of wing max. lift coefficient we calculated the value of

airfoil gross maximum lift coefficient using the following equation

⁄ (24)

Next step is to calculate the wing airfoil net maximum lift coefficient, which is given

as

(25)

From the calculated values of airfoil ideal lift coefficient and airfoil net

maximum lift coefficient I referred figure 5.23 [17]

and found out the matching airfoil

section.

20


Table 10: IN MIT AD TOOL Airfoil selection data

21

3.11 Wing design

The wing we selected was tapered wing. The process is divided in to 4 subdivisions

Wing geometry design

Flap & Aileron parameters

Fuel tank parameters

3.11.1 Wing geometry design

For the design of this we already have the values of required wing area and

aspect ratio, from these values we calculated the value of wing span, using the

following equation

√ (26)

We also required inputting the value of taper ratio. To get the geometry of the wing

following calculations are done.

For Root chord length,

( ) (27)

For Tip chord length,

(28)

For Span wise Location of mean aerodynamic chord,

[

] (29)

For Mean aerodynamic chord length,

[

] (30)

After above calculations we got the following figure of wing

Figure 5: Mean Aerodynamic chord and Aerodynamic center in a straight wing

(FROM Wing Design by Mohammed Sadraey page 48)

22

3.11.2 Flap parameters

In this section we calculated the parameters for high lift device. In our design

we selected split flap as the HLD device. To calculate the values of flap span and flap

chord we had to input the values of takeoff speed, lift coefficient during takeoff, wing

span, mean aerodynamic chord length, flap chord in percentage of Wing chord and

flap span in percentage of Wing span. They are given by the equations given below

Tentative value of flap chord,

Tentative value of flap span,

Top view of flap preliminary design is given below

Figure 6: High Lift Device Parameters (from Wing Design by Mohammed Sadraey

page 79)

3.11.3 Aileron Parameters

In this section we use the values of Aileron chord in percentage of Wing chord

and Aileron span in percentage of Wing span. Which are selected from historical

values, in our design the selected values were 20% for both cases. From these values

tentative values of Aileron span and chord were calculated.

Tentative value of flap chord,

Tentative value of flap span,

23

3.11.4 Fuel tank parameters

We decided to place the fuel tank inside the wing. So it is necessary to

calculate the dimensions of fuel tank, for that we needed the values of fuel weight,

specific weight of aviation fuel, volume occupied by 1gallon. From these values

following calculations were done.

Total Fuel needed for mission

(30)

Tank capacity

(31)

Tank capacity on each wing

(32)

So we had to do an iterative effort to calculate fuel tank dimensions so that the

calculated value should be greater than tank capacity on each wing. Top view of fuel

tank placement is given below

Figure 7: Fuel Tank Inclination in the Wing (from Aircraft performance and Design

by Anderson page 432)

24


25

Table 11: IN MIT AD TOOL Wing design, Flap parameters, and Fuel tank

parameters selection data

26

3.12 Fuselage design

For fuselage design calculations following parameters are input into the MIT AD

tool, they are

Number of Passengers & Crew

Number of abreast

Length, width, height of 1 seat

Foot space

Number of Aisle & its width

Cargo compartment height

Thickness of floor & wall

And from above values we found out the values of fuselage length, height, width etc.

the user interface is given below

Table 12: IN MIT AD TOOL Fuselage design parameters selection data

3.13 CG Location calculations

For center of gravity location we need to have the values of weights of

passengers, baggage, and crew. Also approximate distance to their CG from the nose.

All these values we already knew from weight estimation and fuselage design parts.

27

So in MIT AD tool we didn‟t have to input anything. From these values we calculated

the location of CG without wing using following equation.

( ) ( ) ( )

(33)

Then we assumed that the aerodynamic center of the wing & CG of the fuel tank are

placed at above calculated CG, and then we calculated the average weight of the

wing, which is twice of plan form area. Also found out the location of wing ADC &

CG, which are 20% and 40% of the root chord. From these values we calculated the

CG of the aircraft with wing using below equation.

( ) ( ) ( ) ( ) ( ( ( ))

(34)

Figure 8: CG Locations (from Google images)

28

User interface of CG location calculation is given below

Table 13: IN MIT AD TOOL Center of gravity location parameters

29

3.14 Tail design

Tail design is divided into two. Horizontal tail & vertical tail design. In

horizontal tail design section horizontal tail volume ratio, Moment arm from CG to

ADC of Horizontal tail, Aspect ratio & taper ratio of Horizontal tail are input into the

MIT AD tool, the plan form area of the horizontal tail is calculated using below

equation.

(35)

After getting the values of area, aspect ratio and taper ratio, the geometry of

the horizontal tail is calculated same as the main wing.

For horizontal tail span,

√ (36)


( ) (37)


(38)


[

] (39)


[

] (40)

In vertical tail design section vertical tail volume ratio, Moment arm from CG

to ADC of vertical tail, Aspect ratio & taper ratio of vertical tail are input into the

MIT AD tool, the plan form area of the vertical tail is calculated using below

equation.

(41)

After getting the values of area, aspect ratio and taper ratio, the geometry of the

vertical tail is calculated.

30

Figure 9: Plan view Fuselage and Horizontal tail (from Aircraft performance and

Design by Anderson page 439)

For vertical tail span,

√ (42)


( ) (43)


(44)


[

] (45)


[

] (46)

These are the calculations involved in the tail design and after getting all

results we can plot the horizontal tail and vertical tail. Design of horizontal control

surface, Elevator and vertical control surface, Rudder are not included in this this

project since those procedures are come under detailed design.

31

Figure 10: side view Fuselage and Vertical tail (from Aircraft performance and

Design by Anderson page 439)

User interface of tail design is given below

Table 14: IN MIT AD TOOL Horizontal tail design parameters data

32

Table 15: IN MIT AD TOOL Vertical tail design parameters data

3.15 Landing Gear Design

For Landing gear design we had to input the values of static margin and

location of neutral point. The value of static margin is typically 5 to 10% and location

of neutral point is given by the equation below.

( ) (47)

And we assumed that Aerodynamic center of wing-fuselage combination is the same

as the Aerodynamic center of the wing, that is ( ) and Lift slope of

the tail and that for the whole airplane are essentially the same, that is

For structural & space reasons, we locate the main landing gear at the center

of the wing, that is . Locate the nose wheel so that it can be

conveniently folded rearward & upward into the fuselage so the location of nose

landing gear was chosen as 3ft. which is typical value for the aircrafts of this

particular class.

To calculate the diameters and widths of main and nose wheels we have done

the following calculations. The force diagram is given below

33

Figure 11: Force Diagrams for obtaining the load distribution among the tires (from

Aircraft performance and Design by Anderson page 446)

For force on main landing gear,

(48)

For force on nose landing gear

(49)

At this stage the number of nose wheels and number of main wheels are input

into the MIT AD tool, so we got the values of load on each tire. After getting those

values we calculated the diameter and width using below equations.

For diameter of the tire,

(50)

For width of the tire,

(51)

34

Chapter 4

RESULT ANALYSIS

4.1 VALIDATION

For validation of the project we studied 3 other aircrafts of the same class and

compare the critical performance parameters. The aircrafts which we chosen were

NAL Saras

CASA Aviocar

Let L-410

The critical performance parameters which we compared were

Passenger Capacity, Payload

Range

Service Ceiling

Maximum Speed

Maximum takeoff weight

Table 16: Validation parameters comparison

The results are shown below as graphical form,

Figure 12: Capacity comparison graph

35

Figure 13: Service Ceiling comparison graph

Figure 14: Maximum Speed comparison graph

Figure 15: Maximum take-off weight comparison

36

Figure 16: Range Comparison graph

From the comparison of above data the conclusions are given below, on

payload or capacity comparison CASA Aviocar is the highest, it has 26

passenger capacities. Our design has 19 passenger capacities just same as the

Let L-410.

On maximum speed comparison NAL Saras proved to be the fastest but our

design has the moderate and optimum speed.

On range comparison LTA spatz came second in the chart. This is reasonable

for a low cost town to town aircraft.

On service ceiling comparison LTA Spatz came second in the chart. CASA

Aviocar came first and NAL Saras came third.

On maximum takeoff weight comparison LTA Spatz came third in the chart.

CASA Aviocar came forth and NAL Saras came first.

From the above comparison it is clear that the LTA Spatz would be cheaper

than the other three. So that it can enable low cost town to town flying.

37

4.2 Result Analysis

Weight estimation

o Maximum takeoff weight : 14111 lbs

o Empty weight : 7500 lbs

o Fuel weight : 2411 lbs

o Payload weight : 4200 lbs

o Matching Engine : Garrett TPE331-10R-513CS

Sizing

o For stall speed requirements

Wing loading : 20.08 lbs/ft2

Wing loading at Takeoff : 22.07 lbs/ft2

Wing loading at Landing : 34.69 lbs/ft2

o For takeoff distance requirements

Approach speed : 84.5 kts

Landing ground run : 1119.63 ft

Landing distance : 2169.83 ft

o For climb requirements

Wetted wing area : 1879.87 ft2

Equivalent parasite area : 5.639 ft2

Oswald efficiency factor : 0.757

Max. wing loading : 34.69 lbs/ft2

Wing area : 406.826 ft2

Zero-lift drag coefficient : 0.01386

Drag coefficient : 0.1674

Airfoil selection

Cruise speed : 189 kts

Average weight at flight : 12905.5 lbs

Cruise lift coefficient : 0.35772

Wing cruise lift coefficient : 0.37602

Airfoil ideal lift coefficient : 0.41780

Maximum lift coefficient : 2.33530

Airfoil gross lift coefficient : 2.59478

38

Airfoil net lift coefficient : 1.79478

Matching airfoil : NACA 653-218

Wing design

Aspect ratio : 10

Taper ratio : 0.5

Wing span : 63.783ft

Root chord : 8.504ft

Tip chord : 4.252ft

Span wise location of MAC : 14.174ft

MAC length : 6.615ft

Flap chord : 1.323ft

Flap span : 19.135ft

Aileron chord : 1.323ft

Aileron span : 6.3783ft

Volume of fuel : 427.481gal

Height of fuel tank : 1ft

Width of fuel tank at root : 3.6ft

Width of fuel tank at far side : 2.2ft

Length of the fuel tank : 10ft

Tank capacity on each wing : 29ft3

Fuselage design

Width of the cabin : 7ft

Height of the cabin : 9.2ft

Length of the cabin : 19ft

Length of cockpit : 10ft

Length of tail boom : 20ft

Length of the fuselage : 49ft

CG location

CG without wing : 21.773ft

CG with wing : 21.882ft

Weight of the wing : 813.651lbs

Location of wing ADC from LE : 1.6536ft

Location of wing CG from LE : 2.6458ft

39

Tail design

o Horizontal tail

Horizontal tail volume ratio : 0.7

Moment arm from CG to ADC of HT : 27.118ft

Plan form Area of Horizontal tail : 69.462ft2

Aspect ratio of Horizontal tail : 4

Taper ratio of Horizontal tail : 0.5

Span of Horizontal tail : 16.669ft

Root chord of Horizontal tail : 5.556ft

Tip chord of Horizontal tail : 2.778ft

Span wise location of MAC of HT : 3.7042ft

MAC length of Horizontal tail : 4.3211ft

o Vertical tail

Vertical tail volume ratio : 0.04

Moment arm from CG to ADC of VT : 25ft

Plan form Area of Vertical tail : 41.518ft2

Aspect ratio of Vertical tail : 1.5

Taper ratio of Vertical tail : 0.6

Span of Vertical tail : 7.892ft

Root chord of Vertical tail : 6.576ft

Tip chord of Vertical tail : 3.946ft

Span wise location of MAC of VT : 3.617ft

MAC length of Vertical tail : 5.370ft

Landing gear design

Static margin : 10%

Location of neutral point : 22.543ft

Location of ADC of wing body : 21.843ft

Location of LE of root chord : 19.245ft

Location of main landing gear : 23.497ft

Location of nose landing gear : 3ft

Load on main landing gear : 12999.1lbs

Load on nose landing gear : 1111.89lbs

40

Main wheel

Diameter : 32.335inch

Width : 11.0643inch

Load : 6499.555lbs

Nose wheel

Diameter : 17.4603inch

Width : 6.3779inch

Load : 1111.89lbs

Using above values we created a CATIA model of the aircraft.

Figure 17 Catia model isometric view

41

Figure 18: Catia model side view

Figure 19: Catia model sectional side view

Figure 20: Catia model front view

42

Figure 21: Catia model top view

43

Chapter 5

CONCLUSION

So far we have completed weight estimation, initial sizing, wing design,

fuselage design, tail design and fuselage design. We also developed a designing tool

and a CATIA model. We were well within the time limit and achieved all project

goals. Our preliminary aim was to create a low cost town to town aircraft, so after the

preliminary design it is clear that the empty weight ratio of our model is considerably

less than that of NAL Saras, also our model has lesser stalling speed and greater wing

loading. All these parameters make our design much cheaper than NAL Saras. Hence

the primary project goal is completed.

44

Chapter 6

FUTURE SCOPE OF WORK

Detail Design

This phase simply deals with the fabrication aspect of the aircraft to be

manufactured. It determines the number, design and location of ribs, spars, sections

and other structural elements. All aerodynamic, structural, propulsion, control and

performance aspects have already been covered in the preliminary design phase and

only the manufacturing remains. Flight simulators for aircraft are also developed at

this stage.

Performance Analysis

In performance analysis I compute most mission profiles based on defined

mission specifications. The types of mission flight phases (takeoff, climb, cruise,

loiter, descent.) are suitable for Commercial low cost town to town aircraft. The

estimate of empty weight is obtained and is lower than initial estimate of empty

weight used for our design calculations to this point.

45

REFERENCES

[1] Roskam, J., Airplane Design: Part I, Preliminary Sizing of Airplanes,

Roskam Aviation and Engineering Corporation, Ottawa, Kansas, 1985

[2] Raymer, D., Aircraft Design: A Conceptual Approach, 3rd

Edition, AIAA

Education Series, Reston, VA, 1999

[3] Mattingly, J., Aircraft Engine Design, 2nd

Edition, AIAA Education Series,

2002

[4] Anderson, JD., Aircraft Performance & Design,University of Maryland,

McGrew-Hill International Editions, 1999

[5] Kuchemann, J., Aerodynamic Design of Aircraft, Pergammon Press,

1982.

[6] Shevell, R.S., Fundamentals of Flight, Prentice Hall, 1983.

[7] Schlichting H. and Truckenbrodt E., Aerodynamics of the Airplane, McGraw-

Hill, 1979.

[8] Torenbeek, E., Synthesis of Subsonic Airplane Design, Delft Univ. Press,

1982.

[9] Taylor, J., ed., Jane’s All the World's Aircraft, Jane's Publishing Inc., Annual.

[10] Articles in Aviation Week & Space Technology, McGraw-Hill.

[11] Raymer, D., Aircraft Design-A Conceptual Approach, AIAA, 1992.

[12] Roskam, J., Aircraft Design, Published by the author as an 8 volume set, 1985-

1990.

[13] Nicolai, L.M., Fundamentals of Aircraft Design, METS, Inc., 6520 Kingsland

Court, San Jose, CA, 95120, 1975.

[14] Stinton D., The Design of the Airplane, van Nostrand Reinhold, New York,

1983.

[15] Thurston D., Design for Flying, Second Edition, Tab Books, 1995.

[16] M. S. Rice., Handbook of Airfoil Sections for Light Aircraft. Published in

1971, Aviation Publications

Documents

Passenger Aircraft Design