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AIRCRAFT DESIGN PROJECT-1 150 SEATER PASSENGER AIRCRAFT
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ACKNOWLEDGEMENT
I would like to extend my heartful thanks to Prof. ASHOK (Head of Aeronautical Department) for giving me his able support and encouragement. At this juncture I must emphasis point that this DESIGN DESIGN PROJECT PROJECT would not have have been po without the highly informative and valuable guidance by Prof.SARVESWARAN, whose vast knowledge and experie has must us go about this project with great ease. We have gr pleasure in expressing expressing our sincere sincere & whole whole hearted gratitude gratitude them. It is worth mentioning about my team mates, friends and colleagues of the Aeronautical department, for extending thei kind help whenever the necessity arose. I thank one and all w have directly or indirectly helped me in making this design project a great great success.
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ACKNOWLEDGEMENT
I would like to extend my heartful thanks to Prof. ASHOK (Head of Aeronautical Department) for giving me his able support and encouragement. At this juncture I must emphasis point that this DESIGN DESIGN PROJECT PROJECT would not have have been po without the highly informative and valuable guidance by Prof.SARVESWARAN, whose vast knowledge and experie has must us go about this project with great ease. We have gr pleasure in expressing expressing our sincere sincere & whole whole hearted gratitude gratitude them. It is worth mentioning about my team mates, friends and colleagues of the Aeronautical department, for extending thei kind help whenever the necessity arose. I thank one and all w have directly or indirectly helped me in making this design project a great great success.
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INDEX Serial No. Topic
Page No.
1 Aim of the Project 5 2 Abstract 7 3 Introduction 9 4 Comparative DataSheet 16 5 Graphs 20 6 Mean Design Parameters Parameters 39 7 Weight Estimation 41 8 Powerplant Selection 49 9 Fuel Weight Validation 53 10 Wing Selection 55 11 Airfoil Selection 60 12 Lift Estimation 70 13 Drag Estimation 75 14 Landing Gear Arrangement Arrangement 81 15 semesterFuselage Design 87 Master your with Scribd Read Free Foron 30this Days Sign up to vote title 16 Performance 94 & The New York Times Useful Not useful Characteristics Special offer for students: Only $4.99/month. 17 3 View Diagram 100 Cancel anytime.
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ABBREVIATION A.R. - Aspect Ratio B - Wing Span (m) C - Chord of the Airfoil (m) C root - Chord at Root (m) C tip - Chord at Tip (m) - Mean Aerodynamic Chord (m) C Cd - Drag Co-efficient Cd,0 - Zero Lift Drag Co-efficient Cp - Specific fuel consumption (lbs/hp/hr) CL - Lift Co-efficient D - Drag (N) E - Endurance (hr) E - Oswald efficiency L - Lift (N) (L/D)loiter - Lift-to-drag ratio at loiter l oiter (L/D)cruise - Lift-to-drag ratio at cruise Master your semester withofScribd M - Mach number aircraft Read Free Foron 30this Days Sign up to vote title & The New York Timesfuel fraction Useful Not useful Mff Mission Special offer for students: Only $4.99/month. R - Range (km) Cancel anytime.
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Sfr - Free roll Distance (m) Sg - Ground roll Distance (m) T - Thrust (N) Tcruise - Thrust at cruise (N) Ttake-off - Thrust at take-off (N) (T/W)loiter - Thrust-to-weight ratio at loiter (T/W)cruise - Thrust-to-weight ratio at cruise (T/W)take-off - Thrust-to-weight ratio at take-off Vcruise - Velocity at cruise (m/s) Vstall - Velocity at stall (m/s) Vt - Velocity at touch down (m/s) Wcrew - Crew weight (kg) Wempty - Empty weight of aircraft (kg) Wfuel - Weight of fuel (kg) Wpayload - Payload of aircraft (kg) W0 - Overall weight of aircraft (kg) W/S - Wing loading (kg/m²) ρ - Density of air (kg/m³) μ- Dynamic viscosity (Ns/m²) λ - Tapered ratio Master your semester with Scribd Read Free Foron 30this Days Sign up to vote title R/CYork - Rate of Climb & The New Times Useful Not useful Special offer for students: Only $4.99/month.
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AIM OF THE PROJECT
The aim of this design project is to design a150 seater passenger aircraft by comparing the data and and specifications of prese aircrafts in this category and to calculate the performance characteristics. Also necessary graphs need to be plotted and diagra have to be included wherever needed.
The following design requirements and research studies are set for project: Design an aircraft that will transport 150 passengers and th baggage over a design range of 4820 km at a cruise speed about 890 km/h. To provide the passengers with high levels of safety and Master your comfort. semester with Scribd
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To offer a unique and competitive service to existing schedule operations.
To assess the development potential in the primary role of the aircraft. To produce a commercial analysis of the aircraft project.
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ABSTRACT
The purpose of the project is to design a 150 seater Medium Range International passenger aircraft. The aircraft will possess a low win tricycle landing gear and a conventional tail arrangement. Such an aircraft must possess a wide body configuration to provide sufficie seating capacity. It must possess turbofan engines to provide the required amount of speed, range and fuel economy for the operator The aircraft will possess two engines. You're Reading a Preview Unlock full access with a free trial.
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INTRODUCTION
At the instant time there are different types of aircraft with latest technology. Every year there is a great competition for making an aircraft of having higher capacity of members inside the aircraft. So here in thi You're Reading a Preview report, we intend to implant the differentiation among Unlock full access with a free trial. the aircrafts having sitting capacity of 100-180 members. This reportDownload givesWith theFreedifferent aspects of Trial specifications like wing specification, weight specification, power plant specification and performance specification. Airbus started the development of a very large airline Master your semester withby Scribd (termed Megaliner Airbus in the early developme Read Free Foron 30 Days Sign up to vote this title & The New York Times Useful Not useful stages) in the early 1990s, both to complete its own Special offer for students: Only $4.99/month. range of products and to break the dominance that Cancel anytime.
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only one new aircraft to be profitable in the 600 to 80 seat market segment. Each knew the risk of splitting such a niche market, as had been demonstrated by the simultaneous debut of the Lockheed L-1011 and the McDonnell Douglas DC-10: both planes met the market’s needs, but the market could profitably sustai only one model, eventually resulting in Lockheed's departure from the civil airliner business. In January 1993, Boeing and several companies in the Airbus consortium started a joint feasibility study of an aircra known as the Very Large Commercial Transport (VLCT), aiming to form a partnership to share the limited market. Airplanes come in many different You're Reading a Preview shapes and sizes depending on the mission of the Unlock full access with a free trial. aircraft, but all modern airplanes have certain Download With Free Trial components in common. These are the fuselage, wing tail assembly and control surfaces, landing gear, and powerplant.
For any airplane to fly, it must be able to lift the weig Master your semester with Scribd Read Free Foron 30this Days Sign up to vote title of the airplane, and the cargo. & The New York Times its fuel, the passengers, Useful Not useful generate most of the lift to hold the plane i Special offer forThe students:wings Only $4.99/month. the air. To generate lift, the airplane must be pushed Cancel anytime.
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The fuselage is the body of the airplane that holds all the pieces of the aircraft together and many of the oth large components are attached to it. The fuselage is generally streamlined as much as possible to reduce drag. Designs for fuselages vary widely. The fuselage houses the cockpit where the pilot and flight crew sit and it provides areas for passengers and cargo. It may also carry armaments of various sorts. Some aircraft carry fuel in the fuselage; others carry the fuel in the wings. In addition, an engine may be housed in the fuselage. You're Reading a Preview Unlock full access with a free trial.
The wing provides the principal lifting force of an Download With Free Trial airplane. Lift is obtained from the dynamic action of t wing with respect to the air. The cross-sectional shape of the wing as viewed from the side is known as the airfoil section. The planform shape of the wing (the shapesemester of the wing viewed from above) and placeme Master your withasScribd Read Free Foron 30this Days Sign up to vote title of the wing on the fuselage (including the angle of & The New York Times Useful Not useful as well as the airfoil section shape, depend Special offer forincidence), students: Only $4.99/month. upon the airplane mission and the best compromise Cancel anytime.
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The control surfaces include all those moving surface of an airplane used for attitude, lift, and drag control. They include the tail assembly, the structures at the re of the airplane that serve to control and maneuver the aircraft and structures forming part of the tail and attached to the wing.
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PURPOSE AND SCOPE OF AIRPLANE DESIGN OBJECTIVES
To meet the FUNCTIONAL, OPERATIONAL a SAFETY requirements set out OR acceptable to t USER.
ACTUAL PROCESS OF DESIGN
Selection of aircraft type and shape Determination of geometric parameters Selection of power plant You're Reading a Preview Structural design and analysis of various Unlock full access with a free trial. components With Free Trial Determination ofDownload aircraft flight and operational characteristics .
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DESIGN CYCLE PRELIMINARY DESIGN
It consists of the initial stages of design, resulting in t presentation of a BROCHURE containing prelimina drawings and clearly stating the operational capabiliti Reading a Preview of the airplane beingYou're designed. This Brochure has to b full access with a free trial. APPROVED by the Unlock manufacturer and/or the custome The steps involved: Download With Free Trial Layout of the main components Arrangement of airplane equipment and control systems Selection of power plant Master your semester with Scribd Read Free Foron 30this Days Sign up to vote title Aerodynamic and stability calculations & The New York Times Useful Not useful Special offer for students: Only $4.99/month. structural design of MAJOR Preliminary Cancel anytime.
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DESIGN PROJECT Internal discussions Discussions with prospective customers You're Reading a Preview Discussions with Certification Authorities Unlock full access with a free trial. Consultations with suppliers of power plant and Download With Free Trial major accessories Deciding upon a BROAD OUTLINE to start the ACTUAL DESIGN, which will consist of Construction of Mock-up Structural layout of all the individual units, and Master your semester with Scribd Read Free Foron 30this Days Sign up to vote title theirTimes stress analysis & The New York Useful Not useful Special offer for students: Only $4.99/month. Drafting of detailed design drawings
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(a) Conceptual Design (b) Preliminary Design (c) Detail Design
CONCEPTUAL DESIGN:
Conceptual design is a very fluid process. Ne
ideas and problems emerge as a design is investigat
in ever increasing detail. Each time the latest design You're Reading a Preview
full access with trial. analyzed and sized, Unlock it must bea freeredrawn to reflect t
Download With Free Trial new gross weight, fuel weight, wing size, engine siz
and other changes.
Conceptual design will usually begin w
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by the prospective customer or a company generat
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The actual design effort usually begins w
conceptual sketch. A good conceptual sketch w
include the approximate wing and tail geometries, t
fuselage shape, and the internal locations of the ma components
such
as
the
engine,
cockp
payload/passenger compartment, landing gear a fuel tanks. You're Reading a Preview Unlock full access with a free trial.
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PRELIMINARY DESIGN:
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configuration. During this design the specialists in are
such as structures, landing gear, and control system
will design and analyze their portion of the aircra
Testing is initiated in areas such as aerodynami propulsion, structures, and stability and control.
A key activity during this type of design
“LOFTING’. Lofting is the mathematical modeling of t outside skin of the aircraft with sufficient accuracy You're Reading a Preview
insure proper fit between its different parts, even Unlock full access with a free trial.
Download With Free designers Trial they are designed by different and possib
fabricated in different locations. The ultimate objecti
during this design is to ready the company for t
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Assuming a favorable decision for enterin
full-scale development, the detail design phase begin 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 You're Reading a Preview
analyzed.
Unlock full access with a free trial.
Download With Free Trial Another important part of detail design
called production design. Specialists determine ho
the airplane will be fabricated, starting with smalle
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During detail design, the testing effo
intensifies. Actual structure of the aircraft is fabricat
and tested. Control laws for the flight control syste
are tested on an “iron-bird” simulator, a detail
working model of the actuators and flight cont surfaces. Flight simulators are developed and flown both company and customer test pilots.
Detail design ends with fabrication of t You're Reading a Preview
aircraft. Frequently Unlock thefull fabrication begins on part access with a free trial.
With Free Trial the aircraft before Download the entire detail-design effort
completed.
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NAME OF AIRCRAFT
Boeing
Boeing C-40 Boeing
737-100
Clipper
737-200
CAPACITY
124
LENGTH (M)
28.65
WING SPAN (M)
28.35
HEIGHT (M)
11.23
WING AREA(m^2)
102
102
THRUST (kN)
64
77
EMPTYWEIGHT(kg)
28100
31600
121 33.3 34.2 12.55
136
117
30.53
37.7
28.35
28.4
11.23
8.87
82.3
57150 MAX
TAKE
OFF 50300 You're Reading a Preview
WEIGHT
52400
4990
10700
2645
Unlock full access with a free trial.
78000
Download With Free Trial SERVICE SEILING (m) 10670 12500 RANGE (km) 2850 5600 ASPECT RATIO 8
4300 8
8.5
ENDURANCE
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NAME OF AIRCRAFT
AIRBUS
BOEING
Boeing
A 318-100 737-300
737-500
CAPACITY
132
149
132
140
LENGTH (M)
31.44
33.414
31.008
31.2
WING SPAN(M)
34.1
28.9
28.9
35.8
HEIGHT(M)
12.51
11.15
11.15
12.5
90
90
101
EMPTY WEIGHT(Kg) 39500
32700
31300
3637
MAX
62800
60550
6600
11277
1250
4200
4444
5970
9.11
9.46
9.45
WING AREA(M^2) 112.6 THRUST( kN)
TAKE
106
OFF 68200
You're Reading a Preview
WEIGHT (Kg)
Unlock full access with a free trial.
SERVICE SEILING(M) 11887 RANGE
5700
ASPECT RATIO
10
11277
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Comparative Datasheet – 3 NAME OF AIRCRAFT
Boeing
Boeing
ANTONAV
COM
737-700
717-200
AN-10
ARJ
CAPACITY
148
117
100
105
LENGTH (M)
33.63
37.8
34
36.3
WING SPAN (M)
35.8
28.47
38
36.3
HEIGHT (M)
12.55
8.92
9.8
8.44
84.5
121
80
WING AREA (M^2) THRUST (kN)
117
82.3
82.3
EMPTY WEIGHT (Kg) 38147 30618 You're Reading a Preview MAX
TAKE
65700
2630
Unlock full access with a free trial. OFF 66000 49900
4361
Download With Free Trial
WEIGHT (Kg)
SERVICE SEILING (M) 12500
11000
11000
1190
RANGE (kM)
6370
2645
2532
2200
ASPECT RATIO
8
7.8
7
7.9
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Comparative Datasheet – 4 NAME OF AIRCRAFT
FOKKER100 FOKKER100 Boeing TAY620
TAY650
707-020
CAPACITY
122
122
140
179
LENGTH (M)
35.53
35.53
41.25
44.0
WING SPAN (M)
28.08
28.08
39.9
39.9
HEIGHT (M)
8.5
8.5
12.65
12.9
46785
5558
100800
1165
WING AREA (M^2) 93.5
93.5
THRUST (kN)
67.2
61.6
EMPTY WEIGHT(Kg) 24375 24541 You're Reading a Preview MAX
TAKE
Unlock full access with a free trial. OFF 43090 45810
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WEIGHT (Kg) SERVICE SEILING(M)
11000
11000
RANGE (KM)
2450
3170
7040
8704
ASPECT RATIO
8.5
8.5
11
11
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NAME OF AIRCRAFT
Boeing
Boeing
Boeing
An
707-320B
727 – 100
727 – 200
CAPACITY
147
149
189
99
LENGTH (M)
46.61
40.6
46.7
28.91
WING SPAN (M)
44.42
32.9
32.9
28.91
HEIGHT (M)
12.93
10.3
10.3
8.6
An
WING ARE (Sq.M)
87.32
THRUST (kN)
67.0
EMPTY WEIGHT (Kg) 66406 MAX
TAKE
OFF 151320
45360
45360
76818
95028
You're Reading a Preview
WEIGHT (Kg)
Unlock full access with a free trial.
SERVICE SEILING Download With Free Trial
RANGE (kM)
10650
5000
4400
ASPECT RATIO
11
9
9
6.9
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COMPARITIVE GRAPHS:
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LENGTH: 3o
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RANGE:4800KM
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ASPECT RATIO:9.3
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DESIGN DATA SHEET
GENERAL CHARACTERISTICS CREW
4
PASSENGER CAPACITY
100-180
LENGTH
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WING SPAN (b)
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WING AREA (S) 2
ASPECT RATIO(b /S)
EMPTY WEIGHT
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27.2 m (89ft 2 inch) 2
80.1 m (862.18 ft 9.3 38,506 kg (84,891
MAX TAKE OFF WEIGHT
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30 m (98ft 5 inch)
lb)(corrected)
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31620 kg (69,710
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THRUST
60 KN (13,488 lbf)
PERFORMANCE CHARACTERISTICS 940 km/h (0.78
MAX SPEED
mach, 584 mph) 890 km/h (0.74
CRUISE SPEED
mach,553 mph) 4,820 km (n 2,995
RANGE SERVICE CEILING
miles) You're Reading a Preview
11,640 m(38,188
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MISSION SPECIFICATION
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PAYLOAD: 150 passengers at 75kg each and 25kg of bagga each.
CREW: 2 pilots and 2 cabin attendants at 200lbs and 30 baggage each respectively.
RANGE: 2502nm, followed by 1 hour loiter followed by 100nm flight to alternate. ALTITUDE: 33,100 feet (For the design range). CRUISE SPEED: M=0.74, at 38,100 ft. You're Reading a Preview
access with aft. freeat trial.Max. W CLIMB: A dir6-ect climbUnlock to full 33,100 TO is desired
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POWERPLANTS: 2 Turbo-fan engines.
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MISSION SPECIFICATION
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PAYLOAD: 150 passengers at 75kg each and 25kg of bag each.
CREW: 2 pilots and 2 cabin attendants at 75kgs and 25k baggage each respectively.
RANGE: 2502nm, followed by 1 hour loiter followed by 100nm flight to alternate. ALTITUDE: 33,100 feet (For the design range). You're Reading a Preview Unlock access with aft. free trial. CRUISE SPEED: M=0.74, at full 38,100
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CLIMB: A direct climb to 33,100 ft. at Max. W TO is desired. POWERPLANTS: 2 Turbo-fan engines.
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WEIGHT ESTIMATION
You're Readingmeet a Previewvery stringent ran Airplane must normally Unlock full access with a free trial.
endurance, speed and cruise speed objectives while carry Download With Free Trial
a given payload. It is vital in predicting the minimum airpla
weight and fuel weight needed to accomplish a giv
mission. We know the lift acquired is directly proportional
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a given aircraft plays a key role in design analysis.
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Payload Weight Validation (W PL):
WPL =
No. of passengers*(Wt.of passenger+Wt.of
baggage) = WPL =
150*(75+25) 15000 kg
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Crew Weight Validation (W CREW):
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WTO(approx)=
38393 kg [From Design data sheet]
Now, taking into consideration the appropriate mission phases:
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Airplane Type
Take Off
Climb
Descent
Landing
Business Jets Transpor t Military Trainers Superson ic Cruise
0.995
0.980
0.990
0.992
0.970
0.985
1.000
0.995
0.990
0.980
0.990
0.995
0.995
0.92-0.87
0.985
0.992
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For takeoff, segment 0-1 historical data’s shows that,
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For climb, segment 1-2 historical data shows that, W2 = 0.985 W1
Aircraft Design Project - 1 For loiter, segment 3-4 ignoring the fuel consumption during descent we assume,
4/3 = 1
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5/4 = 0.995
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The Brequet’s range equation is used to calculate the value o 32. As we all know that maximum range is covered duri cruise we considering this equation,
R = (∞/)(/)ln (2/3) You're Reading a Preview Unlock full access with a free trial.
Download With Free Trial L/D values of similar type of aircrafts we come to know tha the approximate the value of L/D for our aircraft to be 15. So,
/ =15
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V∞ = 872 km/hr
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So now substituting these values in the Brequet’s range equation, R = (∞/(/)(ln 2/3) = (940/.6)(15)ln( W2/W3) ln(w2/w3)=(4820/23500)= 0.225 (w2/w3)=1.25 W3/w2=0.8 You're Reading a Preview Unlock full access with a free trial.
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MFF =
0.796
Fuel Weight Validation (W FUEL):
WFUEL = WFUEL
=
(1-Mff )* WTO appro
(1-0.796)*38393 =
7832.2 kg
Approximate Operational Operational Weight Validation (WOE(approx) ):
WOE(apporx ) = =
WOE(apporx)- WFUEL- WPL 38393-7832-9180
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WE(tent)
=
WOE(approx) -WTFO - WCREW
Where , WTFO = = WE(tent)
0.5% of WTO(approx) 21381-0.005*38393-378 =
20811 kg
Maximum Takeoff Weight Validation (W TO):
WTO =semester WE(tent)with + WFUEL + WPL+ WCREW Master your Scribd Read Free Foron 30this Days Sign up to vote title & The New York Times Useful Not useful =
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Aircraft Empty Weight Validation (WE) can be furthe split up into:
WE =
Wstruc +WPP + WFE
We estimate the Gross Wt. (Wg) =
0.9* WTO =34381 kg
From the Group Weight Jet Transport data, Wstruc =0.321
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Hence
Wstruc = 11061.45 kg
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Hence
WFE = 5828 kg
WPP = 0.114 Wg
Hence WPP = 3921.5 kgYou're Reading a Preview Unlock full access with a free trial.
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Hence the empty weight
WE =
=
Wstruc +WPP + W
11061.45+3921.54+5828.15
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You're Reading a Preview Unlock full access with a free trial.
Download With Free Trial POWERPLANT SELECTION
As estimated from the design data sheet, the aircraft to
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(kg)
RollsRoyce
1501
69.93
61.60
1200
71.2
64.54
1595
69.93
67.16
TAY 620 GE-CF348 RollsRoyce TAY 650
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RollsRoyce
2104
72.97
66.72
BR710-48
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1856
77.02
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54.00
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Hence, the optimum choice of engine, from those listed
above would be the Rolls-Royce TAY 620 engine which mee
the demand of weight and thrust required at similar payloa such as the one under design.
In our design, we select 2 rear engines, namely Rolls-Royce
TAY 620 Turbo Fan Engine to meet the given design standards.
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Manufacturer
ROLLS-ROYCE
Type of the engine
TAY -620 TURBOFAN
Thrust at SL/ISA
61.6 kN (13850 lbs)
Inlet mass flow
184 Kg/s(408 lb/s)
Bypass Ratio
3.04
Overall Pressure Ratio
16.0
Fan diameter
1.17 m(44 inch)
Engine weight
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14220.57 Kg(3135lbs)
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Engine length
2.4 m(94.8 inch)
Turbine entry Temperature
1305 K
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ROLLS-ROYCE TAY620 You're Reading a Preview Unlock full access with a free trial.
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FUEL WEIGHT VALIDATION
The choice of a suitable engine, having been made, it is now possible to estimate the amount of fuel required for a flight the given cruising speed for the given range. Wfuel = ( You're ∗ Reading a Preview Unlock full access with a free trial.
∗∗∗.)/ Download With Free Trial
The factor of 1.2 is provided for reserve fuel. Thrust at altitude is calculated using the relation:
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Altitude = 10800m = 35433ft
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Number of engines = 3
CALCULATION:
Wfuel = 3×7784.2×7200×0.4×1.2 872
Wfuel = 92,553.42 kg
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WING SELECTION
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AERODYNAMICS
Wing Configuration:
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GEOMETRY OF THE WING:
Wing geometry is described by
a. Plan-form shape b. Aspect ratio (which is already obtained from comparative graphs) c. Wing sweep d. Taper ratio
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With Free Trial the span e. Aerofoil shape andDownload thickness along
f. Geometric twist (change in aerofoil chord incidence angle along the span).
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Wing span, b = 27.2m Wing Area, S = 80.1 m^2 Aspect ratio, AR = b^2/S = 9.3
Taper Ratio λ = 0.36 (from Roskam book – BAE 196-200)
We know that taper ratio, λ = Ct = 0.36Cr You're Reading a Preview Unlock full access with a free trial.
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Where, Ct is the root chord
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Ct =0.36 Cr
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Hence
Cr = 4.33m
Also Ct = 1.56m
A leading edge sweep angle of 15°. Hence we get th plan form as shown
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DETERMINATION OF THE MEAN AERODYNAMIC CHORD:
Mean chord
2
= {(2/3)*Cr (1+λ+ λ )}/ (1+λ)
= 3.16 m
Distance of the mean chord = {b*(1+2 λ)}/{6*(1+λ)} You're Reading a Preview
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RELATIVE LOCATION OF THE WING:
There are three basic vertical locations of the wing relative the fuselage: 1. High wing
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3. Low wing
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Mid wing: 1. Least interference drag. 2. Gives best stability with little dihedral.
Low wing:
1. Landing gear can easily be retracted into the wing box 2. Added
fillet
interference.
will
avoid
undesirable
aerodynam
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In light of all the above considerations, we choose a hi
wing configuration mainly due to structural and Landing ge considerations.
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In order to select an aerofoil the appropriate thickness chord ratio(t/c) is to be determined.
The given formula can be used to determine the thickness chord ratio: Volume of fuel = 2*{((2/3)*(t/c)*c)*c*(b/4)}*0.5*1.33
Volume of fuel = 2*{((2/3)*(t/c)*c)*c*(b/4)}*2 You're Reading a Preview Unlock full access with a free trial.
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Substituting the above weight of fuel in the equation a solving, we get the value of thickness to chord ratio as
10.965 2*{((2/3)*(t/c)*2.945)*2.945*(27.2/4)}*0.5*1.33
t/c = 0.209 (or) 0.21
As our aircraft will fly only at subsonic speeds, we ha Reading a Preview chosen a NACA 6 seriesYou're aerofoil. Unlock full access with a free trial.
Based on the t/c ratio, we have chosen the aerofoil NA Download With Free Trial 664-221.
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FUSELAGE LAYOUT
The fuselage layout is important as the length of the entire aircraft depends on this.
The length and diameter of the fuselage is related to the seating arrangement.
The fuselage of a passenger aircraft is divided into a numbe of a sections: You're Reading a Preview Unlock full access with a free trial.
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Functions of fuselage:
provision of volume for payload
provide overall structural integrity
access with agear free trial. and power plant possible mountingUnlock of full landing
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Once fundamental configuration is established, fuselage layout
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A relation between the gross weight of the aircraft and the length of the aircraft.
b
Lfus = aW Where W is in lbs and Lfus is in ft. For jet transport,a=0.67,b=0.43 Lfus = 19.5m
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With Free Trial The layout of the flightDownload deck and specified pilot window
geometry is often the starting point of the overall fuselage layout.
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Passenger cabin layout
Two major geometrical parameters that specify t
passenger cabin are Cabin Diameter and Cabin Length . The
are in turn decided by more specific details like number seats, seat width, seating arrangement
(number abrea
seat pitch, aisle width and number of aisles.
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airlines. The main decision to be taken is the number of se
abreast and the aisle arrangement. The number of se
across will fix the number of rows in the cabin and there
the fuselage length. Design of the cabin cross section
further complicated by the need to provide different clas like first class, business class, economy class etc.
Cabin length:
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The total number of Download seats With 150FreeisTrial distributed as 2 sea abreast.
Cabin parameters are chosen based on standards for simi airplanes. Master your semester with Scribd Read Free Foron 30this Days Sign up to vote title & The New Times Useful Not useful The York various parameters chosen are as follows Special offer for students: Only $4.99/month.
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No .of aisles =1
Hence, the total cabin length will be
=
pitch*rows
= 1.06*17+additio space = 23.23m
Cabin Diameter You're Reading a Preview
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calculate the
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Internal diameter of the cabin.
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According to the standards prescribed, the structural
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Rear Fuselage:
The rear fuselage profile is chosen to provide a smooth, l
drag shape which supports the tail surfaces. The lower side
the profile must provide adequate clearance for aircr
when rotation during take off. The rear fuselage should a
house the auxiliary power unit(APU).Based on data collect for similar aircraft we choose the ratio L tail/dfus as 4. You're Reading a Preview
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With Free Trial LDownload tail = 11.2m
Total Fuselage Length
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Total
= 27.18m
Centre of gravity location:
The major weight components for which we have some id
of their location are the engine, the passengers and pilot, a
the baggage. Using this information, we can make
preliminary estimate of the location of the centre of gravity
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LIFT ESTIMATION LIFT:
Component of aerodynamic force generated on aircr perpendicular to flight direction.
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Lift Coefficient (CL) • Amount of lift generated depends on: – coefficient (CL) Lift=(1/2 ρ V^2)SCl=qSCl
• CL is a measure of lifting effectiveness and mainly depend upon: – Section shape, compressibility effects (Mach number), viscous effects (Reynolds’ number). Generation of Lift You're Reading a Preview Unlock fullfrom access with a free trial. • Aerodynamic force arises two natural sources: – Variable pressure distribution. Download With Free Trial
– Shear stress distribution. • Shear stress primarily contributes to overall drag force on aircraft.
• Lift mainly due to pressure distribution, especially on mai Master your with Scribd liftingsemester surfaces, i.e. wing . Read Free Foron 30this Days Sign up to vote title & The New York Times Useful Not useful • Require (relatively) low pressure on upper surface and
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Pressure variations with angle of attack – Negative (nose-down) pitching moment at zero-lift
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Download Free Trial Lift Curves of Cambered andWith Symmetrical airfoils
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CALCULATION:
General Lift equation is given by,
Lift=(1/2 ρ V^2)SCl=qSCl Lift at Cruise = 0.3715 (at the cruising altitude of 10800m) V = 242.2 m/s S = 400.72 kg/m2 You're Reading a Preview CL(cruise) = 0.63022 (from wing and airfoil Unlock full the access with a free trial. estimation) Download With Free Trial
Substituting all these values in the general lift equation,
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Lift at Take-Off = 1.225 (at sea altitude)
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L=12×1.225×(0.7×1.2×66.86)^2×400.72×2.5
Lift at Landing = 1.225 (at sea altitude) V = 0.7 x Vt = 0.7 x 1.3 x Vstall S = 400.72 kg/m2 CL(landing) = 3.058 (flaps extended and kept at the landing position of 40)
Substituting all these values in the general lift equation, L =12×1.225×(0.7×1.3×60.55)^2×400.72×3.5 Lift at landing =. You're Reading a Preview Unlock full access with a free trial.
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DRAG ESTIMATION
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DRAG: - Drag is the resolved component of the complete Master your semester with Scribd aerodynamic force which is parallel Read to direction Free Foron 30this Days Signthe flight up to vote title & The New York Times Useful Not useful relative oncoming airflow). Cancel anytime. Special offer for students: Only $4.99/month.
- It always acts to oppose the direction of motion.
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Planform area (S), air density , flight speed (V), drag coefficient (CD) o
CD is a measure of aerodynamic efficiency and mainly depends upon: o Section shape, planform geometry, angle of attack , compressibility effects (Mach number), viscous effects (Reynolds’ number).
Drag Components - Skin Friction: You're Reading a Preview o Due to shear stresses produced in boundary layer . Unlock full access with a free trial.
Significantly more for turbulent than laminar types of Download With Free Trial boundary layers o
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Form (Pressure) Drag o Due to static pressure distribution around body component resolved in direction of motion.
Sometimes considered separately as forebody and rear (base) drag components. o
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Result of both direct shock losses and the influence of shock waves on the boundary layer. o
o
Often decomposed into portions related to: Lift. Thickness or Volume.
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Typical streamlining effect
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Lift induced (or) trailing vortex drag
The lift induced drag is the component which has to be included to account for the 3-D nature of the flow (finite span) and generation of wing lift
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span) and generation of wing lift
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Where τ = (t/c) t / (t/c) r
(r – root, t – tip)
WING WETTED AREA: Wing wetted area
= 2 *80.1 {1+0.25(0.21) 1+
(0.95) (0.36) }
1+0.
S (wing)wet
2
= 168.52m
You're Reading a Preview TAIL WETTED AREA: Unlock full access with a free trial.
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Tail Wetted Area
= 2 x 20.81 {1+0.25(0.21) 1+
(0.95) (0.29)}
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S
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Fuselage wetted area
=Π x Diameter x Length
=Π (2.87) (30) 2
S(fuselage)wet
=270.49m
PARASITE DRAG ESTIMATION(CDO): CDO =
Σ K Cf Swet Sref
Where, K
-
Form factor
Cf
-
Co-efficient of Friction
Sref -
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Unlock full access with a free trial.
Reference Wing Area Download With Free Trial
FORM FACTOR(K):
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caused due to viscous separation. It is estimated for the
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WING & TAIL
Kwing
=
4
0.18
4
0.18
[1+ 0.6 (0.21) + 100 (0.21) ][1.34x0.74
0.28
15)
] 0.45
=
1.8535
You're Reading a Preview
Ktail
=
full access with a free trial. [1+ 0.6 Unlock (0.21) + 100 (0.21) ][1.34x0.74
0.28
45)
]
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0.45
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FUSELAGE
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f (Fineness Ratio)
=
l/d
= 30 / 2.87 Master your semester with Scribd Read Free Foron 30this Days Sign up to vote title & The New York Times Useful Not useful Special offer for students: Only $4.99/month.
=
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FORMULAS USED IN DRAG ESTIMATION
Reynolds Number (at Altitude 40000ft) =
ρVD μ
Co-efficient of friction (Cf )
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Total Parasite Unlock Drag Co-efficient full access with a free trial. CDO =
Σ K CfWith Swet Download Free Trial Sref
Drag Correction for 4% Interference effect
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Induced Drag (CDI)
2
=
CL
ΠeAR
CALCULATIONS: Reynolds Mach No
Cf
CDO
no.
CDO INTERFERENCE CORRECTION(4%)
700000
0.3 0.004747 0.044145 You're Reading a Preview
0.045910773
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1000000
0.4 0.004412 0.041028
0.042669057
3000000
0.5 0.003595 0.033433
0.03477016
6000000
0.6
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0.00317 0.029475
0.030653716
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LANDING GEAR CONFIGURATIONS
Wheel diameter(inch) Wheel width(inch)
A
B
1.63
0.315
0.1043
0.48
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MAIN LANDING GEAR No. of wheels = 4
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= 0.91m (35.95 inch)
Wheel width
= A(Wm)^B
= 0.1043(18,426)^0.48
= 0.29m (11.63inch)
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NOSE LANDING GEAR
No. of wheels = 2
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Wheel width
= A (Wn) ^B = 0.1043(1652.43)^0.48 = 0.0927m (3.65inch)
These calculated values for diameter and width should be increased about 30% if the aircraft is to rough unpaved runways.
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Xacwb = xn-Vht at/a
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Vht = horizontal tail volume ratio
at/a = lift slope ratio of tail to wing
Static margin = (xn – c.g)/chord
Lets assume static margin
= 18%
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= 16.81m
Therefore wing is placed such that the aerodynamic centre the wing is placed at 16.81m behind the nose
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full access with a free trial. To find the distance ofUnlock leading edge of the wing from nose
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Main landing gear is placed at the centre of the wing You're Reading a Preview
Therefore,
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The location of the centre of the wing = distance of the leading edge of the wing from nose + (root chord/2)
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it can be conveniently folded rearward and upward into the fuselage. Set the nose wheel location of 4.43m as shown.
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AIRCRAFT PERFORMANCE THRUST REQUIRED:
The aircraft to be designed is assumed to be in a stead
level flight at an altitude of 30,000ft and at the given cruise velocity. The aircraft’s power plant must produce the net thrust to overcome the drag, in a sense that the net thrust produced is equal to the drag experienced by the aircraft.
The thrust required to obtain this steady velocity is given b the equation,
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Thrust Required,
TR
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=
_W_
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CL/CD
FORMULAE For a given value of velocity (V),
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2
ρV S
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RATE OF CLIMB
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The aircraft is considered to be in steady, unaccelerated, climbing flight.
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denoted in the Fig.
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34654.91 x 9.81 R/C =
14.73 m/s
or
883.9 m/min
RANGE & ENDURANCE
Range is technically defined as the total distance traversed
the aircraft on a tank of fuel. Endurance is the total time t
aircraft stays on air on a tank of fuel. One of the criti
parameters influencing range and endurance is the Thru
Specific Fuel Consumption which amount of fuel consum You're Reading a Preview
full access with a freeand trial. Endurance are fou per unit thrust per unitUnlock time. Range
Download With Free Trial using the Brequet Formula.
ENDURANCE
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E
=
___1
x 1.36 x ln 34654.91
1.9425x10
-5
34654.91 -
7894.22 =
301 min or 5.02 hrs
RANGE R
=
1/2
2
_2
CL
Ct
ρ S CD
1/2
(WO
- W1
1/2
)
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R
=
5
full access a free trial. 2 xUnlock 0.2334 xwith 0.514x10 x 6.937 x 22.57
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Range =
3756.6 km
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further estimation of parameters. For takeoff the pi accelerates with afterburning thrust of 60kN.
Takeoff Distance
=
1.44 W
2
g ρ S CL,max T
=
1.44 x (34654.91 x 9.81)
2
9.81 x 1.225 x 80.1 x 2 x 60000 You're Reading a Preview
Takeoff Distance
=
1440.82 m
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increase in drag. The maximum lift coefficient with flaps fu employed at touchdown is 2.4.
Landing Distance(SL)
=
1.69 W
2
g ρ S CL,max {D + μr(W-L)} where, μr
SL
-
Co-efficient of rolling friction (0.4)
= _________1.69(34654.91 x 2
9.81) __________________ 9.81 x 1.225 x 80.1 x 2.4 x (23101.5 + 0.4(34654.91 x You're Reading a Preview
9.81))
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Landing Distance
=
531.45 m
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BIBLIOGRAPHY
REFERENCES: rd 1. Airplane Design by Dr. Jan Roskam, 3 edition.
2. Aircraft Design: A Conceptual Approach by Daniel P th
Raymer, 4 edition. 3. Introduction to Flight by John D. Anderson, 2
nd
edition.
4. Aircraft Performance and Design by John D. Anders nd
2 edition.
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With Trial 5. Theory of wing Download sections byFree Ira.H.Abbott, Dover
edition.
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