The document summarizes the design process for a passenger aircraft. It begins with selecting the aircraft's purpose and performance requirements. Next, initial parameters like mass and dimensions are estimated. The aircraft's configuration, engines, and components are then selected and optimized. Models of the aircraft and its systems are developed using computer-aided design software. Loads on the aircraft are calculated. Finally, the document discusses the aircraft's interior layout and seating arrangement.

1. УДК 629.7
Інв. №
МІНІСТЕРСТВО ОСВІТИ І НАУКИ, МОЛОДІ ТА СПОРТУ УКРАЇНИ
Національний аерокосмічний університет ім. М. Є. Жуковського
«Харківський авіаційний інститут»
Кафедра проектування літаків i вертольотів
ДО ЗАХИСТУ ДОПУСКАЮ
Завідувач кафедри
д–р техн. наук, проф.
_________О. Г. Гребеніков
(підпис)
PASSENGER AIRCRAFT INTEGRATED DESIGNING AND MODEL ANALYSIS
Пояснювальна записка до випускної роботи магістра,
напрямок 8.100101 — «Авіація та космонавтика»
Фах — «Літаки і вертольоти»
(номер зал. книжки без позначки «№»)
Виконавець студент гр. 16Е-2 KIRUBAGARAN MAZHALAI PRIYAN
(№ групи) (І.Б.П.)
(підпис, дата)
Керівник–консультант з основного розділу
к. техн. наук, доц. S. Trubaev
(науковий ступінь, вчене звання) (І.Б.П.)
(підпис, дата)
Нормоконтролер к. техн. наук, доц. S. Trubaev
(науковий ступінь, вчене звання кєрівника)(І.Б.П.)
(підпис, дата)

2. [1]
Ministry of Science and Education, Youth and Sports of Ukraine
National Aerospace University,named by N.E. Zhukovskiyj
«Kharkov Aviation Institute»
Faculty of aircraft and helicopter construction
Aircraft and helicopter design department
«Approved by»
Head of department №103,
prof.___________ Grebenikov А. G.
«___»_______________ 201__
ASSIGNMENT
FOR A FINAL WORK OF AN APPLICANT
for a masters degree 8.05110101 «Aircraft and helicopter»
group 16 E- 2 students name KIRUBAGARAN MAZHALAI
(Name)
SUBJECT OF GRADUATION PROJECT
«PASSENGER AIRCRAFT INTEGRATED DESIGNING AND MODEL
ANALYSIS»
Initial data for design: Vmax – _835____ km/h; Vkr – __760__ km/h; Vу – __14___ m/s; Hmax
– _11____ km; Hcr – __10___ km; L – ___2000__ km; Lp – __1970___ m; npas. – _ 47____
men; ncrew – ___4__ men.; mp/l – _39___ t;Т – __4____h; Кcr – __0.10___.
Graduation Project. Table of Contents
Abstract
Design section
1. Computer–aided general design of aircraft
Introduction, design goal – setting and tasking
1.1. Purpose, aircraft performance requirements, conditions of production and
operation, limitations imposed by aviation regulations in design of an

3. [2]
aircraft.
1.2. Statistical data collection, processing and analysis. Selection of aircraft main
relative initial parameters (Characteristics).
1.3. Selection and grounds of aircraft configuration, type of its power plant.
1.4. Selection of engines and examination of take off run.
1.5. Determination optimization of aircraft components and design parameters.
1.6. Development of design–structural configuration, aircraft center of gravity.
1.7. Standard specification of designed aircraft.
Realization of calculations, models and drawings:
 master–geometry of aircraft surface, outline drawing (format А1);
 Design–structural layout of aircraft (format А1).
2. Impact analysis in changes of aircraft component design parameters under their
optimization in aerodynamic and weight characteristics of aircraft
2.1. Determination of designed aircraft drag.
2.2. Lift force, induced drag, aircraft polar curve, aircraft lift–drag ratio, aircraft
polar.
2.3. Longitudinal moment and location of aerodynamic center of aircraft.
2.4. Influence of aircraft design parameters on its aerodynamic and weight
characteristics.
______________________________________________________________
___
3. Integrated designing and computer–aided modeling __________SURFACE
MODEL__________ of designed aircraft
3.1. Development of unit master–geometry.
3.2. Determination of loads acting on unit.
Realization of calculations, models and drawings:
 unit master–geometry;

4. [3]
4. Integrated designing and computer–aided modeling of aircraft systems
4.1. Hydraulic system designing and modeling.
4.2. Maintenance Manual of designed system.
________________________________________________________________
_
Realization of calculations, models and drawings:
 system schematic diagram (format А2);
Technological Section
5. Development of aircraft unit manufacturing technique
5.1. Development of enlarged production (manufacturing) methods in assembly of
units: selection of tools and equipment, specifications for delivery of parts
and assembly units, development of production charts for assembly
procedure, standardization, assembly cycle schedule.
_________________________________________________________________
Realization of calculations, models and drawings:
Economical section
6. Calculation of economic efficiency characteristics
6.1. Business plan: companys history, aircraft characteristic, product market,
marketing, personnel and management, risk analysis and their prevention.
6.2. Project financing: sources of financing, receipts and expenditures –
calculation of expenditures for designing and manufacturing, calculation of
cost value, price income, calculation of companys minimal internal funds,
determination of point of make out, calculation of direct and indirect costs.
6.3. Total transportation cost value and company's revenue.
6.4. Income from project.
6.5. Influence in change of aircraft and its units design parameters on aircraft
efficiency criteria.
7. Special assignment
Cabin layout and interiors design of the aircraft. Seating arrangement with high comfort
level ______________________________________________________

5. [4]
2. Explanatory note contents (list of questions subjected to development):
in compliance with assignment. Design-explanatory
note with Figures, Tables involved in text – up to 120 pages.
3. List of Graph materials (with obligatory drawings clearly specified):
graph material and presentation in strict correspondence to the assignment
Information on CD–R or DVD+/–R medium installed in department
computer network prior to defense
4. Date of assignment issue:
5. Date of final project presentation:
Project supervisor
(Date, signature)
Assignment accepted to fulfillment
« » 200
(Date, students signature)

6. [5]
2012
CONTENTS
ABSTRACT ………………………………………………………………………… 4
INTRODUCTION………………………………………………………………… 5
AIRCRAFT DESIGN PROCESS………………………………………………… 7
GENERAL DESIGNING OF AIRCRAFT
PURPOSE OF THE AIRCRAFT………………………………………………….11
REQUIREMENTS FOR FLIGHT PERFORMANCES………………………….14
DESIGN CHART OF THE DESIGNED AIRCRAFT……………………………15
PROTOTYPE DATAS………………………………………………………………17
SELECTION OF AIRCRAFT MAIN RELATIVE INITIAL PARAMETERS…..24
CALCULATION OF AIRCRAFT MASSES THROUGH THE SOFTWARE AND ITS
RESULTS……………………………………………………………………………..25
ZERO APPROXIMATION…………………………………………………………31
STATISTICAL COMPUTATION OF MASSES OF AIRCRAFT………………..32
AIRCRAFT OPTIMIZATION AND DESIGN PARAMETERS…………………33
SELECTION AND GROUNDS OF AIRCRAFT CONFIGURATION…………..43
SELECTION OF ENGINE…………………………………………………………..55
AVERAGE BETWEEN GRAPHICAL, SEMI-EMPIRICAL & STATISTICAL
METHOD……………………………………………………………………………….63
MAXIMUM TAKE-OFF MASS………………………………………………………66
CENTER OF GRAVITY………………………………………………………………68
DESIGN STRUCTURAL CONFIGURATION …………………………………….83
AERODYNAMICS
AIRCRAFT DESIGN PARAMETERS ON AERODYNAMIC CHARACTERISTICS
………………………………………..83
CALCULATION OF AERODYNAMIC PARAMETERS USING THE SOFTWARE
…………………………………………92
CALCULATION OF ZERO DRAG COEFFICIENT FOR TAKE-OFF AND
LANDING…………………………….109

7. [6]
INTEGERATED DESIGN OF AIRCRAFT AND LOAD CALCULATION
AIRCRAFT MASTER GEOMETRY USING UNIGRAPHICS………………….114
WING LOAD CALCULATION……………………………………………………..118
CALCULATION OF THE DISTRIBUTED FUEL LOAD ON A PLANE WING…119
THE WING STRUCTURE MASS LOAD ALLOCATION…………………………125
SHEAR FORCE, BENDING MOMENT AND REDUCED MOMENT……………127
CALCULATION SCHEME OF REDUCED MOMENT FROM CONCENTRATED
LOADS AND FROM ALL LOADS……………………………………………………132
AIRCRAFT SYSTEMS DESIGN AND SCHEMATIC LAYOUT
AIRCRAFT HYDRAULIC SYSTEM……………………………………….134
HYDRAULIC FLUID…………………………………………………………135
COMPONENTS INVOLVED IN HYDRAULIC SYSTEM…………………137
HYDRAULIC DESCRIPTION OF THE DESIGNED AIRCRAFT……….151
HYDRAULIC SYSTEM PANEL……………………………………………..154
HYDRAULIC SYSTEM MAINTENANCE………………………………….157
MANUFACTURING TECHNOLOGY OF VERTICAL RIB
AIRCRAFT RIB……………………………………………………………………..159
RIB CONSTRUCTION……………………………………….……………………..163
PRODUCTION METHOD OF PARTS OF THE RIB……………………………166
ASSEMBLY PROCEDURE OF THE RIB…………………………………………168
STAGES OF FORMATION OF RIB DIMENSIONS USING TEMPLATES……171
AIRCRAFT VERTICAL ASSEMBLY JIG DESIGN LAYOUT DIAGRAM…….173
ECONOMICAL SECTION
ECONOMIC EFFICIENCY CHARACTERISTICS CALCULATION………..174

8. [7]
SPECIAL ASSIGNMENT
INTERIOR CABIN LAYOUT AND SEATING ARRANGEMENT……..…………..179
FULFILLING REQUIREMENTS OF THREE ABREAST SEATING LAYOUT….179.
CABIN DIMENSIONING FOR 3- ABREAST SEATING…………………………..…179
INTERIOR ARRANGEMENT – CROS SSECTION (TYPICAL)……………………180
DETERMINATION OF DESIGNED AIRCRAFT CABIN CROSS-SECTION…......181
DETERMINATION OF CABIN LENGTH FOR HIGH COMFORT LEVEL………183
LIST OF DIAGRAMS
3-VIEW DRAWING OF THE AIRCRAFT …………………………………189
DESIGN STRUCTURAL LAYOUT………………………………………….188
CENTRE OF GRAVITY LOCATION OF THE AIRCRAFT……………..188
HYDRAULIC SYSTEM SCHEMATIC OF THE AIRCRAFT…………….190
CABIN SEATING LAYOUT OF THE AIRCRAFT………………………..189
CONCLUSION …………………………………………………………………. 191
REFERENCE ………………………………………………………………………. 192

9. [8]
ABSTRACT
The General design of the aircraft is carried out on basis of collection of aircraft
statistical data and in accordance with the pilot project development task and finally the
general view of an aircraft is presented. The main purpose of the aircraft design requirement
is fulfilled according to the aviation rules and regulations.
The main project is categorized into five part, in the first part aircraft take-off mass in
zero approximation is determined and followed by designing the weight of main units, fuel,
equipment, control system, geometrical dimensions of a wing, tail units, fuselage, landing
gear, location of their center of masses, calculating the aircraft’s center-of-gravity. Finally
the design specifications for the aircraft are presented.
The second part is focused on determining aerodynamic forces acting upon the
designed aircraft.
The third part is the development of aircraft unit structure using Computer aided
designing. The loads acting on the designed aircrafts unit structure is calculated and the
materials for unit structure are selected.
The fourth part is concerned about the systems developed for designed aircraft. The
schematic layout of the hydraulic system and its purpose are briefed following the operations
& maintenance manual of the designed system.
The last part is the technological section where the development of production charts
for assembly of designed aircraft vertical rib structure is done.
In special assignment the seating arrangement of the designed aircraft is sketched and
interiors of aircraft components are briefed with the current industry techniques.

10. [9]
INTRODUCTION
The purpose of designing a new aircraft is the creation of a structure with unique
characteristics, which should be reliable, economical fulfilling the conditions of operation,
performance requirements and its primary goal should be attained. To perform the
preliminary design structure of the aircraft it is necessary to be knowledgeable in the field of
general arrangement of aircraft and helicopters, design of power units and systems,
construction of elements of assembly structures and units of the aircraft, aero
hydrodynamics, durability, technologies, material science, and economics. The purpose of
design is to develop a project, realization of which, being limited to a certain extent, would
ensure the most efficient reaching of the defined goals of the design.
In designing a new aircraft the following should be considered,
 fulfillment of targeted tasks
 stability and controllability of an aircraft on a specified trajectory
 control and navigation in various flight conditions
 life support
 Performance characteristics
 Characteristics of technological level of the serial aircraft and its economic efficiency
 The special equipment
 Standardization and unification level
 Requirements to reliability and maintenance system
 Power plant and its systems
 Perspective of development of the aircraft and its basic systems
An aircraft is an element of the aviation complex, which seamlessly unites human and
material resources and carries out certain useful functions. The functional-structural diagram
of the aviation complex is shown on Figure. The aviation complex is an element of state
transport or defense system. All this defines necessity to use systematic approach to aircraft
design.
To implement the process of aircraft design, there was necessity to create specialized
development design offices, which include complicated laboratory and manufacturing
research. The activities of development design officers are based on work of branch-wise
research institutes, which research the prospects of aviation development in various
directions, and on experience of aircraft production and operation.

11. [10]
FOUR STAGES OF DESIGNING
1) External designing: At this stage the research of complicated organization-technical
systems including an aircraft or aircraft family as an element is carried out.
2) The second stage — the development of a technical proposal: At this stage, the
scheme is selected and optimal combination of basic aircraft parameters, composition
and structure of systems ensuring fulfillment of required functions is determined.
3) Third stage – front end engineering: In the process of design arrangement the
aircraft center-of-gravity is specified. The calculation of center-of-gravity is followed
by making weight reports on the basis of strength and weight calculations of airframe
and power unit, lists of equipment, outfit, cargo etc.
4) The fourth stage – working draft: The purpose of this stage is issuing all technical
documentation required for production, assembly, mounting of separate units and
systems and the whole aircraft as well. At this stage, on the basis of design-
technological elaboration the drawings with general view of aircraft units, assembly
and working-out drawings of separate parts of the aircraft.
MAIN STAGES OF AIRCRAFT PROJECT DEVELOPMENT
AIRCRAFT DESIGN PROCESS
The aircraft design process is the steps by which aircraft are designed. These depend on
many factors such as customer and manufacturer demand, safety protocols, physical and
economic constraints etc. For some types of aircraft the design process is regulated by
national airworthiness authorities. This article deals with powered aircraft such as airplanes
and helicopter designs.

12. [11]
Aircraft design is a compromise between many competing factors and constraints and
accounts for existing designs and market requirements to produce the best aircraft.
DESIGN CONSTRAINTS IN DESIGNING PROCESS
A. Aircraft regulations
Another important factor that influences the design of the aircraft are the regulations put
forth by national aviation airworthiness authorities.
Airworthiness Certificates-An airworthiness certificate is an FAA document which grants
authorization to operate an aircraft in flight.
Standard Airworthiness Certificate-A standard airworthiness certificate (FAA form 8100-2
displayed in the aircraft) is the FAA's official authorization allowing for the operation of type
certificated aircraft in the following categories:
 Normal
 Utility
 Acrobatic
 Commuter
 Transport
 Manned free balloons
 Special classes
FUNCTIONAL-STRUCTURAL CHART OF THE AVIATION COMPLEX
A standard airworthiness certificate remains valid as long as the aircraft meets its approved
type design, is in a condition for safe operation and maintenance, preventative maintenance,
and alterations are performed in accordance with 14 CFR parts 21, 43, and 91.

13. [12]
Airworthiness Certification Process-The FAA requires several basic steps to obtain an
airworthiness certificate in either the Standard or Special class.
The FAA may issue an applicant an airworthiness certificate when:
o Registered owner or operator/agent registers aircraft,
o Applicant submits application (PDF) to the local FAA office, and
o FAA determines the aircraft is eligible and in a condition for safe operation
A. Environmental factors
An increase in the number of aircraft also means greater carbon emissions. Environmental
scientists have voiced concern over the main kinds of pollution associated with aircraft,
mainly noise and emissions. Aircraft engines have been historically notorious for creating
noise pollution and the expansion of airways over already congested and polluted cities have
drawn heavy criticism, making it necessary to have environmental policies for aircraft noise.
Noise also arises from the airframe, where the airflow directions are changed. Improved
noise regulations have forced designers to create quieter engines and airframes. Emissions
from aircraft include particulates, carbon dioxide (CO2), Sulphur dioxide (SO2), Carbon
monoxide (CO), various oxides of nitrates and unburnt hydrocarbons. To combat the
pollution, ICAO set recommendations in 1981 to control aircraft emissions.[11] Newer,
environmentally friendly fuels have been developed and the use of recyclable materials in
manufacturing have helped reduce the ecological impact due to aircraft. Environmental
limitations also affect airfield compatibility. Airports around the world have been built to suit
the topography of the particular region. Space limitations, pavement design, runway end
safety areas and the unique location of airport are some of the airport factors that influence
aircraft design.
B. Safety
The high speeds, fuel tanks, atmospheric conditions at cruise altitudes, natural hazards
(thunderstorms, hail and bird strikes) and human error are some of the many hazards that
pose a threat to air travel. Airworthiness is the standard by which aircraft are determined fit
to fly.[19] The responsibility for airworthiness lies with national aviation regulatory bodies,
manufacturers, as well as owners and operators.
The International Civil Aviation Organization sets international standards and recommended
practices for national authorities to base their regulations on The national regulatory
authorities set standards for airworthiness, issue certificates to manufacturers and operators
and the standards of personnel training. Every country has its own regulatory body such as
the Federal Aviation Authority in USA, DGCA (Directorate General of Civil Aviation) in
India, etc.

14. [13]
C. Design optimization
Aircraft designers normally rough-out the initial design with consideration of all the
constraints on their design. Historically design teams used to be small, usually headed by a
Chief Designer who knew all the design requirements and objectives and coordinated the
team accordingly. As time progressed, the complexity of military and airline aircraft also
grew.
D. Design aspects
The main aspects of aircraft design are:
1. Aerodynamics
2. Propulsion
3. Controls
4. Mass
5. Structure
All aircraft designs involve compromises of these factors to achieve the design mission.
E. Computer-aided design of aircraft
In the early years of aircraft design, designers generally used analytical theory to do the
various engineering calculations that go into the design process along with a lot of
experimentation. These calculations were labor intensive and time consuming. In the 1940s,
several engineers started looking for ways to automate and simplify the calculation process
and many relations and semi-empirical formulas were developed. Even after simplification,
the calculations continued to be extensive. With the invention of the computer, engineers
realized that a majority of the calculations could be done by computers, but the lack of
design visualization and the huge amount of experimentation involved kept the field of
aircraft design relatively stagnant in its progress.
F. Financial factors and market
Budget limitations, market requirements and competition set constraints on the design
process and comprise the non-technical influences on aircraft design along with
environmental factors. Competition leads to companies striving for better efficiency in the
design without compromising performance and incorporating new techniques and
technology.

15. [14]
DESIGN SECTION PART-1
COMPUTER AIDED GENERAL DESIGNING OF AIRCRAFT
1. PURPOSE OF THE AIRCRAFT
2. REQUIREMENTS FOR FLIGHT PERFORMANCES
3. DESIGN CHART OF THE DESIGNED AIRCRAFT
4. PROTOTYPE DATAS
5. SELECTION OF AIRCRAFT MAIN RELATIVE INITIAL PARAMETERS
6. CALCULATION OF AIRCRAFT MASSES THROUGH THE SOFTWARE
AND ITS RESULTS
7. ZERO APPROXIMATION
8. STATISTICAL COMPUTATION OF MASSES OF AIRCRAFT
9. AIRCRAFT OPTIMIZATION AND DESIGN PARAMETERS
10. SELECTION AND GROUNDS OF AIRCRAFT CONFIGURATION
11. SELECTION OF ENGINE
12. DETERMINATION OF CENTER OF GRAVITY OF THE AIRCRAFT
13. AVERAGE BETWEEN GRAPHICAL, SEMI-EMPIRICAL & STATISTICAL
METHOD
14. MAXIMUM TAKE-OFF MASS
15. CENTER OF GRAVITY
16. DESIGN STRUCTURAL CONFIGURATION
17. DESIGN SPECIFICATION

16. [15]
PURPOSE OF THE DESIGNING AIRCRAFT
A. Aircrafts Intended Purpose - Commercial usage
Commercial usage denotes using the aircraft for a business purpose or getting
directly/indirectly financial gain from it.
B. Payload category - Passengers
Aircraft adapted for carrying passengers.
C. Type - Regional jet
The term regional jet describes a range of short to medium-haul turbofan powered aircraft,
whose use throughout the world expanded after the advent of airline deregulation in the
United States in 1978.
Example
PRIMARY USERS MANUFACTURER ROLE
Aeroflot Yakolev Yak-40 regional sized mini-jet
airliners
SkyWest Airlines Bombardier CRJ100 Regional jet/Business jet
Pinnacle Airlines
ExpressJet
Comair
Aerosvit Airlines Antonov An-148 regional jet
Rossiya
D. Range - Short-range
short range refers to distance travelled is between 2500.2 km (less than 1350nm) and Time
taken to travel is less than 5 hours
Example
FROM & TO DISTANCE in km TIME
New York-Miami 2051.914 2 hours 49 mins
Tokyo-Seoul 1,159.04 1.5 to 2 hours
Denver-Boston 2800 3hrs 42 mins
G. Special Requirements - Cargo Carrying capability
Can be used to carry cargos and can be used as a cargo variant

17. [16]
H. Mode of Class - Economy class
Economy class refers to the seating arrangement of the aircraft which is usually reclined and
include a fold-down table. The seats pitch range from 29 to 36 inches (74 to 91 cm), usually
30–32 in (76–81 cm), and 30 to 36 in (76 to 91 cm) for international economy class seats.
Domestic economy classes range from 17 to 18.25 in (43 to 46.4 cm).
GENERAL REQUIREMENTS
1. The aircraft, its engines, equipment and other parts, and operational publications shall
meet the following requirements:
 aviation requirements АП-25 and additional requirements for airworthiness of
"AIRCRAFT NAME" aircraft, in consideration of its design and operational features,
forming the "Certification basis of aircraft of "AIRCRAFT NAME" type" together with
mentioned requirements;
 engine - aviation requirements АП-33;
 APU - aviation requirements АП-ВД.
2. As for engine emission the aircraft shall meet the requirements of Appendix 16 to
International Aviation Convention (Volume II «Aviation engines emission», Edition
1981, Revisions 1 to 4) and requirements of Aviation Regulations АП34.
3. As for protection against hijacking the aircraft shall meet the requirements ICAO
Appendix 6,8,17 (with Revisions 97 and 98)Ukrainian Air Law (Section 8).
4. Processing and analysis of flight data using the ground personal computer shall be
provided to control the correctness of maintaining of preset flight modes and the pilot
technique, to evaluate the pilots' professional level, technical state of the aircraft, its
equipment and functional systems in monitoring of operation conditions within life
time limits.
The system shall include:
 aircraft removable data carrier, receiving the information from corresponding aircraft
signal transmitters;
 personal computer with printer, input and reproducing device and specific software.
5.Ground facilities and repair equipment shall correspond to this performance
specification.
6.Simulators and training devices should be designed for aircraft according to individual

18. [17]
performance specifications. The programs for training of flight and technical staff should be
developed up to completion of certification tests.
SPECIFIC AIRCRAFT STRUCTURE REQUIREMENTS
 . The airplane should be designed and manufactured by a principle of ―fail-safe
structure‖.
 Weight layout and airplane center-of-gravity should ensure a capability of operational
both with total and short number of passengers at all possible operational versions of
loading and fuelling according to the instruction of loading and centre-of-gravity not
using ballast. Limit of on-ground tail-heavy center of gravity be no less than 5 % of
MAC.
 The capability of creation of convertible and transport versions should be provided on
the basis of this airplane according to special performance specification.
REQUIREMENTS FOR FLIGHT PERFORMANCES
Maximum passenger capacity with
distance between the seats 750 person 55
(762) mm,
Maximum payload
kg 5000
Cruise speed:
at long range cruise km/h 835
maximum
Cruise altitude,
km 10.5
Required length of RWY (SA, Н =
0, dry concrete), m 1950
for takeoff: 2250
for landing:
Applied flight range (emergency
fuel reserve for 0.75 hour of flight; km 2500
takeoff in SA conditions; Н = 0)
with maximum payload
Fuel consumption for 1 pass/km g 340
while flying for technical range
with maximum payload
Maintenance and overhaul, 8.8

19. [18]
manhour
REQUIREMENTS FOR ENVIRONMENTAL PROTECTION
 . As for perceived noise the aircraft should meet the requirements of Chapter 4 of
"Environmental protection" International Standards, Appendix 16 to the International
Civil Aviation Convention (Volume I «Air noise», 2001) and to requirements of Part
36 of Aviation Regulations АП-36.
 To decrease atmospheric pollution and reduce fuel flow at ground operation the
capability of fulfillment of taxiing before take-off and after landing with one
operating engine should be worked out on airplane.
DESIGN CHART OF THE DESIGNED AIRCRAFT
• Collection and process of statistical data
General
• Design specification and three view diagram
Design
Aerodyna
mic • Designed aircraft drag
Characteri
stics
• Calculation of loads acting on unit
Design
structural unit • Modelling of designed unit
Systems
• Schematic layout of the hydraulic system
Design
Technological
• Design of assembly jigs for developed unit
Activity
• Cabin layout
Special activity

20. [19]
STATISTICAL DATA COLLECTION
Statistical data collection is the process of collecting flight, mass, power plant and
geometrical data’s of required prototypes for the design project. In this project I have
collected four different aircraft data’s and their features are explained and tabulated. These
aircrafts are selected based upon the design requirements and design specification mentioned
below,
TACTICAL TECHNICAL REQUIREMENTS OF THE DESIGNING AIRCRAFT
Maximum speed , Vmax 835 km/h
Cruising speed, Vcruise 760 km/h
Cruising height, Нcruise 11 km
Number of passengers, npass 47
Number of crew members, ncrew 4
Range, L 2000 km
Take-off distance, Lр 1970 m
Vertical speed, Vy 14 m/s
Maximum take-off weight, Ммах 40 tons
DESIGN SPECIFICATION OF THE AIRCRAFT
Type of the aircraft - Transport category with capacity to carry 47 to 55 passengers
including crew
Aerodynamic configuration Normal configuration with horizontal stabilizer on tail
section
Wing Low wing with Dihedral and wing sweep
Tail T-tail configuration
Fuselage Cylindrical shape
Power plant type Turbofan located at aft part of the fuselage
Landing gear Tricycle configuration with nose wheel
Based on the tactical technical requirements and the general design specification of the
designing aircraft we are gathering the similar aircrafts and their detailed specification is
tabulated. From the critical parameters of the aircraft are listed. With the obtained results we
now ready to input all parameters in the software which would give all the relative masses
and some important parameters for further calculation.
Aircrafts data are gathered from various sources which include books, magazines, websites,
etc., Some of the missing parameters are found manually by calculations or it can be found

21. [20]
by scaling the three view picture of the collected aircraft. In obtaining the details it is
important to have the three view pictures of each aircraft for simplification further in drawing
the designed aircraft three view it is very helpful .A short brief of the aircraft is provided for
each of the aircraft with its variant and their three view picture.
Upon the four aircrafts selected we can take any one from that as a main prototype for further
simplification. I have selected the following aircrafts for comparison, 1.EMBRAER ERJ 145
2. BOMBARDIER CRJ100 3. TUPOLEV 134-A 4. BOEING 717-200
My main prototype is EMBRAER ERJ 145
AIRCRAFTS SELECTED FOR STATISTICAL DATA COLLECTION AND THEIR
PARAMETERS
EMBRAER ERJ 145
The Embraer ERJ 145 family is a series of regional jets produced by Embraer, a Brazilian
aerospace company. Family members include the ERJ 135 (37 passengers), ERJ 1
(44passengers), and ERJ 145 (50 passengers). The key features of the production design
included:
1. Rear fuselage-mounted engines
2. Swept wings (no winglets)
3. "T"-tail configuration
4. Range of 2500 km

22. [21]
Civilian models
 ERJ 135ER - Extended range, although this is the Baseline 135 model. Simple shrink
of the ERJ 145, seating thirteen fewer passengers, for a total of 37 passengers.
 ERJ 135LR - Long Range - increased fuel capacity and upgraded engines.
 ERJ 140ER - Simple shrink of the ERJ 145, seating six fewer passengers, for a total
of 44 passengers.
 ERJ 140LR - Long Range (increased fuel capacity (5187 kg) and upgraded engines.
 ERJ 145STD - The baseline original, seating for a total of 50 passenger
Military models
 C-99A - Transport model
 EMB 145SA (R-99A) - Airborne Early Warning model
 EMB 145RS (R-99B) - Remote sensing model
BOMBARDIER CRJ100
The Bombardier CRJ100 and CRJ200 are a family of regional airliners manufactured by
Bombardier, and based on the Canadair Challenger business jet.
The CRJ100 was stretched 5.92 meters (19 feet 5 inches), with fuselage plugs fore and aft of
the wing, two more emergency exit doors, plus a reinforced and modified wing. Typical
seating was 50 passengers, the maximum load being 52 passengers. The CRJ100 featured a
Collins ProLine 4 avionics suite, Collins weather radar, GE CF34-3A1 turbofans with
41.0 kN (4,180 kgp / 9,220 lbf), new wings with extended span, more fuel capacity, and

23. [22]
improved landing gear to handle the higher weights. It was followed by the CRJ100 ER
subvariant with 20% more range, and the CRJ100 LR subvariant with 40% more range than
the standard CRJ100. The CRJ 100 SE sub-variant was produced to more closely meet the
needs of corporate and executive operators.
Variants
Several models of the CRJ have been produced, ranging in capacity from 40 to 50
passengers. The Regional Jet designations are marketing names and the official designation
is CL-600-2B19.
CRJ100 -The CRJ100 is the original 50-seat version. It is equipped with General Electric
CF34-3A1 engines. Operators include Jazz Aviation, Comair and more.
CRJ200 -The CRJ200 is identical to the CRJ100 except for its engines, which were upgraded
to the CF34-3B1 model, offering improved efficiency.
CRJ440 -Certified up to 44-seat, this version was designed with fewer seats in order to meet
the needs of some major United States airlines.
Challenger 800/850 - A business jet variant of the CRJ200
TUPOLEV 134-A

24. [23]
The Tupolev Tu-134 (NATO reporting name: Crusty) is a twin-engined airliner, similar to
the French Sud Aviation Caravelle and the later-designed American Douglas DC-9, and built
in the Soviet Union from 1966–1984. The original version featured a glazed-nose design and,
like certain other Russian airliners (including its sister model the Tu-154), it can operate
from unpaved airfields.
Design and development
Following the introduction of engines mounted on pylons on the rear fuselage by the French
Sud Aviation Caravelle, airliner manufacturers around the world rushed to adopt the new
layout. Its advantages included clean wing airflow without disruption by nacelles or pylons
and decreased cabin noise. At the same time, placing heavy engines that far back created
challenges with the location of the center of gravity in relation to the center of lift, which was
at the wings. To make room for the engines, the tailplanes had to be relocated to the tail fin,
which had to be stronger and therefore heavier, further compounding the tail-heavy
arrangement.
Variants
Tu-134 The glass nosed version. The first series could seat up to 64 passengers, and
this was later increased to 72 passengers. The original designation was Tu-
124A.
Tu-134A Second series, with upgraded engines, improved avionics, seating up to 84
passengers. All Tu-134A variants have been built with the distinct glass nose
and chin radar dome, but some were modified to the B standard with the radar
moved to the nose radome.
Tu-134B Second series, 80 seats, radar moved to the nose radome, eliminating the
glazed nose. Some Tu-134B models have long-range fuel tanks fitted under the
fuselage; these are visible as a sizeable bulge.
Tu- Bomber aircrew training version.
134UBL
Tu134UBK Naval version of Tu-134UBL. Only one was ever built.
BOEING 717-200
Boeing 717 was specifically designed for the short-haul, high frequency 100-passenger
airline market. The highly efficient 717 concluded its production run in May 2006, though
the airplane will remain in service for years to come.
Final assembly of the 717 took place at the Boeing plant in Long Beach, Calif. The airplane
was originally part of the McDonnell Douglas airplane family and designated the MD-95

25. [24]
prior to merger with The Boeing Co. in 1997. The program produced 156 717s and pioneered
breakthrough business and manufacturing process for Boeing Commercial Airplanes
The
.
The standard 717 has a two-class configuration with 106 seats. Its passenger-pleasing interior
features a five-across-seating arrangement in economy class, with illuminated handrails and
large overhead stow bins.
The two-crew flight deck incorporates six interchangeable liquid-crystal-display units and
advanced Honeywell VIA 2000 computers.
Flight deck features include an Electronic Instrument System, a dual Flight Management
System, a Central Fault Display System, and Global Positioning System. Category IIIb
automatic landing capability for bad-weather operations and Future Air Navigation Systems
are available.
Two advanced Rolls-Royce 715 high-bypass-ratio engines power the 717. The engine is
rated at 18,500 to 21,000 pounds of takeoff thrust, with lower fuel consumption and
significantly lower noise and emission levels than the power plants on comparable airplanes.
DESIGN
The 717 features a two-crew cockpit that incorporates six interchangeable liquid-crystal-
display units and advanced Honeywell VIA 2000 computers. The cockpit design is called
Advanced Common Flight deck (ACF) and is shared with the MD-11.

26. [25]
STATISTICAL TABLE
FLIGHT DATA:
Flight includes Vmax – the maximum speed of flight; HV max – flight altitude with the
maximum speed; Vcruise –cruise speed; Нcruise –cruise altitude; Vland – landing speed; Vto –
take-off speed; VY– rate of climb; Hclg– static ceiling; L– flight range; Ltor – distance of the
take-off run; Lto – take-off distance; Lroll–landing roll distance; Lland– landing distance;
1 No 1 2 4 3
2 Name of the EMBRAER Bombardier TUPOLEV BOEING
aircraft 145 CRJ200 134-A 717-200
Producer Embraer Bombardier Boeing
Country Brazil Canada Tupolev United stated
Year of 1989-present 1992 Soviet union 1998–2006
production 1966–1984
3 Source Janes all the world aircraft and Wikipedia
FLIGHT DATA
4 Vcruise, km/h 833 850 850 811
5 Vmax, km/h 679 785 950 629
6 Нcruise, km 11.277 11 11 10.400
7 HV max, km 9.753 11 11.5 11.280
8 Vto, km/h 170 155 - 150
9 Vland, km/h 233 250 - 244
10 VY, km/h 6.5 6 - 6
11 Hclg, km 11.27 12.49 12.1 11
12 L(mf Max ) , km 3037 2500 - 2645
13 L(mcargo max) , 2963 1800 1020 3800
km
14 Lto, km 1.97 1527 2.4 1.7
15 Lland, km 1.3 1423 2.2 1.52
MASS DATA : This includes take-off mass(m0), maximum take-off mass(m0max), payload
mass(mpld), number of passengers(npass), landing mass(mland), empty mass(mempty), mass of
crew(mc), mass of fuel(mf), empty equipped mass(mempt.eqpd) and total mass(mtotal ).

27. [26]
S.NO MASS EMBRAER Bombardier TUPOLEV BOEING 717-
145 CRJ200 134-A 200
16 m0 (mto), kg 19200 21636 47000 49895
17 m0max, kg 20000 22000 47200 22000
MASS DATA
18 mpld, kg 5640 6240 8200 12000
19 npass 47 52 84 100
20 mland, kg 18700 20000 43000 43359
21 mempty, kg 11585 19,958 27,960 30000
22 mc, kg 5,284 6,124 8,200 12400
23 mf, kg 2865 4300 - 8500
24 mempt.eqpd, kg 17,100 13730 29050 43545
25 mtotal, kg 20,100 24,041 47,600 49,900
POWERPLANT DATA: This includes engine thrust (P0), mass of engine (meng), number of
engines and its type, specific fuel consumption (Cp) and bypass ratio(Y).
S.NO ENGINE EMBRAER Bombardier TUPOLEV BOEING 717-
SPECS 145 CRJ200 134-A 200
26 P0 (N0), 31.3 31 103 97.9
POWERPLANT DATA
daN (kN)
27 meng, kg 1438 2305 4640
28 No of 2 2 2 2
engines Twin-spool
Type of non-
engine afterburning
turbofan
29 Cp, lb/lbf·hr 0.39 - 0.498 -
30 Y, Bypass 3:1 5:1 - -
ratio
GEOMETRICAL DATA: This includes wing area(S), wing span(L), sweep angle(),
aspect ratio of wing(), thickness ratio at chord( c 0 ) and at tip( c tip ),taper ratio(), length of

28. [27]
fuselage(Lf), diameter of fuselage(df), area of aileron( S ail ), relative fuselage mid section
area(  S mcs ), wing loading(P0) and thrust to weight ratio(t0).
S.NO GEOMETRICAL EMBRAER Bombardier TUPOLEV BOEING
PARAMETERS 145 CRJ200 134-A 717-200
31 S, m2 51.18 54.54 127.3 92.97
32 L, m 20.04 20.52 29.00 28.45
33  22.73 24.75 35.00 24.50
GEOMETRICAL DATA
34  7.85 7.72 6.61 8.7
35 c0 4.09 5.13 - -
c tip 1.04 1.27 - -
36  4 3.4 0.255 5.10
37 Lf, m 29.87 24.38 37.10 33
38 df, m 2.28 2.69 2.9 3.34
39 f 12.25 9.06 11.45 4.30
40 S ail 1.70 1.93 - -
, m2
2
41  S mcs , m 7.56 8.38 10.5416 18.84
42 P0=m0g/10S, 375.15 394.3 369.21 556
daN/m2
43 t0=10P0/m0g 0.3326 0.3884 0.289 0.3806
DERIVATIVE VALUE: This includes specific fuel weight (eng), effective load factor (
K eff .load ), relative aileron area ( S ail ), relative horizontal ( S HT ) and vertical stabilizer area (
S VT ).
S.NO DERIVATIVE EMBRAER Bombardier TUPOLEV BOEING 717-
PARAMETERS 145 CRJ200 134-A 200
DERIVATIVE
VALUES
44 eng, kg/daN2 306.51 263 - 294
45 m c arg o 0.2752 0.288 0.1744 0.2405
K eff .load 
m0
46 K mcs  m 0  S mcs , 2539 2600 4459 2627

29. [28]
daN/m2
47 S ail  S ail S 0.0332 0.0353 - -
48 S HT  S HT S 0.219 0.173 0.241 0.205
49 S VT  S VT S 0.141 0.168 0.167 0.210
SELECTION OF AIRCRAFT MAIN RELATIVE INITIAL PARAMETERS
Thus finally tabulating all the required values it is necessary to find the main relative initial
parameters of wing, fuselage and the tail unit. The obtained result is used in the software.
WING PARAMETER
aspect ratio,  7.85
sweep angle,  22.73
taper ratio,  4
relative width of airfoil, c 18 or 0.18
relative chord of flap, b f  b f / b wing 0.25
deflection angles of flap, f 18
relative area of ailerons, S ail  S ail / S 0.06
FUSELAGE PARAMETER
fineness ratio f 12.25
fuselage diameter Df 2.50
TAIL UNIT PARAMETER
relative area of horizontal stabilizer, S HT  S HT S 0.219
relative area of vertical stabilizer, S VT  S VT S 0.141
aspect ratio of horizontal surface, HS 4.077
aspect ratio of vertical surface, VS 1.36
Sweep angle of horizontal surface, HS 20
Sweep angle of horizontal surface, HS 32
Relative thickness of horizontal surface, c HS 12 or 0 .12
Relative thickness of vertical surface, c VS 10 or 0.10

30. [29]
CALCULATION OF AIRCRAFT MASSES THROUGH THE SOFTWARE AND ITS
RESULTS
Aircraft masses in zero approximation are calculated using software by entering necessary
parameters taken from statistical data and the initial parameters. First the relative masses for
fuselage, wing, power plant, tail unit, fuel and landing gear are found with respect to the
aspect ratio of wing taken as 4 and aircrafts wing loading attained from graphical result as
600 N/m2. The graphs obtained from this result help us to select the desired wing loading and
from the wing the lowest value of takeoff mass is taken as the final one. The relative masses
are changed to direct masses by multiplying it with the finally obtained take-off mass, for me
it is 39.11tons. Therefore my relative masses are multiplied with 39.11 ton to get direct mass
Graphs for each lab are plotted versus each parameter and from that the final wing loading is
obtained followed by the takeoff mass. Other parameters obtained include engine
performance data like thrust to weight ratio at take-off, climbing and cruise. My main
comparative parameter is Aspect ratio which I took in three variations as
ASPECT RATIO 2 , 4 AND 6
AIM OF THE LAB AND ITS RESULT TO FIND THE RELATIVE MASS OF
AIRCRAFT
LAB 5: In this part we are finding the relative mass of power plant respect to the aspect
ratio of wing taken as 4 and aircrafts wing loading 600 N/m2.

31. [30]
RESULT: In table P,denotes wing loading Tk,aspect ratio and SU is the RELATIVE MASS
OF POWER PLANT which is 0.086
LAB 7A:In this part we are finding the relative mass of wing respect to the aspect ratio of
wing taken as 4 and aircrafts wing loading 600 N/m2.
RESULT: In table p, denotes wing loading and Tk, aspect ratio and Mkp is the RELATIVE
MASS OF wing which is 0.048.

32. [31]
LAB 7B:In this part we are finding the relative mass of fuselage respect to the aspect ratio
of wing taken as 4 and aircrafts wing loading 600 N/m2.
RESULT: In table DF, refers to diameter of the fuselage (3.84m) and Lf refers to aspect ratio
of the fuselage (12.25). By comparing both the values we get the relative value of fuselage
equals to 0.350.
LAB 7G:In this part we are finding the relative mass of tail unit with respect to the aspect
ratio of wing taken as 4 and aircrafts wing loading 600 N/m2.
RESULT: In table P, denotes wing loading and MOP, denotes RELATIVE MASS OF THE
TAIL UNIT which is 0.0182

33. [32]
LAB 7V:In this part we are finding the relative mass of landing gear with respect to the
aspect ratio of wing taken as 4 and aircrafts wing loading p,600 N/m2.
RESULT: RELATIVE MASS OF THE LANGING GEAR is 0.062
LAB 8: In this part we are finding the mass of equipment, crew and payload.
RESULT: MASS OF THE Equipment, crew and payload is 9502.98kg
LAB 9: In this part we are finding the take off mass of the aircraft which is equal to
39.11tons.

34. [33]
RESULT: Take off MASS OF is 39110 kg or 39.11 tons
RESULTS FROM GRAPH WITH RESPECT TO THE RELATIVE PARAMETER
LAB PARAMETERS RESULTS
NO
3 Lift to drag ratio 12.20
4 Thrust to weight ratio at Take-off 0.251
4 Thrust to weight ratio at Landing 0.271
4 Thrust to weight ratio at Cruising 0.179
5 Mean Thrust to weight 0.271
5 Relative mass of powerplant 0.086
7a Relative mass of wing 0.048
7b Relative mass of fuselage 0.350
7g Relative mass of Tail unit 0.0182
7B Relative mass of Landing gear 0.062
6 Relative mass of Fuel 0.224
8 MEQ = MCREW+MPAYLOAD+MEQUIPMENT 9502.98 kg
9 Take-off mass relative to wing loading 39110 kg

35. [34]
DIRECT MASS
Mass of fuselage 13688.5 kg
Mass of wings 1877.28 kg
Mass of tail unit 711.802 kg
Mass of powerplant 3363.46 kg
Mass of landing gear 2424.82 kg
Mass of fuel 760.64 kg
COMPUTATION OF AIRPLANE TAKE-OFF MASS IN ZERO APPROXIMATION
DETERMINATION OF MASS FROM LAB RESULTS
Mass of fuselage = Relative mass of fuselage * Take off mass
= 0.350 * 39110 kg = 13688.5 kg
Mass of wings = Relative mass of wing * Take off mass
= 0.048* 39110 kg = 1877.28 kg
Mass of tail unit = Relative mass of tail unit * Take off mass
= 0.0182* 39110 kg = 711.802 kg
Mass of power plant = Relative mass of power plant * Take off mass
= 0.086* 39110 kg = 3363.46 kg
Mass of landing gear = Relative mass of landing gear * Take off mass
= 0.062* 39110 kg = 2424.82 kg
Mass of Fuel = Relative mass of fuel * Take off mass
= 0.224* 39110 kg = 8760.64 kg
Mass of crew = 4 * 80 kg = 320kg
Mass of payload = 47 * 90 kg = 4230kg
Mass of Equipment and control systems = 4952kg
ZERO APPROXIMATION
Take-off mass of the airplane for zero approximation is determined by the formula received
from the equation of mass ratio with statistical data.

36. [35]
m 0  m st  m p . p  m f
 m pl  m crew  m eq ;
Here, m 0 = Take-off mass, m st = Structural mass of the aircraft, m p. p = Power plant
mass,
m f = Fuel mass, m pl = Payload mass, m crew = Crew mass, m eq = Equipment mass
m pl  m crew
Mass Ratio (dimensionless) equation is, 1  m st  m p. p m f  m eq 
m0
Re-arranging we get final takeoff mass as,
m pl  m crew
m0 
1  ( m st  m p . p  m f  m eq )
m st - Relative airframe mass
= Relative mass of fuselage+ Relative mass of wing+ Relative mass of tail unit+
landing gear
= 0.350+0.048+0.0182+0.062 = 0.4782
m p. p - Relative mass of power plant = 0.086
m f - Relative mass of fuel = 0.224
m eq - Relative mass of Equipment = 0.1266
4230  320
m0 
1  ( 0 . 4782  0 . 086  0 . 224  0 . 1244 )
m 0 = 53403.755 kg
STATISTICAL COMPUTATION OF MASSES OF AIRCRAFT
When the airplane takeoff mass in zero approximation is determined it is necessary to
m airfr m wing m fus
calculate airframe mass and its components (mass of the wing , fuselage ,
m tail unit m fuel m pow . pl
tail unit , landing gears), and also mass of fuel , power plant and
m
engines eng . Relative masses of airframe, power plant, equipment and control system, and
also of the aircraft performing normal take-off and landing are given in Table below
Plane Purpose m airfr m pow . pl m ctl . sys m fuel
Subsonic light 0.30…0.32 0.12…0.14 0.12…0.14 0,18…0,22
passenger long- medium 0.28…0.30 0.10…0.12 0.10…0.14 0,26…0,30
distance heavy 0.25…0.27 0.08…0.10 0.09…0.11 0,35…0,40
Multipurpose for local airlines 0.29…0.31 0.14…0.16 0.12…0.14 0.12…0.18

37. [36]
Take off mass from lab 9 = 39110 kg
Relative mass of Airframe = 0.30
Mass of Airframe = Relative mass of airframe * Take off mass
= 0.30 * 39110 kg = 12515.2 kg
Relative mass of power plant = 0.14
Mass of power plant = Relative mass of power plant * Take off mass
= 0.14* 39110 kg = 5475.4 kg
Relative mass of control systems and equipments = 0.11209
Mass of control sys = Relative mass of control systems and equipments * Take off mass
= 0.11209* 39110 kg = 4383.839 kg
Relative mass of fuel = 0.18
Mass of fuel = Relative mass of fuel * Take off mass
= 0.18* 39110 kg = 7039.8 kg
AIRCRAFT OPTIMIZATION AND DESIGN PARAMETERS
Geometrical parameters for designed aircraft are calculated by formulas taken from pilot
project book and rest is determined statistically by comparing with the prototypes. After
calculating the geometrical parameters we are drawing the theoretical drawing. The
geometrical parameters are calculated and obtained satisfying the general requirements of the
aircraft.
The stages of aircraft optimization include the following:
 Determination of Wing Parameters
 Determination of Fuselage Parameters
 Determination of Tail Unit Parameters
 Determination of Position of Center of Mass of the Airplane
 Determination of Landing gear parameters
Determination of Wing Parameters:
In determining Wing parameters its plan form shape is very important in obtaining number of
useful relations that apply to a trapezoidal shape. These are based on knowing the wing area,
aspect ratio, taper ratio, and leading-edge sweep angle.
Before finding the wing area it is necessary to determine the wing loading corresponding to
take off mass of 39110kg, which is found in LAB 9 as given below,

38. [37]
RESULT: The wing loading is found to be 600 daN/m2
Wing statistical parameter
aspect ratio,  sweep angle,  taper ratio, 
7.85 22.73 4
WING AREA
m0  g
S  Where m 0  39110 ( kg ) , g  9 . 8 ( m / s 2 ) , p 0  600 ( dN / m )
2
10  p 0
39110  9 . 8
S   63 . 87 ( m )
2
10  600
Wing Span ( l )
L  S Where λ= 7.85 (choosing from table)
L 7 . 85  63 . 87 = 22.39m

39. [38]
Wing Chords (b)
S  2 
b root  b 0  
 
 Where  = 4
L  1
63 . 87  2  4 
b root  b 0     4 .5 ( m )
22 . 39  4  1 
b 4 .5
b tip  0   1 . 1( m )
 4
Quarter chord line sweep angle of the Wing (  0 .25 )
 0 .25  22 . 73
0
(Choosing from the statistic’s table)
Leading edge sweep angle of the Wing (  0 )
 1 4 1
tg  0  tg  0 . 25   tg 22 . 73   0 . 4953
0
    1  7 . 85   4  1 
 0  arctg 0 . 4953   26 . 349 0
Mean Aerodynamic Chord of the Wing (MAC = b Aw )
  1
2
2
bA   b0 
3     1 
4  4 1
2
2
bA   4 .5   3 . 15 ( m )
3 4  4  1
Vertical distance between horizontal central line to MAC ( z A )
L  2 22 . 39 4  2
zA      4 . 478 ( m )
6  1 6 4 1
Horizontal distance between wing root tips to MAC root
l  2 22 . 39 4  2
xA    tg  0  . 0 . 4953  2 . 2179 ( m )
6  1 6 4 1
Determination of Fuselage Parameters
The size and shape of subsonic commercial aircraft are generally determined by the
number of passengers, seating arrangements and cargo requirements. Seating arrangements
on commercial passenger aircraft vary depending on the size and range.
Fuselage fineness ratio f fuselage diameter Df Nose section Rear section
fineness ratio ,  N fineness ratio,  T
12.25 2.5 m 1.5 2.5

40. [39]
Overall Fuselage length, Lf
L f   f  D f  12 . 25  2 . 5  30 . 62 ( m )
Fuselage nose length, L f .n
L f . n   N  D f  1 . 5  2 . 5  3 . 75 ( m )
Fuselage rear length, L f .r
L f . r   T  D f  2 . 5  2 . 5  6 . 25 ( m )
Fuselage middle section length, L f .m
Lfm = Lf – Lf.n – Lf.r = 30.62 - 3.75 - 6.25 = 20.62 m
Determination of Tail Unit Parameters
Tail unit parameters include Horizontal and Vertical stabilizer. Their geometrical parameters
are determined by the same formulae which were used for the wing.
Horizontal stabilizer parameter
Horizontal stabilizer statistical parameter
aspect ratio, sweep angle,  taper ratio,  Relative horizontal
 stabilizer area, S h .t
4.077 20 2 0.219
Horizontal stabilizer area
S h .t  S h .t  S  0 . 219  63 . 87 ( m )  13 . 98 m
2 2
Where S =63.89, wing area
Length of the Horizontal stabilizer
L h .t   h .t  S h .t  4 . 077  13 . 98  7 . 55 ( m )
Chords of the Horizontal tail Unit
S ht 2  13 . 98 2  2.
b root  b 0 . ht      2 . 48 ( m )
L ht   1 7 . 55 2 1
b 0 . ht 2 . 48
b tip . ht    1 . 24 ( m )
 ht 2 .0

41. [40]
Quarter chord line sweep angle of the Horizontal tail unit (  0 .25 . ht )
 0 . 25 . ht  20 (Choosing from the static’s table 1.1)
0
Mean Aerodynamic Cord of the Horizontal Tail ( b A . ht )
   ht  1 2.  2.  1
2 2
2 ht 2
b A . ht   b 0 . ht    2 . 48   1 . 93 ( m )
3  ht   ht  1  3 2 2  1
Vertical Distance between horizontal central line to MAC ( z A . ht )
L ht  ht  2 7 . 55 2  2
z A . ht      1 . 68 ( m )
6  ht  1 6 2 1
Leading edge sweep angle of the Horizontal stabilizer (  0 )
 1 2 1
tg  0  tg  0 . 25   tg 20   0 . 44572
0
    1  4 . 077   2  1 
 0  arctg 0 . 44572   24 . 023 0
Horizontal distance between wing root tip to MAC root ( x A . h .t )
L ht  ht  2 7 . 55 2  2
x A . ht    tg  0  . . 0 . 4452  0 . 7488 ( m )
6  ht  1 6 2 1
x A . ht  0 . 30 ( m )
Vertical stabilizer parameter
Vertical stabilizer statistical parameter
aspect ratio, sweep angle,  taper ratio, vs Relative horizontal
 stabilizer area, S h .t
1.36 32 1 0.141
Vertical stabilizer area
S v .t  S v .t  S  0 . 141  63 . 87 ( m )  9 . 01 m
2 2
Where S =63.89, wing area

42. [41]
Length of the Vertical stabilizer
L v .t   v .t  S v .t  1 . 36  9 . 01  3 . 50 ( m )
Chords of the Vertical tail Unit
S vt 2  9 . 01 1 2.
b root  b 0 . vt      2 . 57 ( m )
L vt   1 3 .5 11
b 0 .vt 2 . 57
b tip .vt    2 . 57 ( m )
 vt 1
Quarter chord line sweep angle of the Vertical tail unit (  0 .25 .vt )
 0 . 25 . vt  32
0
(Choosing from the static’s table 1.1)
Mean Aerodynamic Cord of the Vertical Tail ( b A .vt )
   vt  1 1 11
2 2
2 vt 2
b A . vt   b 0 . vt    2 . 57   2 . 57 ( m )
3  vt   vt  1  3 11  1 
Vertical Distance between horizontal central line to MAC ( y A .v .t )
L v .t  v .t  2 3 . 50 1  2
y A . v .t      1 . 76 ( m )
3  v .t  1 3 11
Leading edge sweep angle of the Vertical stabilizer (  0 )
 1 11
tg  0  tg  0 . 25   tg 32   0 . 62
0
    1  1 . 36  1  1 
 0  arctg 0 . 62   31 . 79 0
Horizontal distance between wing root tip to MAC root ( x A .v .t )
x A .v .t  Lvs  tg  0  3 . 50  tg  0  3 . 50  0 . 62  2 . 17 ( m )
DETERMINATION OF LANDING GEAR PARAMETERS
For nose-wheel tri-cycle landing gear the following parameters are considered
Wheel base (b ) : distance between axels of nose wheel and main landing gear wheels in
side view. It depends on fuselage length.
b  (0.30.5)l fus,
b = 0.48 * 30.62 = 14.6976m
Nose wheel offset (a ) : It is distance between vertical line passing through the airplane center
of gravity and nose wheel axis (or axis of several wheels whenever);
a = 0.968 * 14.697 = 14.235m

43. [42]
Main Landing Gear offset (e) : It is distance (on side view) between vertical line passing
through the airplane center of gravity and axis (or center line of several wheels, bogie) of
MLG;
e = 0.031 * 14.697 = 0.462m
Static ground angle() : It is the angle between fuselage construction plane and runway
surface. It is generally between
2 to +2. So it is taken as
1.
Angle of wing setting sett : It is the angle formed between wing construction plane to the
fuselage axis. It is generally between sett  (04). So it is taken be 2.
Angle of overturning(): It is the angle appearing when fuselage tail part or its tail bump
touches the runway surface;
Ψ=2 parking angle
amax = 12 maximum angle of attack
aw = -1 angle between a wing chord and longitudinal axis of fuselage.
Ф = amax –(-1) – 2 = 12+1-2 = 11
As a rule  = 1018, smaller values are accepted for non-maneuverable subsonic aircraft.
Offset angle(): It is the angle of offset for wheels of MLG relatively to airplane CG.It
prevents airplane overturning backward during landing.
 =  + (12).
 =11+ 2 = 13.
Wheel track(В ) : It is the distance (on front view) between planes of symmetry of MLG
wheels. This can be found by
B  ( 0 . 15 .... 0 . 35 ) LW
B  0 . 21  22 . 39  4 . 7 m
Height of airplane center of gravity(Н): It is the distance from airplane CG to the ground.
H=
H= 0.462 / tan (13)
H= 2.1 m
Height of the landing gear (h): Distance from leg attachment fittings to the runway surface
when shock absorber and tiers compression is of parking state (at take-off mass).
DF = 2.5 m
h =H – DF/2
h = 2.1 – (2.5/2)

44. [43]
h = 0.85 m
Determination of Position of Center of Mass of the Airplane
Position of the airplane center of mass is determined relative to nose part of the wing mean
aerodynamic chord (MAC).
The recommended distance for the center of mass from the nose part of mean aerodynamic
chord x m as follows:
For airplanes with swept wing:
x m  0 . 23  b A  0 . 23  3 . 15  0 . 7245 ( m )
Determination of tail arms
Vertical tail unit arm:
It is the distance measured from the aircraft center of massup to the vertical tail unit centre of
pressure. It is selected statistically , Tvs = 12m. For T-Tail Tvs is not equal to Ths.
Horizontal tail unit arm :
It is the distance measured from the aircraft center of mass up to the horizontal tail unit
centre of pressure. Horizontal tail unit arm is obtained after drawing theoretical diagram.
Calculation of high lifting devices parameters:
Flap configuration:
We will consider relative value of flap span from statistical data L flap  0 . 5 to 0 . 8
I considered relative length of flap = 0.6
The length of the flap is calculated by the formula
L span  D fuselage  2  L flap
L flap  L flap
2
ΔLflap- is the gap between flap and fuselage = 100mm
22 . 73  2 . 5  2 * 0 . 100
L flap   0 .6  6 m
2
Flap root chord is calculated by the following formula
   1 D fuselage  2  flap 
b 0 flap  b flap  b 0   1 
  

  L 
Here b flap is relative chord of flap is 0.2 to 0.4
 4  1 2 . 5  2  0 . 100 
There for b 0 flap  0 . 2  4 . 5   1     0 . 818 m
 4 22 . 73 
Flap tip chord length is calculated by fallowing formula

45. [44]
   1 D fuselage  2  flap  2 l flap 
b кflap  b flap  b 0   1 
  

  L 
 4  1 2 . 5  2  0 . 100  2  6 
b кflap  0 . 2   4 . 5   1     0 . 46 m
 4 22 . 73 
Slats configuration:
We will consider relative value of slat span from statistical data L slat  0 .6 to 0 .85
I considered relative length of slat = 0.65
The length of the slat span is calculated by below formula
L span  D fuselage  2  L slat
L slat   L slat
2
ΔLslat- is the gap between slat and fuselage = 300mm
There for
22 . 73  2 . 5  2 * 0 . 3
L slat   0 . 65  6 . 3 m
2
slat root chord is calculated by the fallowing formula
   1 D fuselage  2  slat 
b 0 slat  b slat  b 0   1 
  

  L 
Here b slat is relative chord of slat is 0.08 to 0.15 = 0.09
 4  1 2 .5  2  0 .3 
There for b 0 slat  0 . 09  4 . 5   1     0 . 363 m
 4 22 . 73 
Slat tip chord length is calculated by fallowing formula
   1 D fuselage  2  slat  2 l slat 
b кslat  b slat  b 0   1 
  

  L 
 4  1 2 .5  2  0 .3  2  6 .3 
b кslat  0 . 09   4 . 5   1     0 . 19 m
 4 22 . 73 
Calculation of control surfaces parameters:
Aileron configuration:
AREA ( SAIL) : It is found by formula S AIL  S AIL  S , from statistical data Relative are of
aileron is 0.06, S AIL  0 .06  63 .87  3 .83 m2
Length of the aileron is calculated by the following formula

46. [45]
L  D fuselage
l aileron   l flap   flap
  AF   Aw , м,
2
here  AF – the gap between flap and aileron= 0.02
 Aw – the gap between aileron and wing tip = 1.2
22 . 73  2 . 5
l aileron   6  0 . 10  0 . 02  1 . 2  2 . 7 m
2
  1 
Aileron root chord b оaileron  b aileron  b 0   1 
  Z оaileron ,
 м,
  
Here b aileron 0.25…0.3 – from the statistical data = 0.25
D fuselage  2  l flap   Af   Aw 
Here Z оaileron  .
L
2 . 5  2  6  0 . 02  1 . 2 
Z оaileron   0 . 74
22 . 73
 4 1 
There for b оaileron  0 . 25  4 . 5   1   0 . 74   0 . 50 m
 4 
  1 
Aileron tip chord b кaileron  b aileron  b 0   1 
  Z кaileron ,
 м,
  
D fuselage  2  l flap  l aileron   af   Aw 
Here Z кaileron  .
L
2 . 5  2  6  2 . 7  0 . 02  1 . 2 
Z кaileron   0 . 98
22 . 73
 4 1 
There for b кaileron  0 . 25  4 . 5   1   0 . 98   0 . 29 m
 4 
Elevator configuration:
Length of the elevator
l H .S
l elevator    HT   Ht , м,
2
Here  HT – is the operating gap of elevator =0.02m
 Ht – is the gap between elevator stabilizer to horizontal stabilizer tip=0.6m
7 . 55
l elevator   0 . 02  0 . 6  3 . 155 m
2
Elevator root chord
   1 2  HT 
b 0 er  b er  b 0 H . S   1  H . S  , м,
  H .S 
 l H .S 

47. [46]
Here b er  0.25…0.35 – relative length of elevator chord = 0.25
 2  1 2  0 . 02 
b 0 er  0 . 25  2 . 48   1     0 . 617 m
 2 7 . 55 
Elevator tip chord:
  H , S  1 l H . S  2  Ht 
b к er  b er  b 0 H . s   1 
  ,

m.
  H .S l H .S 
 2  1 7 . 55  2  0 . 6 
b к er  0 . 25  2 . 48   1     0 . 359 m
 2 7 . 55 
ELEVATOR AREA
S EL  S EL / S HS  0 , 2  0 , 4 (lower values – for supersonic aircraft);
SEL= S EL × SHS = 0.3 × 13.98 m2 = 4.194 m2
Rudder configuration:
rudder area
S RUD  0 ,20  0 ,45 (lower values – for supersonic aircraft);
SRD = S RD × SVS = 0.40 × 9.01m2 = 3.64 m2
Length of the rudder
l rudder  lV . S   Rf   RT , м,
Here  RF – the gap between fuselage to rudder root chord= 0.015m
 RT – the gap between vertical tail tip to rudder tip chord.=0.6m
l rudder  3 . 5  0 . 015  0 . 6  2 . 885 m
Rudder root chord
   1  RF 
b 0 rudder  b rudder  b 0 V . S   1  V . S  , м,
  V .S 
 lV . S 
Here b rudder  0.25…0.4 – relative chord of rudder = 0.25
 1  1 0 . 02 
b 0 rudder  0 . 25  2 . 57   1     0 . 642 m
 1 9 . 56 
For T-tail tip and root chord of rudder is same there for bkrudder= 0.642 m

48. [47]
SELECTION AND GROUNDS OF AIRCRAFT CONFIGURATION
AERODYNAMIC CONFIGURATION
The ―normal" classical configuration applies to my aircraft. The advantages of this
configuration are:
 wing is in the pure, undisturbed airflow and is not shadowed by stabilizers;
 nose section of a fuselage is short and does not create destabilizing moment
relatively to the vertical axis; this allows to reduce area and mass of vertical
stabilizer;
 Crew has better observation of the front semi-sphere.
Selection of wing position relatively to the fuselage
Low-wing aircraft. Advantages:
 due to ground shield effect (aerodrome surface) Ycr increases, Vto, Vland; decreases
 height of the landing gear struts and their mass is less, their retraction becomes
simpler;
 high-lift devices can also be located on ventral wing parts;
 Safety of passengers and crew increases during emergency landing – the wing
provides additional protection;
 Floating capacities during emergency landing on water are higher, that allows to
evacuate passengers and crew.

49. [48]
SELECTION OF WING EXTERIOR SHAPE
Swept wings are applied at M = 0,82.With increase of sweep angle:
 the shockwave drag on moderate subsonic and supersonic speeds (Fig. 2.6) is
drastically decreased.
2
M cr  Ì cr   0
 1  cos  is increased.
 Critical values of flutter speed Vfl increase, divergence speed Vdiv (at swept wings),
lateral stability increases.
WING SHAPE ON FRONT VIEW
It is characterized by wing dihedral angle
Dihedral angle is defined by  angle between wing chords plane and plane
perpendicular to aircraft of symmetry plane passing through the inboard chord. The
following types are distinguished:
At = 0+7 dihedral angle (for straight wings)

50. [49]
SHAPES OF WING CROSS-SECTIONS AND TAIL UNIT STABILIZER CROSS
SECTION
Double convex asymmetrical - high Cy max, smaller Cxр, stable position of the centre of
pressure. These airfoils find wide application in various subsonic aircraft;
The airfoil should have low profile drag in a range of Cy factors characteristic for cruise
flight;
 It is necessary that the airfoil with the extended flap has small Cxp at Cmax, especially
during climb;
 Tip wing cross-sections at Cmax should have smooth performances of shock stall;
 Internal wing cross-sections should have high values of Cmax with extended flaps;
 It is necessary to ensure a high value of Mcr above 0,65;
Airfoils for my aircraft are selected as mentioned below
The NACA four-digit wing sections define the profile by
1. One digit describing maximum camber as percentage of the chord.
2. One digit describing the distance of maximum camber from the airfoil leading edge in
tens of percent's of the chord.
3. Two digits describing maximum thickness of the airfoil as percent of the chord.
WING AIRFOIL - NACA 2415 airfoil has a maximum camber of 2% located 40% (0.4
chords) from the leading edge with a maximum thickness of 15% of the chord. Four-digit
series airfoils by default have maximum thickness at 30% of the chord (0.3 chords) from the
leading edge.

51. [50]
HORIZONTAL STABILIZER AIRFOIL - NACA 2412 airfoil has a maximum camber of
2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of the
chord. Four-digit series airfoils by default have maximum thickness at 30% of the chord (0.3
chords) from the leading edge.
Selection of the scheme of ailerons
b ail  b ail / b  0 ,25  0 ,3 ;  ail .upw  20  25  ; S ail  S ail / S  0 ,03  0 ,08 ;
L ail  L ail / L  0 ,2  0 ,4 ;  ail .downw  10  15  .
Requirements to ailerons:
 Minimum yawing moment (an aircraft yaw motion relatively to OY axis) in bank, with
aircraft turning in the side of bank angle.
 Full weight balancing with least weight of balance weights.
 Provision of efficiency on all flight phases.

52. [51]
 Critical speed of reverse thrust should be sufficient.
SELECTION OF FUSELAGE STRUCTURE
The basic geometrical sizes of the fuselage are selected statistically by comparing the
prototypes and are listed below
Lf – length of a fuselage; 30.62m
Df – diameter of the greatest mid-section, 2.5m
Smcs – the area of fuselage mid-section;
Lns, length of nose section of a fuselage. 3.75m
Lts – tail section of a fuselage. 6.25m
n.f – Fitness ratio of fuselage nose section – 1.5
t.f – Fitness ratio of tail section – 2.5
NOSE SECTION
TAIL SECTION
Shapes of nose and tail sections are also determined generally from conditions of
aerodynamics, layout, technology, purpose. For the nose section, an important condition is
ensuring the demanded observation from the cockpit that leads to smaller fineness and
smaller sharpness. It is reasonable to deflect the tail section of a fuselage upwards to ensure

53. [52]
the landing angle  during take-off. Many cargo aircraft have a large door in the tail section
with the cargo ramp lowered on ground for loading and an unloading of cargos. For modern
aircraft, for aerodynamic drag decrease considerations, the whole tail section is lengthened
and bended. Some cargo aircraft have the cargo door in the fuselage nose (An-124, C-5А,
Boeing 747F). The lower part is capable to open back and upwards that simplifies loading
and unloading of transported vehicles and cargos.
CONTOURS OF FUSELAGE NOSE & TAIL SECTION
Yn.f =  a (Хn.fdn.f/4n.f)
m
.
Хn.f – Length of nose section – 3.75m
dn.f – Diameter of fuselage nose section – 2.5m
n.f – Fitness ratio of fuselage nose section – 1.5
m – 0.5 from table
a – 1 from table
Yn.f =  1 (3.75* 2.5 / 4 * 1.5) 0.5 =  1.25
Factors m and a are presented in Tab
m 0,35 0,4 0,45 0,5 0,55 0,6 0,65
а 1,1293 1,08845 1,04136 1 0,96026 0,9221 0,88546
n
 n 
 
Y   в  Х d  Х  d / 4 
f .ts   f .ts f

 f f .ts 

.
в – 1 from table
Хt.f – Length of tail section – 6.25 m
df. – Diameter of the fuselage – 2.5m
t.f – Fitness ratio of tail section – 2.5
n – 0.50 from table

54. [53]
x – Total length of the fuselage – 30.62m
Y   1  6 . 25 * 2 . 5  30 . 62  0 . 50 2 . 5 / 4 * 2 . 5  0 . 50 = 0.9839
f .ts 
 

Factors n and в are presented in Table
n 0,30 0,35 0,40 0,45 0,50 0,55 0,60 0,65 0,70
в 18,5396 8,9707 4,3174 2,0728 1 0,48122 0,23162 0,11147 0,053649
The point of reference of the tail section is located at distance of l f.ns+l f.ts from a fuselage
nose. Coordinates of Yf.ns and Yf.ts axis are turned counter-clockwise on n.f and t.f angles. If
m and n are selected equal to 0,5 of the contours are usual quadratic parabolas. If smaller
values of m and n are selected, shapes of nose and tail sections will be fuller, if greater values
are selected, the shapes will be more concentrated.
SELECTION OF STABILIZERS STRUCTURE
T-shaped stabilizers: HS is placed off the zone of wing wash on all flight phases,VS is
loaded additionally, its mass is increased
.
DORSAL FIN
1. Dorsal fin installation (fig. 2.35). The dorsal fin improves VS washing, activates
at high speeds, favorably influences side-slip, increases the effective area of VS. Allows to
reduce SVS, mVS, VS.