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.
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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.
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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
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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.
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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
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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.
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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
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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
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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,
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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
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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.
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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.
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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)
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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
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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)
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User interface of MIT AD tool for the above calculations shown below
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.
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User interface of MIT AD tool for the above calculations shown below
Table 10: IN MIT AD TOOL Airfoil selection data
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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)
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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,
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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)
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User interface of MIT AD tool for the above calculations shown below
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Table 11: IN MIT AD TOOL Wing design, Flap parameters, and Fuel tank
parameters selection data
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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.
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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)
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User interface of CG location calculation is given below
Table 13: IN MIT AD TOOL Center of gravity location parameters
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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)
For Root chord length,
( ) (37)
For Tip chord length,
(38)
For Span wise Location of mean aerodynamic chord,
[
] (39)
For Mean aerodynamic chord length,
[
] (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.
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Figure 9: Plan view Fuselage and Horizontal tail (from Aircraft performance and
Design by Anderson page 439)
For vertical tail span,
√ (42)
For Root chord length,
( ) (43)
For Tip chord length,
(44)
For Span wise Location of mean aerodynamic chord,
[
] (45)
For Mean aerodynamic chord length,
[
] (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
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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
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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)
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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.
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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
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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