ADP PROJECT - final
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INTRODUCTION In today’s world there are different types of aircrafts with the latest technology. Every year there is a great competition for making an aircraft capable of carrying a large no:of passengers in the aircraft. So, in this report we intend to implant the differentiation among the aircrafts having a sitting capacity of 60-70 passengers. This report gives the different aspects of specifications like wing configuration, weigh specification, power plant specification and performance specification.
In 2007 Hindustan Aeronautics Limited (HAL) and the National Aerospace Laboratories (NAL) were planning to jointly design and develop a 70-seater civil regional aircraft. NAL had held discussions with Pratt and Whitney (Canada) and General Electric (U.S.) for an engine. The NAL-designed RTA-70 is meant to ply short-haul routes and compete with planes of French- Italian aircraft maker Avions de Transport Régional (ATR), a leading exporter of turbo-prop aircraft to the Indian sub-continent.
In 2008, the Indian government through the Ministries of Defence and Civil Aviation have approved the plan and have asked HAL to prepare a roadmap for the project. It will not be an indigenous venture as the government is planning to enter into a memorandum of understanding with major names in the industry like Embraer, Bombardier Aerospace or United Aircraft Corporation. The aircraft was expected to fly in six to seven years.
In 2010 at the India Aviation exhibition held in Hyderabad, a proposed cabin was on display and more details on the specifications of the aircraft have been revealed.
On 23 December 2010, it was announced that the Indian government had asked NAL to consider the use of turbofan engines on the RTA-70. According to an NAL official, the use of a jet engine was seen as "a stepping stone to the high end" by the government
Airplanes come in many different shapes and sizes depending on the mission of the aircraft, but all modern airplanes have certain components in common. An aircraft design is a separate discipline of aeronautical engineering. It is very different from the analytical aspect such as aerodynamics, structures, control and propulsion.
Actually the design is an iterative process as shown in the design wheel. Requirements are set by the prior design trade studies. Concepts are developed to meet the requirements. Design analysis frequently point towards the new concepts and technologies which can initiate a new design effort. However, once a particular design under progress, all these activities are equally important in producing a good aircraft concept.
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AIRCRAFT DESIGN CYCLE
Figure 1 Aircraft design is a separate discipline of aeronautical engineering- different from the analytical disciplines such as aerodynamics, structures, controls and propulsion. An aircraft designer needs to be well versed in these and many such analyses. A good aircraft designs seems to miraculously glide through subsequent evaluations by specialists without major changes being required.
Design is not just the actual layout, but also the analytical processes used to determine what should be designed and how the design should be modified to better meet the requirements. Sometimes a design will begin as an innovative idea rather than as a response to a given requirements.
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THE DESIGN WHEEL
Figure 2 Those involved in design can never quite agree as to just where the design process begins. The designer thinks it starts with a new airplane concept. The sizing specialist knows that nothing can begin until an initial estimate of the weight is made. The customer, civilian or military feels that the design with requirements. All concept over above is mean to be correct.
Actually, design is an iterative effort, as shown in “Design Wheel”. Concepts are developed to
meet requirements. Design analysis frequently points toward new concepts and technologies, which can initiate a whole new design effort.
CYCLES OF DESIGN PROCESS:
Figure 3
Aircraft design can be broken into three major phases, 1.Conceptual design 2.Preliminary design 3.Detail design
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CONCEPTUAL DESIGN:
Conceptual design is a very fluid process. New ideas and problems emerge as a design is investigated in ever increasing detail. Each time the latest design is analyzed and sized, it must be redrawn to reflect the new gross weight, fuel weight, wing size, engine size, and other changes.
Conceptual design will usually begin with either a specific set of design requirements established by the prospective customer or a company -generated guess as to what future customers need. Design requirements include aircraft range and payload, take-off and landing distances, and maneuverability and speed requirements.
The actual design effort usually begins with conceptual sketch. A good conceptual sketch will include the approximate wing and tail geometries, the fuselage shape, and the internal locations of the major components such as the engine, cockpit, payload/passenger compartment, landing gear and fuel tanks.
PRELIMINARY DESIGN:
It can be said to begin when the major changes are over. The big questions such as whether to use a canard or an aft tail have been resolved. At some point late in preliminary design, even minor changes are stopped when a decision is made to freeze the configuration. During this design the specialists in areas such as structures, landing gear, and control systems will design and analyze their portion of the aircraft. Testing is initiated in areas such as aerodynamics, propulsion, structures, and stability and control.
A key activity during this type of design is “LOFTING’. Lofting is the mathematical modeling of the outside skin of the aircraft with sufficient accuracy to insure proper fit between its different parts, even if they are designed by different designers and possibly fabricated in different locations. The ultimate objective during this design is to ready the company for the detail stage, also called “FULL-SCALE DEVELOPMENT”.
DETAIL DESIGN:
Assuming a favorable decision for entering full-scale development, the detail design phase begins in which the actual pieces to be fabricated are designed. For example, during conceptual and preliminary design the wing box will be designed and analyzed as a whole. During detail design, that whole will be broken down into individual ribs, spars, and skins, each of which must be separately designed and analyzed.
Another important part of detail design is called production design. Specialists determine how the airplane will be fabricated, starting with smallest and simplest subassemblies and building upto the final assembly process. Production designers frequently wish to modify the design for ease of manufacture; that can have a major impact on performance or weight. Compromises are inevitable, but the design must still meet the original requirements.
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During detail design, the testing effort intensifies. Actual structure of the aircraft is fabricated and tested. Control laws for the flight control system are tested on an “iron-bird” simulator, a detailed working model of the actuators and flight control surfaces. Flight simulators are developed and flown by both company and customer test pilots. Detail design ends with fabrication of the aircraft. Frequently the fabrication begins on part of the aircraft before the entire detail-design effort is completed.
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DATA COLLECTION AND REQUIREMENTS
In this section we present the design specifications of various aircrafts which comes under the same category as of our project. The data has been organized in tabular form and graphs have been made to depict the required value for our project. The various aircrafts taken into considerations are:
1. XAC Y-7 100 2. IPTN N-250-100 3. ATR-72-200 4. ATR-72-500 5. ILYUSHIN II-114 6. SAAB 2000 7. ANTONOV AN-140 8. De Havilland Dash 8 Q300
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Graphs showing the chosen values..
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Max
.Fue
l(kgf
) M
ax. p
ay lo
ad(k
gf)
7000
6000
5000
4000
3000
Series1
2000
1000
0 0 200 400 600 800
Max.Cruise Speed(kmph)
8000
7000
6000
5000
4000
3000
Series1
2000
1000
0 0 100 200 300 400 500 600 700 800
Max. Cruise Speed(kmph)
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Max
.To
Wei
ght(
Kgf)
M
ax.L
andi
ng W
eigh
t(Kg
f) 30000
25000
20000
15000 Series1
10000
5000
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
30000
25000
20000
15000 Series1
10000
5000
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
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Prop
elle
r Di
a.(m
) Ai
lero
ns(m
2)
6
5
4
3
Series1
2
1
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
4
3.95
3.9
3.85
3.8
3.75
3.7
Series1
3.65
3.6
3.55
0 100 200 300 400 500 600 700 800 Max.Cruise Speed(kmph)
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Ove
rall
Heig
ht(m
) O
vera
ll Le
ngth
(m)
30
25
20
15
Series1
10
5
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
10
9
8
7
6
5
4 Series1
3
2
1
0
0 100 200 300 400 500 600 700 800 Max.Cruise Speed(kmph)
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Fuse
lage
Max
.Wid
th(m
)
Fuse
lage
Ma x
.Dep
th(m
)
3.5
3
2.5
2
1.5 Series1
1
0.5
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
3.5
3
2.5
2
1.5 Series1
1
0.5
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
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Rang
e(km
)
Land
ing
run(
m)
1400
1200
1000
800
600
Series1
400
200
0 0 100 200 300 400 500 600 700 800
Cruise Speed(Kmph)
2500
2000
1500
1000 Series1
500
0 0 100 200 300 400 500 600 700 800
Cruise Speed(kmph)
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Whe
el T
rack
(m)
W/S
450
400
350
300
250
200
150
Series1
100
50
0
0 100 200 300 400 500 600 700 800 Max.Cruise Speed(kmph)
9
8
7
6
5
4 Series1
3
2
1
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
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Ope
ratin
g em
pty
wei
ght(
kgh)
W
heel
Bas
e(m
)
12
10
8
6
Series1
4
2
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
18000
16000
14000
12000
10000
8000
6000
Series1
4000
2000
0
0 100 200 300 400 500 600 700 800 Max.Cruise Speed(kmph)
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Win
gs g
ross
Are
a(m
2)
Win
g Sp
an(m
)
35
30
25
20
15 Series1
10
5
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
100
90
80
70
60
50
40 Series1
30
20
10
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
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Tape
r Rat
io(W
ing)
As
pect
Rat
io
35
30
25
20
15 Series1
10
5
0
0 100 200 300 400 500 600 700 800 Max.Cruise Speed(kmph)
0.7
0.6
0.5
0.4
0.3 Series1
0.2
0.1
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(kmph)
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Max
.Zer
o fu
el w
eigh
t(kg
f)
Max
.pay
load
(kgf
) 8000
7000
6000
5000
4000
3000
Series1
2000
1000
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(m)
25000
20000
15000
10000 Series1
5000
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(m)
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Max
.Win
g lo
adin
g(kg
/m2)
M
ax.p
ower
load
ing(
kgf/
kw)
450
400
350
300
250
200
Series1
150
100
50
0 0 100 200 300 400 500 600 700 800
Max.Cruise speed(m)
7
6
5
4
3 Series1
2
1
0 0 100 200 300 400 500 600 700 800
Max.Cruise speed(m)
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Mai
n W
heel
Size
(m)
Serv
ice
Ceili
ng(m
) 10000
9000
8000
7000
6000
5000
4000
Series1
3000
2000
1000
0
0 100 200 300 400 500 600 700 800 Max.Cruise speed(m)
300
250
200
150 Series1
100
50
0 0 100 200 300 400 500 600 700 800
Max.Cruise Speed(m)
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Nos
e W
heel
Siz
e(m
)
160
140
120
100
80
Series1 60
40
20
0 0 100 200 300 400 500 600 700 800
Max.Cruise speed(m)
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1 . Initial design specifications 1.1 Introduction
In this section, we present an application of the preliminary design procedure. A
60 seater airplane cruising at M = 0.45, at 4.5 km altitude and having a gross still
air range (GSAR) of 2000 km is considered. The presentation is divided into eight
sections
• Data collection
• Preliminary weight estimation
• Optimization of wing loading and thrust loading
• Wing design
• Fuselage design, preliminary design of tail surface and preliminary layout
• CG calculation
• Control surface design
• Performance estimation and presentation of results
1.1 Design Philosophy In this report the aim is to design a 60 seater aircraft for regional transportation purpose. The infrastructural growth, in road and the rail transportation is not in the pace with the country’s GDP growth. The only alternative is air transportation. A sector of upper middle class people would prefer to commute at fast between cities. For this sector of people the airfare which is near first class AC train is comfortable. So the aircraft should be efficient to reduce the per passenger cost, should able to operate in small airports and should be reasonably fast. Considering all the facts in the present design projects we are concentrating on the aircrafts with turboprops.
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The requirements are as mentioned bellow
1. Gross Still Air Range = 2000 km
2. No. of Passenger = 60
3. Flight Cruise speed = 500 km/h
4. Service ceiling = 9000 m
5. Takeoff Distance = 900 m
6. Landing Distance = 900 m
Certification requirements
All commercial aircrafts must satisfy the airworthy requirements to fly in various
countries. Typically each country has its own aviation authority to qualify.
India - DGCA (Director General of Civil Aviation)
UK - CAA (Civil Aviation Authority)
USA - FAA (Federal Aviation Authority)
Russia- CIS
In all the regulation the following aspects should be covered, the severity may
vary.
1. Flight :- This includes performance like stall, take off, climb, cruise,
descent, landing, response to gust etc. Also included are requirements of
stability, controllability and maneuverability.
2. Structural :- Flight loads, ground loads, emergency landing condition,
fatigue evaluation, damage tolerant design and failsafe designs.
3. Power plant :- Fire protection, auxiliary power unit, air intake/exhaust, fuel
system, cooling
4. Others :- Material quality regulations, bird strike, Propeller blade
dismantling and hitting the fuselage etc.
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All these regulation and test are meant for the at most safety of the passengers. Apart
from above said qualifications the environment concerns like emission and noise
pollution needs to be met.
1.2 Preliminary Design
To begin with data is collected for the existing commercial aircrafts available in
service. The following preliminary configurations are taken.
Power plant - To meet the short range, medium speed, short take off and
landing requirements it is preferred to choose turboprops.
Wing Mounting - The wing is mounted on the top of the fuselage. This
configuration is best for the turboprops. The engine can be mounted on the
bottom surface of the wing. This configuration is highly efficient because bottom
surface of the wing generates small amount of lift as compared to the top.
Anything on the top surface of the wing reduces lift considerably.
Landing gear - The aircraft has a retractable tricycle landing gear
Wing and Empennage - The conventional tapered wing configuration will be
used. The T-tail configuration is good from aerodynamic point of view and
conventional tail configuration is good from structural point of view. The tail plane
surfaces are kept well out of the airflow behind the wing, giving smoother flow,
more predictable design characteristics, and better pitch control. This is
especially important for planes operating at low speed, where clean airflow is
required for control.
The effective distance between wing and tailplane can be increased without a
significant increase in the weight of the aircraft. The distance between the two planes
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gives the "leverage" by which the tailplane can control the aircraft's pitch attitude -
with a greater distance, smaller, lighter tailplanes and elevators can be used.
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The tail surfaces are mounted well out of the way of the rear fuselage, permitting this
site to be used for the aircraft's engines. This is why the T-tail arrangement is also
commonly found on airliners with rear-mounted engines.
The horizontal stabilizer is kept farther away from the ground, which helps reduce
damage to it by objects on the ground when taking off or landing.
1.2.1 Preliminary Weight Estimate
For the given number of passenger the pay load estimation is done as follows
1. One crew member for every 30 passengers. Total of 2-crew member for
60 passengers.
2. Flight pilot and co-pilot
3. As per the practiced standards 102 kg per passenger( 80 kg passenger
weight and 22 Kg check in luggage)
Thus the total payload becomes 64 x 102 = 6528 Kg
This 102 kg is considered after referring to similar airplanes.
A database is prepared referring to various aircrafts similar to the aircraft under
design and is presented in Table.1. In this report we will be referring to this table very
often. The design will follow ATR-72 aircraft. Referring to Table.1 we can consider
WTo = 22000 Kg and
WGross=22200 Kg
Calculation
S = W/(W/S) W/S =344 kg/m2 from data collection.
S = 22000/344 ≈ 64 m2
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b = √(AxS) b=27.71 ≈ 28 m
The Root chord Cr =
Taper ratio λ = 0.4 (Ref. from the data table for Aircrafts) Cr
= 3.26 m
Ct = λ X Cr =1.31 m
Referring to Table.1 we can consider
Sv / S =0.21 and Sh / S = 0.25
Sv = 0.21 x 64
= 13.44 m2 Sh = 0.25 x 64
=
16.0 m2
To find the aspect ratio of the vertical tail and horizontal tail
Referring to Table 4.3 of Ref (Raymer Ref.2). for taper ratio and aspect ratio. Horizontal tail Vertical
tail Ah λh Av λv
Fighter 3 -4 0.2 – 0.4 0.6 – 1.4 0.2 – 0.4
Sail plane 6-10 0.3 – 0.5 1.5 – 2.0 0.4 – 0.6
Others 3 - 5 0.3 – 0.6 1.3 – 2.0 0.3 – 0.6
T-tail - - 0..7 – 1.2 0.6 – 1.0
Aspect ratio Av = 1.5 Taper ratio λv = 0.0.35
Ah =
5.5 λh = 0.5
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bv = 4.5 m
bh = 9.4 m
Crv =
Crv = 4.42 m
Ctv = λvCrv = 1.55 m
Crh = 2.27m
Cth = λhCrh = 1.13m
Control Surfaces:
A number of aircrafts and their 3-view drawings are studied and the following
parameters are chosen.
S_flap / S = 0.2
b_flap / b = 0.4
S_ele / S_ht = 0.33
S_rud / S_vt = 0.32
Therefore the following parameters are calculated:
S_ele = 5.28 m2
S_rud = 4.30 m2
b_flap = 11.2 m
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Fuselage:
Length: Referring to ATR-72-200 aircraft whose capacity is 66 passengers and the
length is 27.17.
The overall length of the aircraft is taken as 27 m (our Aircraft capacity is of 64)
lf =27 m
Diameter:
Referring to ATR-72-200, the width of the fuselage is taken as 2.6 m. df
=2.6 m and height of 1.9 m
Overall height:
Based on the dimensions of different aircrafts the overall height is taken as 7.7 m
Engine Location:
Engines will be mounted on the bottom surface of the wing. The upper surface
contributes more for generation of lift. So clean upper surface is an advantage.
Landing gear:
Tricycle retracting type landing gear will be located on the belly of the fuselage.
Main wheel size 863x250 mm
Nose Wheel size 450x190 mm
Wheel Track (m) 4.1 m
Wheel Base (m) 10.77 m
Power Plant:
2 Pratt and Witney 123D Each rated 1604 kW (According to the table.1) and
referring to company website.
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Fig 1.1Antonov- AN-140
Fig 1.2 XAC Y-7 100
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Fig 1.3 ATR 72-500
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Fig 1.4 Proposed Aircraft under design KARAN G
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2. Revised weight estimation
In the previous section ,an initial estimate for the aircraft parameters has been done
.The weight estimate is being revised using refined estimates of fuel weight and empty
eight. The fuel fractions for various phases are worked out in the following steps .The fuel
fractions for warm up, take off ,climb and landing are taken from Raymer Ref.2(4),chapter 3 .
FUEL FRACTION ESTIMATION
The fuel weight depends on the mission profile and the fuel required as reserve. The mission
profile for a civil turbo prop transport aircraft involves Take off
Climb
Cruise
Loiter before landing
Descent and landing
2.1.1 Warm up and Take off
The value for this stage is taken by following the standards given in Raymer Ref.2(4)
,chapter 3
W 1 = 0.97 W 0
W 0 is the weight at take-off and W 1 is the weight at the end of the take-off phase.
2.1.2 Climb
The weight ratio for this stage is chosen by following the standards given in Raymer
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Ref.2(4), chapter 3.
W 2 = 0.985 W 1
2.1.3 Cruise
Derivation for Range:-
The weight ratio for the cruise phase of flight is calculated using the following
Breguet equation Where,
BSFC = 3 N/kW-hr = 0.5 lb/BHP-hr
Gross still air range = 2000 km. Hence
Cruise safe air range = GSAR = 2000 = 1333.33 km 1.5 1.5
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From figure 3.5 of Raymer Ref.2 (4), the Swet Sref value is 5.2.
Therefore the corresponding wetted Aspect ratio is
Aspect ratio/( Swet
Sref ) = 12/5.2=2.3
Corresponding to this value of wetted aspect ratio, ( L D )max is taken as 17
from figure 3.6 of Raymer Ref.2(4). This corresponds to the average value for Civil Turbo
prop aircrafts.
As prescribed by Raymer Ref.2(4) ,chapter 3
( L D )cruise = ( L
D )max
To account for allowances due to head wind during cruise and provision for diversion to
another airport, we proceed as follows.
Head wind is taken as 15 m/s. The time to cover the cruise safe range of 1333.33 km at cruise
velocity of 500km/hr is
Time = 1333.3
500
= 2.67 hrs
Therefore with a head wind of 15m/s or 54km/hr, the additional distance that has to
be accounted for
Additional distance = 54 x 2.67 = 144 km
The allowance for diversion to another airport is taken as 300 km. From the information
available the air distance between nearest airport is about 300 km. The corresponding map
is given in this report.
The total distance during cruise, R = 1333.3 + 144 + 300 = 1777.3 =1780(approx).
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Substituting the above values in the Breguet equation, we get
2.1.4 Loiter
Derivation for endurance:-
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Therefore the
weight ratio for
Loiter phase of
flight is
calculated using
the following
expression
W 1 =3374 N/ m 2
S
from chapter 1.
W 3 = 3374 x 0.985 x 0.897 =2982 N/ m 2
S , η p = 0.7
We design for a loiter time of 30 min, so endurance E = 0.5 hrs
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W
Therefore we get
W 4 = 0.9878 W 3
2.1.5 Landing
Following the standards specified by Raymer Ref.2(4), chapter 3, we take this ratio as
W 5 = 0.995 W 4
Therefore ,
W 5 = 0.995 x 0.9878 x 0.897 x 0.985 x 0.97 = 0.8423 Wo
Allowing a reserve fuel of 6%, we obtain the fuel fraction as
Wf = ς = 1.06(1
- Wo
W 5 ) = 0.1671 Wo
2.2 Empty Weight Fraction To determine the empty weight ratio, we follow the method in Raymer Ref.2(4),
chapter 3, which gives a relation between We o
and W o
as follows.
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W
We = 0.92 Wo
− 0.05
o
(eq 1)
Hence,
W 0 = Wpayload = 6528
(eq 2) 1 − Wf Wo − We Wo 1 − 0.1671 − We Wo
From chapter 1 ,Wpayload =(102x64)=6528 kgf
We solve this equation by iteration
Table 2 : Iterative procedure for W o
Wo(guess) We Wo (from eq 1) W o (from eq 2)
22200 0.5577 23728.6
23728.6 0.5559 23566.7
23566.7 0.5562 23500
23500 0.5562 23500 Hence, the gross weight W o
Wo = 23500 Kg is obtained as
The critical weight ratios are:- We
= 0.5562 , Wg
Wf = 0.1671 , Wo
Wpay
Wo
= 0.2778 K
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Fig 2.1 This picture shows major airports in India.
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Fig 2.2 This picture shows Air network between major airports in India.
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THP = TRe q *V
= 15553*138.9
= 2160.3Kw 1000 1000
η p = 0.8
BHP = THP
= 2160.3
= 2700Kw 0.8 0.8
BHP per engine
BHP = BHP =
2700 = 1350Kw =1822 hp per engine
2 2 The power required per engine is 1350kW
Engine selection
P req= BHP per engine=1350 kW
From Pratt and Whitney engine data base
Model Max SHP TO RPM Max continuous power
PW 120 1491 kW 1212 1268 kW
PW 121A 1640 kW 1212 1417 kW
PW 123D 1604 kW 1212 1454 kW ◄
PW 127F 2051 kW 1212 1864 kW
Considering the power requirement, PW123D engine is selected.
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Propeller selection N = 1212 RPM information from website for PW123D engine
n= 1212/60 = 20.2 RPS
ρ =0.77 kg/m3 = 0.001508 slug/ft3
=455.6(0.001508/(1002100*20.22))1/5
=2.36 For a four bladed propeller with Cs =2.36
J for ηmax = 1.64
J=V/nd
Propeller dia D=V/n J = 455.6/20.2*1.64
= 13.75 ft = 4.2 m Referring to the similar airplanes data,
The propeller diameter is chosen as 3.9 m. K
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Fig 3.1 Drag polar compared with Fokker-50
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4.0 WING DESIGN
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4.1 Introduction
The weight and the wing loading of the airplane have been obtained in sections 2 and 3.
The details are as below
Weight = 23500 Kg = 230535 N
Wing loading = 3600 N/m2
Wing area is obtained as 64.04 m2
The wing design involves choosing the following parameters.
1. Airfoil selection
2. Aspect ratio
3. Sweep
4. Taper ratio
5. Twist
6. Incidence
7. Dihedral
8. Vertical location
4.2 Airfoil selection
The airfoil shape influences Many aerodynamic parameters. I has an influence
on stalling speed, fuel consumption during cruise, turning performance and weight of the
airplane.
After referring to the existing similar airplanes, NACA 653618 is chosen. For
NACA 653618
Details:
Design lift co-efficient 0.6
Thickness ratio 18 percent
4.3 Aspect ratio
The aspect ratio affects CL, Cdi and wing weight.
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At present stage of design, we chose aspect ratio A = 12 based on the data
collection. (Table 1).
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4.4 Sweep Referring to Raymer Ref.2,
Generally for low subsonic speed, sweep will be taken as zero. Based
on the data of similar airplanes, sweep = 0 deg.
4.5 Taper ratio
Based on the data collected, taper ratio selected as 0.4. 4.6 Twist
Based on the data of similar airplanes, twist = 0 deg.
4.7 Wing incidence
The wing incidence angle is the angle between wing reference chord and fuselage
reference line. Wing incidence angle is chosen to minimize drag at some operating
conditions, usually at cruise.
The wing incidence angle is the angle between wing reference chord and fuselage reference line. Wing incidence angle is chosen to minimize drag at some operating conditions, usually at cruise.
W
CLopt = S = qcr
3600 0.5 * 0.77 *138.92
= 0.48466
CLopt = 0.48466 = CLcruise
CLcruise = CLα (iw − α oL )
t/c = 18 From Raymer Ref.2 Λ = 0
β 2 = 0.8319
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Λ = 0
Clα Slope of the aerofoil NACA 653618
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w
Clα =6.07 per radian
CLα= 5.62 per radian = 0.098 per deg
CLcruise = CLα (iw − α oL )
α oL = 4.1 for NACA 653618 aerofoil.
iw = 0.04125 rad
i = 2.360
Based on the data collected, the wing incidence angle is chosen as iw= 2o
4.8 Vertical location of wing.
The vertical location of wing for the designed airplane has been chosen to be a High
wing configuration which is typical of similar airplanes.
High Wing configuration:
Advantages:
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i) Allows placing fuselage closer to ground, thus allowing loading and unloading
without special ground handling equipment.
ii)Jet engines & propeller have sufficient ground clearance without excessive landing gear
length leading to lower landing gear weight.
iii) For low speed airplanes, weight saving can be effected by strut braced wing.
iv)For short take off and landing (STOL) airplanes with high wing configuration have
following specific advantages. (a) Large wing flaps can be used (b) Engines are away
from the ground and hence ingestion of debris rising from unprepared runways is avoided
(c) Prevents floating of wing due to ground effect which may occur for low wing
configuration.
4.9 Dihedral
Based on the data of similar airplanes, dihedral is chosen as 3 deg.
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Parameters
XAC-Y7
ATR-42
ATR-72
IPTN-N250-100
IL 114 SAAB 2000
An140
1
Aspect ratio
11.7
11.1
12
12.1
11
11
2
Wing Area (m2)
75.26
54.5
61
65
81.9
55.74
3
W (Kgf)
21800
18600
21500
24800
23500
22800
19150
4 W/S (Kg/m2) 289.7 341.3 352.5 381.5 286.9 409
5
Wing location
High Mounted
High Mounted
High Mounted
Low Wing
Low Wing
High
6
Aerofoil
NACA 43 Series
MS 0317
(MS0313)
7
t/c
18 % (At root)
17% (At root)
16%
8 Sweep 0 0
9
Taper Ratio
0.313
0.548
0.618
0.518
0.524
10 Twist (Degree) 3
11 Wing incident (i) Deg
3 (At root)
2 (At root)
2 (At root)
2
2 (At root)
12
Dihedral Angle
2.5 deg
2.5 deg
3
7
6
13
High Lift device
Fowler Flaps Single and
double slotted
Double slotted
flaps
Double slotted
flaps
Double slotted
fowler flaps
Double
Slotted Flaps
Single Slotted Flaps
Table: 4.1 Comparison of different parameters of similar aircrafts
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5.Fuselage And Tail Sizing Fuselage sizing
lf = aWoC
For a twin turbo prop (From Raymer Ref.2 book)
a= 0.169
C=0.51
We have Wo= 23500 Kg Therefore lf = (0.169)*(23500)0.51 =28.65 m
Therefore the length of the fuselage is 28.65 m.
Length of Nose
=0.03 therefore lnose= 0.86m
lcockpit = 2.5 m . It is standard for a 2 pilot cockpit.
Cabin Length:
Economic classs. No. of passengers = 48 (12 rows)
Business class. No. of passengers = 12 (3 rows)
Parameter Economy class Business class
Seat pitch (inches) 32 38
Seat width (inches) 22 28
Aisle width (inches) 22 56
Seat abreast 4 2
No. of aisle 1 1
Max height (m) 2.2 2.2
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Class No. of seats No. of rows Seat pitch Cabin length (m)
Economy 52 13 32 10.56
Business 8 4 38 3.86
Cabin diameter
df(internal) = (22*4 +22*1)*2.54/100 = 2.8
m t = (110)(0.02) +1”
= 3.2 inch
= 0.0813 m
External diameter of the fuselage = 2.8+0.0813*2 = 2.96 m
Rear fuselage
= 0.25
= 0.25* 28.65 = 7.1625 m
Total fuselage length (m) Nose length = 0.86
Cockpit = 2.5
Passenger seating = 14.42
Rear fuselage = 7.162
Toilets, and other = 1.3
Total length = 28.5
Cargo door = 1.3
Tail sizing
1. Aspect ratio
Aspect ratio of horizontal wing Ah = 5.5
Aspect ratio of vertical tail Av = 1.5
2. Area ratio
Sh/S = 0.25 Sv/S= 0.21
Sh= 16 m2 Sv= 13.44 m2
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3. Span
bh= 9.4 m
bv= 4.5 m
Crh= 2.27 m
Cth=1.13m
Crv= 3.73 m
Ctv=2.24 m
4. Engine location
The type of Engine mounting and it’s location play a major role in deciding the overall
drag coefficient of the airplane. A conventional wing mounted engine is chosen as
it facilitates periodic maintenance in an industry where an
attached to the lower side of the wing using pylons to reduce drag. The other reason for
choosing a wing mounted engine is the fuel is stored in the wings itself, thereby
reducing the length of the fuel line. From the data collection of similar airplanes, the
engineFrom referring to the data of similar airplanes the engines are located at 33.4 % of
wing semi-span.
5. Landing gear
Hydraulically retractable tricycle type, nose unit retract forward, main units
inward into fuselage and large under fuselage fairing.
Minimum turning radius on ground is 17.08m
Wheel base 12 m
Track 4.10 m
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Cargo Door Business class Economy class Passenger door
Fig 5.1 Cabin Layout
(The design of ATR-72 aircraft is considered for Cabin layout.)
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6. Estimation of component weights and CG
location Aircraft weight is a common factor which links different design activities (aerodynamics,
structures, propulsion, layout, airworthiness,environmental, economic and operational
aspects).To this end, at each stage of the design, a check is made on the expected total mass
of the completed aircraft. A separate design organization(weights department)is employed
to assess and control weight. In the preliminary design stage, estimates have to made from
historical statistical data of all the component parts of the aircraft from similar airplanes. As
parts are manufactured and the aircraft prototype reaches completion it is possible to check
the accuracy of the estimates by weighing each component and where necessary instigate
weight reduction programmes.
6.1 Aircraft mass statement
The weight of the entire airplane can be sub-divided into empty weight and useful
load. The empty weight can be further subdivided into-
• Structures group
• Propulsion group
• Equipment group
DCPR(Defense Contractor Planning Report) weight is taken as the weight obtained
after deducting weights of wheels, brakes, tires, engines, starters, batteries, equipments,
avionics etc from the empty weight. DCPR weight is important for cost estimation, and can
be viewed as the weight of the parts of the airplane that the manufacturer makes as
opposed those of items bought and installed.
It has become normal practice in aircraft design to list the various components of
aircraft mass in a standard format.
The components are grouped in convenient subsections as shown below.
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6.1.1 Structures Group
1. Wing(including control surfaces)
2. Tail(horizontal and vertical including controls)
3. Body(or fuselage)
4. Nacelles
5. Landing gear (main and nose units)
6. Surface controls
6.1.2 Propulsion Group
1. Engine(s)(dry weight)
2. Accessory gearbox and drives
3. Induction system
4. Exhaust system
5. Oil system and cooler
6. Fuel system
7. Engine controls
8. Starting system
9. Thrust reversers
6.1.3 Fixed equipment group 1. Auxiliary power unit
2. Flight control systems (sometimes included in structural group)
3. Instruments and navigation equipment
4. Hydraulic systems
5. Electrical systems
6. Avionics systems
7. Furnishing
8. Air conditioning and anti-icing
9. Oxygen system
10. Miscellaneous (e.g. fire protection and safety systems)
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6.2 weights of various components
After making the classification between various groups and listing the
components in each group, we next proceed to determine the weights of these
components. In the preliminary design stages it is not possible to know the size of
individual aircraft components in great detail but it is possible to use prediction
methods that progressively become more accurate as the aircraft geometry is
developed. Most aircraft design textbooks contain a set of equations empirically
derived based on existing aircraft. For the present design, we choose to follow
equations prescribed in Appendix 8.1 of [5]. Using these equations, the weights of
various individual components are calculated.
6.3 C.G Location and C.G Travel
6.3.1 Wing Location on Fuselage
The wing longitudinal location is decided based on the consideration the C.G of the
entire airplane with full payload and fuel is around the quarter chord of the m.a.c. We
tabulate the weights and the corresponding C.G locations of various components and then
apply moment equilibrium about the nose of the airplane in order to solve for Xl.e (the
distance of leading edge of root chord of the wing from the nose).In tabulating the results,
we assume that the C.G locations of wing, horizontal tail and vertical tail are at 40% of the
respective m.a.c. The fuselage C.G is taken to be at 40% of it’s length. The engine C.G
location was taken to be at 50% of it’s length. All other components were taken to have a
net C.G location at 40% of the fuselage length. The tabulated values are given below. The
nose wheel was placed referring to ATR-72 aircraft, main landing gear position was
determined based on load distribution.
• Using data for equivalent trapezoidal wing in section 4, the location of wing c.g. is at 5.34
m behind the leading edge of the root chord. The quarter chord of m.a.c is at
4.76 m behind the leading edge of root chord.
• Noting that the tail arm is 14.85 m and that the c.g of tail is 15 % behind the a.c., the
distance of horizontal tail c.g. from leading edge of root chord of wing is 20.05 m. In a
similar way, c.g. of vertical tail is at 19.56 m behind leading edge of the root chord of
wing
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Components Wt (Kgf) CG (m) Wt*CG (Kgf*m) % Wt
Wing 2678.34 12.92 34613.15 11.40
Fuselage 3290.00 11.40 37506.00 14.00
HT 432.00 25.72 11112.92 1.84
VT 362.88 25.07 9097.90 1.54
Engine 1170.00 11.91 13930.68 4.98
Fixed Equipment 3995.00 11.40 45543.00 17.00
Nose Wheel 167.00 1.74 290.41 0.71
Main Landing Gear 949.93 13.74 13051.09 4.04
Pay Load 6528.00 12.36 80686.08 27.78
Fuel 3926.85 12.56 49319.70 16.71 295150.93
Table 6.1 Weights and CG location
By applying moment equilibrium about the nose of the airplane, we obtain
location of wing leading edge at the root to be 11.74 m from the nose of the airplane.
The C.G of the airplane lies at 12.56 m from the nose.
6.4 C.G Travel for Critical Cases
6.4.1 Full Payload and No Fuel
For the case of full payload and no fuel, the fuel contribution to the weight is
not present. However, since we have assumed that the c.g of the fuel to be at the
quarter chord of the m.a.c of the wing (where the c.g of the entire airplane has been
positioned) there will be no c.g shift in this case.
Hence, the C.G shift is 0%.
6.4.2 No Payload and No Fuel
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For this case, the fuel as well as the payload contribution are not present. Since the
c.g of payload is not at the c.g of the entire airplane, the c.g is bound to shift by a
certain amount in this case. Calculations showed that, the C.G shift is 4%.
6.4.3 No Payload and Full fuel
For this case, since there is no payload, the c.g is bound to shift. On performing
calculations, we obtain the new c.g location. The shift in CG is about 3.13 % 6.5 Summary
• Wing location(leading edge of root of trapezoidal wing) – 11.74 m
• c.g location with Full payload and full fuel - 12.56 m
• c.g travel for No Payload and No Fuel – 4.0%
• c.g travel for No Payload and full Fuel – 3.13%
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7 Control Surfaces
7.1 Stability and Controllability The ability of a vehicle to maintain its equilibrium is termed stability and the influence which the pilot or control system can exert on the equilibrium is termed its controllability. The basic requirement for static longitudinal stability of any airplane is a negative slope of the curve of the pitching moment coefficient, Cmcg, versus lift coefficient, CL. Dynamic stability requires that the vehicle be not only statically stable, but also that the motions following a disturbance from equilibrium be such as to restore the equilibrium. Even though the vehicle might be statically stable, it is possible that the oscillations following a disturbance might increase in magnitude with each oscillation, thereby making it impossible to restore the equilibrium.
7.2 Static Longitudinal Stability and Control
7.2.1 Specifications
• The horizontal tail must be large enough to insure that the static longitudinal
stability criterion, dCmcg/dCL will be negative for all anticipated center of gravity
positions.
• An elevator should be provided so that the pilot will be able to trim the airplane(maintain
Cm = 0) at all anticipated values of CL.
• The tail should be large enough and its elevator powerful enough to enable the pilot rotate
the airplane during the take-off run to the required angle of attack. This condition is
termed as the Nose wheel Lift-off condition.
7.2.2 Aft Center of gravity limit
For the “stick free” case and for small angles of attack ,the following expression for the aft
center of gravity limit in terms of the tail-size parameter, V we have the following equation.
The value of xc.g from above equation is termed the “stick-free neutral point”, since it is the
c.g location at which the static stability is neutral.
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7.2.3 Forward center of Gravity Limit
The forward c.g. limit is not generally dependent on maintaining stability. As the c.g
is moved forward ,the stability contribution xc.g −xa.c of the wing becomes more and
more negative ,thereby increasing the static stability. In order to keep the airplane in
equilibrium as the c.g is moved forward, the elevator must be capable of trimming out the
resulting negative pitching moment. The pitching moment will be the greatest when the
airplane is at CLmax when the airplane is landing and ground effects decrease the
downwash at the tail.
aw=5.62 /radian = 0.098 /degree
from the value obtained in section 4 on wing design.
From Raymer Ref.2 Fig 16.4, M=0.5
Cmα = -0.7 / radian
Therefore,
Referring to Nelson book
We get Cmα = -0.68 / radian for Navion
Cmα = -0.78 / radian for STOL
Therefore the Cmα = -0.7 / radian in appropriate.
Cmα (fuse) =0.223 for Navion.
For Jet planes
Therefore for turboprop aircraft can be considered.
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= 0.52 from Perkins and Hage
Horizontal Tail
Aspect Ratio : 5.5
Landing Mach number= 0.17 β2 =1 - M2 = 0.971, A=5.5
From Equation in section 4,
aw Land =4.354/ radian = 0.076 /deg
-0.125=0.06+0.06-(0.076/0.098) X V X 1 X (1-0.2979) X (1-(-0.0066/-0.0114)X0.52) V =
0.6438
V = therefore
The value is closer to the value from similar aircrafts.
We can take
Therefore Sht= 12.81 m2
Vertical Tail
Aspect ratio= 1.5
From Equation in section 4,
at Land =2.064 / radian = 0.036/deg
Cnβ desirable =0.08 / radian = 0.0014 /deg
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Cnβ (fuselage) = =0.0025
-0.0014=0+0.0025-0.036 V V
= 0.102
V = therefore
The value is closer to the value from similar aircrafts. We
can take
Therefore Svt= 13.4 m2
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8. Performance Estimation
The details regarding overall dimensions, engine details, weights, geometric
parameters of wing, fuselage, horizontal tail, vertical tail, vertical tail and other
details like CLmax in various conditions and maximum load factor are given in section
8.2 - 8.10. The details of flight condition for estimation of drag polar are as follows
Altitude : 4500 m = 14760 ft
Mach number : 0.45 Kinematic Viscosity : 2.12355 ×10−5 m2/s Density : 0.777 kg/m3
Speed of Sound : 322.6 m/s
Flight Speed : 138.9 m/s
Weight of the Airplane : 23500 kgf
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Wing Loading Calculations
The power-to-weight ratio (P/W) and the wing loading (W/S) are the two most
important parameters affecting aircraft performance. Optimization of these
parameters forms a major part of the design activities conducted after initial weight
estimation. For example, if the wing loading used for the initial layout is low, then the area
would be large and there would be enough space for the landing gear and fuel tanks.
However it results in a heavier wing.
Wing loading and thrust-to-weight ratio are interconnected for a number of critical
performance items, such as take-off distance, maximum speed etc. These are often the
design drivers. A requirement for short takeoff can be met by using a large wing (low W/S)
with a relatively low P/W. On the other hand, the same takeoff distance could be met with
a high W/S along with a higher P/W.
In this section, we use different criteria and optimize the wing loading and thrust
loading. Wing loading affects stalling speed, climb rate, takeoff and landing distances,
minimum fuel required and turn performance.
Similarly, a higher thrust loading would result in more cost which is undesirable.
However it would also lead to enhanced climb performance. Hence a trade-off is needed
while choosing W/S and P/W. Optimization of W/S and P/W based on various
considerations is carried out in the following subsections.
1. Landing Distance Consideration
SLand = 1200 m
Initially landing distance was considered as 900m. But it is found that for all similar
aircrafts the landing distance is around 1200 m . so it is decided to consider
landing distance as 1200m
Va = SLand ( feet)
0.3
Va = (1200)(3.28)
0.3
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2
V a = 114.54 K nots V a = 5 9 .0 4 m / s Stall Speed , Vs = Va/1.3 = 59.04/1.3
Vs = 45.4 m/s
Now (W/S)Land = 1 ρ σV 2C
2 0 s L max
1 (1.225) * (45.4)2 * 2.7 2
(W/S)Land = 3408.64 N/m2
WLand/WTO = 0.97
(W/S) TO = 3514.065 N/m2
With 10% Variation in Vs
2761.0 N/m2 < (W/S)Land < 4124.46 N/m2
2. Maximum Speed Consideration
Vmax = 1.1 Vcr Vcr = 500 Kmph
Vmax = 1.1*500
Vmax = 550 Km/h = 152.8 m/s
C D =
K =
C D 0
1
+ K C L
CD 0
π Ae = C fe
Swet
Sref
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2
2
A Swet
Sref
= 12 5.2
= 2.3
1
= 1 + 1 + 0.05
e ewing e fuge
1 = 1 + 1 + 0.05 e 0.84 0.1
e = 0 .7 4 6
K = 1 π *12 * 0.746
K = 0.03555
CD 0 = 1
4K (L D )
2 Max
CD 0 = 1
4 * 0.03555 (17)2
CD 0 = 0.02433
C D = C D 0 + K C L
CD = 0.02433 + 0.03555CL
This is the Drag polar for the aircraft under design
CD 0 = C fe * Swet
Sref
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2 2
0.02433 = C fe * 5.2
C fe = 0.004679
F1 = 0.004679 *1.87 * (1 + 0.25 + 0.21)
F1 = 0.01277
F = CD 0 − F1 2 W
S
F2 = 0.02433 − 0.01277 3440
F = K
3 q2
F2 = 3.36046E − 6
F3 = 0.03555
( 1 * 0.77 *152.82 )
F3 = 4.39972 E − 10
F1 Popt =
F3
Popt = 0.01277
4.39972 E − 10
Popt = 5387.449 N / m2
tv max = 0.07282
Considering +5% of tvmax (0.076461)
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3573.65 N/m2 < W/S < 8121.815 N/m2
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CL =
2W ρ SV 2
C = 2 * 5387.446
L 0.77 *152.82
CL = 0.59934
C D = 0.02433 + 0.03555*0.59934 2
CD = 0.03709986
T = WCD
CL
T = 23500 * 9.81* 0.03709986 0.59934
PRe q =
T = 14 270 .3 9 N TV
1000
PRe q = 14270.39 *152.8
1000
PRe q = 2180.515 Kw (Total power)
To convert it to sea level static thrust
PRe q = 2180.515
= 3634.19Kw 0.6 0.6
Power per engine is 3634.19 = 1817.1 kw
2
P = 3634.19 W 23500 *9.81
= 0.01576
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)
3. (R/C)Max Consideration
(CL
)min P =
3CD 0
K
(CL )min P =
3 * 0.02433 0.03555
(CL )
V
min P =
= 1.432886
2W S
min P ρ (CL
min P
Vmin P = 2 * 3440
1.225 *1.432886
Vmin P = 62.61m / s
q = 1 *1.225 * 62.612
2
q = 2401.01
F = K
3 q2
F = 0.03555
3 2401.012
F3 = 6.16668E − 9
F1 Popt =
F3
Popt = 0.01277
6.16668E − 9
Popt = 1439.03N / m2
Since the value is not appropriate, it is not considered
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4. Based on the Range consideration (R) Vcr =500 Kmph = 138.9 m/s
q = 1 * 0.77 *138.92
2
q = 7427.886
F = K
3 q2
F = 0.03555
3 7427.8862
F3 = 6.4433E − 10
F1 Popt =
F3
Popt = 0.01277
6.4433E − 10
Popt = 4451.86N / m2
2997.12 N/m2 < W/S < 6612.69 N/m2
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Summary of above considerations.
Popt (N/m2) Wing loading N/m2) power loading
S land consideration 3514 2761- 4125
Vmax consideration 5387.4 3573.6- 8121.8 0.01576
Range consideration 4452 2998- 6613
R/C consideration
Take off consideration 0.08112
The wing loading is chosen as 3600 N/m2
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ENGINE CHARACTERISTICS Fig. 9.1 to 9.6 shows the engine characteristics at different altitude during different flight conditions.
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Fig 9.1
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Fig 9.3
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Fig 9.6
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9 features of the designed aircraft
9.1 Three View Drawing The 3-view drawing of the airplane designed is given in figure
9.2 Overall Dimensions
Length : 34.32
Wing Span : 27.72 m
Wheel base : 12.0 m
Wheel track : 4.1 m
9.3 Engine details
Pratt and Whitney PW 123D Engine
Maximum SHP TO 1604 kW
Maximum continuous power 1454 kW
8.4 Weights
Gross Weight : 23500 kgf
Empty Weight : 13045 kgf
Fuel Weight : 3927 kgf
Payload : 6528 kgf
9.5 Wing Geometry
Planform Shape : Tapered wing with no sweep
Span : 27.72 m Area : 64.04 m2
Airfoil : NACA - 653618, t/c = 18%, Clopt = 0.6
Root Chord : 3.3 m
Tip Chord : 1.32 m
Mean Aerodynamic Chord : 2.451 m
Quarter chord Sweep : 0 deg
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Dihedral : 3 deg
Twist : 0 deg Incidence : 2 deg
Taper Ratio : 0. 4 (Equivalent Trapezoidal wing)
Aspect Ratio : 12
9.6 Fuselage Geometry
Length : 28.5 m
Maximum Diameter : 2.96 m
9.7 Horizontal Tail Geometry Span : 9.4 m
Area : 12.81 m2
Mean Aerodynamic Chord : 1.74 m
Root Chord : 1.81 m
Tip chord : 0.91 m
Taper Ratio : 0.5
Aspect Ratio : 5.5
9.8 Vertical Tail Geometry
Span : 4.5 m
Area : 13.44 m2
Root Chord : 4.26 m
Tip chord : 1.7 m
Mean Aerodynamic Chord : 2.15 m
Quarter Chord Sweep : 0 deg
Taper Ratio : 0.4
Aspect Ratio : 1.5
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9.9 Other details
CLmax without flap : 1.5
CLmax with landing flaps : 2.7
9.10 Crew and Payload
Flight crew : 2 (pilot and co-pilot)
Cabin crew : 2
Passenger seating : 52 economy and 08 business class
9.11 Performance The detailed performance estimation is given in section 9. The highlights are as
follows.
• The performance is worked for a gross weight of 23500 kgf and wing
loading of 3600 N/m2 except for landing where the landing weight is taken
as 85% of take-off weight.
• Maximum Mach No. at 10000 ft Mmax = 0.45
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Fig 8.1 Three view drawing of Aircraft under design
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BIBLIOGRAPHY References:
1) Dr. E G Tulapurkara, A. Venkattraman, V. Ganesh, “ An Example of Airplane
Preliminary Design Procedure- Jet Transport, AE TR 2007-4, April 2007
2) Raymer .D.P. Aircraft design a conceptual approach.
AIAA’ educational series, 2006
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of Aerospace Engineering I.I.T Madras, 2008
4) Roskam J. Methods of estimating drag polars of subsonic airplanes
Roskam Aviation & Engineering Corporation, Ottawa, Kansas,1983
5) Jenkinson L.R., Simpkin P. and Rhodes D.
Civil Jet Aircraft Design, Arnold, 1999
6) Wood K.D. Aerospace vehicle design, Volume 1, Johnson
publishing company, Boulder, Colorado, 1966
7) Perkins C.D. & Hage A.E. Airplane performance stability & control,
McGraw Hill, 1963
8) Abbot I.H. and Doenhoff A.E. Theory of wing sections,
Dover publications, 1959
9) Roskam J. Aircraft design,
Roskam Aviation & Engineering Corporation, Ottawa, Kansas, 1990
10) Kroo, Ilan & Shevel, Richard - Aircraft Design, Synthesis and Analysis
11) Lloyd R. Jenkinson, James F. Marchman III
Aircraft Design Projects For Engineering Students
12) BRUHN- Airplane Design
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INTERNET REFERENCES 1. www.wikipedia.org
2. www.cessena.com
3. www.diamondaircraft.com
4. www.airliners.net
5. www.boeing.com
6. www.google.com
7. www.scribd.org
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