УДК 629.7Інв. № 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)
 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 masters degree 8.05110101 «Aircraft and helicopter»group 16 E- 2 students 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
 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. 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: companys 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 companys minimal internal funds, determination of point of make out, calculation of direct and indirect costs. 6.3. Total transportation cost value and companys 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 ______________________________________________________
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 defense4. Date of assignment issue:5. Date of final project presentation: Project supervisor (Date, signature) Assignment accepted to fulfillment « » 200 (Date, students signature)
 2012 CONTENTSABSTRACT ………………………………………………………………………… 4INTRODUCTION………………………………………………………………… 5AIRCRAFT DESIGN PROCESS………………………………………………… 7 GENERAL DESIGNING OF AIRCRAFTPURPOSE OF THE AIRCRAFT………………………………………………….11REQUIREMENTS FOR FLIGHT PERFORMANCES………………………….14DESIGN CHART OF THE DESIGNED AIRCRAFT……………………………15PROTOTYPE DATAS………………………………………………………………17SELECTION OF AIRCRAFT MAIN RELATIVE INITIAL PARAMETERS…..24CALCULATION OF AIRCRAFT MASSES THROUGH THE SOFTWARE AND ITSRESULTS……………………………………………………………………………..25ZERO APPROXIMATION…………………………………………………………31STATISTICAL COMPUTATION OF MASSES OF AIRCRAFT………………..32AIRCRAFT OPTIMIZATION AND DESIGN PARAMETERS…………………33SELECTION AND GROUNDS OF AIRCRAFT CONFIGURATION…………..43SELECTION OF ENGINE…………………………………………………………..55AVERAGE BETWEEN GRAPHICAL, SEMI-EMPIRICAL & STATISTICALMETHOD……………………………………………………………………………….63MAXIMUM TAKE-OFF MASS………………………………………………………66CENTER OF GRAVITY………………………………………………………………68DESIGN STRUCTURAL CONFIGURATION …………………………………….83 AERODYNAMICSAIRCRAFT DESIGN PARAMETERS ON AERODYNAMIC CHARACTERISTICS………………………………………..83CALCULATION OF AERODYNAMIC PARAMETERS USING THE SOFTWARE…………………………………………92CALCULATION OF ZERO DRAG COEFFICIENT FOR TAKE-OFF ANDLANDING…………………………….109
 INTEGERATED DESIGN OF AIRCRAFT AND LOAD CALCULATIONAIRCRAFT MASTER GEOMETRY USING UNIGRAPHICS………………….114WING LOAD CALCULATION……………………………………………………..118CALCULATION OF THE DISTRIBUTED FUEL LOAD ON A PLANE WING…119THE WING STRUCTURE MASS LOAD ALLOCATION…………………………125SHEAR FORCE, BENDING MOMENT AND REDUCED MOMENT……………127CALCULATION SCHEME OF REDUCED MOMENT FROM CONCENTRATEDLOADS AND FROM ALL LOADS……………………………………………………132 AIRCRAFT SYSTEMS DESIGN AND SCHEMATIC LAYOUTAIRCRAFT HYDRAULIC SYSTEM……………………………………….134HYDRAULIC FLUID…………………………………………………………135COMPONENTS INVOLVED IN HYDRAULIC SYSTEM…………………137HYDRAULIC DESCRIPTION OF THE DESIGNED AIRCRAFT……….151HYDRAULIC SYSTEM PANEL……………………………………………..154HYDRAULIC SYSTEM MAINTENANCE………………………………….157 MANUFACTURING TECHNOLOGY OF VERTICAL RIBAIRCRAFT RIB……………………………………………………………………..159RIB CONSTRUCTION……………………………………….……………………..163PRODUCTION METHOD OF PARTS OF THE RIB……………………………166ASSEMBLY PROCEDURE OF THE RIB…………………………………………168STAGES OF FORMATION OF RIB DIMENSIONS USING TEMPLATES……171AIRCRAFT VERTICAL ASSEMBLY JIG DESIGN LAYOUT DIAGRAM…….173 ECONOMICAL SECTIONECONOMIC EFFICIENCY CHARACTERISTICS CALCULATION………..174
 SPECIAL ASSIGNMENTINTERIOR CABIN LAYOUT AND SEATING ARRANGEMENT……..…………..179FULFILLING REQUIREMENTS OF THREE ABREAST SEATING LAYOUT….179.CABIN DIMENSIONING FOR 3- ABREAST SEATING…………………………..…179INTERIOR ARRANGEMENT – CROS SSECTION (TYPICAL)……………………180DETERMINATION OF DESIGNED AIRCRAFT CABIN CROSS-SECTION…......181DETERMINATION OF CABIN LENGTH FOR HIGH COMFORT LEVEL………183LIST OF DIAGRAMS3-VIEW DRAWING OF THE AIRCRAFT …………………………………189DESIGN STRUCTURAL LAYOUT………………………………………….188CENTRE OF GRAVITY LOCATION OF THE AIRCRAFT……………..188HYDRAULIC SYSTEM SCHEMATIC OF THE AIRCRAFT…………….190CABIN SEATING LAYOUT OF THE AIRCRAFT………………………..189CONCLUSION …………………………………………………………………. 191REFERENCE ………………………………………………………………………. 192
 ABSTRACT The General design of the aircraft is carried out on basis of collection of aircraftstatistical data and in accordance with the pilot project development task and finally thegeneral view of an aircraft is presented. The main purpose of the aircraft design requirementis fulfilled according to the aviation rules and regulations. The main project is categorized into five part, in the first part aircraft take-off mass inzero approximation is determined and followed by designing the weight of main units, fuel,equipment, control system, geometrical dimensions of a wing, tail units, fuselage, landinggear, location of their center of masses, calculating the aircraft’s center-of-gravity. Finallythe design specifications for the aircraft are presented. The second part is focused on determining aerodynamic forces acting upon thedesigned aircraft. The third part is the development of aircraft unit structure using Computer aideddesigning. The loads acting on the designed aircrafts unit structure is calculated and thematerials for unit structure are selected. The fourth part is concerned about the systems developed for designed aircraft. Theschematic 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 chartsfor assembly of designed aircraft vertical rib structure is done. In special assignment the seating arrangement of the designed aircraft is sketched andinteriors of aircraft components are briefed with the current industry techniques.
 INTRODUCTIONThe purpose of designing a new aircraft is the creation of a structure with uniquecharacteristics, which should be reliable, economical fulfilling the conditions of operation,performance requirements and its primary goal should be attained. To perform thepreliminary design structure of the aircraft it is necessary to be knowledgeable in the field ofgeneral arrangement of aircraft and helicopters, design of power units and systems,construction of elements of assembly structures and units of the aircraft, aerohydrodynamics, durability, technologies, material science, and economics. The purpose ofdesign is to develop a project, realization of which, being limited to a certain extent, wouldensure 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 andmaterial resources and carries out certain useful functions. The functional-structural diagramof the aviation complex is shown on Figure. The aviation complex is an element of statetransport or defense system. All this defines necessity to use systematic approach to aircraftdesign.To implement the process of aircraft design, there was necessity to create specializeddevelopment design offices, which include complicated laboratory and manufacturingresearch. The activities of development design officers are based on work of branch-wiseresearch institutes, which research the prospects of aviation development in variousdirections, and on experience of aircraft production and operation.
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 DEVELOPMENTAIRCRAFT DESIGN PROCESSThe aircraft design process is the steps by which aircraft are designed. These depend onmany factors such as customer and manufacturer demand, safety protocols, physical andeconomic constraints etc. For some types of aircraft the design process is regulated bynational airworthiness authorities. This article deals with powered aircraft such as airplanesand helicopter designs.
Aircraft design is a compromise between many competing factors and constraints andaccounts for existing designs and market requirements to produce the best aircraft.DESIGN CONSTRAINTS IN DESIGNING PROCESSA. Aircraft regulationsAnother important factor that influences the design of the aircraft are the regulations putforth by national aviation airworthiness authorities.Airworthiness Certificates-An airworthiness certificate is an FAA document which grantsauthorization to operate an aircraft in flight.Standard Airworthiness Certificate-A standard airworthiness certificate (FAA form 8100-2displayed in the aircraft) is the FAAs official authorization allowing for the operation of typecertificated aircraft in the following categories: Normal Utility Acrobatic Commuter Transport Manned free balloons Special classes FUNCTIONAL-STRUCTURAL CHART OF THE AVIATION COMPLEXA standard airworthiness certificate remains valid as long as the aircraft meets its approvedtype 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.
Airworthiness Certification Process-The FAA requires several basic steps to obtain anairworthiness 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 operationA. Environmental factorsAn increase in the number of aircraft also means greater carbon emissions. Environmentalscientists have voiced concern over the main kinds of pollution associated with aircraft,mainly noise and emissions. Aircraft engines have been historically notorious for creatingnoise pollution and the expansion of airways over already congested and polluted cities havedrawn heavy criticism, making it necessary to have environmental policies for aircraft noise.Noise also arises from the airframe, where the airflow directions are changed. Improvednoise regulations have forced designers to create quieter engines and airframes. Emissionsfrom aircraft include particulates, carbon dioxide (CO2), Sulphur dioxide (SO2), Carbonmonoxide (CO), various oxides of nitrates and unburnt hydrocarbons. To combat thepollution, ICAO set recommendations in 1981 to control aircraft emissions. Newer,environmentally friendly fuels have been developed and the use of recyclable materials inmanufacturing have helped reduce the ecological impact due to aircraft. Environmentallimitations also affect airfield compatibility. Airports around the world have been built to suitthe topography of the particular region. Space limitations, pavement design, runway endsafety areas and the unique location of airport are some of the airport factors that influenceaircraft design.B. SafetyThe 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 thatpose a threat to air travel. Airworthiness is the standard by which aircraft are determined fitto fly. 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 recommendedpractices for national authorities to base their regulations on The national regulatoryauthorities set standards for airworthiness, issue certificates to manufacturers and operatorsand the standards of personnel training. Every country has its own regulatory body such asthe Federal Aviation Authority in USA, DGCA (Directorate General of Civil Aviation) inIndia, etc.
C. Design optimizationAircraft designers normally rough-out the initial design with consideration of all theconstraints on their design. Historically design teams used to be small, usually headed by aChief Designer who knew all the design requirements and objectives and coordinated theteam accordingly. As time progressed, the complexity of military and airline aircraft alsogrew.D. Design aspectsThe main aspects of aircraft design are: 1. Aerodynamics 2. Propulsion 3. Controls 4. Mass 5. StructureAll aircraft designs involve compromises of these factors to achieve the design mission.E. Computer-aided design of aircraftIn the early years of aircraft design, designers generally used analytical theory to do thevarious engineering calculations that go into the design process along with a lot ofexperimentation. These calculations were labor intensive and time consuming. In the 1940s,several engineers started looking for ways to automate and simplify the calculation processand many relations and semi-empirical formulas were developed. Even after simplification,the calculations continued to be extensive. With the invention of the computer, engineersrealized that a majority of the calculations could be done by computers, but the lack ofdesign visualization and the huge amount of experimentation involved kept the field ofaircraft design relatively stagnant in its progress.F. Financial factors and marketBudget limitations, market requirements and competition set constraints on the designprocess and comprise the non-technical influences on aircraft design along withenvironmental factors. Competition leads to companies striving for better efficiency in thedesign without compromising performance and incorporating new techniques andtechnology.
 DESIGN SECTION PART-1 COMPUTER AIDED GENERAL DESIGNING OF AIRCRAFT1. PURPOSE OF THE AIRCRAFT2. REQUIREMENTS FOR FLIGHT PERFORMANCES3. DESIGN CHART OF THE DESIGNED AIRCRAFT4. PROTOTYPE DATAS5. SELECTION OF AIRCRAFT MAIN RELATIVE INITIAL PARAMETERS6. CALCULATION OF AIRCRAFT MASSES THROUGH THE SOFTWARE AND ITS RESULTS7. ZERO APPROXIMATION8. STATISTICAL COMPUTATION OF MASSES OF AIRCRAFT9. AIRCRAFT OPTIMIZATION AND DESIGN PARAMETERS10. SELECTION AND GROUNDS OF AIRCRAFT CONFIGURATION11. SELECTION OF ENGINE12. DETERMINATION OF CENTER OF GRAVITY OF THE AIRCRAFT13. AVERAGE BETWEEN GRAPHICAL, SEMI-EMPIRICAL & STATISTICAL METHOD14. MAXIMUM TAKE-OFF MASS15. CENTER OF GRAVITY16. DESIGN STRUCTURAL CONFIGURATION17. DESIGN SPECIFICATION
 PURPOSE OF THE DESIGNING AIRCRAFT A. Aircrafts Intended Purpose - Commercial usageCommercial usage denotes using the aircraft for a business purpose or gettingdirectly/indirectly financial gain from it. B. Payload category - PassengersAircraft adapted for carrying passengers. C. Type - Regional jetThe 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 theUnited States in 1978.ExamplePRIMARY USERS MANUFACTURER ROLEAeroflot Yakolev Yak-40 regional sized mini-jet airlinersSkyWest Airlines Bombardier CRJ100 Regional jet/Business jetPinnacle AirlinesExpressJetComairAerosvit Airlines Antonov An-148 regional jetRossiya D. Range - Short-rangeshort range refers to distance travelled is between 2500.2 km (less than 1350nm) and Timetaken to travel is less than 5 hoursExampleFROM & TO DISTANCE in km TIMENew York-Miami 2051.914 2 hours 49 minsTokyo-Seoul 1,159.04 1.5 to 2 hoursDenver-Boston 2800 3hrs 42 minsG. Special Requirements - Cargo Carrying capabilityCan be used to carry cargos and can be used as a cargo variant
H. Mode of Class - Economy classEconomy class refers to the seating arrangement of the aircraft which is usually reclined andinclude a fold-down table. The seats pitch range from 29 to 36 inches (74 to 91 cm), usually30–32 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 performancespecification. 6.Simulators and training devices should be designed for aircraft according to individual
performance specifications. The programs for training of flight and technical staff should bedeveloped 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 PERFORMANCESMaximum passenger capacity withdistance between the seats 750 person 55(762) mm,Maximum payload kg 5000Cruise speed:at long range cruise km/h 835maximumCruise altitude, km 10.5Required length of RWY (SA, Н =0, dry concrete), m 1950for takeoff: 2250for landing: Applied flight range (emergencyfuel reserve for 0.75 hour of flight; km 2500takeoff in SA conditions; Н = 0)with maximum payloadFuel consumption for 1 pass/km g 340while flying for technical rangewith maximum payloadMaintenance and overhaul, 8.8
manhourREQUIREMENTS 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 layoutSpecial activity
 STATISTICAL DATA COLLECTIONStatistical data collection is the process of collecting flight, mass, power plant andgeometrical data’s of required prototypes for the design project. In this project I havecollected four different aircraft data’s and their features are explained and tabulated. Theseaircrafts are selected based upon the design requirements and design specification mentionedbelow,TACTICAL TECHNICAL REQUIREMENTS OF THE DESIGNING AIRCRAFTMaximum speed , Vmax 835 km/hCruising speed, Vcruise 760 km/hCruising height, Нcruise 11 kmNumber of passengers, npass 47Number of crew members, ncrew 4Range, L 2000 kmTake-off distance, Lр 1970 mVertical speed, Vy 14 m/sMaximum take-off weight, Ммах 40 tonsDESIGN SPECIFICATION OF THE AIRCRAFTType of the aircraft - Transport category with capacity to carry 47 to 55 passengers including crewAerodynamic configuration Normal configuration with horizontal stabilizer on tail sectionWing Low wing with Dihedral and wing sweepTail T-tail configurationFuselage Cylindrical shapePower plant type Turbofan located at aft part of the fuselageLanding gear Tricycle configuration with nose wheelBased on the tactical technical requirements and the general design specification of thedesigning aircraft we are gathering the similar aircrafts and their detailed specification istabulated. From the critical parameters of the aircraft are listed. With the obtained results wenow ready to input all parameters in the software which would give all the relative massesand 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
by scaling the three view picture of the collected aircraft. In obtaining the details it isimportant to have the three view pictures of each aircraft for simplification further in drawingthe designed aircraft three view it is very helpful .A short brief of the aircraft is provided foreach 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 furthersimplification. I have selected the following aircrafts for comparison, 1.EMBRAER ERJ 1452. BOMBARDIER CRJ100 3. TUPOLEV 134-A 4. BOEING 717-200My main prototype is EMBRAER ERJ 145AIRCRAFTS SELECTED FOR STATISTICAL DATA COLLECTION AND THEIRPARAMETERSEMBRAER ERJ 145The Embraer ERJ 145 family is a series of regional jets produced by Embraer, a Brazilianaerospace company. Family members include the ERJ 135 (37 passengers), ERJ 1(44passengers), and ERJ 145 (50 passengers). The key features of the production designincluded: 1. Rear fuselage-mounted engines 2. Swept wings (no winglets) 3. "T"-tail configuration 4. Range of 2500 km
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 passengerMilitary models C-99A - Transport model EMB 145SA (R-99A) - Airborne Early Warning model EMB 145RS (R-99B) - Remote sensing modelBOMBARDIER CRJ100The Bombardier CRJ100 and CRJ200 are a family of regional airliners manufactured byBombardier, 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 ofthe wing, two more emergency exit doors, plus a reinforced and modified wing. Typicalseating was 50 passengers, the maximum load being 52 passengers. The CRJ100 featured aCollins ProLine 4 avionics suite, Collins weather radar, GE CF34-3A1 turbofans with41.0 kN (4,180 kgp / 9,220 lbf), new wings with extended span, more fuel capacity, and
improved landing gear to handle the higher weights. It was followed by the CRJ100 ERsubvariant with 20% more range, and the CRJ100 LR subvariant with 40% more range thanthe standard CRJ100. The CRJ 100 SE sub-variant was produced to more closely meet theneeds of corporate and executive operators.VariantsSeveral models of the CRJ have been produced, ranging in capacity from 40 to 50passengers. The Regional Jet designations are marketing names and the official designationis CL-600-2B19.CRJ100 -The CRJ100 is the original 50-seat version. It is equipped with General ElectricCF34-3A1 engines. Operators include Jazz Aviation, Comair and more.CRJ200 -The CRJ200 is identical to the CRJ100 except for its engines, which were upgradedto the CF34-3B1 model, offering improved efficiency.CRJ440 -Certified up to 44-seat, this version was designed with fewer seats in order to meetthe needs of some major United States airlines.Challenger 800/850 - A business jet variant of the CRJ200TUPOLEV 134-A
The Tupolev Tu-134 (NATO reporting name: Crusty) is a twin-engined airliner, similar tothe French Sud Aviation Caravelle and the later-designed American Douglas DC-9, and builtin 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 operatefrom unpaved airfields.Design and developmentFollowing the introduction of engines mounted on pylons on the rear fuselage by the FrenchSud Aviation Caravelle, airliner manufacturers around the world rushed to adopt the newlayout. Its advantages included clean wing airflow without disruption by nacelles or pylonsand decreased cabin noise. At the same time, placing heavy engines that far back createdchallenges with the location of the center of gravity in relation to the center of lift, which wasat 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-heavyarrangement.VariantsTu-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.134UBLTu134UBK Naval version of Tu-134UBL. Only one was ever built.BOEING 717-200Boeing 717 was specifically designed for the short-haul, high frequency 100-passengerairline market. The highly efficient 717 concluded its production run in May 2006, thoughthe 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 airplanewas originally part of the McDonnell Douglas airplane family and designated the MD-95
prior to merger with The Boeing Co. in 1997. The program produced 156 717s and pioneeredbreakthrough business and manufacturing process for Boeing Commercial Airplanes The.The standard 717 has a two-class configuration with 106 seats. Its passenger-pleasing interiorfeatures a five-across-seating arrangement in economy class, with illuminated handrails andlarge overhead stow bins.The two-crew flight deck incorporates six interchangeable liquid-crystal-display units andadvanced Honeywell VIA 2000 computers.Flight deck features include an Electronic Instrument System, a dual Flight ManagementSystem, a Central Fault Display System, and Global Positioning System. Category IIIbautomatic landing capability for bad-weather operations and Future Air Navigation Systemsare available.Two advanced Rolls-Royce 715 high-bypass-ratio engines power the 717. The engine israted at 18,500 to 21,000 pounds of takeoff thrust, with lower fuel consumption andsignificantly lower noise and emission levels than the power plants on comparable airplanes.DESIGNThe 717 features a two-crew cockpit that incorporates six interchangeable liquid-crystal-display units and advanced Honeywell VIA 2000 computers. The cockpit design is calledAdvanced Common Flight deck (ACF) and is shared with the MD-11.
 STATISTICAL TABLEFLIGHT DATA:Flight includes Vmax – the maximum speed of flight; HV max – flight altitude with themaximum 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 thetake-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 WikipediaFLIGHT 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 1.52MASS DATA : This includes take-off mass(m0), maximum take-off mass(m0max), payloadmass(mpld), number of passengers(npass), landing mass(mland), empty mass(mempty), mass ofcrew(mc), mass of fuel(mf), empty equipped mass(mempt.eqpd) and total mass(mtotal ).
 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 22000MASS 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 49,900POWERPLANT DATA: This includes engine thrust (P0), mass of engine (meng), number ofengines 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 - - ratioGEOMETRICAL 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
 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 200DERIVATIVE 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
 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 PARAMETERSThus finally tabulating all the required values it is necessary to find the main relative initialparameters of wing, fuselage and the tail unit. The obtained result is used in the software. WING PARAMETERaspect ratio, 7.85sweep angle, 22.73taper ratio, 4relative width of airfoil, c 18 or 0.18relative chord of flap, b f b f / b wing 0.25deflection angles of flap, f 18relative area of ailerons, S ail S ail / S 0.06FUSELAGE PARAMETERfineness ratio f 12.25fuselage diameter Df 2.50TAIL UNIT PARAMETERrelative area of horizontal stabilizer, S HT S HT S 0.219relative area of vertical stabilizer, S VT S VT S 0.141aspect ratio of horizontal surface, HS 4.077aspect ratio of vertical surface, VS 1.36Sweep angle of horizontal surface, HS 20Sweep angle of horizontal surface, HS 32Relative thickness of horizontal surface, c HS 12 or 0 .12Relative thickness of vertical surface, c VS 10 or 0.10
CALCULATION OF AIRCRAFT MASSES THROUGH THE SOFTWARE AND ITSRESULTSAircraft masses in zero approximation are calculated using software by entering necessaryparameters taken from statistical data and the initial parameters. First the relative masses forfuselage, wing, power plant, tail unit, fuel and landing gear are found with respect to theaspect ratio of wing taken as 4 and aircrafts wing loading attained from graphical result as600 N/m2. The graphs obtained from this result help us to select the desired wing loading andfrom the wing the lowest value of takeoff mass is taken as the final one. The relative massesare changed to direct masses by multiplying it with the finally obtained take-off mass, for meit is 39.11tons. Therefore my relative masses are multiplied with 39.11 ton to get direct massGraphs for each lab are plotted versus each parameter and from that the final wing loading isobtained followed by the takeoff mass. Other parameters obtained include engineperformance data like thrust to weight ratio at take-off, climbing and cruise. My maincomparative parameter is Aspect ratio which I took in three variations asASPECT RATIO 2 , 4 AND 6AIM OF THE LAB AND ITS RESULT TO FIND THE RELATIVE MASS OFAIRCRAFTLAB 5: In this part we are finding the relative mass of power plant respect to the aspectratio of wing taken as 4 and aircrafts wing loading 600 N/m2.
RESULT: In table P,denotes wing loading Tk,aspect ratio and SU is the RELATIVE MASSOF POWER PLANT which is 0.086LAB 7A:In this part we are finding the relative mass of wing respect to the aspect ratio ofwing 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 RELATIVEMASS OF wing which is 0.048.
LAB 7B:In this part we are finding the relative mass of fuselage respect to the aspect ratioof 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 ratioof the fuselage (12.25). By comparing both the values we get the relative value of fuselageequals to 0.350.LAB 7G:In this part we are finding the relative mass of tail unit with respect to the aspectratio 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 THETAIL UNIT which is 0.0182
LAB 7V:In this part we are finding the relative mass of landing gear with respect to theaspect ratio of wing taken as 4 and aircrafts wing loading p,600 N/m2.RESULT: RELATIVE MASS OF THE LANGING GEAR is 0.062LAB 8: In this part we are finding the mass of equipment, crew and payload.RESULT: MASS OF THE Equipment, crew and payload is 9502.98kgLAB 9: In this part we are finding the take off mass of the aircraft which is equal to39.11tons.
RESULT: Take off MASS OF is 39110 kg or 39.11 tonsRESULTS FROM GRAPH WITH RESPECT TO THE RELATIVE PARAMETERLAB PARAMETERS RESULTSNO3 Lift to drag ratio 12.204 Thrust to weight ratio at Take-off 0.2514 Thrust to weight ratio at Landing 0.2714 Thrust to weight ratio at Cruising 0.1795 Mean Thrust to weight 0.2715 Relative mass of powerplant 0.0867a Relative mass of wing 0.0487b Relative mass of fuselage 0.3507g Relative mass of Tail unit 0.01827B Relative mass of Landing gear 0.0626 Relative mass of Fuel 0.2248 MEQ = MCREW+MPAYLOAD+MEQUIPMENT 9502.98 kg9 Take-off mass relative to wing loading 39110 kg
DIRECT MASSMass of fuselage 13688.5 kgMass of wings 1877.28 kgMass of tail unit 711.802 kgMass of powerplant 3363.46 kgMass of landing gear 2424.82 kgMass of fuel 760.64 kgCOMPUTATION OF AIRPLANE TAKE-OFF MASS IN ZERO APPROXIMATIONDETERMINATION OF MASS FROM LAB RESULTSMass of fuselage = Relative mass of fuselage * Take off mass = 0.350 * 39110 kg = 13688.5 kgMass of wings = Relative mass of wing * Take off mass = 0.048* 39110 kg = 1877.28 kgMass of tail unit = Relative mass of tail unit * Take off mass = 0.0182* 39110 kg = 711.802 kgMass of power plant = Relative mass of power plant * Take off mass = 0.086* 39110 kg = 3363.46 kgMass of landing gear = Relative mass of landing gear * Take off mass = 0.062* 39110 kg = 2424.82 kgMass of Fuel = Relative mass of fuel * Take off mass = 0.224* 39110 kg = 8760.64 kgMass of crew = 4 * 80 kg = 320kgMass of payload = 47 * 90 kg = 4230kgMass of Equipment and control systems = 4952kg ZERO APPROXIMATIONTake-off mass of the airplane for zero approximation is determined by the formula receivedfrom the equation of mass ratio with statistical data.
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 plantmass,m f = Fuel mass, m pl = Payload mass, m crew = Crew mass, m eq = Equipment mass m pl m crewMass Ratio (dimensionless) equation is, 1 m st m p. p m f m eq m0Re-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 = 0.4782m p. p - Relative mass of power plant = 0.086m f - Relative mass of fuel = 0.224m eq - Relative mass of Equipment = 0.1266 4230 320m0 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 fuscalculate airframe mass and its components (mass of the wing , fuselage , m tail unit m fuel m pow . pltail unit , landing gears), and also mass of fuel , power plant and mengines eng . Relative masses of airframe, power plant, equipment and control system, andalso 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
Take off mass from lab 9 = 39110 kgRelative mass of Airframe = 0.30Mass of Airframe = Relative mass of airframe * Take off mass = 0.30 * 39110 kg = 12515.2 kgRelative mass of power plant = 0.14Mass of power plant = Relative mass of power plant * Take off mass = 0.14* 39110 kg = 5475.4 kgRelative mass of control systems and equipments = 0.11209Mass of control sys = Relative mass of control systems and equipments * Take off mass = 0.11209* 39110 kg = 4383.839 kgRelative mass of fuel = 0.18Mass of fuel = Relative mass of fuel * Take off mass = 0.18* 39110 kg = 7039.8 kg AIRCRAFT OPTIMIZATION AND DESIGN PARAMETERSGeometrical parameters for designed aircraft are calculated by formulas taken from pilotproject book and rest is determined statistically by comparing with the prototypes. Aftercalculating the geometrical parameters we are drawing the theoretical drawing. Thegeometrical parameters are calculated and obtained satisfying the general requirements of theaircraft.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 parametersDetermination of Wing Parameters:In determining Wing parameters its plan form shape is very important in obtaining number ofuseful 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 totake off mass of 39110kg, which is found in LAB 9 as given below,
RESULT: The wing loading is found to be 600 daN/m2Wing statistical parameter aspect ratio, sweep angle, taper ratio, 7.85 22.73 4WING AREA m0 gS Where m 0 39110 ( kg ) , g 9 . 8 ( m / s 2 ) , p 0 600 ( dN / m ) 2 10 p 0 39110 9 . 8S 63 . 87 ( m ) 2 10 600Wing Span ( l )L S Where λ= 7.85 (choosing from table)L 7 . 85 63 . 87 = 22.39m
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 .5b tip 0 1 . 1( m ) 4Quarter 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 1tg 0 tg 0 . 25 tg 22 . 73 0 . 4953 0 1 7 . 85 4 1 0 arctg 0 . 4953 26 . 349 0Mean Aerodynamic Chord of the Wing (MAC = b Aw ) 1 2 2bA b0 3 1 4 4 1 2 2bA 4 .5 3 . 15 ( m ) 3 4 4 1Vertical distance between horizontal central line to MAC ( z A ) L 2 22 . 39 4 2zA 4 . 478 ( m ) 6 1 6 4 1Horizontal distance between wing root tips to MAC root l 2 22 . 39 4 2xA tg 0 . 0 . 4953 2 . 2179 ( m ) 6 1 6 4 1Determination of Fuselage Parameters The size and shape of subsonic commercial aircraft are generally determined by thenumber of passengers, seating arrangements and cargo requirements. Seating arrangementson 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
Overall Fuselage length, LfL f f D f 12 . 25 2 . 5 30 . 62 ( m )Fuselage nose length, L f .nL f . n N D f 1 . 5 2 . 5 3 . 75 ( m )Fuselage rear length, L f .rL f . r T D f 2 . 5 2 . 5 6 . 25 ( m )Fuselage middle section length, L f .mLfm = Lf – Lf.n – Lf.r = 30.62 - 3.75 - 6.25 = 20.62 mDetermination of Tail Unit ParametersTail unit parameters include Horizontal and Vertical stabilizer. Their geometrical parametersare determined by the same formulae which were used for the wing.Horizontal stabilizer parameterHorizontal stabilizer statistical parameteraspect ratio, sweep angle, taper ratio, Relative horizontal stabilizer area, S h .t 4.077 20 2 0.219Horizontal stabilizer areaS h .t S h .t S 0 . 219 63 . 87 ( m ) 13 . 98 m 2 2 Where S =63.89, wing areaLength of the Horizontal stabilizerL 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 . 48b tip . ht 1 . 24 ( m ) ht 2 .0
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 2b A . ht b 0 . ht 2 . 48 1 . 93 ( m ) 3 ht ht 1 3 2 2 1Vertical Distance between horizontal central line to MAC ( z A . ht ) L ht ht 2 7 . 55 2 2z A . ht 1 . 68 ( m ) 6 ht 1 6 2 1Leading edge sweep angle of the Horizontal stabilizer ( 0 ) 1 2 1tg 0 tg 0 . 25 tg 20 0 . 44572 0 1 4 . 077 2 1 0 arctg 0 . 44572 24 . 023 0Horizontal distance between wing root tip to MAC root ( x A . h .t ) L ht ht 2 7 . 55 2 2x A . ht tg 0 . . 0 . 4452 0 . 7488 ( m ) 6 ht 1 6 2 1x A . ht 0 . 30 ( m )Vertical stabilizer parameterVertical stabilizer statistical parameteraspect ratio, sweep angle, taper ratio, vs Relative horizontal stabilizer area, S h .t 1.36 32 1 0.141Vertical stabilizer areaS v .t S v .t S 0 . 141 63 . 87 ( m ) 9 . 01 m 2 2 Where S =63.89, wing area
Length of the Vertical stabilizerL 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 11 b 0 .vt 2 . 57b tip .vt 2 . 57 ( m ) vt 1Quarter 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 11 2 2 2 vt 2b A . vt b 0 . vt 2 . 57 2 . 57 ( m ) 3 vt vt 1 3 11 1 Vertical Distance between horizontal central line to MAC ( y A .v .t ) L v .t v .t 2 3 . 50 1 2y A . v .t 1 . 76 ( m ) 3 v .t 1 3 11Leading edge sweep angle of the Vertical stabilizer ( 0 ) 1 11tg 0 tg 0 . 25 tg 32 0 . 62 0 1 1 . 36 1 1 0 arctg 0 . 62 31 . 79 0Horizontal 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 PARAMETERSFor nose-wheel tri-cycle landing gear the following parameters are consideredWheel base (b ) : distance between axels of nose wheel and main landing gear wheels inside view. It depends on fuselage length.b (0.30.5)l fus,b = 0.48 * 30.62 = 14.6976mNose wheel offset (a ) : It is distance between vertical line passing through the airplane centerof gravity and nose wheel axis (or axis of several wheels whenever);a = 0.968 * 14.697 = 14.235m
Main Landing Gear offset (e) : It is distance (on side view) between vertical line passingthrough the airplane center of gravity and axis (or center line of several wheels, bogie) ofMLG;e = 0.031 * 14.697 = 0.462mStatic ground angle() : It is the angle between fuselage construction plane and runwaysurface. It is generally between2 to +2. So it is taken as1.Angle of wing setting sett : It is the angle formed between wing construction plane to thefuselage axis. It is generally between sett (04). So it is taken be 2.Angle of overturning(): It is the angle appearing when fuselage tail part or its tail bumptouches the runway surface;Ψ=2 parking angleamax = 12 maximum angle of attackaw = -1 angle between a wing chord and longitudinal axis of fuselage.Ф = amax –(-1) – 2 = 12+1-2 = 11As a rule = 1018, 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.Itprevents airplane overturning backward during landing. = + (12). =11+ 2 = 13.Wheel track(В ) : It is the distance (on front view) between planes of symmetry of MLGwheels. This can be found byB ( 0 . 15 .... 0 . 35 ) LWB 0 . 21 22 . 39 4 . 7 mHeight 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 mHeight of the landing gear (h): Distance from leg attachment fittings to the runway surfacewhen shock absorber and tiers compression is of parking state (at take-off mass).DF = 2.5 m h =H – DF/2h = 2.1 – (2.5/2)
h = 0.85 mDetermination of Position of Center of Mass of the AirplanePosition of the airplane center of mass is determined relative to nose part of the wing meanaerodynamic chord (MAC).The recommended distance for the center of mass from the nose part of mean aerodynamicchord 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 armsVertical tail unit arm:It is the distance measured from the aircraft center of massup to the vertical tail unit centre ofpressure. 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 unitcentre 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 . 8I considered relative length of flap = 0.6The length of the flap is calculated by the formula L span D fuselage 2 L flap L flap L flap 2ΔLflap- is the gap between flap and fuselage = 100mm 22 . 73 2 . 5 2 * 0 . 100 L flap 0 .6 6 m 2Flap 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
 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 .85I considered relative length of slat = 0.65The length of the slat span is calculated by below formula L span D fuselage 2 L slat L slat L slat 2ΔLslat- is the gap between slat and fuselage = 300mmThere for 22 . 73 2 . 5 2 * 0 . 3 L slat 0 . 65 6 . 3 m 2slat 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 ofaileron is 0.06, S AIL 0 .06 63 .87 3 .83 m2Length of the aileron is calculated by the following formula
 L D fuselage l aileron l flap flap AF Aw , м, 2here 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, 2Here 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 2Elevator root chord 1 2 HT b 0 er b er b 0 H . S 1 H . S , м, H .S l H .S
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 AREAS 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 m2Rudder configuration:rudder areaS RUD 0 ,20 0 ,45 (lower values – for supersonic aircraft);SRD = S RD × SVS = 0.40 × 9.01m2 = 3.64 m2Length 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 mRudder 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
 SELECTION AND GROUNDS OF AIRCRAFT CONFIGURATIONAERODYNAMIC CONFIGURATIONThe ―normal" classical configuration applies to my aircraft. The advantages of thisconfiguration 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 fuselageLow-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.
SELECTION OF WING EXTERIOR SHAPESwept wings are applied at M = 0,82.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 VIEWIt is characterized by wing dihedral angle Dihedral angle is defined by angle between wing chords plane and planeperpendicular to aircraft of symmetry plane passing through the inboard chord. Thefollowing types are distinguished:At = 0+7 dihedral angle (for straight wings)
SHAPES OF WING CROSS-SECTIONS AND TAIL UNIT STABILIZER CROSSSECTIONDouble convex asymmetrical - high Cy max, smaller Cxр, stable position of the centre ofpressure. These airfoils find wide application in various subsonic aircraft;The airfoil should have low profile drag in a range of Cy factors characteristic for cruiseflight; 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 belowThe 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 percents 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% (0.4chords) from the leading edge with a maximum thickness of 15% of the chord. Four-digitseries airfoils by default have maximum thickness at 30% of the chord (0.3 chords) from theleading edge.
HORIZONTAL STABILIZER AIRFOIL - NACA 2412 airfoil has a maximum camber of2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of thechord. Four-digit series airfoils by default have maximum thickness at 30% of the chord (0.3chords) 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.
 Critical speed of reverse thrust should be sufficient.SELECTION OF FUSELAGE STRUCTUREThe basic geometrical sizes of the fuselage are selected statistically by comparing theprototypes and are listed belowLf – length of a fuselage; 30.62mDf – diameter of the greatest mid-section, 2.5mSmcs – the area of fuselage mid-section;Lns, length of nose section of a fuselage. 3.75mLts – tail section of a fuselage. 6.25mn.f – Fitness ratio of fuselage nose section – 1.5t.f – Fitness ratio of tail section – 2.5 NOSE SECTION TAIL SECTION Shapes of nose and tail sections are also determined generally from conditions ofaerodynamics, layout, technology, purpose. For the nose section, an important condition isensuring the demanded observation from the cockpit that leads to smaller fineness andsmaller sharpness. It is reasonable to deflect the tail section of a fuselage upwards to ensure
the landing angle during take-off. Many cargo aircraft have a large door in the tail sectionwith the cargo ramp lowered on ground for loading and an unloading of cargos. For modernaircraft, for aerodynamic drag decrease considerations, the whole tail section is lengthenedand 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 loadingand unloading of transported vehicles and cargos.CONTOURS OF FUSELAGE NOSE & TAIL SECTION Yn.f = a (Хn.fdn.f/4n.f) m .Хn.f – Length of nose section – 3.75mdn.f – Diameter of fuselage nose section – 2.5mn.f – Fitness ratio of fuselage nose section – 1.5m – 0.5 from tablea – 1 from tableYn.f = 1 (3.75* 2.5 / 4 * 1.5) 0.5 = 1.25Factors 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 tableХt.f – Length of tail section – 6.25 mdf. – Diameter of the fuselage – 2.5mt.f – Fitness ratio of tail section – 2.5n – 0.50 from table
x – Total length of the fuselage – 30.62mY 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 0,053649The point of reference of the tail section is located at distance of l f.ns+l f.ts from a fuselagenose. Coordinates of Yf.ns and Yf.ts axis are turned counter-clockwise on n.f and