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Aircraft Design Project Eight Seater Short Range Business Jet Aircraft
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EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT
AN AIRCRAFT DESIGN
P.VASANTHA PRABHU (30609101059)
J.SELVA KUMAR (30609101051
in partial fulfillm
BACHELOR OF ENGINEERING
JEPPIAAR ENGINEERING COLLEGE, CHENNAI 600 119
ANNA UNIVERSITY: CHENNAI 600 025
EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT
AN AIRCRAFT DESIGN PROJECT-I REPORT
Submitted by
S.VIGNESH (30609101062)
P.VASANTHA PRABHU (30609101059)
J.SELVA KUMAR (30609101051)
N.VIGNESH (30609101061)
in partial fulfillment for the award of the degree
of
BACHELOR OF ENGINEERING
in
AERONAUTICAL
JEPPIAAR ENGINEERING COLLEGE, CHENNAI 600 119
ANNA UNIVERSITY: CHENNAI 600 025
APRIL/MAY 2012
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EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT
REPORT
degree
JEPPIAAR ENGINEERING COLLEGE, CHENNAI 600 119
ANNA UNIVERSITY: CHENNAI 600 025
JEPPIAAR ENGINEERING COLLEGE
JEPPIAAR NAGAR, RAJIV GANDHI SALAI, CHENNAI
Certified that this AIRCRAFT DESIGN PROJECT
RANGE BUSINESS JET AIRCRAFT
carried out the project work
HEAD OF THE DEPARTMENT INTERNAL EXAMINER
JEPPIAAR ENGINEERING COLLEGE
JEPPIAAR NAGAR, RAJIV GANDHI SALAI, CHENNAI – 600 119.
BONAFIDE CERTIFICATE
AIRCRAFT DESIGN PROJECT-1 report “EIGHT SEATER SHORT
RANGE BUSINESS JET AIRCRAFT” is the bonafide work of
on 28/04/12 under my supervision.
HEAD OF THE DEPARTMENT
INTERNAL EXAMINER
SUPERVISOR
EXTERNAL EXAMINER
2
JEPPIAAR ENGINEERING COLLEGE
EIGHT SEATER SHORT
the bonafide work of “ S.VIGNESH ” who
SUPERVISOR
EXAMINER
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ACKNOWLEDGEMENT
It gives us immense pleasure in expressing our sincere gratitude to Honourable Dr.Jeppiaar, M.A., B.L., Ph.D., founder and Chairman of Jeppiaar Engineering College for bestowing us with an opportunity to bring out this project as a successful one.
We are very much grateful to our principal Dr.Susil Lal Das, M.Sc., Ph.D., for their encouragement and moral support.
We are very much indebted to Mr.G.Prabakaran (HOD) Aeronautical Department for giving me his able support and encouragement.
At this juncture I must emphasis the point that this AIRCRAFT DESIGN PROJECT-I would not have been possible without the highly informative and valuable guidance by our respected preceptor Mr.G.Rajeuv and Mr.Nelson Maxwell whose vast knowledge and experience has must us go about this project with great ease. We have great pleasure in expressing our sincere & whole hearted gratitude to them. It is worth mentioning about my team mates, friends and colleagues of the Aeronautical department, for extending their kind help whenever the necessity arose. I thank one and all who have directly or indirectly helped me in making this design project a great success.
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INDEX
Sl.No. Topic Page No
Bonafide ii Acknowledgement iii Index iv List of symbols & abbreviation v List of figure vi Aim of project viii Abstract ix 1 Introduction 1 2 Design Methodology 3 3 Comparative data & Mean design parameter 14 4 Weight estimation 29 5 Power plant Selection 37 6 Wing Selection 42 7 Airfoil Selection 45 8 Flap selection 59 9 Fuselage and Cabin layout 63
10 Centre of gravity 67 11 Tail selection 69 12 Lift Estimation 72 13 Drag Estimation 77 14 Landing Gear Arrangement 81 15 Performance Characteristics 91 16 Three views of Aicraft 101
Conclusion 102 Future 104 Bibliography 105 Website references
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LIST OF SYMBOLS & ABBREVIATION A.R - Aspect Ratio b - Wing Span (m) C - Chord of the Airfoil (m) C root - Chord at Root (m) C tip - Chord at Tip (m) Cm - Mean Aerodynamic Chord (m) C CD - Drag Co-efficient ��� - Zero Lift Drag Co-efficient Cp - Specific fuel consumption (lbs/hp/hr) CL - Lift Co-efficient D - Drag (N) E - Endurance (hr) e - Oswald efficiency L - Lift (N) M - Mach number of aircraft Mff - Mission fuel fraction R - Range (km) Re - Reynolds Number S - Wing Area (m²) Sref - Reference surface area Swet - Wetted surface area Sa - Approach distance (m) Sf - Flare Distance (m) Sfr - Free roll Distance (m) Sg - Ground roll Distance (m) T - Thrust (N) Tcruise - Thrust at cruise (N) Ttake-off - Thrust at take-off (N) Vcruise - Velocity at cruise (m/s) Vstall - Velocity at stall (m/s) Vt - Velocity at touchdown (m/s) Wcrew - Crew weight (kg) Wempty - Empty weight of aircraft (kg) Wfuel - Weight of fuel (kg) Wpayload- Payload of aircraft (kg) W0 - Overall weight of aircraft (kg) W/S - Wing loading (kg/m²)
� - Density of air (kg/m³) µ - Dynamic viscosity (Ns/m²) λ - Tapered ratio
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LIST OF FIGURES
SL.NO FIGURE DESCRIPTION
Fig.1 Business jet aircraft Fig.2 Phase of design Fig.3 The preliminary design flowchart Fig.4 Student view of design Fig.5 The real design process Fig.6 Wing lift and drag to lift of an Airfoil Fig.7 Flowchat of Preliminary Design and Opimization Fig.8 Cruise Speed Vs Service Ceiling Fig.9 Cruise Speed Vs Take Off Distance Fig.10 Cruise Speed Vs Thrust Fig.11 Cruise Speed Vs Wing Area Fig.12 Cruise Speed Vs Wing Span Fig.13 Cruise Speed Vs Aircraft Length Fig.14 Cruise Speed Vs Aspect Ratio Fig.15 Cruise Speed Vs Mach No Fig.16 Cruise Speed Vs Range Fig.17 Normal mission profile for business aircraft Fig.18 CutView of TFE73 1-20 Fig.19 Pictorial View of TFE73 1-20 Fig.20 Airfoil Fig.21 NACA 23020 Airfoil Fig.22 Aerodynamic characterstic of NACA 23024 Fig.23 Drag Polar for NACA 23024 Fig.24 Lift for NACA 23024 Fig.25 NACA 23021 Airfoil
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Fig.26 Aerodynamic characterstic of NACA 23021 Fig.27 Drag Polar for NACA 23021 Fig.28 Lift for NACA 23021 Fig.29 Types of FLAP Fig.30 DOUBLE FLOWER-SLOTTED Fig.31 Cabin Layout of 8 Seater Business Aircraft Fig.32 Dimensional View of Fuselage Fig.33 Layout of Aircraft Centre of Gravity Fig.34 Types of Aircraft Tail Fig.35 Longitudinal stable due to Horizontal Tail Fig.36 Lift Diagram Fig.37 CAMBERED AIRFOIL at Positive LIFT
Fig.38 Schematic of Pressure and shear stress distribution over a body surface
Fig.39 Pressure Distribution on an airfoil Fig.40 Lift Curves of Cambered and Symmetrical airfoils Fig.41 Drag separation Fig.42 Wave Drag Fig.43 Typical streamlining effect Fig.44 Types of Landing gear Fig.45 Landing gear Fig.46 Prop clearance Fig.47 Main gear and forward Cg Fig.48 Examination of fuselage clearance during take –off rotation Fig.49 Wheel track (front view) Fig.50 Calculation of the overturn angle for the aircraft Fig.51 Take-off of Aircraft Fig.52 Weight has rearward component Fig.53 Thrust versus Climb Angle Fig.54 Gliding of Aircraft
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AIM OF THE PROJECT
The aim of this design project is to design an EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT by comparing the data and specifications of present aircrafts in this category and to calculate the performance characteristics. Also necessary graphs need to be plotted and diagrams have to be included wherever needed. The following design requirements and research studies are set for the project:
Ø Design an aircraft that will transport 8 passengers and their baggage over a design range of 5200 km at a cruise speed of about 850 km/h.
Ø To provide the passengers with high levels of safety and comfort.
Ø To operate from regional and international airports.
Ø To use advanced and state of the art technologies in order to reduce the operating costs.
Ø To offer a unique and competitive service to existing scheduled operations.
Ø To assess the development potential in the primary role of the aircraft.
Ø To produce a commercial analysis of the aircraft project.
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ABSTRACT
The aim of this design project is to design an 8 Seater Short Range Executive Aircraft by comparing the data and specifications of present executive aircrafts and to calculate performance details. The aircraft designed is such that the landing and takeoff field lengths they require are accordingly shorter than those for the larger transport aircraft minimum drag and maximum thrust is also taken into consideration. Then the necessary graphs have to be plotted for further performance calculation. Required diagrams are also drawn.
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INTRODUCTION
Name : VYMANAAS SV3
Ever since the revolutionary introduction of the Learjet 23 in 1964, corporate chief executives and wealthy travelers have been flying in style on custom jet aircraft. The business or executive jet has become so common that most passengers anymore are middle management types. In fact, the executive jet industry is now a buyer's market. Do cost-benefit analyses before you whip out the bucks for a private business jet, however? Aviation experts tell us those 350 to 400 hours of flight time per year is justification for owning an executive jet. If you are not flying the friendly skies that often, then you should look into fractional ownership.
Consider the hidden costs involved in ownership of a jet. In addition to a price tag that ranges from $6 million to $50 million for a new private jet, consider necessities such as insurance, fuel, catering, and pilots (and there aren't that many of them to go around for executive jets). Aircraft management companies will take care of these needs for about $100,000 to $200,000 per year, depending on the size and usage of the jet. You will also want to determine the size and flying range you're going to require.
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Light jets (which cost in the range of $3 million to $8 million) can take five to eight passengers roughly 2000 miles; midsize executive jets (which cost in the range of $9 million to $16 million) can take up to nine passengers from 2000 to 3000 miles; and large executive jets (costing from $17 million to $45 million) can carry 2 passengers over 4000 miles. When you're ready to buy, begin contacting private jet manufacturers and ask around about aircraft specifications and pricing. In addition, surf the Net and hit places such as CharterAuction.com which sell new and used jets, including repossessed aircraft at deep discounts. If you feel that it is worth considering, the way fractional ownership works involves you purchasing a share in a jet plane from a management company, for which you get a tax deduction, and then paying a monthly fee and hourly operation costs. The management company sends out whichever jet is most conveniently located to service you, even on a moment's notice. A used jet may not be such a great deal. For permission to land at many U.S. Airports, a jet must be compliant with Stage 3 Federal Aviation Administration regulations. Converting a private jet to comply with regulations takes several hundred thousand dollars and many months of repair time, and most private jet aircraft repair centers already have more business than they can handle.
üüüü AOG/aircraft on ground üüüü Aerospace products üüüü Avionics and instruments üüüü Aerospace wet wipes üüüü Pilot Supplies
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DESIGN METHODOLOGY The start of the design process requires the recognition of a ‘need’. This normally comes from a ‘project brief’ or a ‘request for proposals (RFP)’. Such documents may come from various sources:
« Established or potential customers. « Government defence agencies. « Analysis of the market and the corresponding trends from aircraft demand. « Development of an existing product (e.g. aircraft stretch or engine change). « Exploitation of new technologies and other innovations from research and development.
It is essential to understand at the start of the study where the project originated and to recognize what external factors are influential to the design before the design process is started. At the end of the design process, the design team will have fully specified their design configuration and released all the drawings to the manufacturers. In reality, the design process never ends as the designers have responsibility for the aircraft throughout its operational life. This entails the issue of modifications that are found essential during service and any repairs and maintenance instructions that are necessary to keep the aircraft in an airworthy condition. The design method to be followed from the start of the project to the nominal end can be considered to fall into three main phases. These phases are illustrated in Figure 2.0.
Detail design
Manufacturing
Fig. 2 Phase of Design Testing
Project design
Preliminary design
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2.1 PRELIMINARY DESIGN
The preliminary phase (sometimes called the conceptual design stage) starts with the project brief and ends when the designers have found and refined a feasible baseline design layout. In some industrial organisations, this phase is referred to as the ‘feasibility study’. At the end of the preliminary design phase, a document is produced which contains a summary of the technical and geometric details known about the baseline design. This forms the initial draft of a document that will be subsequently revised to contain a thorough description of the aircraft. This is known as the aircraft ‘Type Specification’.
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2.2 PROJECT DESIGN The next phase (project design) takes the aircraft configuration defined towards the end of the preliminary design phase and involves conducting detailed analysis to improve the technical confidence in the design. Wind tunnel tests and computational fluid dynamic analysis are used to refine the aerodynamic shape of the aircraft. Finite element analysis is used to understand the structural integrity. Stability and control analysis and simulations will be used to appreciate the flying characteristics. Mass and balance estimations will be performed in increasingly fine detail. Operational factors (cost, maintenance and marketing) and manufacturing processes will be investigated.
The design process has been described in detail in the previous chapters. All the steps that are necessary to successfully complete the preliminary design stages have been identified. The amount of effort and time spent in each stage depends on the overall schedule for the project. It is essential to complete the process with a feasible baseline design, therefore it is necessary to programme and manage the work in association with all other commitments. Although the design method has been shown as a sequential process, it is possible to run some of the steps in parallel. It is also possible to do some preparation work (e.g. develop estimating methods and spreadsheets) ahead of the later stages. This is particularly useful if the project is to be done by a group, or team, of people. In such cases, it would be essential to allocate all tasks and to set a rigid timetable for the completion of the work some of the case studies that follow are laid out in the standard format shown below. This format mirrors the sequence of the work to be done in the preliminary design of any aircraft.
2.2.1 Introduction to the project
1) Project brief 2) Problem definition 3) Design concepts 4) Initial sizing and layout 5) Initial estimates 6) Constraint analysis and trade-offs 7) Revised baseline layout 8) Further work 9) Study review
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Design project work, as taught at most universities, concentrates on the preliminary phase of the design process. The project brief, or request for proposal, is often used to define the design problem. Alternatively, the problem may originate as a design topic in a student competition sponsored by industry, a government agency, or a technical society. Or the design project may be proposed locally by a professor or a team of students. Such design project assignments range from highly detailed lists of design objectives and performance requirements to rather vague calls for a ‘new and better’ replacement for existing aircraft. In some cases student teams may even be asked to develop their own design objectives under the guidance of their design professor.
To better reflect the design atmosphere in an industry environment, design classes at most universities involve teams of students rather than individuals. The use of multidisciplinary design teams employing students from different engineering disciplines is being encouraged by industry and accreditation agencies. The preliminary design process presented in this text is appropriate to both the individual and the team design approach although most of the cases presented in later chapters involved teams of design students. While, at first thought, it may appear that the team approach to design will reduce the individual workload, this may not be so. The interpersonal dynamics of working in a team requires extra effort. However, this greatly enhances the design experience and adds team communications, management and interpersonal interaction to the technical knowledge gained from the project work. It is normal in team design projects to have all students conduct individual initial assessments of the design requirements, study comparable aircraft, make initial estimates for the size of their aircraft and produce an initial concept sketch.
The full team will then begin its task by examining these individual concepts and assessing their merits as part of their team concept selection process. This will parallel the development of a team management plan and project timeline. At this time, the group will allocate various portions of the conceptual design process to individuals or small groups on the team. At this point in this chapter, a word needs to be said about the role of the computer in the design process. It is natural that students, whose everyday lives are filled with computer usage for everything from interpersonal communication to the solution of complex engineering problems, should believe that the aircraft design process is one in which they need only to enter the operational requirements into some supercomputer and wait for the final design report to come out of the printer (Figure 4).
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Indeed, there are many computer software packages available that claim to be ‘aircraft design programs’ of one sort or another. It is not surprising that students, who have read about new aircraft being ‘designed entirely on the computer’ in industry, believe that they will be doing the same. They object to wasting time conducting all of the basic analyses and studies recommended in this text and feel that their time would be much better spent searching for a student version of an all-encompassing aircraft design code. They believe that this must be available from Airbus or Boeing if only they can find the right person or web address.
While both simple aircraft ‘design’ codes and massive aerospace industry CAD programs do exist and do play important roles, they have not yet replaced the basic processes outlined in this text. Simple software packages which are often available freely at various locations on the Internet, or with many modern aeronautical engineering texts, can be useful in the specialist design tasks if one understands the assumptions and limitations implicit in their analysis. Many of these are simple computer codes based on the elementary relationships used for aircraft performance, aerodynamics, and stability and control calculations. These have often been coupled to many simplifying assumptions for certain categories of aircraft (often home-built general aviation vehicles). The solutions which can be obtained from many such codes can be obtained more quickly and certainly with a much better understanding of the underlying assumptions, by using directly the well-known relationships on which they are based. In our experience, if students spent half the time they waste searching for a design code (which they expect will provide an instant answer) on
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thinking and working through the fundamental relationships with which they are already supposedly familiar, they would find themselves much further along in the design process.
The vast and complex design computer programs used in the aerospace industry have not been created to do preliminary work. They are used to streamline the detail design part of the process. Such programs are not designed to take the initial project requirements and produce a final design. They are used to take the preliminary design, which has followed the step-by-step processes outlined in this text, and turn it into the thousands of detailed CAD drawings needed to develop and manufacture the finished vehicle.
It is the task of the aircraft design students to learn the processes which will take them from first principles and concepts, through the conceptual and preliminary design stages, to the point where they can begin to apply detailed design codes (Figure 5).
At this point in time, it is impossible to envisage how the early part of the design process will ever be replaced by off-the-shelf computer software that will automatically design novel aircraft concepts. Even if this program were available, it is probably not a substitute for working steadily through the design process to gain a fundamental understanding of the intricacies involved in real aircraft design.
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2.3 DETAIL DESIGN The process of designing an aircraft, generally divided into three distinct phases: conceptual design, preliminary design, and detail design. Each phase has its own unique characteristics and influence on the final product. These phases all involve aerodynamic, propulsion, and structural design, and the design of aircraft systems.
2.3.1. Design phases:
Conceptual design activities are characterized by the definition and comparative evaluation of numerous alternative design concepts potentially satisfying an initial statement of design requirements. The conceptual design phase is iterative in nature. Design concepts are evaluated, compared to the requirements, revised, reevaluated, and so on until convergence to one or more satisfactory concepts is achieved. During this process, inconsistencies in the requirements are often exposed, so that the products of conceptual design frequently include a set of revised requirements. During preliminary design, one or more promising concepts from the conceptual design phase are subjected to more rigorous analysis and evaluation in order to define and validate the design that best meets the requirements. Extensive experimental efforts, including wind-tunnel testing and evaluation of any unique materials or structural concepts, are conducted during preliminary design. The end product of preliminary design is a complete aircraft design description including all systems and subsystems. During detail design the selected aircraft design is translated into the detailed engineering data required to support tooling and manufacturing activities.
2.3.2. Requirements
The requirements used to guide the design of a new aircraft are established either by an emerging need or by the possibilities offered by some new technical concept or invention. Requirements can be divided into two general classes: technical requirements (speed, range, payload, and so forth) and economic requirements (costs, maintenance characteristics, and so forth).
2.3.3. Aerodynamic design
Initial aerodynamic design centers on defining the external geometry and general aerodynamic configuration of the new aircraft. The aerodynamic forces that determine aircraft performance capabilities are drag and lift. The basic, low-speed drag level of the aircraft is conventionally expressed as a term
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at zero lift composed of friction and pressure drag forces plus a term associated with the generation of lift, the drag due to lift or the induced drag. Since wings generally operate at a positive angle to the relative wind (angle of attack) in order to generate the necessary life forces, the wing lift vector is tilted aft, resulting in a component of the lift vector in the drag direction (see illustration).
Aircraft that fly near or above the speed of sound must be designed to minimize aerodynamic compressibility effects, evidenced by the formation of shock waves and significant changes in all aerodynamic forces and moments. Compressibility effects are mediated by the use of thin airfoils, wing and tail surface sweepback angles, and detailed attention to the lengthwise variation of the cross-sectional area of the configuration. The size and location of vertical and horizontal tail surfaces are the primary parameters that determine aircraft stability and control characteristics. Developments in digital computing and flight-control technologies have made the concept of artificial stability practical. See also Stability augmentation.
2.3.4. Propulsion design
Propulsion design comprises the selection of an engine from among the available models and the design of the engine's installation on or in the aircraft. Selection of the best propulsion concept involves choosing from among a wide variety of types ranging from reciprocating engine-propeller power plants through turboprops, turbojets, turbofans, and ducted and unducted fan engine developments. The selection process involves aircraft performance analyses comparing flight performance with the various candidate engines installed. In the cases where the new aircraft design is being based on a propulsion system which is still in development, the selection process is more complicated. Once an engine has been selected, the propulsion engineering tasks are to design the air inlet for the engine, and to assure the satisfactory physical and aerodynamic
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integration of the inlet, engine, and exhaust nozzle or the engine nacelles with the rest of the airframe. The major parameters to be chosen include the throat area, the diffuser length and shape, and the relative bluntness of the inlet lips.
2.3.5. Structural design
Structural design begins when the first complete, integrated aerodynamic and propulsion concept is formulated. The process starts with preliminary estimates of design air loads and inertial loads (loads due to the mass of the aircraft being accelerated during maneuvers). During conceptual design, the structural design effort centers on a first-order structural arrangement which defines major structural components and establishes the most direct load paths through the structure that are possible within the constraints of the aerodynamic configuration. An initial determination of structural and material concepts to be used is made at this time, for example, deciding whether the wing should be constructed from built-up sheet metal details, or by using machined skins with integral stiffeners, or from fiber-reinforced composite materials. During preliminary design, the structural design effort expands into consideration of dynamic loads, airframe life, and structural integrity. Dynamic loading conditions arise from many sources: landing impact, flight through turbulence, taxiing over rough runways, and so forth., Airframe life requirements are usually stated in terms of desired total flight hours or total flight cycles. To the structural designer this translates into requirements for airframe fatigue life. Fatigue life measures the ability of a structure to withstand repeated loadings without failure. Design for high fatigue life involves selection of materials and the design of structural components that minimize concentrated stresses. Structural integrity design activities impose requirements for damage tolerance, the ability of the structure to continue to support design loads after specified component failures. Failsafe design approaches are similar to design for fatigue resistance: avoidance of stress concentrations and spreading loads out over multiple supporting structural members.
2.3.6. Aircraft systems design
Aircraft systems include all of those systems and subsystems required for the aircraft to operate. Mission systems are those additional systems and subsystems peculiar to the role of military combat aircraft. The major systems are power systems, flight-control systems, navigation and communication systems, crew systems, the landing-gear system, and fuel systems.
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Design of these major subsystems must begin relatively early in the conceptual design phase, because they represent large dimensional and volume requirements which can influence overall aircraft size and shape or because they interact directly with the aerodynamic concept (as in the case of flight-control systems) or propulsion selection (as in the case of power systems). During preliminary design, the aircraft system definition is completed to include additional subsystems. The installation of the many aircraft system components and the routing of tubing and wiring through the aircraft are complex tasks which are often aided by the construction of partial or complete aircraft mock-ups. These are full scale models of the aircraft, made of inexpensive materials, which aid in locating structural and system components.
2.4 MANUFACTURING What is the Aircraft, Engine & Parts Manufacturing Industry?
Businesses in this industry do one or more of the following: manufacture complete aircraft; manufacture aircraft engines, propulsion units and other related equipment or parts; develop and make prototypes of aircraft; aircraft conversions (i.e. major modification to systems); and complete aircraft overhaul and rebuilding (i.e. periodic restoration of aircraft to original design specifications).
Ø Industry Products Ø Aircraft Ø Aircraft engines and engine parts Ø Other aircraft parts and auxiliary equipment
Ø Industry Activities Ø Manufacturing and rebuilding of aircraft Ø Developing and producing prototypes for aircraft Ø blimps, gliders, hand gliders, ultra light aircraft and helicopters Ø Manufacturing aircraft engines and engine parts Ø Developing and producing prototypes for aircraft engines and engine parts Ø Manufacturing aircraft assemblies, subassemblies, propellers, joints, and other
parts Ø Manufacturing aircraft auxiliary parts Ø Developing and producing prototypes for aircraft parts and auxiliary
equipment
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2.5 TESTING
Flight testing is a branch of aeronautical engineering that develops and gathers data during flight of an aircraft and then analyzes the data to evaluate the flight characteristics of the aircraft and validate its design, including safety aspects.
The flight test phase accomplishes two major tasks:
1) Finding and fixing any aircraft design problems and then 2) Verifying and documenting the aircraft capabilities for government certification or
customer acceptance.
The flight test phase can range from the test of a single new system for an existing aircraft to the complete development and certification of a new aircraft. Therefore the duration of a flight test program can vary from a few weeks to many years.
Examples of some subsystems we have performed aerospace testing on include:
• Airframes: Structural, Fatigue,
Limit Loads • Antennas • Avionics • Power Inverters, TRU's • Entertainment Systems • Communications • Drive Systems • Escape Slides, Rafts and Vests • Flight Control Surfaces, Winglets
• Gear Trains • Hydraulic motors and
components • Interior and Exterior Lighting
Systems • Landing Gear • Oxygen Systems • Passenger Service Units (PSU's) • Rotor Systems • Safety belts • Windows and doors
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3.1 WING CHARACTERISTICS
S.NO.
NAME OF THE AIRCRAFT
WING SPAN (m)
AIRCRAFT LENGTH
(m)
WING AREA (��)
ASPECT RATIO
(s)
TAKE OFF DISTANCE
1. BOMBARDIER BD-100 CHALLENGER 300 19.46 20.92 48.5 7.8 1509
2. CESSNA 560 CITATION ENCORE 16.49 16 29.94 8.54 1085
3. CESSNA 560 XL
CITATION EXCEL AND XLS
17.17 19.37 34.35 7.8 1109
4. CESSNA 680 CITATION SOVEREIGN 19.24 14.9 47.93 9.09 1086
5. CESSNA 750 CITATION X 19.5 22 49 7.76 1567
6. GULFSTREAM G150 16.94 17.3 17.3 8.7 1524
7. HAWKER 4000 HORIZON 18.82 21.08 49.3 7.1 1545
8. LEARJET 25 12.84 17.6 21.5 7.3 1440
9. LEARJET 45 14.57 17.8 29 7.2 1661
10. LEARJET 60 13.4 14.5 24.6 7.66 1581
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3.2 WEIGHT CHARACTERISTICS
S.NO. NAME OF THE AIRCRAFT EMPTY WEIGHT (kg)
MAXIMUM TAKE-OFF WEIGHT (kg)
1. BOMBARDIER BD-100 CHALLENGER 300 10591 17622
2. CESSNA 560 CITATION ENCORE 4627 7544
3. CESSNA 560 XL CITATION EXCEL AND XLS 5402 9163
4. CESSNA 680 CITATION SOVEREIGN 7893 13743
5. CESSNA 750 CITATION X 8618 16374
6. GULFSTREAM G150 6849 13716
7. HAWKER 4000 HORIZON 9877 17917
8. LEARJET 25 3465 6804
9. LEARJET 45 6146 9163
10. LEARJET 60 6641 10660
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3.3 POWER PLANT CHARACTERISTICS
S.NO. NAME OF THE AIRCRAFT TYPE OF ENGINE
NO. OF ENGINE
THRUST (KN)
1. BOMBARDIER BD-100 CHALLENGER 300 TURBO FAN 2 30.4
2. CESSNA 560 CITATION ENCORE TURBO FAN 2 13.6
3. CESSNA 560 XL CITATION EXCEL AND XLS TURBO FAN 2 18.43
4. CESSNA 680 CITATION SOVEREIGN TURBO FAN 2 25.3
5. CESSNA 750 CITATION X TURBO FAN 2 28.7
6. GULFSTREAM G150 TURBO FAN 2 19.7
7. HAWKER 4000 HORIZON TURBO FAN 2 28.9
8. LEARJET 25 TURBO FAN 2 13.1
9. LEARJET 45 TURBO FAN 2 15.6
10. LEARJET 60 TURBO FAN 2 23.2
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3.4 PERFORMANCE CHARACTERISTICS
S.NO. NAME OF THE AIRCRAFT MACH NO.
RANGE (km)
SERVICE CEILING
(M)
CRUISE SPEED (kph)
1. BOMBARDIER BD-100 CHALLENGER 300 0.89 5741 13716 870
2. CESSNA 560 CITATION ENCORE 0.75 3297 13716 793
3. CESSNA 560 XL CITATION EXCEL AND XLS 0.75 3852 13716 815
4. CESSNA 680 CITATION SOVEREIGN 0.8 5223 14326 846
5. CESSNA 750 CITATION X 0.92 5686 15548 1128
6. GULFSTREAM G150 0.85 5463 13716 850
7. HAWKER 4000 HORIZON 0.84 5287 13716 859
8. LEARJET 25 0.81 2663 13716 878
9. LEARJET 45 0.81 3925 15548 859
10. LEARJET 60 0.81 4441 15548 887
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3.6 MEAN DESIGN PARAMETERS
Parameters Values Unit
Vcruise 850 Km/h
Mach no. .715 (no unit)
Length 19.5 m
Wing span 12.84 m
Aspect ratio 7.55 (no unit)
Wing area 21.84 m2
Max Take-off distance 1500 kg
Wing loading 585.89 Kg/m2
Range 5200 km
Service ceiling
13700 m
Thrust 16 N
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4
WEIGHT ESTIMATION
4.1 THE WEIGHT OF AN AIRCRAFT AND ITS FIRST ESTIMATE
Let us discuss the nature of the weight of an airplane in detail. There are various types’ ways to subdivide and categorize the weight components of an airplane. The following is a common choice.
1. Crew weight Wcrew. The crew comprises the people necessary to operate the air plane in flight. For our airplane, the crew is simply the pilot.
2. Payload weight Wpayload . The payload is what the airplane is intended to transport –passenger, baggage, freight, etc. If airplane is intended for military combat use, the payload includes bombs, rockets, and other disposable ordnance.
3. Fuel weight Wfuel. This is the weight of the fuel in the fuel tanks. Since fuel is consumed during the course of the flight, Wfuel is a variable, decreasing with time during the course of the flight.
4. Empty weight Wempty. This is the weight of everything else-the structure, engines( with all accessory), electronic equipment (including radar computers, communication device, etc.),landing gear, fixed equipment(seats, galleys, etc.), and anything else that is not crew, payload, or fuel.
The sum of these weights is the total weight of the airplane W. Again, W varies throughout the fight because fuel is being consumed, and for a military combat airplane, ordnance may be dropped or expended, leading to a decrease in the payload weight.
The design takeoff gross weight W0 is the weight of airplane at the instant it begins its mission. It includes the weight of all the fuel on board at the beginning of the flight. Hence,
W0 = Wcrew + Wpayload + Wfuel + Wempty [4.1]
In Eq. (4.1), Wfuel is the weight of the full fuel load at the beginning of the flight.
40
In Eq. (4.1), W0 is the important quantity for which we want a first estimate; W0 is the desired result from graph. To help make this estimate, Eq. (4.1) can be rearranged as follows. If we denote Wfuel by Wf and Wempty by We (for notational simplicity), Eq. (4.1) can be written as
W0 = Wcrew + Wpayload + Wf + We [4.2]
[4.3]
Solving Eq. (4.3) for W0, we have
[4.4]
The form of Eq. (4.4) is particularly useful. Although at this stage we do not have a value of W0, we can fairly readily obtain values of the ratios Wf/W0 and We/W0, as we will see next. Then Eq. (4.4) provides a relation from which W0 can be obtained in an irerative fashion. [The iteration is required because in Eq.(4.4) Wf/W0 and We/W0 may themselves be functions of W0.]
4.2 ESTIMATION OF We/W0
Most airplane design are evolutionary rather than revolutionary; that is, a new de- sign is usually an evolutionary change from previously existing airplanes. For this reason, historical, statistical data on previous airplanes provides a starting point for the conceptual design of a new airplane. We will use such data here. In particular, Graph of We/W0 versus W0 for a number of Turbofan engine, jet aircrafts.
As a result of the data shown in graph. we choose for our first estimate
��
= 0.58 [4.5]
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4.3 ESTIMATION OF Wf/ W0
The amount of fuel required to carry out the mission depends critically on the efficiency of the propulsion device-the engine specific fuel consumption and the propeller efficiency. It also depends critically on the aerodynamic efficiency-the lift-to-drag ratio. These factors are principal players in the Brequet range equation, represented here:
R = �� �
�� ln �
�� [4.6]
Equation (4.6) is very important in our estimation of Wf/W0, as defined below. The total fuel consumed during the mission is that mission is that consumed from the moment the engines are turned on at the airport to the moment they are shut down at the end of the flight. Between these times, the flight of the airplane can be described by a mission profile, a conceptual sketch of altitude versus time such as shown in (figure 4.1). As stated in the specifications. The mission profile is that for a simple cruise from one location to another. This is the mission profile shown in Figure. It starts at the point labeled 0, when the engines are first turned on. The takeoff segment is denoted by the line segment 0-1, which includes warm-up, taxing, and takeoff. Segment 1-2 denotes the climb to cruise altitude (the use of a straight line here is only schematic and is not meant to imply a constant rate of climb to altitude). Segment 2-3 represents the cruise, which is by far the largest segment of the mission. Segment 2-3 shows an increase in altitude during cruise, consistent with an attempt to keep CL (and hence L/D) constant as the aircraft weight decreases because of the consumption of fuel. Segment 3-4 denotes the descent, which generally includes loiter time to account for air traffic delays; for design purposes, a loiter time of 20 min is commonly used. Segment 4-5 represents landing. The mission profile shown in Figure is particularly simple. For other types of missions, especially those associated with military combat aircraft, the mission profile with include such aspects as combat dogfighting, weapons drop, in-flight refueling etc. For a discussion of such combat mission profiles, see, for example, Raymer book. For our purpose, we will deal only with the simple cruise mission profile sketched in Figure (Fig. 4.1)
42
The mission profile is a useful bookkeeping tool to help us estimate fuel weight. Each segment of the mission profile is associated with a weight fraction, defined as the aircraft weight at the end of the segment divided by the weight at the beginning of the segment.
Mission segment weight fraction = ��
����
For example, the cruise weight fraction is W3/W2, where W3 is the aircraft weight at the end of the cruise and W2 is the weight at the beginning of cruise. The fuel weight ratio Wf/W0, can be obtained from the product of the mission segment weight fractions as follows. Using the mission profile in Figure, the ratio of the aircraft weight at the end of the mission to the initial gross weight is W5/W0. In turn,
[4.7]
The right side of Eq. (4.7) is simply the product of the individual mission segment weight fractions. Also, keep in mind that for the simple cruise mission shown in Figure, the change in weight during each segment is due to the consumption of fuel. It, at the end of the flight, the fuel tanks were completely empty, then
or
43
[4.8]
However, at the end of the mission, the fuel tanks are not completely empty-by design. There should be some fuel left in reserve at the end of the mission in case weather conditions or traffic problems require that the pilot of the aircraft divert to another airport, or spend a longer-than-normal time in a holding pattern. Also, the geometric design of the fuel tanks and the fuel system leads to some trapped fuel that is unavailable at the end of the flight. Typically, a 6% allowance is made for reserve and trapped fuel. Modifying Eq. (4.8) for this allowance, we have
[4.9]
Hence, the sequence for the calculation of Wf/W0 that appears in the denominator of Eq. (4.9) is as follows:
a. Calculate each individual mission segment weight fraction W1/ W0, W2/ W1 etc., that appears in Eq. (4.7).
b. Calculate W5/ W0 from Eq. (4.7). c. Calculate Wf / W0 from Eq. (4.9).
Let us proceed to make this calculation for our business jet eight seater aircraft.
For takeoff, segment 0-1, historical data show that W1/ W0 are small, on the order of 0.97. Hence, we assume
���
� � ��� [4.10]
For climb, segment 1-2. we again rely on historical data for a first estimate which indicate that W2/ W1 is also small, on the order of 0.985. Hence, we assume
����
� � � [4.11]
For cruise, segment 2-3, we make use of the Brequet range equation. This requires an estimate of L/D. At this stage of our design process, we cannot carry out a detailed aerodynamics analysis to predict L/D- we have not even laid out the shape of the airplane
44
yet. Therefore, we can only make a crude approximation, again based on data from existing aircraft. Loftin has tabulated the values of (L/D)max for a number of famous aircraft over the past century.
Hence, a reasonable first approximation for our aircraft is
(L/D)max = 13 [4.12]
Also needed in the range equation, are the specific fuel consumption c and velocity Vcr.
A typical value of specific fuel consumption for aircraft turbo fan engine is 0.6 lb of fuel consumed per horsepower per hour. In consistent units, noting that 1 hp = 550 ft-lb/s, we have
c = 0.7 ��������
� � [4.13]
A reasonable value for the velocity, assuming a variable- pitch engine
Vcr = 527 mi/hr [4.14]
The ratio W0/W1 in that equation is replaced for the mission segment 2-3 by W2/W3. Hence, for range equation
R = �� �
�� ln ��
� [4.15]
Solving Eq. (4.15) for W2/W3, we have
� ��
� �� � [4.16]
The loiter segment 3-4 in figure is essentially the descent from cruise altitude to the landing approach. For our approximate calculation here, we will ignore the detail of fuel consumption during descent is part of the required 3221.13-mi range, Hence, for this assumption
�!�
� � �� [4.17]
45
Finally, the fuel consumed during the landing process, segment 4-5, is also based on historical data. The amount of fuel used for landing is small, and based on previous aircraft, the value of W5/W4 is approximately 0.995. Hence, we assume for our airplane that
���!
� � ��� [4.18]
Collecting the various segment weight fractions form Eq. (4.10), (4.11), (4.16), (4.17), and (4.18), we have from Eq. (4.7)
= 0.995×0.98×0.719×0.99×0.992
���
� � ��� [4.19]
Inserting the of W5/W0 from Eq. (4.19) into Eq. (4.9), we have
=1.06 (1 - 0.689)
���
� � [4.20]
4.4 CALCULATION OF W0 Return to Eq. (4.4) for the design takeoff gross weight W0. We have obtained a value for We/W0 giver by Eq. (4.5). We have also obtained a value for Wf/W0
given by Eq. (4.20). All we need to obtain W0 from Eq. (4.4) are values for the crew and payload weights Wcrew and Wpayload, respectively. Corning suggests the average passenger weight of 242.50 lb with baggage per passenger. For our aircraft, there are eight passengers, two stewards and two pilots, 14 people in total. Let us assume the average weight per person is 242.50 lb. Hence, since the 4 crew is the pilots and stewards, we assume
46
Wcrew = 220.46 lb [4.21] The payload is the 8 passenger, plus the baggage for all 8 people. The type of short business trip for which this aircraft will most likely be used would require less baggage than a longer, intercontinental trip. Thus, including the pilot’s baggage, we have
Wpayload = 8(220.46)
= 1940.066 [4.22] Inserting the values form Eq. (4.5) and (4.20) to (4.22) into Eq. (4.4), we have
This is only the first estimation. Now by doing iterations, we can get a fairly accurate value of the Maximum Take Off Weight (W0).
ØØØØ ITERATION PROCESS (W0):
For the iteration process, we use the given formula,
We/""""0 = 1.02×""""0−0.06 [4.23]
v FIRST: We/"0 = 1.02× 31354.59−0.06 We/"0 = 1.241201 lb W0 = 31354.59 lb
v SECOND: W0 = 26373.01 lb
v THIRD: W0 = 28219.4
v FOURTH: W0 = 16599.49
47
v FIFTH:
W0 = 28219.14
v SIXTH: W0 = 28219.14
After doing sixth iterations, we can take the value W0 =28219.14 as the final estimate of the W0.
Max Takeoff Weight (W0) = 28219.14 lb [4.24]
We know that, W####/W0 = 0.57
So, substituting the value of W0, we get the first estimation value of Wf, Wf = 28219.14 × 0.57 Wf = 9313.92 lb Weight of the Fuel Wf = 9313.92 lb [4.25] The weight of aviation gasoline is 5.64 lb/gal. Hence, the capacity of the fuel tank (or tanks) should be
Tank capacity = � � ���
��$!
Tank capacity = 1651.4042 gal [4.26]
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5
POWER PLANT SELECTION
5.1 ENGINE SELECTION
The literature survey indicated a thrust to weight ratio of 0.25 was appropriate. Hence: Engine total take-off thrust = 0.25 × 28219.14 × 9.81 = 32.00 kN (7193.8862 lb) [5.1] With two engines this equates to 16.00 kN per engine (3596.943 lb) [5.2] A choice of engines from different manufacturers is always the preferred commercial position for the airframe manufacturer. This ensures that the engine price and availability is more competitive. It also provides the potential airline customer with more bargaining power when selecting the aircraft/engine purchase. There are several available engines that would suit our requirement. All of them are currently used on civil aircraft operations therefore considerable experience is available. The engines below are typical options:
Sl.no Name of the engine type Thrust (KN)
1 Honey well TFE731-20(Garrett)
turbofan 15.6
2 General Electric CF700 turbofan 18.7
3 Pratt & Whitney JTIA8 turbofan 20
4 Honey well TFE731-40R-200G
turbofan 18.9
5.2 REQUIRED ENGINE
Calculated thrust and weight of the engine are satisfied with the General
HONEY WELLHONEY WELLHONEY WELLHONEY WELLTFE731TFE731TFE731TFE731----22220000 therefore chosen this engine
REQUIRED ENGINE
Calculated thrust and weight of the engine are satisfied with the General
therefore chosen this engine.
49
Calculated thrust and weight of the engine are satisfied with the General
The TFE731 family of engines has a legacy of proven reliability. The first TFE731-20 was certified in 1972. Since that time, more than 11,000 engines have been produced, logging an astounding 100 million plus hours of service on modifferent aircraft applications. The TFE731with proven experience, performance, and reliability. The newest member of the TFE73120 family is currently in development
The TFE731 family of engines has a legacy of proven reliability. The first was certified in 1972. Since that time, more than 11,000 engines have been
produced, logging an astounding 100 million plus hours of service on modifferent aircraft applications. The TFE731-20 gives you everything you want in an engine with proven experience, performance, and reliability. The newest member of the TFE731
development.
51
The TFE731 family of engines has a legacy of proven reliability. The first was certified in 1972. Since that time, more than 11,000 engines have been
produced, logging an astounding 100 million plus hours of service on more than 27 gives you everything you want in an engine
with proven experience, performance, and reliability. The newest member of the TFE731-
52
5.3 ENGINE SPECIFICATION :
HONEY WELL TFE731-20
Type : TURBOFAN (fuselage mounted)
Length : 129.54cm (51 in.)
Diameter : 71.628 cm (28.2 in.)
Dry Weight: 401.42Kg (885 lb)
Stages :
COMPRESSOR - FAN; 4 HP axial & 1 HP centrifugal
TURBINE - 1 HP & 3 LP
Combustors : ANNULAR
Max. Thrust : 15.6 KN (3500 lbf)
Overall Pressure ratio : 13:1
Specific fuel consumption: 0.5 lb/lb-hr
Thrust/weight : 4.7:1.
Note: engine manufacturers commonly quote values in Imperial units. These details will be enough for initial performance and layout purposes but as the design progresses it will be necessary to periodically review the choice of engines to be used on the aircraft.
53
6 WING SELECTION
6.1 INTRODUCTION :
After the final weight estimation of the aircraft, the primary component of the aircraft to be designed is the wing. The wing weight and its lifting capabilities are in general, a function of the thickness of the airfoil section that is used in the wing structure. The first step towards designing the wing is the thickness estimation. The thickness of the wing, in turn depends on the critical mach number of the airfoil or rather, the drag divergence Mach number corresponding to the wing section.
The critical Mach number can well be delayed by the use of an appropriate Sweep-back angle to the wing structure. The natural choice of the standard series is the 65 series which is designed specifically for use in high-speeds.
6.2 WING GEOMETRY DESIGN
« The geometry of the wing is a function of four parameters, namely the
üüüü Wing loading (W/S), üüüü Sweepback angle at quarter chord (Λqc).
« The Take-off Weight that was estimated in the previous analysis is used to find the
Aspect Ratio (b2/S), « The value of S also enables us to calculate the Taper ratio (λ)
Form Raymer book we choose our, Taper Ratio (λ) = 0.33 [6.1]
6.2.1. WING AREA:
Wing Area (S) = �%&'()�&*' +,-.&*' [6.2]
= ��/
�/��/�
S = 21.84 m2 (or) 235.083 ft2 [6.3]
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6.2.2. WING SPAN (b):
From Our Graph we found mean value of Aspect Ratio (AR) = 7.55 [6.4]
Wing Span (b) = 01� �� 2 ��� �! [6.5]
b = 12.84 m [6.6]
6.3 WING CHORD DESIGN
6.3.1. ROOT CHORD Cr
[6.7]
= �2���/!���/!2��3� �
Cr =2.55 m [6.8]
6.3.2. TIP CHORD Ct
[6.9]
= 0.33 × 2.55 [6.10]
Ct = 2.55 m [6.11]
6.3.3. Distance of the Mean Chord from the Aircraft Centre line
= 4�2��3�5�6
$2��35� [6.12]
= ���/!2��3��� ��
$2��� �
=2.67 m [6.13]
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6.3.4. SWEEP ANGLE :
Sweep back angle at leading angle = )-*�� 78 � 9 [6.14]
= 11.130 [6.15]
6.3.5. DIHEDRAL ANGLE
For our Value of Sweep back angle to shows the tabulated value of Dihedral Angle
Dihedral Angle = 2.70 [6.16]
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7 AIRFOIL SELECTION
7.1 INTRODUCTION: The airfoil is the main aspect and is the heart of the airplane. The airfoils affects the cruise speed landing distance and take off, stall speed and handling qualities and aerodynamic efficiency during the all phases of flight Aerofoil Selection is based on the factors of Geometry & definitions, design/selection, families/types, design lift coefficient, thickness/chord ratio, lift curve slope, characteristic curves.
The following are the airfoil geometry and definition: Chord line: It is the straight line connecting leading edge (LE) and trailing edge (TE). Chord (c): It is the length of chord line. Thickness (t): measured perpendicular to chord line as a % of it (subsonic typically 12%). Camber (d): It is the curvature of section, perpendicular distance of section mid-points from chord line as a % of it (sub sonically typically 3%). Angle of attack (α): It is the angular difference between chord line and airflow direction. The following are airfoil categories: Early it was based on trial & error.
57
NACA 4 digit is introduced during 1930’s. NACA 5-digit is aimed at pushing position of max camber forwards for increased CL max. NACA 6-digit is designed for lower drag by increasing region of laminar flow. Modern it is mainly based upon need for improved aerodynamic characteristics at speeds just below speed of sound.
NACA 4 Digit 1st digit: maximum camber (as % of chord). 2nd digit (x10): location of maximum camber (as % of chord from leading edge (LE)). 3rd & 4th digits: maximum section thickness (as % of chord).
NACA 5 Digit 1st digit (x0.15): design lift coefficient. 2nd & 3rd digits (x0.5): location of maximum camber (as % of chord from LE). 4th & 5th digits: maximum section thickness (as % of chord).
NACA 6 Digit 1st digit: identifies series type. 2nd digit (x10): location of minimum pressure (as % of chord from leading edge (LE)). 3rd digit: indicates acceptable range of CL above/below design value for satisfactory low drag performance (as tenths of CL). 4th digit (x0.1): design CL. 5th & 6th digits: maximum section thickness (%c)
The airfoil that is to be used is now selected. As indicated earlier during the calculation of the lift coefficient value, it becomes necessary to use high speed airfoils, i.e., the 6x series, which have been designed to suit high subsonic cruise Mach numbers.
7.2 ESTIMATION OF THE CRITICAL PERFORMANCE PARAMETERS
We now move to pivot point 3, namely, an estimation of critical performance (CL) max,
L/D, W/S, and T/W. These parameters are directed by the requirements; that is, they will be determined by such aspects as maximum speed, range, and ceiling, rate of climb, stalling speed, landing gear, and takeoff distance.
58
7.3.1. Maximum Lift Coefficient This is the stage in the design process where we make an initial choice for the airfoil shape for the wing. Historically, general aviation airplanes have employed the NACA four-digit, and 6-series airfoil sections-the laminar-flow series.
L=W=0.5×ρ×V2stall×S×CL max [7.1]
VStall = 0.25 × Vcruise [7.2]
VStall = 0.25× 236.111
VStal l = 59.027 m/s [7.3]
sub, the value Eq.(7.3) in (7.1)
�: = 0.5 × 0.23884 × (59.02)2 × CL max
CL max = 1.40 [7.4] 7.3 SELECTION OF AIRFOIL
The NACA five-digit airfoils have been particularly favored by the general aviation industry in the United States. The airfoils, such as the NACA 23024 and the NACA 23021, respectively, were designed in the middle 1930s. The maximum chamber was placed closer to the leading edge (at 2.8c for the two airfoils shown) than was the case for the earlier NACA five-digit airfoils. A benefit of this design is a higher (CL) max compared to the earlier airfoils. A disadvantage is the very sharp stalling behavior.
For many airplanes, including some general aviation aircraft, one airfoil section is used at the wing root, and another airfoil shape is used at the wing tip, with the airfoil sections between the root and tip sections. Several examples from existing general aviation airplanes are tabulated below.
In these examples, the airfoil section is relatively thick and the wing airfoil shape tapers to thinner section at the tip. There are good reasons for this. Structurally, the wing bending moment is greatest at the root; a thicker airfoil readily allows the design for greater
59
structural strength at the root. Aerodynamically, an 18% airfoil will stall at a lower angle of attack than a 12% airfoil. Hence, a wing which has airfoil sections which has tapper from 18% thick at the root to 12% thick at the tip will tend to stall first at the wing root, with attached flow still at the tip. For all these reasons, we make an initial choice of the airfoil section for our airplane design as follows: at the root, an NACA23024 from Airfoil Data Base and at the tip, an NACA23024 from Airfoil Data Base
7.3.2. ROOT AIRFOIL NACA23024
Thickness: 24.0% Camber: 2.8% Trailing edge angle: 32.3o Lower flatness: 14.6% Leading edge radius: 7.4% Max CL: 1.36 Max CL angle: 15.0
Max L/D: 27.038
Max L/D angle: 6.0 Max L/D CL: 0.761 Stall angle: 1.5 Zero-lift angle: -1.0
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LIFT TO DRAG RATIO OF NACA 23024
a). α = 00 CL =0.1 (from above fig.) Cd =0.0072 (from above fig.)
;<;=
= >�?>�>>@A
= 13.88 b). α = 60
CL =0.57 (from above fig.) Cd =0.0085 (from above fig.)
;<;=
= >�B@>�>>CB
= 67.05 c). α = 120
CL =0.6 (from above fig.) Cd =0.007 (from above fig.)
;<;=
= >�D>�>>@
= 85.71
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7.3.3. TIP AIRFOIL
NACA 23021
Thickness: 21.0%
Camber: 1.9%
Trailing edge angle:
28.3o
Lower flatness: 22.2%
Leading edge radius: 6.2%
Max CL: 1.244
Max CL angle: 15.0
Max L/D: 25.219
Max L/D angle: 7.0
Max L/D CL: 0.873
Stall angle: 3.0
Zero-lift angle: -1.0
66
LIFT TO DRAG RATIO OF NACA 23021
a). α = 00 CL =0.1 (from above fig.) Cd =0.007 (from above fig.)
;<;=
= >�?>�>>@
= 14.28 b). α = 60
CL =0.7 (from above fig.) Cd =0.0085 (from above fig.)
;<;=
= >�@>�>>CB
= 82.35 c). α = 120
CL =0.75 (frm above fig.) Cd =0.0087 (frm above fig.)
;<;=
= >�@B>�>>C@
= 86.206
69
7.4 AVERAGE CL max Average CL max with NACA 23024 and NACA 23021
= ?�ED3?�AFF
A
CL max = 1.302 [7.5]
70
8 FLAP SELECTION
8.1 INTRODUCTION
During takeoff and landing the airplane's velocity is relatively low. To keep the lift high (to avoid objects on the ground!), airplane designers try to increase the wing area and change the airfoil shape by putting some moving parts on the wings' leading and trailing edges. The part on the leading edge is called a slat, while the part on the trailing edge is called a flap. The flaps and slats move along metal tracks built into the wings. Moving the flaps aft (toward the tail) and the slats forward increases the wing area. Pivoting the leading edge of the slat and the trailing edge of the flap downward increases the effective camber of the airfoil, which increases the lift. In addition, the large aft-projected area of the flap increases the drag of the aircraft. This helps the airplane slow down for landing.
8.2 TYPES OF FLAP
Types of flap systems include:
§ Krueger flap: hinged flap on the leading edge. Often called a "droop". § Plain flap: rotates on a simple hinge. § Split flap: upper and lower surfaces are separate, the lower surface operates like a
plain flap, but the upper surface stays immobile or moves only slightly. § Gouge flap: a cylindrical or conical aerofoil section which rotates backwards and
downwards about an imaginary axis below the wing, increasing wing area and chord without affecting trim. Invented by Arthur Gouge for Short Brothers in 1936.
§ Fowler flap: slides backwards before hinging downwards, thereby increasing both camber and chord, creating a larger wing surface better tuned for lower speeds. It also provides some slot effect. The Fowler flap was invented by Harlan D. Fowler .
§ Fairey-Youngman flap: moves body down before moving aft and rotating. § Slotted flap: a slot (or gap) between the flap and the wing enables high pressure air
from below the wing to re-energize the boundary layer over the flap. This helps the airflow to stay attached to the flap, delaying the stall.
§ Blown flaps: systems that blow engine air over the upper surface of the flap at certain angles to improve lift characteristics.
72
8.3 SELECTED FLAP
A wing designed for efficient high-speed flight is often quite different from one designed solely for take-off and landing. Take-off and landing distances are strongly influenced by aircraft stalling speed, with lower stall speeds requiring lower acceleration or deceleration and correspondingly shorter field lengths. It is always possible to reduce stall speed by increasing wing area, but it is not desirable to cruise with hundreds of square feet of extra wing area (and the associated weight and drag), area that is only needed for a few minutes. It is also possible to reduce stalling speed by reducing weight, increasing air density, or increasing wing CLmax. The latter parameter is the most interesting. One can design a wing airfoil that compromises cruise efficiency to obtain a good CLmax, but it is usually more efficient to include movable leading and/or trailing edges so that one may obtain good high speed performance while achieving a high CLmax at take-off and landing. The primary goal of a high lift system is a high CLmax; however, it may also be desirable to maintain low drag at take-off, or high drag on approach. It is also necessary to do this with a system that has low weight and high reliability.
73
8.4 CL max INCREASES DUE TO FLAP
« Our flap is Double fowler flap the required value is at above.
8.4.1. TAKE-OFF CL max DUE TO FLAP
During Take-off Flap deflection upto 200
∆ (CL max ) = 0.5 + 1.302 [8.1] = 1.802
8.4.2. LANDING CL max DUE TO FLAP During Landing Flap deflection upto 500
∆ (CL max ) = 0.9 + 1.302 [8.2] = 2.202
74
9 FUSELAGE AND CABIN LAYOUT
9.1 INTRODUCTION The interiors of business aircraft are laid out more flexibly than are commercial transports. Interior appointments often cost millions of dollars and can be very luxurious, especially for the large long range aircraft such as the Gulfstream V or Global Express. Business aircraft based on commercial transports such as Boeing Business Jet provide even greater possibilities.
Cabin parameters obtained from similar business aircrafts
Ø Seat pitch = 0.9652m Ø Seat width = 0.7m Ø Aisle width=0.61m Ø Seats abreast=2 Ø No. of aisles=1
9.2 FUSELAGE LAYOUT-INTRODUCTION The fuselage layout is important as the length of the entire aircraft depends on this. The length and diameter of the fuselage is related to the seating arrangement. The fuselage of a passenger aircraft is divided into a number of sections:
a. Nose b. Cockpit c. Cabin d. Tail fuselage
75
9.2.1. Functions of fuselage: - provides of volume for payload - provide overall structural integrity - Possible mounting of landing gear and power plant Once fundamental configuration is establishment, fuselage layout proceeds almost independent of other design aspects.
9.2.2. Pressurization: - If required, it has a major impact upon the overall shape. - Overall effect depends on the level of pressurization.
9.3 FUSELAGE SIZING: The required value of Fuselage size is taken from the graph
Lfuselage = 19.5 m [9.1]
9.4 NOSE AND COCKPIT-FRONT FUSELAGE:
The layout of the flight deck and specified pilot window geometry is often the starting point of the overall fuselage layout. For the current design, flight decks of various airplanes are considered and the value of
�GH���I�
is found to be 0.03 [9.2]
Lnos = 0.03 × 19.5
Lnos = 0.58 m [9.3]
The cockpit length for a 2 member crew is given by RAYMER
9.5 Passenger cabin layout:
Two major geometric parameters that specify the passenger cabin are Cabin Diameter and Cabin Length. These are in turn decided by more specific details like number of seats, seat width, seating arrangement (number abreast), seat pitch, aisle width and number of aisles.
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We choose a circular cross section for the fuselage. The overall size must be kept small to reduce aircraft weight and drag, yet the resulting shape must provide a comfortable and flexible cabin interior which will appeal to the customer airlines. The main decision to be taken is the number of seats abreast and the aisle arrangement. The number of seats across will fix the number of rows in the cabin and thereby the fuselage length. Design of the cabin cross section is further complicated by the need to provide different classes like first class, business class, economy class etc.
9.6 Cabin length:
The total number of seats (8) is distributed as 2 seats abreast. Cabin parameters are chosen based on standards of similar airplanes.
The various parameters chosen are as follows
ª Seat pitch=0.9653m ª Seat width=0.7m ª Aisle width=0.61m ª Seats abreast=2 ª No. of aisles=1
Hence, the total cabin length will be = seat pitch × rows
=0.9652 × 6 + additional space
Total cabin length =8m [9.4]
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9.7 Cabin diameter:
Using the number of seats abreast, seat width, aisle width we calculate the internal diameter of the cabin.
dfus (internal) = 0.61×1+0.7×2
= 2.01m [9.5]
According to the standards prescribed by Raymer, chapter 9, the structural thickness is given by
t = 0.02df + 1 inch = 0.02 × 2.01 + 0.0254 = 0.06m [9.6]
Therefore the external diameter of the fuselage is obtained as
= 2.01 + 0.06×2
External diameter = 2.18m [9.7]
9.8 Rear fuselage:
The rear fuselage profile is chosen to provide a smooth, low drag shape which supports the tail surfaces. The lower side of the provide adequate clearance for aircraft when rotation during takeoff. The rear fuselage should also house the auxiliary power unit (APU). Based on data collected for similar aircraft we choose the ratio Ltail / dfus as 4.
Ltail = 6m [9.8]
Total fuselage length:
Various parts of the fuselage are indicated below
Cockpit length = 2.54m
Cabin length = 8m
Total = 10.54m [9.9]
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10
10.1 CENTRE OF GRAVITY
The center-of-gravity (CG) is the point at which an aircraft would balance if it were possible to suspend it at that point. It is the mass center of the aircraft, or the theoretical point at which the entire weight of the aircraft is assumed to be concentrated. Its distance from the reference datum is determined by dividing the total moment by the total weight of the aircraft. The center-of-gravity point affects the stability of the aircraft. To ensure the aircraft is safe to fly, the center-of gravity must fall within specified limits.
Center of gravity is calculated as follows:
ª Determine the weights and arms of all mass within the aircraft. ª Multiply weights by arms for all mass to calculate moments. ª Add the moments of all mass together. ª Divide the total moment by the total weight of the aircraft to give an overall arm.
The arm that results from this calculation must be within the arm limits for the center of gravity. If it is not, weight in the aircraft must be removed, added (rarely), or redistributed until the center of gravity falls within the prescribed limits.
For the sake of simplicity, centre of gravity calculations are usually performed along only a single line from the zero point of the reference datum.
Weight is calculated simply by adding up all weight in the aircraft. This weight must be within the allowable weight limits for the aircraft.
First estimate weight components for which we have some idea of their location of the engine, the passengers and pilot, and the baggage.
Using this information, we can make a preliminary estimate of the location of the centre of gravity.
Considering the forces to be acting at middle each part, and hence taking moment about the nose, we get the centre of gravity.
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CG (a) = �JH� K�LM� 2JH� �GL�M �3�8N��G K�LM�28N��G �GL�M�3 �K�GL K�LM� 2K�GL �GL�M�3�GL�G K�LM�2GL�G �GL�M
JH� K�LM�38N��G K�LM�3K�GL K�LM� 3GL�G K�LM�
= 7.71 m
10.2 Layout of Aircraft Centre of Gravity :
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11 TAIL SELECTION
11.1. TAIL SURFACES:
The type and area of the tail surfaces are important in determining the stability of the airplane. A conventional tail arrangement is chosen. Some of the important parameters
that decide the aerodynamic characteristics of the tail are area ratio (St/S), tail volume ratio
(VH and Vv), tail arm, tail span etc. All this parameters have to be decided for both the horizontal and vertical tail.
From the above list of tail types, the T-tail unit type is chosen which the most suitable configuration for the current design.
11.2. T-TAIL
A T-tail is an aft tail configuration (see figure. 34) that looks like the letter “T”; which implies the vertical tail is located on top of the horizontal tail. The T-tail
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configuration is another aft tail configuration that provides a few advantages, while it has a few disadvantages. The major advantage of a T-tail configuration is that it is out of the regions of wing wake, wing downwash, wing vortices, and engine exit flow (i.e. hot and turbulent high speed gas). This allows the horizontal tail to provide a higher efficiency, and a safer structure. The lower influence from the wing results in a smaller horizontal tail area; and the lower effect from the engine leads in a less tail vibration and buffet. The less tail vibration increases the life of the tail with a lower fatigue problem. Furthermore, another advantage of the T-tail is the positive influence of horizontal tail on the vertical tail. It is referred to as the end-plate effect and results in a smaller vertical tail area.
On the other hand, the disadvantages that associated with a T-tail are: 1. heavier vertical tail structure, 2. deep stall. The bending moment created by the horizontal tail must be transferred to the fuselage through the vertical tail. This structural behavior requires the vertical tail main spar to be stronger; which cause the vertical tail to get heavier.
Aircraft with T-tail are subject to a dangerous condition known as the deep stall (Ref. 6); which is a stalled condition at an angle of attack far above the original stall angle. T-tail Aircraft often suffer a sever pitching moment instability at angles well above the initial stall angle of about 13 degrees, without wing leading edge high lift device, or about 18 degrees, with wing leading edge high lift device. If the pilot allows the aircraft to enter to this unstable region, it might rapidly pitch up to a higher angle of about 40 degrees. The causes of the instability are fuselage vortices, shed from the forward portion of the fuselage at high angles of attack, and the wing and engine wakes. Thus the horizontal tail contribution on the longitudinal stability is largely reduced. Eventually, at a higher angle of attack, the horizontal tail exits the wing and nacelle wakes and the aircraft become longitudinally stable (see figure. 35).
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This condition may be assumed as a stable condition, but it accompanies an enormous drag along with a resulting high rate of descent. At this moment, the elevator and aileron effectiveness have been severely reduced because both wing and horizontal tail are stalled at the very high angle of attack. This is known as a locked-in deep stall, a potentially fatal state. The design solutions to avoid a deep stall in a T-tail configuration are to: 1. Ensure a stable pitch down at the initial stall, 2. Extend the horizontal tail span substantially beyond the nacelles, and 3. Employ a mechanism to enable full down elevator angles if a deep stall occurs. In addition, the aircraft must be well protected from the initial stall by devices such as stick shaker, lights, and stall horn.
11.3. HORIZONTAL & VERTICAL TAIL CALCULATION:
From the data collected on similar business airplanes, we choose the following values for the tail parameters.
Parameter
Horizontal Tail
Vertical Tail
Area ratio (St/S)
0.31
0.21
Aspect ratio
5
1.7
Taper ratio 0.26 0.31
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TAIL AREA:
The areas of the horizontal and vertical tail (SH and Sv) are calculated as,
SH = 0.31 × 34.3
SH= 10.63 m2 [11.1]
Sv = 0.21 × 34.3
SV = 7.2 m2 [11.2]
TAIL SPAN:
The span of the horizontal and vertical tail (bh and bv) are given as,
bh = (AhSh)0.5 [11.3]
bv = (AvSv)0.5 [11.4]
Taking ARH = 5 and ARv = 1.7, we get
bh = 7.29 m [11.5]
bv = 3.5 m [11.6]
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12 LIFT ESTIMATION
12.1. LIFT: Component of aerodynamic force generated on aircraft perpendicular to flight direction
12.2. Lift Coefficient (CL)
. Amount of lift generated depends on:
_ Planform area (S), air density (p), flight speed (V), lift coefficient (CL)
L = ½ ×ρ×V2×S× OPQ [12.1]
. CL is a measure of lifting effectiveness and mainly depends upon:
_ Section shape, planform geometry, angle of attack (α), compressibility effect
(mach number), viscous effects (Reynolds number)
12.3. Generation of Lift
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12.3.1. Aerodynamic force arises from two natural sources:
§ Variable pressure distribution.
§ Shear stress distribution.
• Shear stress primarily contributes to overall drag force on aircraft.
• Lift mainly due to pressure distribution, especially on main lifting surfaces, i.e. wing.
• Require (relatively) low pressure on upper surface and higher pressure on lower surface.
• Any shape can be made to produce lift if either cambered or inclined to flow direction.
• Classical aerofoil section is optimum for high subsonic lift/drag ratio. 12.4. Pressure variations with angle of attack
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– Negative (nose-down) pitching moment at zero-lift (negative α).
– Positive lift at α = 00.
– Highest pressure at LE stagnation point, lowest pressure at crest on upper surface.
– Peak suction pressure on upper surface strengthens and moves forwards with increasing α.
– Most lift from near LE on upper surface due to suction.
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12.5. Lift Curves of Cambered and Symmetrical airfoils
12.6. CALCULATION:
CL cruise � ��� 2�� �2: [12.2]
where, ρ = 0.23884 (at 13700 m)
= A2?AC>>2R�C?>�AECCF2�AED�??2>�AECCF�S2A?�CF
CL cruise = 0.8636 [12.3]
12.7. LIFT AT TAKE-OFF
CL Take-off = CL cruise × ∆ CL flap [12.4]
where, Take off flap at 200 is 0.5
=0.8636 + 0.5
CL Take-off = 1.364 [12.5]
VR = 1.1 V stall [12.6]
= 1.1 × 59.02
VR = 64.922 m/sec [12.7]
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L Take-off = ½ ×ρ×V2×S× OQT-U%�,VV
where, ρ = 1.225 (at sea level)
= ½ ×1.225 × (64.922)2×21.84×1.364
L Take-off =76905.4 N [12.8]
12.8. LIFT AT LANDING
CL Landing = CL cruise + ∆ CL flap [12.9]
= 0.8636 + 0.9
CL Landing = 1.7636 [12.10]
VR = 0.7 V stall [12.11]
= 0.7 × 59.02
VR = 41.314m/sec [12.12]
where, ρ = 1.225 (at sea level)
L Landing = ½ ×ρ×V2×S× OQQ-*.&*'
= ½ ×1.225 × (41.314)2×21.84×1.7636
L Landing = 32871.32 N [12.13]
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13 DRAG ESTIMATION
13.1. DRAG:
Drag is the resolved component of the complete aerodynamic force which is parallel to the flight direction (or relative oncoming airflow).
It always acts to oppose the direction of motion.
It is the undesirable component of the aerodynamic force while lift is the desirable component.
13.2. Drag Coefficient (CD)
Amount of drag generated depends on:
Planform area (S), air density, flight speed (V), drag coefficient (CD)
CD is a measure of aerodynamic efficiency and mainly depends upon:
Section shape, planform geometry, angle of attack, compressibility effects (Mach number), viscous effects (Reynolds’ number).
13.3. Drag Components
• Skin Friction.
• Due to shear stresses produced in boundary layer.
• Significantly more for turbulent than laminar types of boundary layers.
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13.4. Form (Pressure) Drag Due to static pressure distribution around body - Drag is the resolved
component of the complete aerodynamic force which is parallel to the flight direction (or relative oncoming airflow).
It always acts to oppose the direction of motion.
It is the undesirable component of the aerodynamic force while lift is the desirable component. Amount of drag generated depends on:
Planform area (S), air density, flight speed (V), drag coefficient (CD)
CD is a measure of aerodynamic efficiency and mainly depends upon: Section shape, planform geometry, angle of attack, compressibility effects (Mach number), viscous effects (Reynolds’ number).
• Skin Friction. • Due to shear stresses produced in boundary layer. • Significantly more for turbulent than laminar types of boundary layers. • Due to static pressure distribution around body. • Component resolved in direction of motion. • Sometimes considered separately as forebody and rear (base) drag components.
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13.5. Wave Drag Due to the presence of shock waves at transonic and supersonic speeds.
Result of both direct shock losses and the influence of shock waves on the boundary layer.
Often decomposed into portions related to: • Lift. • Thickness or Volume.
13.6. Typical streamlining effect
The lift induced drag is the component which has to be included to account for the 3-D nature of the flow (finite span) and generation of wing lift.
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13.7. DRAG CALCULATION:
13.7.1. DRAG AT TAKE-OFF
CD Take-off = OP + K8� T-U%�,VV � [13.1]
K = �
�W%XY� [13.2]
K = �
��!2�1!!�21���
K = 0.057 [13.3]
where CD0 = 0.03 and apply eq. (13.4) and (12.5) in 13.1
= 0.03 + (0.057 × (1.364)2)
CD Take-off = 0.136 [13.4]
D Take-off = ½ ×ρ×V2×S× OPT-U%�,VV [13.5]
where, apply eq. (12.7) and (13.4) in (13.5)
= ½ ×1.225 × (64.922)2×21.84×0.136
D Take-off = 7667.9 N [13.6]
13.7.2. DRAG AT LANDING
CD Landing = OP + K8� Q-*.&*' � [13.7]
where CD0 = 0.03 and apply eq. (13.3) and (12.10) in (13.7)
= 0.03 + (0.057 × (1.7636)2)
CD Landing = 0.207 [13.9]
D Landing = ½ ×ρ×V2×S× OPQ-*.&*' [13.10]
where apply eq. (12.12) and (13.9) in (13.10)
= ½ ×1.225 × (41.314)2×21.84×0.207
D Landing = 4726.32 N [13.11]
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14
LANDING GEAR SELECTION
14.1 INTRODUCTION Every aircraft maintained in today’s Aerospace Company is equipped with a landing
gear system. Most Aerospace company aircraft also use arresting and catapult gear. The landing gear is that portion of the aircraft that supports the weight of the aircraft while it is on the ground. The landing gear contains components that are necessary for taking off and landing the aircraft safely. Some of these components are landing gear struts that absorb landing and taxiing shocks; brakes that are used to stop and, in some cases, steer the aircraft; nose wheel steering for steering the aircraft; and in some cases, nose catapult components that provide the aircraft with carrier deck takeoff capabilities.
LANDING GEAR DESIGN REQUIREMENTS
The following design requirements are identified to be satisfied: ground clearance requirement, tip-back (or tip-forward angle if tail gear) angle requirement, take-off rotation requirement, overturn angel requirement, structural integrity, aircraft ground stability, aircraft ground controllability, low cost, maintainable, and manufacturable.
14.2 LANDING GEAR CONFIGURATION
This is a transport aircraft, and the passenger’s comfort is an important requirement. So, the tail gear, bicycle, single main configurations would not satisfy this requirement.
Three viable configurations are:
1. Tricycle or nose-gear,
2. Quadricycle, and
3. Multi-bogey.
4. Ski type gear
5. Float type gear
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Since the aircraft weight is not very high, both quadricycle, and multi-bogey configurations are set aside due to their cost and weight. Therefore the best landing gear configuration for this aircraft is Nose gear or tricycle. An attractive feature for this configuration is that the aircraft will be horizontal at the ground. The passengers do not have to climb during boarding period. The nose gear also decreases the take-off run, and at the same time, the aircraft will take-off sooner.
14.3 FIXED OR RETRACTABLE
The aircraft must compete with other transport aircraft in the market, and it must have a fairly high performance, so a retractable landing gear (see figure) is the best option. The cost of this configuration covered by the customers (passengers). Then, this will reduce the aircraft drag during flight and therefore the aircraft will feature a higher performance. The higher landing gear weight due to retraction system will be paid off compared with the other advantages of a retractable landing gear.
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14.4 Landing gear height
Based on figure 1, the lowest point of the aircraft is the propeller tip. There must be a reasonable clearance between the prop and the ground. Due to the fact that the aircraft engine is turboprop, and for the sake of passengers’ safety considerations, a 1.2 m ground clearance for the propeller is considered need. This distance may be revised in the later phases of design.(i.e. ∆Hclear = 1.2 m). Hence, the distance between aircraft center of gravity and the ground would be:
Hcg = ∆H clear + Z[\][
A [14.1]
where, our propeller diameter is 1m = 1.2 + ½ Hcg = 1.7 m [14.2]
This clearance is shown in Fig 46. The aircraft cg is at the same height as the wing mid-plane. The landing gear height is a function of its attachment location. The nose gear will be naturally attached to the fuselage. But, the main gear attachment tends to have two main alternatives: 1. Attach to the fuselage, 2. Attach to the wing. As soon the wheel track is determined, we are able to decide about landing gear attachment; and then the landing gear height may be determined.
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14.5 ATTACH MAIN GEAR TO THE FUSELAGE:
HLG = Hcg - Z^_`a
A [14.3]
apply eq.(14.2) and (9.5) in (14.3)
= 1.7 – A�>?
A HLG = 0.695 m [14.4] 14.6 ATTACH MAIN GEAR TO THE WING Wing mean aerodynamics chord is :
bc = de [14.5]
apply eq.(6.3) and (6.6) in (14.5)
= A?�CF?A�CF
bc= 1.701m [14.6] The wing airfoil is NACA 23021 and NACA 23024, so the wing thick-to-ratio is 14
percent. Thus the wing thickness is:
tw= (t/c)max bc [14.7]
apply (14.6) in (14.7)
= 0.14 × 1.701
tw = 0.238 m [14.8] In this case, the landing gear height will be:
HLG = Hcg – fgA [14.9]
= 1.7 – >�AECA
HLG = 1.581 m [14.10]
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These two landing gear heights are shown in figure 9.32. When wheel track and wheel base are determined, the main gear attachment will be finalized. Furthermore, in the later steps, other landing gear requirements will be checked to make sure this clearance does not violate any other design requirements.
14.7 THE DISTANCE BETWEEN MAIN GEAR TO AIRCRAFT FORWARD Cg
Now we determine the location of main landing gear. Take-off rotation requirement will be employed to obtain this distance. The aircraft is required to be able to rotate during transition segment of the take-off operation with the amount of 9deg/s2. This requirement must be examined for the aircraft critical cg location which is the most forward cg.
Since the aircraft forward cg is at 18% MAC, and the wing-fuselage aerodynamic center is located at 22% MAC, we can write the following relationship for the wing-fuselage lift moment arm:
X L wf = Xmg- (hi- hicg fwd) bc [14.11]
apply eq.(14.6) in (14.11)
= Xmg – (0.22-0.18) ×1.701
=Xmg- 0.06804 [14.12]
Furthermore, the distance between horizontal tail aerodynamic center to the wing-fuselage aerodynamic center is 12 m, hence the tail moment arm would be:
Xh = Xac h -Xmg [14.13]
= lh + (hih- hicg fwd) bc-Xmg [14.14]
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= 9.5 + (0.22 -0.18) 1.701 - Xmg
= 9.56 - Xmg [14.15]
where Xmg is measured from main gear to the forward cg in meters.
From Figure.46, we can extract the following dimension: hd = hcg = hT = 1.7 m
Lh = ½ ρ VR2 ShCLh [14.16]
= ½ × 1.225 × (41.314)2 × 14 × (-0.8)
Lh = -11714.5 N [14.17]
Mac wf = ½ ρ VR2 C m ac wf Sref bc [14.18]
= -1165.14 N [14.19]
Lwf = LTO - Lh [14.20]
apply Eq.(12.8) and (14.17) in (14.20)
= 76905.4 – (-11714.5)
= 88619.9 N [14.21]
Friction Force:
Ff = µ (W- LTO) [14.22]
= 0.04{(12800 × 9.81) – 76905.4)}
Ff = 1946.504 N [14.23]
In value (5.1) our require Engine Thrust is given by
Engine Total Thrust:
T = 32000 N [14.24]
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Aircraft linear acceleration at the time of take-off rotation
a = j���kl
m [14.25]
= EA>>>�@DD@�R�?RFD�B>F
?AC>>
= 1.748 m/s2 [14.26]
Contributing moments:
Mw = W (Xmg – Xcg) [14.27]
= 12800 × 9.81× Xmg [14.28]
MD = D (ZD – Zmg) [14.29]
= 13035.43 Nm [14.30]
MT = T (ZT-Z mg) [14.31]
MT = 54400 [14.32]
ML wf = Lwf (Xmg –Xac wf to cg)
apply eq. (14.12) and (14.21) in (14.30)
= 88619.9 × (Xmg- 0.06804)
ML h = Lh (Xac - Xmg) [14.33]
= -11714.5 (9.56- Xmg) [14.34]
Xa = 12800 × 1.748 × 1.7
= 38036.48 [14.35]
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where, for the sake of simplicity, the x-reference line is considered to be the aircraft forward cg. Now all moments are substituted into the equation.
[14.36] By substituting the force and moment,
Xmg = 2.01 m [14.37]
So far, the fact that, the given mass moment of inertia is about aircraft cg, was ignored. Hence, the calculation must be repeated with the revised equation to include the parallel axis theorem. Thus, the aircraft mass moment of inertia about cg (y-axis) must be transferred to the main gear contact
point (Iyymg) by employing the parallel axis theorem
[14.38]
where Xmg and hcg are shown in Figure.47 Now, The Iyymg from equation 9.53a is inserted into equation 9.54. The result would be a nonlinear equation with only one unknown parameter (X mg) as follows:
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[14.39]
The solution of this revised equation would be: Xmg = 1.961 m [14.40] 14.8 CHECK TIP-BACK REQUIREMENT
In order to check the tip back angle, we have to obtain the distance between aft cg
and the main gear. Based on the problem statement, the forward cg is located at 18%MAC, while the aft cg is at 0.3C. Thus, the distance between aircraft aft cg and forward cg is:
∆Xcg =Xcg fwd – Xcg aft [14.41]
= (0.30 – 0.18) bc ∆Xcg = 0.204 m [14.42]
So the distance between main gear and after Cg is: Xcg aft = Xmg - ∆Xcg
= 1.757 m
The distance indicates that the aircraft has the following tipback angle:
af = tan-1nopqrsq
t [14.43]
= tan-1 7?�@B@?�@ 9
= 45.50 [14.44]
This tipback angle is greater than aircraft take-off rotation angle (140) 45.5 u 14 + 5 [14.45] Therefore the distance between main gear and aft Cg satisfies the tip back angle requirement. This Xmg is the distance between main gear and the aircraft forward Cg just to satisfy take-off rotation requirement as well as tip back angel requirement. In the upcoming steps, this value must be examined again to ensure it meets the other design requirements.
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14.9 Check the take-off rotation clearance requirement The take-off rotation ground clearance requirement to prevent a fuselage hit as
follows: a v aTO [14.46]
In order to determine the clearance angle (ac), two distances should be obtained: 1. height between lowest points of the fuselage to the ground, 2. distance between the main gears to the fuselage upsweep point. Figure 46 illustrates that the fuselage height (Hf) is 1.7m. On the other hand, from figure 46, the length between nose and the fuselage upsweep point is 8.66m; and the distance between wing leading edge and the fuselage nose in 5m. Thus, the distance between fuselage upsweep point and the wing leading edge is
= 8.66m Furthermore, the distance main gear and the wing leading edge is: Xmg-LE = Xmg fwd + 0.18 bc [14.47] = 1.961 + (0.18 × 1.701) = 2.267m [14.48]
Thus, the distance between the main gear to the fuselage upsweep point (figure 48), the fuselage is:
ac = tan-1nopqrsq
t [14.49]
= tan-1 7 ?�@D�ERA9
= 14.89 0 [14.50]
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Since the clearance angle (ac = 14.890) is greater than the aircraft rotation angle
(aTO = 120), the fuselage will not hit the ground during take-off operation.
14.10 WHEEL BASE: Due to the ground controllability requirement, the nose gear must not carry less than about 5 percent of the total load and also must not carry more than about 20 percent of the total load (e.g. aircraft weight). Thus, main gear carries about 80% to 95% of the aircraft load. To meet this requirement, it is decided that the nose gear would carry 15% of the total load and the main gear carries 90% of the total load. The wheel base is determined using equation (14.51)
Fn = wxw W [14.51]
where Fn is selected to be 10% of the weight, so:
[14.52]
where Bm is obtained previously as 1.916 m. Thus, the wheel base (B) is: B = 6.667 × 1.961 = 13.07 m [14.53]
When the Cg is at after location, the nose wheel will carry less than 15 percent of the aircraft weight. This value for wheel base could be revising later for optimization when examining other requirements.
14.11 WHEEL TRACK: Three main design requirements which drive the wheel track (T) are: 1. Ground lateral control, 2. Ground lateral stability, and 3. Structural integrity. The overturn angle is the angle which is critical to the aircraft overturn. There are two overturn angles (Fig. 47); in that the smaller one is considered in this technique. The minimum allowable value for the wheel track must satisfy the overturn angle requirements The maximum allowable value for the wheel track must satisfy the structural integrity requirements. In the first place, to determine the wheel track, we use the criterion of overturning prevention. The lateral distance between each main gear to the Cg must be greater than 25 degrees. Here, we consider 30 degrees. Fig. 49 illustrates the front
104
view of the aircraft; showing one of the main wheels relative to the aircraft Cg.
In the step 4, the height of cg from the ground was determined to be 1.7 m. Using the traiangle shown in figure 49., the whel track is determined as follow:
[14.54]
apply Eq.(14.2) in (14.54) then we get.,
= 1.97 m [14.55]
Now, we need to examine the overturn angle based on the top view. Fig 50 shows the top view of the aircraft and a triangle to determine φot based on top-view.
105
In order to determine overturn angle (φot), for this aircraft, we first need to determine the parameter Yot as shown in Fig. 50-1. This parameter is calculated by using the sine law in the triangle ADE (Fig. 50-2) using the angle φ1. However, this angle is obtained via triangle ACF. In the triangle ABC, the side AC is the wheel base, and the side FC is a one half of the wheel track. Thus, in triangle ACF:
[14.56]
= ��/
� �1
= 0.07 ═> Ø = tan-1 (0.07) = 4.290 [14.57]
The overturn angle is greater than 25 degrees, so the wheel track satisfies the rule of
thumb for overturn prevention requirement. In the later steps, ground lateral control, ground lateral stability, and structural integrity must be examined to validate the wheel track.
14.12 Landing gear attachment As a natural selection, the nose gear is attached to the fuselage nose. However, for the
main gear, we need to compare the fuselage diameter with the wheel track. It is observed that the fuselage diamter (1 m) is smaller than the wheel track (31.97 m). Hence, the main gear cannot be attached to the fuselage. Thus, main gear may be either attached directly to the wing; or attached under the nacelle. In order to determine the best location, several design requirements must be examined, which is beyond the scope of this example. For the time being, it is decided to attach the landing gear to the wing. Thus, the landing gear height will be:
HLG = 1.581m [14.4]
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15 PERFORMANCE CHARACTERISTICS
16.1 TAKE-OFF PERFORMANCE
Distance from rest to clearance of obstacle in flight path and usually considered in two parts:
« Ground roll - rest to lift-off (SLO) « Airborne distance - lift-off to specified height (35 ft FAR, 50 ft others). « The aircraft will accelerate up to lift-off speed (Vlo = about 1.2 x Vstall) when it will
then be rotated.
16.2 A first-order approximation for ground roll take-off distance may be made from:
SLO = ��!!��L�:8� �Nyz [15.1]
This shows its sensitivity to W (W2) and ρ (1/ ρ2 since T also varies with
ρ). Slo may be reduced by increasing T, S or CL max (high lift devices relate to latter two). An improved approximation for ground roll take-off distance may be made by including drag, rolling resistance and ground effect terms.
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[15.2]
The bracketed term will vary with speed but an approximation may be made by using an instantaneous value for when V = 0.7 x Vlo. In the above equation:
[15.3]
Where h = height above ground, b = wing span.
µr = 0.02 for smooth paved surface, 0.1 for grass.
Ø = 0.861315 [15.4]
16.3 TAKE OF RUNWAY DISTANCE
[15.5]
SLo = ��!! 2���/2��/�����/�2�����2���/!2�/��2{ ��|��� $3�����/����������}~
= 1007.46 m [15.6]
16.4 CLIMBING Increasing the power by advancing the throttle produces a marked difference in the rate of climb. Climb depends upon the reserve power or thrust. Reserve power is the available power over and above that required to maintain horizontal flight at a given speed. Thus, if an airplane is equipped with an engine which produces 200 total available horsepower and the airplane requires only 130 horsepower at a certain level flight speed, the power available for climb is 70 horsepower.
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Although we sometimes use the terms "power" and "thrust" interchangeably, erroneously implying that they are synonymous, it is well to distinguish between the two when discussing climb performance. Work is the product of a force moving through a distance and is usually independent of time. Work is measured by several standards; the most common unit is called a "foot-pound." If a 1 pound mass is raised 1 foot, a work unit of 1 foot-pound has been performed. The common unit of mechanical power is horsepower; one horsepower is work equivalent to lifting 33,000 pounds a vertical distance of 1 foot in 1 minute. The term "power," implies work rate or units of work per unit of time, and as such is a function of the speed at which the force is developed. "Thrust," also a function of work, means the force which imparts a change in the velocity of a mass. This force is measured in pounds but has no element of time or rate. It can be said then, that during a steady climb, the rate of climb is a function of excess thrust.
When the airplane is in steady level flight or with a slight angle of climb, the vertical component of lift is very nearly the same as the actual total lift. Such climbing flight would exist with the lift very nearly equal to the weight. The net thrust of the power plant may be inclined relative to the flight path but this effect will be neglected here for the sake of simplicity. Although the weight of the airplane acts vertically, a component of weight will act rearward along the flight path (Fig.52). If it is assumed that the airplane is in a steady climb with essentially a small inclination of the flight path, the summation of forces along the flight path resolves to the following:
Forces forward = Forces after
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The basic relationship neglects some of the factors which may be of importance for airplanes of very high climb performance. (For example, a more detailed consideration would account for the inclination of thrust from the flight path, lift not being equal to weight, a subsequent change of induced drag, etc.) However this basic relationship will define the principal factors affecting climb performance. This relationship means that, for a given weight of the airplane, the angle of climb depends on the difference between thrust and drag, or the excess thrust (Fig.53). Of course, when the excess thrust is zero, the inclination of the flight path is zero and the airplane will be in steady, level flight. When the thrust is greater than the drag, the excess thrust will allow a climb angle depending on the value of excess thrust. On the other hand, when the thrust is less than the drag, the deficiency of thrust will allow an angle of descent.
16.5 RATE OF CLIMB
Rate of climb = [ T×V stall – D × V stall ] [15.7] = [ (32×103) (59.027) ] – [ 2996.301 × 59.027 ] =13.634 KN [15.8]
Sin γC = z� - �
��� [15.9]
= -0.0744 γC = 1.298 × 10-3 m/s [15.10]
16.6 MANOEUVRES / TURNING
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An aircraft is capable of performing many different types of turns and manoeuvres. Three of the more common turns will be considered here in simplistic terms:
Ù Constant altitude banked turn.
Ù Vertical pull-up manoeuvre.
Ù Vertical pull-down manoeuvre.
In the case of a commercial transport aircraft, it is capable of performing only a constant altitude banked turn and not any vertical pull-up or pull-down manoeuvre.
16.7 CONSTANT ALTITUDE BANKED TURN In steady condition: T = D Force balance gives: W = mg = Lcos θ
Fr = MV2/r = Lsin θ tan θ = V2/(Rg)
So for given speed and turn radius there is only one correct bank angle for a co-ordinate (no sideslip) turn. Maneuverability equations simplified through use of normal load factor (n) = L/W. In the turn, n = L/W = sec θ > 1 and is therefore determined by bank angle. Turn radius (R) and turn rate (θ) are good indicators of aircraft maneuverability. W = L cos θ [15.11] let, θ = 300 W = 189309.97 × cos (30)
W = 163947.24 [15.12]
n = �� [15.13]
apply (16.12) in (16.13)
n = ?CRE>R�R@?DERF@�AF
= 1.1547 [15.14]
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RT = AED�????
R�C? ��?�?BF@�S�? [15.15]
Turn Radius, RT = 9844.33 m
16.8 GLIDING Similar to the steady unaccelerated case but with T = 0.
Force balance gives:
Ø = tan -1 { ?��
~ [15.16]
Ø = tan -1 [ ??E]
Ø = 4.230 [15.17]
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16.9 LANDING PERFORMANCE APPROACH & LANDING Consists of three phases:
Ø Airborne approach at constant glide angle (around 3o) and constant speed.
Ø Flare - transitional manoeuvre with airspeed reduced from about 1.3 Vstall down to touch-down speed.
Ø Ground roll - from touch-down to rest. Ground roll landing distance (s3 or sl) estimated from:
[15.18]
Where, Vav may be taken as 0.7 x touch-down speed (Vt or V2) and Vt is assumed as 1.3 x Vstall.
µr is higher than for take-off since brakes are applied – use µr = 0.4 for paved surface.
If thrust reversers (Tr) are applied, use: Landing Runway distance = 710.3 m [15.19]
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CONCLUSION:
After all analysis and design the first part of the aircraft design project has came for landing. The result of the output specifications tells about features, advantages, disadvantages, future requirements and all amenities in and about the aircraft.
The overview of the SHORT RANGE BUSINESS JET aircraft are very need-of-hour for today corporate and entrepreneurs who are travelling across trans-continent and domestically with less span of time with hassle-free.
The main advantage of this class aircraft is pacifying from commercial jet operators, less-operating cost and fuel consumption, less purchase cost and least time of total travel time enveloped. Since it takes fraction of kilometer for landing it can operate at any airport rather large aircraft whose need mile and more run for take-off. To fly a light business jet just PPL is sufficient so pay scale for the pilot will be low compare to passenger jets.
And always there will be existence of negative if present of positive. Those
disadvantages are it cannot fly inter-continental. Since it design for short range. The specifications of the designed aircraft as follows:
S.No DESIGN PARAMETER MAGNITUDE UNIT
1. Cruising speed 850 Kph
2. Wing span 12.84 m
3. Aircraft Length 19.5 m
4. Wing area 21.84 m2
5. Height 4.8 m
6. Aspect ratio 7.55 (No unit)
7. Wing loading 585.89 Kg/m2
8. Empty weight 7296 Kg
9. Maximum take-off weight 12800 Kg
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10. Pay load 1280 kg
11. No. of engines 2 (No unit)
12. Thrust power 32.00 kN
13. Range 5200 km
14. Service ceiling 13700 m
15. Mach no. 0.715 (No unit)
16. Thrust/weight ratio 0.25 (No unit)
17. Gliding angle 4.23 o
18. Seating capacity 8 (No unit)
19. Fuselage 10.54 m
20. Take-off distance 1007.46 m
21. Landing distance 710.3 m
22. Rate of climb 1.298 × 10-3 m/s
Design is a fine blend of science, creativity, presence of mind and the application of each one of them at the appropriate time. Design of anything needs experience and an optimistic progress towards the ideal system. The scientific society always looks for the best product design. This involves the strong fundamentals in science and mathematics and their skilful applications, which is a tough job endowed upon the design. We have enough hard work for this design project. A design never gets completed in a flutter sense but it is one step further towards ideal system. But during the design of this aircraft, we learnt a lot about aeronautics and its implications when applied to an aircraft design.
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FUTURE WORK In Future project work, in the AIRCRAFT DESIGN PROJECT-II will
consist of following work to develop and complete a Perfect Aircraft by the way of Structural Analysis:
1. V-n diagram for the design study
2. Gust and maneuverability envelopes
3. Critical loading performance and final V-n graph calculation
4. Structural design study – Theory approach
5. Load estimation of wings
6. Load estimation of fuselage.
7. Balancing and Maneuvering loads on tail plane, Aileron and Rudder loads.
8. Detailed structural layouts
9. Design of some components of wings, fuselage
10. Preparation of a detailed design report with drawings.
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BIBLIOGRAPHY
i. Lloyd R. Jenkinson and James F.Marchman III., “Aircraft Design project”,
Butterworth-Heinemann., Burlington, 2003.
ii. Anderson, John D., Jr:Aircraft Performance and Design, McGraw-Hill, Boston,
1999.
iii. Jenkinson, L.R., Page., G.J.,Marchmann, J.F., ‘A model for international teaming in
aircraft design education’, Journal of Aircraft Design, vol 3, No.4, Elsevier,
December 2000.
iv. Jane’s, “All the world’s Aircraft 2004-2005”, Jane’s Information group limited.,
2004.
v. Daniel P.Raymer, “Aircraft Design: A conceptual Design”, American institute of
aeronautics and astronautics, Washington.D.C, 1989.
vi. Norman S. Currey, “Aircraft Landing gear: Principle and Practices”, American
institute of aeronautics and astronautics, Washington.D.C, 1988.
vii. Ira h. Abbott, Albert e. Von Doenhoff,and Louis S. Stivers, Jr,”Summary of Airfoil
Data”, National advisory committee for aeronautics, 1947.
viii. John T.Lowry., “Performance of Light Aircraft”, American institute of aeronautics
and astronautics, Washington.D.C, 1935.
ix. Dr.lng.S.F.Hoerner,”Fluid –Dynamic Drag”,Hoerner fluid dynamics., 1992.
x. J.B. Russell,” Performance and Stability of Aircraft”. ISBN 0-340-63170-8. Arnold
1996.
xi. Mark D. Ardema, Mark C. Chambers, Anthony P. Patron, Andrew S. Hahn, Hirokazu
Miura, and Mark D. Moore, “Analytical Fuselage and Wing Weight Estimation of
Transport Aircraft”,1996.
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WEBSITE REFERENCES
i. www.faculty.dwc.edu
ii. http://www.worldofkrauss.com/
iii. http://www.businessjet.com/aircraft-data/
iv. http://www.wkap.nl/natopco/pco_aga.htm
v. http://www.arnoldpublishers.com/aerodata/appendices/data
vi. http://www.aiaa.org/
vii. http://www.arnoldpublishers.com/aerodata/appendices/data
viii. http://www.utsi.edu/Academic/graduate.html
ix. http://www.honeywell.com/sites/portal?smap=aerospace&page=Turb
ofan-Products&theme=T6
x. http://en.wikipedia.org/wiki/Main_Page
xi. http://adg.stanford.edu/aa241/wingdesign/wingparams.html
xii. http://www.airliners.net/aviation-
forums/tech_ops/read.main/147672/
