acd 2510 conceptual aircraft design report

8/10/2019 ACD 2510 Conceptual Aircraft Design Report

1/66

MSRSAS - Postgraduate Engineering and Management Programme - PEMP

Aircraft Conceptual Design

i

FINAL REPORT

Module Code ACD 510

Module Name Aircraft Conceptual Design

Course M.Sc in Aircraft Design

Department Automotive and Aeronautical Engg.

Name of the Student ACD FT-11

Batch Full-Time / Part-Time 2011.

Module Leader Dr. H K Narahari

POSTGRADUATEENGIN

EERING

ANDMANAGEMENTPROGRA

MME(PEMP)

M.S.Ramaiah School of Advanced StudiesPostgraduate Engineering and Management Programmes(PEMP)

#470-P Peenya Industrial Area, 4th Phase, Peenya, Bengaluru-560 058

Tel; 080 4906 5555, website: www.msrsas.org


2/66


ii


Declaration SheetStudent Name ACD FT-11

Reg. No

Course MSc [Egg] Aircraft Design Batch Full-Time 2011.Batch FT 11

Module Code ACD 510

Module Title Aircraft Conceptual Design

Module Date 06.08.2012 to 1.09.2012

Module Leader Dr. H K Narahari

Extension requests:Extensions can only be granted by the Head of the Department in consultation with the module leader.

Extensions granted by any other person will not be accepted and hence the assignment will incur a penalty.

Extensions MUST be requested by using the Extension Request Form, which is available with the ARO.A copy of the extension approval must be attached to the assignment submitted .

Penalty for late submissionUnless you have submitted proof of mitigating circumstances or have been granted an extension, the

penalties for a late submission of an assignment shall be as follows:

Up to one week late: Penalty of 5 marks

One-Two weeks late: Penalty of 10 marks

More than Two weeks late: Fail - 0% recorded (F)

All late assignments: must be submitted to Academic Records Office (ARO). It is your responsibility to

ensure that the receipt of a late assignment is recorded in the ARO. If an extension was agreed, the

authorization should be submitted to ARO during the submission of assignment.

To ensure assignment reports are written concisely, the length should be restricted to a limit

indicated in the assignment problem statement. Assignment reports greater than this length may

incur a penalty of one grade (5 marks). Each delegate is required to retain a copy of the

assignment report.

DeclarationThe assignment submitted herewith is a result of my own investigations and that I have conformed to the

guidelines against plagiarism as laid out in the PEMP Student Handbook. All sections of the text and

results, which have been obtained from other sources, are fully referenced. I understand that cheating and

plagiarism constitute a breach of University regulations and will be dealt with accordingly.

Signature of the student Date

Submission date stamp(by ARO)

Signature of the Module Leader and date Signature of Head of the Department and date


3/66



iii

Abstract____________________________________________________________________________

Aircraft design mainly depends on the existing historical data. Drastic changes in design cannot

be made hence it is usually the up gradation or the enhancement of the existing design. This

report is such an effort to Design of a Passenger Aircraft with a Range of 3000 nautical miles

with a passenger Capacity of 80. This report is made by a group of five members working on

different areas of design.

The first Chapter deals with the initial weight estimation which is usually the first step in the

design process of any aircraft. This is only a rough estimate rather than detailed weight

estimation because the weights must be varied in order to meet various customer requirements

in the later part of the design phase. This is because the design is a compromise of several

parameters.

Chapter 2 deals with the design of wing and high lifting devices. Various parameters like Wing

area, aspect ratio, airfoil, geometry, sweep angle, taper ratio is actually assumed based on the

historical data available. This is a good starting point because a lot of time is saved and most

importantly near close or meaningful results can be obtained in first set of assumption itself.

CFD analysis is carried out with the final geometry to validate the results obtained. Similar

procedure is carried out with flaps deflected at a particular angle calculated using Javafoil in

order to obtain max Cl needed as per the requirement.

Chapter 3 deals with the fuselage design which includes the entire layout and seating

arrangement which usually depends on the number of passengers. Several considerations are

carried out in design process to meet all FAA regulation like vision of the pilot and space

between the seats and also the dimension of the seats.

Chapter 4 deals with the selection of a propulsion system and the integration of the same. This

process depends upon the gross weight and the thrust requirements.Chapter 5 deals with the design of horizontal and vertical tail and various performance

parameters are calculated to check it meets with the customer requirement.


4/66


iv


Contents____________________________________________________________________________

Declaration Sheet ......................................................................................................................... iiAbstract ....................................................................................................................................... iii

Contents ........................................................................................................................................iv

List of Figures ..............................................................................................................................vi

List of Nomenclature ....................................................................................................................ix

CHAPTER 1 .............................................................................................................................. 10

1.1 Introduction ....................................................................................................................... 10

1.2 Initial Weight Estimate ...................................................................................................... 10

1.2.1 Payload Weight .......................................................................................................... 11

1.2.2 Crew Weight ............................................................................................................... 11

1.2.3 Empty weight Fraction ............................................................................................... 11

1.2.4 Fuel weight Fraction ................................................................................................... 12

2.1 Wing Design ...................................................................................................................... 16

2.1.1 Wing loading .............................................................................................................. 17

2.1.2 Aspect Ratio ............................................................................................................... 17

2.1.3 Wing Sweep and Taper .............................................................................................. 18

2.1.4 Wing Geometry and Planform ................................................................................... 19

2.1.5 Number of wings ........................................................................................................ 20

2.1.6 Wing vertical location on the fuselage [1] ................................................................. 20

2.1.7 Steps for selection of the Airfoil for the wing [4] ...................................................... 23

2.1.8 Wing Twist ................................................................................................................. 27

2.2 Wing high lift devices ....................................................................................................... 30

2.2.1 Calculation of Takeoff and landing distance .............................................................. 31

2.2.2 Javafoil ....................................................................................................................... 32

2.2.3 Finite 3D Wing ........................................................................................................... 34

3.1Fuselage layout [6] ............................................................................................................. 36

3.2 Fuselage nose section ........................................................................................................ 38

3.3 Fuselage Mid-section ........................................................................................................ 39

3.4 Fuselage tail section .......................................................................................................... 39

3.5 Galley and Toilet configuration ........................................................................................ 40

3.6 Passenger loading and emergency exits ............................................................................ 40


5/66



v

3.7 Structural consideration ..................................................................................................... 42

3.8 Weight estimation for fuselage .......................................................................................... 43

3.9 Landing gear layout: .......................................................................................................... 45

4.1Propulsion System .............................................................................................................. 474.2Engine Selection ................................................................................................................. 47

4.3 Why to go for CFM56-5B3 Engine ................................................................................... 49

4.4 Engine integration ............................................................................................................. 51

5.1 Empennage ........................................................................................................................ 52

5.2 Empennage types ............................................................................................................... 52

5.3 Empennage design ............................................................................................................. 53

5.4 Tail geometry .................................................................................................................... 53

5.5 Tail sizing .......................................................................................................................... 55

5.6 Tail layout .......................................................................................................................... 56

6.1 Empty weight build up and C.G location .......................................................................... 59

6.2 Performance ....................................................................................................................... 61

7.1 Assembly ........................................................................................................................... 64

8.1 Conclusion ......................................................................................................................... 65

REFERENCES .......................................................................................................................... 66


6/66



vi

List of Figures____________________________________________________________________________

Figure1. 1 Empty weight fraction Vs Wo[1] ............................................................................... 11

Figure1. 2 Mission Profile ........................................................................................................... 12

Figure1. 3 Aircraft type based on Range[2] ................................................................................ 13

Figure1. 4 Selection of SFC[1] .................................................................................................... 13

Figure1. 5 Selection of L/D Ratio[1] ........................................................................................... 13

Figure1. 6 Selection of L/D based on type of Aircraft[2] ........................................................... 14

Figure1. 7 No. of Passenger Vs MTOM[3] ................................................................................. 15

Figure1. 8 Range Vs MTOM/Passenger[3] ................................................................................. 16

Figure2. 1 Selection of Wing Loading[2][3] ............................................................................... 17

Figure2. 2 Selection of Aspect Ratio[2] ...................................................................................... 18

Figure2. 3 Mach no Vs Leading edge sweep angle[1] ................................................................ 19

Figure2. 4 Maximum and ideal lift co-efficient plot [5] ............................................................. 25

Figure2. 5 Cl Vs Cd plot for 631-412 [5] .................................................................................... 26

Figure2. 6 Airfoil created using ICEM CFD ............................................................................... 28

Figure2. 7 CATIA model of the wing (isometric view) .............................................................. 28

Figure2. 8 Top view of the wing ................................................................................................. 29

Figure2. 9 Domain around the wing ............................................................................................ 29

Figure2. 10 3-D domains around the wing .................................................................................. 29

Figure2. 11 Unstructured mesh for the wing ............................................................................... 30

Figure2. 12 Creation of Airfoil geometry ................................................................................... 32

Figure2. 13 Flap deflection ......................................................................................................... 33

Figure2. 14 Computation of Cl .................................................................................................... 33

Figure2. 15 Generation of curves using ICEM CFD from point data ......................................... 34Figure2. 16 CATIA model of 3D wing ....................................................................................... 34

Figure2. 17 Domain sketch ......................................................................................................... 35

Figure2. 18 Domain surface ........................................................................................................ 35

Figure2. 19 Mesh generated in ICEM CFD ................................................................................ 35

Figure3. 1 Seat pitch and height. [3] ........................................................................................... 36

Figure3. 2 Fuselage Interior detail (mm). .................................................................................... 37

Figure3. 3 Fuselage Plan view (mm). .......................................................................................... 37

Figure3. 4 Fuselage side view (mm). .......................................................................................... 38


7/66



vii

Figure3. 5 Nose section layout. [6] ............................................................................................. 38

Figure3. 6 Pilot Vision designed in Catia (mm). ......................................................................... 38

Figure3. 7 Pilot seat dimensions as per FAR. [3] ........................................................................ 39

Figure3. 8 Galley and toilet layout. [6] ....................................................................................... 40Figure3. 9 CATIA model for fuselage view-1 ............................................................................ 41

Figure3. 10 CATIA model fuselage view-2 ................................................................................ 41

Figure3. 11 Mesh geometry of fuselage and domain. ................................................................. 42

Figure3. 12 Mesh geometry of fuselage. ..................................................................................... 42

Figure3. 13 Stringers layout. [1] ................................................................................................. 43

Figure3. 14 Mass fractions for rapid mass estimation. [3] .......................................................... 44

Figure3. 15 Landing gear forward retract. [8] ............................................................................ 45

Figure3. 16 Nose landing gear load calculation. [8] ................................................................... 46

Figure3. 17 Tires used in typical aircraft. ................................................................................... 46

Figure 4. 1 Effect of flight speed on engine efficiency. [6] ........................................................ 48

Figure 4. 2 Effect of Mach number and specific thrust on thrust lapse rate. [6] ......................... 48

Figure 4. 3 Inlet locations-podded engines. [1] ........................................................................... 51

Figure 4. 4 Location of the nacelle compared to the wing .......................................................... 51

Figure 5. 1 Tail variations [1] ...................................................................................................... 52

Figure 5. 2 Tail aspect ratio and taper ratio for various types of aircraft [1] .............................. 54

Figure 5. 3 Tail volume coefficient [1] ....................................................................................... 55

Figure 5. 4 NACA 0012 airfoil coordinates with deflected control surface ............................... 57

Figure 5. 5 Velocity distribution and Cl values with deflected control surface .......................... 58

Figure 5. 6 NACA 0012 airfoil with deflected control surface in CATIA ................................. 58

Figure 5. 7 CATIA model-Horizontal tail ................................................................................... 59

Figure 5. 8 CATIA model-Vertical tail ....................................................................................... 59

Figure 6. 1 Approximate empty weight build up and c.g. location [1] ....................................... 60

Figure 7. 1 Front View ................................................................................................................ 64

Figure 7. 2 Top view ................................................................................................................... 64

Figure 7. 3 Side view ................................................................................................................... 64

Figure 7. 4 Isometric view ........................................................................................................... 65


8/66



viii

List of Tables

____________________________________________________________________________

Table1. 1Iterations ....................................................................................................................... 16

Table1. 2 2D Clincrement for leading edge flaps[2] .................................................................. 31

Table2. 1 Different set of airfoils selected from the graph [5] .................................................... 25

Table2. 2 Geometric twist for different aircrafts ......................................................................... 27

Table3. 1 Typical seat width and pitch for different class of travel. [6] ..................................... 36

Table3. 2 Typical guidelines for fuselage front and aft closure ratio. [3] ................................... 37

Table3. 3 Emergence exits requirements. [6] .............................................................................. 40

Table3. 4 Dimensions for types of exits. [6] ............................................................................... 41

Table3. 5 Mass amended for different configuration. [6] ........................................................... 43

Table 4. 1 Specifications and applications of CFM56-5 series engines. [10] ............................. 50

Table 5. 1 Design data of wing and fuselage .............................................................................. 54

Table 5. 2 Geometrical data supposed for tail design ................................................................. 54

Table 5. 3 Tail Design configuration ........................................................................................... 57

Table 6. 1 Empty weight build up and c.g. location .................................................................... 61

Table 6. 2 Derived data from design ........................................................................................... 62

Table 6. 3 Designed aircraft characteristics ................................................................................ 63


9/66



ix

List of Nomenclature____________________________________________________________________________

Acronym Expansion

C l Coefficient of lift

Cd Coefficient of Drag

TWR Thrust to Weight Ratio

WL Wing Loading

LDR Lift to Drag Ratio

Ct Specific Fuel Consumption

AR Aspect ratio

Wf

MTOM

Cdo

DSL

Fuel Weight

Maximum Takeoff weight

Zero lift Drag co efficient

Density at Sea Level

.


10/66



CHAPTER 1

Design of a Passenger Aircraft with a Range of 3000 nautical miles with a passenger

Capacity of 80

1.1 Introduction

This report deals with the conceptual design of an aircraft. Any type of aircraft has the same

parts like fuselage, wing, engine etc but analyzing it a bit deeper it is evident that they are only

same by parts not on the geometry or the size. Fighter planes are smaller but much faster when

compared to that of the commercial transport plane. The root cause for such a big difference in the

design of aircraft is mainly due to the change in the functionality that has to be met. So in order to

start the conceptual design, proper understanding of the purpose of the aircraft to be designed is

required. Design in any field is a compromise of various factors, especially in the field of aircraft

where changing one influencing parameter affects many other influencing parameters. Since there

is a correlation between various influencing parameters the final modified design cannot be

achieved by changing only one influencing parameter. It is usually a combination of various

parameters to meet the final customer requirements and moreover other performance parameter as

per FAA (Federal Aviation Administration) regulation has to be satisfied for the aircraft to be

certified.An aircraft with a range of 3000 (nautical miles) and a passenger capacity of 80 comes under

the small Jet transport category. Looking into the available statistical data of similar type of aircraft

is regarded as the good starting point for a new design. The design process begins with the rough

estimate of the maximum take off mass. In this process several decisions has to be made on initial

assumptions to be used for calculations. Since most of the processes in aircraft design are iterative

in nature it is important to start with a meaningful assumption. By doing so reasonable results can

be achieved in a shorter span. This is done with the help of available statistical data.

1.2 Initial Weight Estimate

First step is to come up with a rough estimate of the total takeoff weight which is denoted by

Wo. It is the sum of crew weight, payload weight, fuel weight and empty weight.


11/66



11

1.2.1 Payload Weight

Here the payload weight can be estimated with the number of passenger and the amount of baggage

that they are permitted to carry.

W payload = N * (Average weight of a passenger + Permitted baggage per individual)

Where N is the number of passengers and the initial assumption for average weight of an

individual passenger is taken as 80 kg (173.36 pounds) and the permitted baggage as 40 kg (88.14

pounds). So

W payload = 80 * (80+40) = 9600 kg (21165 pounds)

1.2.2 Crew Weight

Crew consists of a pilot, a co pilot, a flight engineer and for a passenger capacity of 80 we need two

air hostesses which all together make a crew as five members and assuming the crew members are

allowed to carry a baggage of 40 kg. Then

W crew = 5 * (120) = 600 kg (1323 pounds)

1.2.3 Empty weight Fraction

Empty weight is denoted as We, rather than calculating the empty weight as such, empty weight

fraction is calculated. Empty weight fraction is the ratio between the empty weight and the total

takeoff weight. This can be calculated from the statistical data available.

Empty weight fraction = (We/ Wo) and is given by We/ Wo = A* WoC* Kvs

Figure1. 1 Empty weight fraction Vs Wo[1]


12/66



12

Since the aircraft to be designed comes under Jet transport appropriate values for A and C are

selected. Therefore

(We/Wo) = 1.02 Wo-0.06 *

1.04

(We/Wo) = 1.0608 Wo-0.06

1.2.4 Fuel weight Fraction

Fuel weight as such cannot be estimated from statistical data because amount of fuel to be carried

varies according to the mission profile. The mission profile is usually given as the customer

requirement. Statistical data are available to find out the fuel fraction at various stages of the

mission profile hence the total fuel fraction can be estimated.

Figure1. 2 Mission Profile

Where

1-2 Warm up and Takeoff

2-3 Climb

3-4 Cruise

4-5 Loiter and descent

5-6 Landing phase

Let W1, W2, W3, W4, W5and W6be the weights after respective phases of the mission profile.

1.2.4.1 Calculation

Amount of fuel utilized for warm up and takeoff phase is estimated as 0.97

(W1/W0) = 0.97

Amount of fuel utilized for climb phase with respect to the amount of fuel left over after thetakeoff phase is estimated as 0.985


13/66



13

(W2/W1) = 0.985

Fuel utilized for Cruise depends upon the range to be achieved and can be calculated usingRange equation

Therefore Wi-1/ Wi in our case it is given by (W3/W2) is given by

Where R is the Range, C is the specific Fuel Consumption, V is the velocity, L/D is the lift to

drag ratio

Figure1. 3 Aircraft type based on Range[2]

Range for this mission is 3000 nautical miles (18228341.6 ft)

Specific fuel consumption for a high bypass turbo fan engine for cruise was found to be 0.5

1/hr (0.0001389 1/s)

Figure1. 4 Selection of SFC[1]

The cruise velocity was assumed to be 0.85 M (845.58 ft/sec).

In a similar fashion the L/D ratio for jet engine at cruise is given by L/D = 0.866 (L/D) max.

The selection of L/D and specific fuel consumption comes from the statistical data

available. (L/D) max was assumed to be 16. So L/D for cruise = 0.866*16 = 13.9

Figure1. 5 Selection of L/D Ratio[1]


14/66



14

Figure1. 6 Selection of L/D based on type of Aircraft[2]

Substituting all the values we get

(W3/W2) = e-0.2154

= 0.8062

In case the intended airport is closed due to some unavoidable reason like worst climatic

condition, the aircraft has to loiter for extra duration or can have an increased range to reach

the nearest airport. According to FAA regulation an additional fuel for loitering atleast for

30 min has to be provided. In this case the additional time for endurance is taken as 45 min

which includes both loiter and descent. Fuel ratio calculation for endurance is as follows.

It can be obtained from endurance formula.

Therefore Wi-1/ Wi in our case it is given by (W4/W3) is given by

Where

E is the endurance in seconds. In this case it is 45 min (2700 sec)

But again the selection of specific fuel consumption and L/D ratio is based on statistical data

but it varies because it is a loitering phase. Specific fuel consumption for jet engines during

loiter phase is estimated as 0.4 1/ hr (0.0001111 1/s)

L/D for jet engines for loitering phase is the L/D max hence the value is 16 from statistical

data which says L/D selection can lie between (15 - 18.2).

Substituting all the values we get

(W4/W3) = e-0.01874

= 0.9814


15/66



15

Fuel fraction used for landing is estimated as 0.995

(W5/W4) = 0.995

Therefore W5/W0= (W1/WO) * (W2/W1) * (W3/W2) * (W4/W3) * (W5/W4)

= 0.97 * 0.985 * 0.8062 * 0.9814 * 0.995

W5/W0= 0.7521

But the Total fuel fraction (Wf/ Wo) is given by the formula

(Wf/ W

o) = 1.06 * (1- W

5/W

0)

(Wf/ Wo) = 1.06 * (1- 0.7521) = 0.2626

We know that

Simplifying the equation, we get

77

It is an iterative process so the initial guess should be reasonable so that a lot of time can be saved.

Figure1. 7 No. of Passenger Vs MTOM[3]


16/66



16

From the Figure 1.3 a initial estimate of 40000 kg approximately 80000 pounds is taken

Table1. 1Iterations

No ofteration nitial Gess ofWo Obtained ValeWo1 80000 113148.22 113150 107162.13 107000 1080754 108000 107922.45 107900 1079376 107933 107932.6

So the initial estimate of Maimm takeoff mass is 48957 kg (107933 ponds). The allatedale an be ross heked ith the range VS MTOM / passenger

Figure1. 8 Range Vs MTOM/Passenger[3]

So for a range of 3000 nm and a passenger capacity of 80 we have MTOM as 80*600 we get 48000

kg which is comparable to the estimated value.

2.1 Wing Design

The different parameters that needs to be found during the wing design is

Wing area

Aspect ratio

Taper ratio

Root chord

Tip chord

Mean aerodynamic chord

Span

Number of wings

Wing position on the fuselage


17/66



17

Airfoil cross section

Twist angle

Sweep angle

Dihedral angle Wing incidence angle for Cruise

The step followed in the selection of wing area is discussed here in a sequential manner

2.1.1 Wing loading

Once the weight estimation is done, the next step is the proper selection of wing loading. The

selection of wing loading depends upon the mission requirement of an aircraft. Wing loading is the

ratio of total takeoff weight to that of the wing area. It is denoted by W/S and its units are Kg/m2or

lbs/ft2. It is found that for a short/ medium range aircraft the wing loading is about 110 lbs/ft

2

(537.06 kg/m2).

W/S = 537.06 Kg/m2(110 lbs/ft

2)

Figure2. 1 Selection of Wing Loading[2][3]

2.1.2 Aspect Ratio

Aspect ratio is defined as the ratio between the wing span to the wing mean aerodynamic. The

selection of aspect ratio in conceptual design phase is mainly on the historical data but what leads to


18/66



18

the selection of particular aspect ratio is based on its effects on the various flight features such as

aircraft performance, stability, control, cost, and manufacturability.

The effects of aspect ratio are

As the aspect ratio increases, the aerodynamic features of the three-dimensional wing suchas CL, 0, CLmax, CDmin will get closer to the two-dimensional airfoil properties. This is

because of reduction of the influence of wing tip vortex, so it is desired have high AR.

As the AR increases, the maximum lift co-efficient for a particular angle of attack increases,

because the wing effective angle of attack increases, so it is desired to have a high aspect

ratio wing.

As the aspect ratio increases the weight of the wing increases, this needs more stiffer wing,

hence more stress on the root to hold the wing, hence it is desired to have a short wing to

reduce the weight of the wing.

From statistical data it is found that the aircraft that comes under the jet transport category has an

optimum Aspect ratio which lies between (7 - 9.5) hence a approximated value of aspect ratio 9 is

assumed. It is denoted by A.R and is a dimensionless quantity.

Aspect ratio = b2/S = 9

Figure2. 2 Selection of Aspect Ratio[2]

2.1.3 Wing Sweep and Taper

The next step in the design process is to come up with the wing geometry which includes the Root

chord and tip chord dimension, taper ratio, Sweep angle. The sweep back angle is provided to

reduce the effective Mach number at the leading edge. By doing so the loss of lift associated with

supersonic flow can be reduced. The sweep angle usually depends on the cruise Mach number.


19/66



19

Figure2. 3 Mach no Vs Leading edge sweep angle[1]

From Figure 1.6 for a design cruise Mach no of 0.85 the leading edge sweep angle is chosen as 30

degrees. Taper ratio is defined as the ratio of the tip chord to that of the root chord. Tapered wing is

used in order to reduce the lift induced drag. A commercial transport aircraft with swept back wings

has a taper ratio that lie between 0.2 - 0.3. The maximum permitted taper ratio (0.3) is taken in this

design. So

Sweep Angle = 30 degrees

Taper Ratio = 0.3

2.1.4 Wing Geometry and Planform

Wing geometry includes calculation of Wing span, tip chord length, root chord length, Mean

aerodynamic chord length and span wise location of mean aerodynamic chord. All these parameters

can be calculated by the previously assumed parameters. The assumptions were Wing area = 90 m2,

Aspect ratio = 9, Wing sweep = 30 degrees, Taper ratio = 0.3.

Wing span b =

. =

9 9 0

Aspect Ratio =

=

Root chord length Cr = 4.864 m

For calculating Tip chord length, We have Taper Ratio = 0.3


20/66



20

Tip chord length = 1.459 m

Mean Aerodynamic chord length

Span wise location of mean aerodynamic chord

2.1.5 Number of wings

In the olden days because of manufacturing limitations, more number of wings was used to

generate the required lift, now with modern manufacturing technologies and with the development

of new materials such as Aluminum and its alloys and composite have enabled the manufacturing

of single with long span wing.

With the modern technologies and the materials, now a days only single wing is used.

2.1.6 Wing vertical location on the fuselage [1]

The parameter that could be found during the initial stage of the design is the vertical location of

the wing with respect to fuselage centerline. This parameter will directly affect the other parameters

such as the tail location, landing gear design etc

Generally there are four types of configurations available, they are High wing

Mid wing

Low wing

Parasol wing

As seen from the figure that, the most of the cargo aircrafts have a high wing, where as fighter

aircrafts will have the mid wing and the long rage passenger aircrafts will have the low wing. The

advantages and disadvantages of the different configuration are


21/66



21

High wing

The high wing has some advantages as well as some disadvantages for a particular mission, they are

Advantages

It eases the loading and unloading of cargo in to and out of the aircraft, and the trucks tounload and load can move easily around the aircraft.

The clearance from the ground for this configuration is more compared to low wing, which

facilitates the installation of engine on the wing.

It facilitates the aircraft to take-off and land from the sea in case of amphibian aircrafts, thus

preventing the spilling of water in to the engine during take-off, which may shut down the

engine.

The wing will produce more lift compared to low wing and mid-wing, because part of the

fuselage near the connection between two parts of the wing also contributes for the lift

produced by the wing.

Since the CL produced by the wing the wing is high the aircraft can fly at a lower stall speed

compared to high and the low-wing.

Disadvantages

The aircraft in this configuration will have more frontal area, which increases the drag of the

aircraft.

The ground effect will be lower compared to low wing, this will influence on landing and

take-off distances.

Landing gear is longer if connected to the wing. This makes the landing gear heavier and

requires more space inside the wing for retraction system. This will further make the wing

structure heavier.

The wing will produce more induced drag (Di), due to higher lift coefficient.

A high will be structurally 20% more heavier than the low wing.

Low Wing

In this section, advantages and disadvantages of a low wing configuration are discussed

Advantages

The aircraft take off performance will be better compared with a high wing configuration

due to ground effect.

The pilot will have a better view above the horizon, since the wing the wing is below the

pilot


22/66



22

There will be option for the landing gear retracting system in the wing as well as fuselage.

Landing gear will be shorter; this makes the landing gear system lighter and requires less

space inside the wing for retracting system.

The aircraft is lighter compared to the high wing structure, aircraft frontal area in this caseis less. Since the frontal area is less, it has less induced drag.

Disadvantages

The wing generates less lift, compared with a high wing configuration since the wing has

two separate sections because of this the wing has less induced drag.

Because of the first reason the aircraft will have higher stall speed compared with a high

wing configuration due to a lower CLmax because of that the take-off run is longer.

The dihedral effect by the wing is less compared to the high wing, thus the aircraft is

laterally dynamically less stable.

The wing below the pilot will obstruct the view of the pilot below the horizon.

Mid Wing

In general, the features of the mid-wing configuration stand between features of high-wing

configuration and features of low-wing configuration. Some of the new features of a mid-wing

configuration are as follows:

The aircraft structure is heavier, due to the necessity of reinforcing wing root at the

intersection with the fuselage.

The mid wing is more expensive compared with high and low-wing configurations.

The mid wing is more attractive compared with two other configurations.

The mid wing is aerodynamically streamliner compared with two other configurations.

The strut is usually not used to reinforce the wing structure.

The pilot can get into the cockpit using the wing as a step in a small GA aircraft.

The mid-wing has less interference drag than low-wing and high-wing.


23/66



23

2.1.7 Steps for selection of the Airfoil for the wing [4]

The design of airfoil is time consuming as well more cost is involved, for a conceptual design it is

better to select the best available airfoil from the data base. The steps involved in selection of the

airfoil are as follows.

The cruise altitude considered is 12 km at which the air speed is given by, temperature at 12 km is

216

= 7

The cruise number given for the design is 0.85 mach at an altitude of 12 km, therefore the cruise

velocity is given by

Vcruise = 249.8 m/s

First step is to determine the average weight in the cruising flight

7

Where Wiis the initial aircraft weight at the beginning of cruise and Wf is the final aircraft weight

at the end of cruise.

Calculation of aircraft ideal lift co-efficient (CLc).

In the cruising flight the aircraft weight is equal to the lift force.


24/66



24

C= 2 VS

C=.

..

C= 2425299.810.3194249.8 90 =0.465Where V is the aircraft cruise speed, is the air density at cruising altitude, and S is the wing

planform area.

Calculate the wing cruise lift co-efficient (CLcw).

We consider that the wing the only component responsible for the generation of lift, but other

aircraft components such as tail, fuselage etcwill contribute to the total lift negatively or

positively. Thus the relation between aircraft cruise lift coefficient and wing cruise lift coefficient is

a function of aircraft configuration. The contribution of fuselage, tail and other components will

determine the wing contribution to the aircraft lift co-efficient. In the preliminary design phase

where the other components are been decided, then the following relation is used to calculate the

Wing cruise lift co-efficient.

C = .C = 0.4650.95 =0.489

Later in the design process when the other components are decided, this should be validated by

CFD simulations.

Calculation of wing airfoil ideal lift coefficient (Cli).

The wing is a 3-dimensional body whereas the airfoil is 2-D section, therefore the airfoil ideal lift

coefficient is different from the 3-D wing because the wing has a finite span and different chord

lengths and sweep angle results in this variation from the airfoil lift co-efficient, this variation can

be approximated using the relation.

C = C0.9

C = .. =0.54


25/66



25

As per the statistical data the thickness to chord ratio generally used is 10-12%, but the cruise Cl

0.54 is not achieved in any of the airfoil with this thickness range with in the drag bucket. So the

wing area is modified to decrease the Cl cruise.

So the calculation are redone to adjust the Cl by considering wing area as 120 m2

, then the valueslift co-efficient values are

=

1Later in the design process when the other components are decided, this should be validated by

CFD simulations

Table2. 1 Different set of airfoils selected from the graph [5]

No NACA Cdmin Cmo s

(deg)

o

(deg)

Stall

quality

1 631-412 0.006 -0.08 14 -1.5 moderate

2 641-412 0.004 -0.040 12 -1.2 Moderate

3 651-412 0.004 -0.060 12 -1.5 Sharp

4 652-415 0.0035 -0.028 14 -1.2 Sharp

5 642-415 0.0045 -0.028 16 -1.3 Moderate

Figure2. 4 Maximum and ideal lift co-efficient plot [5]


26/66



26

From the ideal lift co-efficient and the maximum lift co-efficient values required, the airfoil

with corresponding values of lift co-efficient are selected from the figure, where the figure

represents the collection of airfoils with different lift co-efficient values. If there is no airfoil ofparticular values then the airfoil that is nearest to the design point is selected.

If the wing is designed for a high subsonic passenger aircraft, select the thinnest airfoil (the

lowest (t/c max). The reason is to reduce the critical Mach number (Mcr) and drag-divergent9 Mach

number (Mdd). This allow the aircraft fly closer to Mach one before the drag rise is encountered. In

general, a thinner airfoil will have a higher Mcr than a thicker airfoil.

Figure2. 5Cl Vs Cd plot for 631-412 [5]

We can notice that the 631-412 is the airfoil which has the maximum Cl which is equal to the

calculated value and the cruise Cl of 0.41 according to the calculation, and the stall is moderate,

which is acceptable so 631-412 airfoil is chosen for the design.

Since the cruise Cl obtained is 0.41 to achieve that the angle of attack of 1.50 is required so the

wing is set at 1.50 angle for cruise. The result is validated using CFD


27/66



27

2.1.8 Wing Twist

Twist is the difference in angle of attack between wing tip and the root, if the wing tip has lower

angle of attack than the root, the wing is said to have negative incidence and if the wing root is at

lower angle of attack then the tip, the wing is said to have positive incidence. In most of the cases

negative incidence is employed, the main reason for it is to have stall at the root first than at the tip,

because in case of stall if the root stalls first, then the pilot can still have a control on the ailerons to

control the aircraft, since the tip is at negative incidence, it will lower the overall lift produced by

the wing. There are two types of twists employed, they are

Aerodynamic twist

Geometric twist

If the different airfoil cross-section used in the root and tip, which will have different zero lift angle

of attack then, it s called as aerodynamic twist. If the tip and root have the same airfoil cross section

and if the incidence is not same then, it is referred to as Geometric twist.

In most of the cases aerodynamic twist is employed because it is easy to manufacture, where as

geometric twist is difficult to manufacture. The negative incidence of -1 to -4 is used in most of the

aircraft, because if the negative incidence is more then it decreases the overall lift produced by the

wing and the twist is also used to obtain the elliptical lift distribution on the wing.

In case of conceptual design phase, the twist is decided based on the historical data available and in

the later stage it can refined based on numerical calculations.

Table2. 2 Geometric twist for different aircrafts

No Aircraft MTOW

(lb)

Wing incidence at

root (iw) (deg)

Wing angle at

tip (deg)

Twist

(deg)

1 Fokker 50 20,800 +3.5 +1.5 -2

2 Cessna 310 4,600 +2.5 -0.5 -3

3 Cessna

Citation I

11,850 +2.5 -0.5 -3

4 Beech King

Air

11,800 +4.8 0 -4.8

5 Beech T-1A

JawHawk

16,100 +3 -3.3 -6.3

6 Beech T-34C 4,300 +4 +1 -3

7 Cessna

StationAir 6

3,600 +1.5 -1.5 -3

8 Gulfstream IV 73,000 +3.5 -2 -5.5

9 Northrop-

Grumman E-

2C Hawkeye

55,000 +4 +1 -3

10 Piper 11,200 +1.5 -1 -2.5


28/66



28

Cheyenne

11 Beech Super

King

12,500 + 3o 48' -1o 7' 4.55'

12 Beech starship 14,900 +3 -5 -3.5

13 Cessna 208 8000 +2o 37' -3o 6' -5o 31'14 Beech 1900D 16,950 +3o 29' -1o 4' -4o 25'

15 Beechjet

400A

16,100 +3 -3o 30' -6o 30'

16 AVRO RJ100 101,500 +3o 6' 0 -3o 6'

17 Lockheed C-

130 Hercules

155,000 +3 0 -3

18 Pilatus PC-9 4,960 +1 -1 -2

19 Piper PA-28-

161 Warrior

2,440 +2 -1 -3

As seen from the table 2.2 for the design weight of 48900, the similar class of aircraft is Northrop-

Grumman E-2C Hawkeye, for which the twist angle of -30is used, since the design weight is close

to it, for the initial design phase the twist angle of -30is being used.

Figure2. 6 Airfoil created using ICEM CFD

Figure2. 7 CATIA model of the wing (isometric view)


29/66



29

Figure2. 8 Top view of the wing

Figure2. 9 Domain around the wing

Figure2. 10 3-D domains around the wing


30/66



30

Figure2. 11 Unstructured mesh for the wing

2.2 Wing high lift devices

The airfoil selected may give a design Cl required for cruise. But the same amount of lift may not

be sufficient during the takeoff and landing condition. Hence it is necessary to design high lifting

devices such as flaps and leading edge slats. The lift force is denoted by L and is given by

The need for high lift devices can be explained through the formula itself. The first thing to be

noticed is that the cruise velocity and the takeoff or landing velocity is not same. The takeoff

velocity is less when compared to that of the cruise velocity. So in order to achieve the desired lift

during takeoff and landing two approaches can be formulated, one is by increasing the C l and the

other way by increasing the surface area. Cl can be increased by increasing the camber of the

airfoil by deflecting the trailing edge flaps or the leading edge slats. There are several types of high

lift devices are available like plain flap, slotted flap, double and triple slotted flap, fowler flaps,

leading edge Krueger flap, slotted leading edge flap (slats) can be used. By using a plain flap, only

the camber of the airfoil can be changed but by using fowler flaps both the effective wing area and

also the camber can be varied. This is achieved by the extension or protrusion of the trailing edge

through some distance and is then deflected. Higher deflection can be achieved in case of fowler

flaps when compared to that of the split flaps because split flaps are prone to flow separation at


31/66



31

higher deflection of the trailing edge but whereas in case of fowler flaps the flow separation does

not occur even at a higher angle of deflection because the effective chord length is also increased.

. Delta Clmax values for different type of leading edge flaps are

Table1. 2 2D Clincrement for leading edge flaps[2]

S.No Type Cl max

1 Fixed Slot 0.2

2 Leading Edge Flap 0.3

3 Kruger Flap 0.3

4 Slats 0.4

The design cruise Cl is 0.44 and if the leading edge slats are used the C l max value will be 0.8

which will also be not sufficient for takeoff and landing purpose. It is roughly estimated that the Cl

max required for takeoff and landing should be somewhere around 1.6 to 2.2 for long/ medium

range aircraft. Hence the desired Cl cannot be achieved just by using only one high lift device but

is usually a combination of both the leading edge slats and the trailing edge flaps. It is roughly

estimated that the leading edge flap deflection is usually 30 to 40 degrees. Since it is only a

conceptual design phase the normal split flap is only considered for initial computation rather than a

slotted flap. The most common flap chord length is 0.25 C from the trailing edge where C is theairfoil chord.

2.2.1 Calculation of Takeoff and landing distance

Landing and takeoff run is usually specified in terms of ground roll. A initial rough approximate of

takeoff distance is given by the formula [2]

Sg

=.(/)

Considering the design cruise Cl of 0.44 (i.e.) without using high lift devices.

Sg=

Sg = 12,726 ft (3,878 m)

So it is clearly evident that the takeoff distance is too large approximately 4 km. The airports cannot

afford to such a big runway. Moreover the takeoff and landing distance comes under customer

specifications and in most cases short take off and landing distance is preferred. The aircraft maynot be certified by FAA if the design does not meet the FAA regulations for takeoff and landing. As


32/66



32

previously stated the Cl max required to meet takeoff and landing should be somewhere around 1.6

to 2.2 for long/ medium range aircraft. So assuming the required Cl max of 2.1, the landing distance

obtained will be 2666.42 fts (812.72 m) which is acceptable. So the takeoff distance with the

implementation of high lift devices is

Sg = 812.72 m

The rough estimate for landing distance for initial calculation is given by the formula [2]

SL= 118 (LP) + 400

Where LP is the landing parameter and is given by

LP =

Substituting the known values in the formula the landing distance is calculated as

SL = 29,900 ft (9113.52 m) for Cl of 0.44 without using high lift devices.

SL =6530 ft (2005.58 m) for Cl of 0.44 with the implementation of high lift devices.

These calculations clearly reveal the importance of lifting devices.

2.2.2 Javafoil

Figure2. 12 Creation of Airfoil geometry


33/66



33

Javafoil is an applet which is available in public domain that can be used to come up with new

airfoil geometry as per the customer requirement. This applet is widely utilised not only to create

the required airfoil co-ordinates but it also provides options to vary the flap deflection angle andalso the location of percentage of flap with respect to the airfoil chord.

Figure2. 13 Flap deflection

The smooth option can be used to smoothen the curve when the flap is deflected.

Figure2. 14 Computation of Cl

It is found that the created airfoil generates a C l of 0.464 at zero degree angle of attack which is

bit higher than the design Cl of 0.44 but taking the 3 D effects into consideration the wing is

attached to the fuselage at an minimum angle not more than 2 degrees because the CL for wing


34/66



34

will be less than 2D airfoil. It is also found that the desired flap deflection angle of 25.5 degrees is

required to achieve the Clmax value of (2.1). This is just an approximation of flap deflection

angle because the results may not be the same in case of a 3D wing hence a bit more deflection of

flap angle of 5 degree may be required to attain the Clmax

2.2.3 Finite 3D Wing

In order to study the variation between the 2D airfoil and a 3D wing, the obtained airfoil co

ordinates from Javafoil applet is converted into point data and is imported in ICEM CFD to

generate the airfoil curve and is then converted into IGES format and is taken to CATIA to create a

3 D wing geometry with all those taper ratio and sweep angle considered. Then the wing geometry

along with the domain is converted into IGES format and is imported to ICEM CFD. Meshing is

carried out in ICEM CFD and the mesh is read in Fluent. The resulting C l value will be less when

compared to that of the Javafoil case because of the 3D effects.

Figure2. 15 Generation of curves using ICEM CFD from point data

Figure2. 16 CATIA model of 3D wing


35/66



35

Figure2. 17 Domain sketch

Figure2. 18 Domain surface

Figure2. 19 Mesh generated in ICEM CFD


36/66



36

3.1Fuselage layout [6]

Structural requirements for pressurization mainly dictate the shape of the fuselage cross section.

Internal pressure loads are better handled by circular cross section by Hoop tension. This makes the

structure stronger and also lighter in comparison to other cross sectional shape. Non circular cross

section is more prone to bending stress for a shell structure. For a fuselage to maximize internal

volume usually interconnecting two or more circular section are considered.

Laying out the fuselage structure mainly depends on the payload specification. The number of

passenger dictates the width and length of the fuselage. The seat arrangement and number of aisle

helps in deciding the fuselage interior structures. The seat configuration considered in this design is

4 abreast single aisle and number of seats along the fuselage helps in fixing the length of the

fuselage. The length to diameter ratio for a given fuselage also needs to be considered at the time of

design since this helps in reducing drag. Low ratio increases the drag penalty where as high ratio

makes the fuselage long and thin which will adversely affect the structural stability of the airplane.

FAR rules have specified the minimum dimensions for different class of passenger seats and are

given in table 1.3.

Table3. 1 Typical seat width and pitch for different class of travel. [6]

Class Seat width (mm) Seat pitch (mm)

Charter 400-420 700-775

Economy 475-525 775-850

Business 575-625 900-950

First class 625-700 950-1050+

The seat width considered for the design is 500mm the seat pitch is 800mm and the aisle width as

500mm. the length of the fuselage for passenger seating is 800x20=16000mm. Fuselage consists of

a nose section, midsection barrel with constant cross section and aft-end closure.

Figure3. 1 Seat pitch and height. [3]


37/66



37

Figure3. 2 Fuselage Interior detail (mm).

Table3. 2 Typical guidelines for fuselage front and aft closure ratio. [3]

Seating abreast Front fuselage

closure ratio. Fcf

Aft fuselage

closure ratio, Fca

Aft closure angle

(deg)

3 1.7-2 2.6-3.5 5-10

4-6 1.5-1.75 2.5-3.75 8-14

7 1.5 2.5-3.75 10-15

Figure3. 3 Fuselage Plan view (mm).


38/66



38

Figure3. 4 Fuselage side view (mm).

Figure3. 5Nose section layout. [6]

3.2 Fuselage nose sectionThe fuselage nose section is from the tip of the nose end to the constant cross section of the mid

fuselage. This section holds the cockpit, flight deck, forward looking radar, nose undercarriage and

the windscreen. The fuselage length to diameter ratio considered is 1.5 as shown in figure 3.3. The

shape of the nose cone is designed such as to reduce drag.

Figure3. 6 Pilot Vision designed in Catia (mm).


39/66



39

Figure3. 7 Pilot seat dimensions as per FAR. [3]

The pilot seat is standardized to have stress free condition for the pilot during takeoff and landing

with ample space to reduce fatigue to the pilot. The wind screen is designed to allow adequate

vision to the pilot for flight maneuver.

3.3 Fuselage Mid-section

This section holds the passenger seating, the length for this section depends on the number of

passenger to be accommodate. The seating are 4 abreast and 20 along with single aisle

configuration with a galley and toilet at both ends. Foldable seats one each at each end of the mid-

section for the crew is provided. The mid-section is 20148mm long as shown in figure3.3

3.4 Fuselage tail section

The tail section is shaped to provide smooth surface to reduce the drag. The tail section also

supports the tail surfaces and in some configuration the engine installation. The lower side of the

profile is tapered to 9deg to provide clearance for the aircraft during takeoff. The tail section length

to diameter ratio considered is 3. The overall fineness ratio of the fuselage is around ~10.


40/66



40

3.5 Galley and Toilet configuration

Figure3. 8Galley and toilet layout. [6]

As per the FAR rules there should be one galley for 10-60 passengers and one toilet for 15-40

passenger. Since the design is for 80 passengers there are two galley and toilets provided at each

end of the midsection of the fuselage. The sizes considered for the galley is 762 mm x 914 mm and

for the toilet is 914mm x 914mm.

3.6 Passenger loading and emergency exits

FAR rules state that during emergency the plane needs to be evacuated within 90 second. This leads

to FAR guidelines for the number of emergency exits that is needed for different number of

passenger.

Table3. 3 Emergence exits requirements. [6]

Seats Emergency exit type

Less than Type I Type II Type III Type IV10 - - - 1

20 - - 1 -

40 - 1 1 -

80 1 - 1 -

110 1 - 2 -

140 2 - 1 -

180 2 - 2 -


41/66



41

Table3. 4 Dimensions for types of exits. [6]

Type Dimension (mm)

Type I 610 x 1219

Type II 508 x 1118

Type III 508 x 914

Type IV 483 x 660

Type A (passenger or service loading door) 1067 x 1829

In consideration to the FAR rules two doors are provided at each end of the fuselage with Type-A

and two emergency exits one with Type I and one with Type III provided at each end of the

fuselage.

Figure3. 9 CATIA model for fuselage view-1

Figure3. 10 CATIA model fuselage view-2


42/66



42

Figure3. 11 Mesh geometry of fuselage and domain.

Figure3. 12 Mesh geometry of fuselage.

3.7 Structural consideration

The fuselage structure should be able to take the bending moment, shear force, torsional loads and

the compressive loads due to self-weight, weight of wings and the weight of engine along with the

thrust force generated by the engine. The structure of the fuselage is constructed with a large

number of stringers distributed along the circumference of the fuselage which helps in resisting

bending of the fuselage. Bulkhead is provided at the ends of midsection of fuselage with frames

along the length of fuselage.


43/66

MSRSAS - Postgraduate E


3.8 Weight estimation for

Even though the weight of fuse

under carriage, initial guess c

considered at 7-12% of the m

formula [2] recommended for

fuselage mass. All the equati

correction factors need to be fa

Where LF= fuselage overall len

DF= fuselage diameter.

VD=aircraft maximum s

Table3

Configurat

For pressurized cabin

For fuselage mounted eng

For fuselage mounted mai

For large cargo door (etc.)

If free from structural disc

For the design considered the

m/s. the fuselage overall length

The fuselage body mass can be

_ = =1

Rapid mass estimation method

gineering and Management Programme - PEMP

Figure3. 13 Stringers layout. [1]

fuselage

age depends on the fuselage size, layout, an

an be assumed from historical data of si

ximum takeoff weight. For a better estima

civil aircraft (50-300 seats) to arrive at i

ns are for all metal (aluminum) constructi

tored in for different material or advanced li

th.

peed.

. 5 Mass amended for different configuration. [6]

on Mass to be a

Increase by

ines Increase by

n undercarriage Increase by

discontinuity Increase by

ontinuity Reduce by

light Mach # 0.85, hence the flight speed i

is 35898 mm, fuselage diameter is 3500 mm

estimated by the Howes formula as

^ ^889 kg.

an also be used to arrive at initial estimate f

43

location of engine and

ilar aircraft which is

ion one can use Howe

itial estimation of the

on only and necessary

hter material.

ended

8%.

4%.

7%.

10%.

4%.

at 340.3 x 0.85 = 289

.

r fuselage weight.


44/66



44

Figure3. 14 Mass fractions for rapid mass estimation. [3]

Roskams [7] suggest few different ways of estimating the fuselage weight,

The general dynamic method,

=10.43(). 100 . 1000 . .Where Kinlet=1.25 for inlets in or on the fuselage, otherwise 1.0

qD=dive dynamic pressure in psf

L=fuselage length

D=fuselage depth

The Torenbeek method,


45/66



45

=0.021 ( + ) . .

Here Kf=1.8 for a pressurized fuselage.

=1.07 for main undercarriage attached to the fuselage

= 1.1 for a cargo aircraft with rear door.

VD=design dive speed in knots equivalent air speed (KEAS)

LH_tail=tail arm of the H-tail

Sfus_gross_area= fuselage shell gross area.

In light of the fuselage construction it is very difficult to predict the weight since the weight

depends on the layout, type of material used and advancement of material technology for better

strength to weight ratio, which the initial estimate needs to be cross checked with detail calculation

and arrive at reasonably accurate estimation. Empirical formulas are available to get an accurate

estimate but are very time consuming since every structural component needs to be accounted for.

3.9 Landing gear layout:

The landing gear used is a tri cycle type. The layout of the landing gear is usually done at the end

after all the weight estimation is made for different section of the aircraft so that the CG is known

for the landing gear location. The landing gear location also depends on the tail-down angle

requirements suited for takeoff and landing attitudes, tipover and general airframe configurations.

The weight estimation also gives an idea whether to use two large wheels or 4 small wheels per

strut. For rough estimation 92% of gross weight is distributed on the main gear at aft CG condition

and 8% of the loads distributed on the nose landing gear at aft CG. The nose landing gear is placed

as far forward as possible to minimize the load on nose landing gear. The gear is designed to retract

forward to have a free fall capability. [8]

Figure3. 15 Landing gear forward retract. [8]


46/66



46

Figure3. 16 Nose landing gear load calculation. [8]

The nose gear load is calculated as, [8]

Max static main gear load (per strut) = W (F-M)/2F

Max static nose gear load = W (F-L)/F

Min static nose gear load = W (F-N)/F

Where W is the gross weight,

For tire selection the nose gear dynamic load is necessary which is calculated as,

Max braking nose gear load = max static load + 10J. W/32.2F

Figure3. 17 Tires used in typical aircraft.


47/66



47

From figure 3.17 the landing gear configuration used is a tri cycle type with two wheel per struts

with tire size 40 x 14 with 155 psi T type.

4.1Propulsion System

The turbofan is a type of air breathing jet engine that is widely used for aircraft propulsion. A

ducted fan which uses the mechanical energy from the gas turbine to accelerate air rearwards. The

ratio of the mass-flow of air bypassing the engine core compared to the mass-flow of air passing

through the core is referred to as the bypass ratio. Current high bypass ratio turbofan engines are

thermodynamically much more efficient than the early turbojet and low bypass types. This has been

largely brought about by the introduction of advanced technologies which have enabled turbine

blades to withstand high centrifugal loads whilst operating in gas temperatures considerably higher

than the melting point of the unprotected blade material. Most commercial aviation jet engines in

use today are of the high-bypass type, and most modern military fighter engines are low-bypass.

In passenger aircraft, efficiency is the main factor rather than performance, large aircraft the fuel

price accounts for some 30% of the aircraft direct operating costs. [3] By increasing the fuel

efficiency the amount of fuel carried will be lesser there by reducing the total weight. A more fuel

efficient engine will require less fuel to fly a given range and hence will lead to a lower take-off

weight.

We are designing 80 seater passenger aircraft having maximum takeoff mass is 48957 kg and range

of 3000 nm keeping this in mind we have to select the appropriate engine which suits the

requirements, the engine should weigh less weight so that maximum takeoff mass will not alter so

much and the engine should have better efficiency as the engine efficiency is one of the key factor

in passenger aircraft as efficiency decreases operating cost increases in order to decrease the

operating cost we need to choose the engine which is having better efficiency.

4.2Engine Selection

The weight estimation can be made by using the statically data of similar plane and plotting the

results of MTOW with number of passenger. Similar plots can also be made with range against

MTOW per passenger and using this number one can arrive at estimating the maximum take-off

weight (MTOW). Figure 1.1 and table 1.1 shows this relation and is used to arrive at first estimate

for the given design.


48/66



48

Figure 4. 1 Effect of flight speed on engine efficiency. [6]

Above shows the difference between turboprop, turbofan and propfan engines, we can see from the

graph that Turboprop engines are having higher efficiency than Propfan and Turbofan engines but

Turboprop engine efficiency is high at lesser mach number (mach 0.5-0.6), but we need the cruise

speed of 0.85, for this cruise speed Turbofan engines are having higher efficiency than Propfan and

Turboprop, so it will be efficient if we install Turbofan engine to our design.

Figure 4. 2 Effect of Mach number and specific thrust on thrust lapse rate. [6]

Our aircraft design has total mass of 48957 kg and we know that thrust to weight ratio of passenger

aircraft is in between 0.3 to 0.4.

= 48957 9.81 =0.3


49/66



49

Required thrust = 144080 N = 144.08kN

So we need to choose the engine which gives minimum thrust of 144.08kN. By consideringall above conditions one can go for Pratt & Whitney PW6000 series or CFM International

CFM56-5 series engine.

CFM International CFM56-5 series engine which gives the thrust of more than 100kN, in

our design the CFM56-5B3 engine can be used which gives thrust of 150kN, has bypass ratio of 5.4

and overall pressure ratio of 35.5, CFM56-5B3 is a dual rotor, axial flow turbofan engine, the

integrated fan and booster (low pressure turbine) is driven by a 4 stage low pressure turbine. A

single stage high pressure turbine drives the 9 stage high pressure compressor, the two rotors are

mechanically independent of each other. Air entering the engine is divided into a primary (inner)

airstream and a secondary (outer) airstream. After the primary airstream has been compressed by

the LPC and HPC, combustion of the fuel in the annular combustion chamber increases the HPC

discharge air velocity to drive the high and low pressure turbines. An accessory drive system off the

N2 rotor drives engine and airplane accessory components. [9]

4.3 Why to go for CFM56-5B3 Engine

CFM56-5B3 engines gives more thrust than the PW6000 series.

CFM56-5B3 engines has higher bypass ratio than PW6000 series engines thus there is

increase in the efficiency.

In CFM56-5B3 engines there is option of a double-annular combustor that reduces

emissions (particularly NOx).

A new fan in a longer fan case, and a new low-pressure compressor with a fourth stage.

It has higher efficiency than PW6000 series.

The CFM56-5B Tech Insertion configuration provides operators with up to 1 percent

improvement in fuel consumption over the life of the product compared to the base CFM56-

5B engine.

Low carbon emission than PW6000 series engines.


50/66



50

Table 4. 1 Specifications and applications of CFM56-5 series engines. [10]

( )

.

() 133.46 137.91 146.81 120.11 97.87 104.54 120.11 96.09 103.65

(/) 4194.65 4252.47 4350.85 3990 3638.6 3754.3 3990 3607.5 3741

5.5 5.5 5.4 5.7 6 5.9 5.7 6 5.9

(10.668 0.80 )

() 28.56 28.56 28.56 25.05 25.05 25.05 28.56 25.05 25.05

35.4 35.4 35.5 32.6 32.6 32.6 35.5 32.6 32.6

.

.

() 25.98 25.98 25.98 22.33 22.33 22.33 25.98 22.33 22.33

() 2.6 2.6 2.6 2.6 2.6 2.6 2.6 2.6 2.6

() 1.734 1.734 1.734 1.734 1.734 1.734 1.734 1.734 1.734

() 23.36 23.36 23.36 23.36 23.36 23.36 23.36 23.36 23.36

: 321 321 321 320 319 319 319 319 319

319

Specific Fuel Consumption of the engine is 1.00034*10-4

N/N-S. [11]

Lapse rate = 5.650 C/km


51/66



51

4.4 Engine integration

The engine integration has a significant impact on the aircraft, affecting safety, structural weight,

flutter, drag, control, maximum lift, propulsive efficiency, maintainability, and aircraft growth

potential. Engines can be mounted on different wing positions,

Figure 4. 3 Inlet locations-podded engines. [1]

Engine location is influenced by many considerations including the interference between the

nacelle and the wing which increases drag. Consequently, nacelles must be sufficiently forward and

low to avoid drag increases. However, to minimize the weight of the landing gear and engine pylon,

a general rule is drawn, the nacelles are usually located as close to the wing lower surface as

possible, without causing undue heating of the wing by the engine exhaust.

Figure 4. 4 Location of the nacelle compared to the wing

From the literature survey, the engine can be placed below the wing as shown in the figure.


52/66



52

From the data of thrust to weight ratio, the minimum required thrust is found to be 144.08kN for

our design and CFM56-5B3 engine is selected for the aircraft and clarification for choosing this

engine is discussed.

5.1 Empennage

The function of an empennage is to stabilize the aircraft and provide control moments

needed for maneuver and trim. The empennage consists of a horizontal and a vertical tail. Together

they stabilize an aircrafts pitch and yaw moments. Trim for a horizontal tail refers to the balancing

of moment created by the wing. For a vertical tail, trim force generated by it is largely unexploited,

since most aircrafts are axis symmetrical. But in the case of an engine failure the vertical tail must

provide for enough trim to sustain the aircraft stable. Though it is possible to build a tailless

aircraft, it often comes with greater compromises in weight, wing area, airfoil selection and narrow

centre of gravity range. The other major function of the tail is to provide control. The tail must be

sized so as to provide adequate control at all critical conditions. For a horizontal tail, this includes

control during takeoff and landing, low speed flight and transonic maneuvering. For a vertical tail,

engine out flight, spin recovery and maximum roll rate are vital control conditions.

5.2 Empennage types

Different types of tail variations are available for various aircrafts depending on their functionality.

Some variations in tail design are as shown in figure 5.1.

Figure 5. 1Tail variations [1]

Each type of tail has its uniqueness owing to its functionality. Some of the designs are explained as

follows.


53/66



53

Conventional tail

About 70% of the aircrafts use a conventional tail. This is due to the fact that a conventional

tail will be able to produce the necessary stability and control at the lightest weight.

T tailThis design is very popular due to its high aerodynamic efficiency. Since the horizontal tail

is above, it avoids wing wake and propwash and also reduces fatigue in both structure and pilot. A

smaller area of the vertical and horizontal tails would suffice compared to a conventional tail to

produce the necessary moments. But it comes with a high weight penalty though.

Cruciform tail

The cruciform tail is a compromise between conventional tail and the T tail. It can function

even at high angles of attack like the T tail but with comparatively lesser weight penalty. However

no reduction in tail area can be made as in the case of a T tail.

H tail

In this design the vertical tail is positioned as such so as to have undisturbed flow of air at

high angles of attack. Also, in the case of twin engines the H tail is positioned as such to be in line

of propwash so as to have better control in the case of engine out.

V tail

The V tail is intended to reduce wetted area. With a V tail the horizontal and vertical tail

forces are the result of horizontal and vertical projection of forces exerted on the V surface. The

resulting wetted area of a V tail would be lesser than that of separate horizontal and vertical tails. V

tails offer reduced interference drag, but with some penalty in control surface actuation complexity

as the rudder and elevator controls must be blended to provide the proper movement of V tail

ruddervators. This also results in adverse roll-yaw coupling.

Hence for the purpose of commercial transport aircraft a conventional tail would be appropriate.

5.3 Empennage design

The tail design is quite similar to the design of a wing. However a smaller area of tail is

enough to compensate for the moments about the aircraft aerodynamic centre due to the distance

between tail and wing.

5.4 Tail geometry

The surface area required by the tail is directly proportional to the area of the wing. Hence

area of the tail cannot be determined without the area of the wing. However other geometric

parameters like aspect ratio and taper ratio are similar over a wide range of aircraft types. These are


54/66



54

obtained through statistical data and can be adhered to at the initial stage of tail design. Statistically

obtained data of tail aspect ratio and taper ratio for various types of aircraft is as shown in figure

5.2.

Figure 5. 2 Tail aspect ratio and taper ratio for various types of aircraft [1]

Other geometric parameters are selected based on the following guidelines:

Tail thickness ratio is similar to wing thickness ratio as per historical guidelines

provided in the wing geometry section. Since the aircraft designed is a commercial

transport aircraft, the airfoil selected for the tail is NACA0012 which is similar to the

thickness of the wing airfoil.

Horizontal tail leading edge sweep is set to about 5 deg more than the wing sweep. It

enables the tail to stall after the wing and provides the tail with higher critical Mach

number. Hence loss of elevator effectiveness due to shock formation is avoided.

Vertical tail sweep varies from 35 to 55 deg for most aircrafts. This is to ensure that the

vertical tail has higher critical Mach number than the wing.

Data obtained from wing design and fuselage design of the given aircraft is as shown in table5.1

Table 5. 1 Design data of wing and fuselage

Wing (a

acd 2510 conceptual aircraft design report

Documents