Adp Report.docx

ABSTRACT

The aim of this project is to design and conceptualize a small passenger aircraft that can cater to a wide range of clientele ranging from business conglomerates to private organizations and individual parties. It is a term describing a type of aircraft for transporting passengers and air cargo. Such aircraft are most often operated by airlines. It is a aircraft usually of smaller size, designed for transporting groups of business people or wealthy individuals. The project involves the design of a airliner that can accommodate about 20 passengers at full seating layout, with level of comfort that an airliner is expected to provide while incorporating the design specifications and performance parameters of a short range commercial airliner. The aircraft allows for short range transport with better efficiency and reduced fuel consumption and noise levels owing to a state of the art engine and design features.

TABLE OF CONTENTS

 INTRODUCTION TO DESIGN

1.1 Defining a new design 1.1.1 Aircraft Purpose 1.2 Design Motivation 1.3 Design Process 1.4 Conceptual Design 1.5 Design Process Breakdown 2. INTRODUCTION TO AIRLINERS 2.1 Classification of Airliners

2.2 Need for Airliners 3. COMMON COMPARATIVE STUDY 4. COMPARATIVE DATA SHEET 5. COMPARATIVE GRAPHS 6. WEIGHT ESTIMATION 7. AIRFOIL SELECTION 8. DRAG ESTIMATION 9. POWERPLANT SELECTION 10. LANDING GEAR DESIGN 11. PERFORMANCE CHARACTERISTICS 12. CENTRE OF GRAVITY ESTIMATION 13. STABILITY AND CONTROL 14. 3-VIEW DIAGRAM 15. FINALIZED DESIGN PARAMETERS

INTRODUCTION TO DESIGN

Modern aircraft are a complex combination of aerodynamic performance, lightweight durable structures and advanced systems engineering. Air passengers

demand more comfort and more environmentally friendly aircraft. Hence many technical challenges need to be balanced for an aircraft to economically achieve its design specification. Aircraft design is a complex and laborious undertaking with a number of factors and details that are required to be checked to obtain optimum the final envisioned product. The design process begins from scratch and involves a number of calculations, logistic planning, design and real world considerations, and a level head to meet any hurdle head on. Every airplane goes through many changes in design before it is finally built in a factory. These steps between the first ideas for an airplane and the time when it is actually flown make up the design process. Along the way, engineers think about four main areas of aeronautics: Aerodynamics, Propulsion, Structures and Materials, and Stability and Control. Aerodynamics is the study of how air flows around an airplane. In order for an airplane to fly at all, air must flow over and under its wings. The more aerodynamic, or streamlined the airplane is, the less resistance it has against the air. If air can move around the airplane easier, the airplane's engines have less 2

Work to do. This means the engines do not have to be as big or eat up as much fuel which makes the airplane more lightweight and easier to fly. Engineers have to think about what type of airplane they are designing because certain airplanes need to be aerodynamic in certain ways. For example, fighter jets maneuver and turn quickly and fly faster than sound (supersonic flight) over short distances. Most passenger airplanes, on the other hand, fly below the speed of sound (subsonic flight) for long periods of time. Propulsion is the study of what kind of engine and power an airplane needs. An airplane needs to have the right kind of engine for the kind of job that it has. A passenger jet carries many passengers and a lot of heavy cargo over long distances so its engines need to use fuel very efficiently. Engineers are also trying to make airplane engines quieter so they do not bother the passengers onboard or the neighborhoods they are flying over. Another important concern is making the exhaust cleaner and more environmentally friendly. Just like automobiles, airplane exhaust contains chemicals that can damage the earth's environment. Structures and Materials is the study of how strong the airplane is and what materials will be used to build it. It is really important for an airplane to be as lightweight as possible. The less weight an airplane has, the less work the engines have to do and the farther it can fly. It is tough designing an airplane that is lightweight and strong at the same time. In the past, airplanes were 3

Usually made out of lightweight metals like aluminum, but today a lot of engineers are thinking about using composites in their designs. Composites look and feel like plastic, but are stronger than most metals. Engineers also need to make sure that airplanes not only fly well, but are also easy to build and maintain. Stability and Control is the study of how an airplane handles and interacts to pilot input and feed. Pilots in the cockpit have a lot of data to read from the airplane's computers or displays. Some of this information could include the airplane's speed, altitude, direction, and fuel levels as well as upcoming weather conditions and other instructions from ground control. The pilot needs to be able to process the correct data quickly, to think about what kind of action needs to be taken, and to react in an appropriate way. Meanwhile, the airplane should display information to the pilot in an easy-to-read and easy-to-understand way. The controls in the cockpit should be within easy reach and just where the pilot expects them to be. It is also important that the airplane responds quickly and accurately to the pilot's instructions and maneuvers. “A beautiful aircraft is the expression of the genius of a great engineer who is also a great artist.”

When you look at aircraft, it is easy to observe that they have a number of common features: wings, a tail with vertical and horizontal wing sections, engines to propel them through the air, and a fuselage to carry passengers or cargo. If, however, you take a more critical look beyond the gross features, you also can see subtle, and sometimes not so subtle, differences. This is where design comes into play. Each and every aircraft is built for a specific task, and the design is worked around the requirement and need of the aircraft. The design is modeled about the aircraft role and type and not the other way around. Thus, this is why airplanes differ from each other and are conceptualized differently. Aircrafts that fall in the same category may

have similar specifications and performance parameters, albeit with a few design changes. Design is a pivotal part of any operation. Without a fixed idea or knowledge of required aircraft, it is not possible to conceive the end product. Airplane design is both an art and a science. In that respect it is difficult to learn by reading a book; rather, it must be experienced and practiced. However, we can offer the following definition and then attempt to explain it. Airplane design is the intellectual engineering process of creating on paper (or on a computer screen) a flying machine to (1) meet certain specifications and requirements established by potential users (or as perceived by the manufacturer) and/or (2) pioneer innovative, new ideas and technology. An example of the former is the design of most commercial transports, starting at least with the Douglas DC-1 in 1932, 5 which was designed to meet or exceed various specifications by an airplane company. (The airline was TWA, named Transcontinental and Western Air at that time.) An example of the latter is the design of the rocket-powered Bell X-1, the first airplane to exceed the speed of sound in level or climbing flight (October 14, 1947). The design process is indeed an intellectual activity, but a rather special one that is tempered by good intuition developed via experience, by attention paid to successful airplane designs that have been used in the past, and by (generally proprietary) design procedures and databases (handbooks, etc) that are a part of every airplane manufacturer. 1. Defining a new design 1.2 The design of an aircraft draws on a number of basic areas of aerospace engineering. These include aerodynamics, propulsion, light- weight structures and control. Each of these areas involves parameters that govern the size, shape, weight and performance of an aircraft. Although we generally try to seek optimum in all these aspects, with an aircraft, this is practically impossible to achieve. The reason is that in many cases, optimizing one characteristic degrades another. There are many performance aspects that can be specified by the mission requirements. These include:  The aircraft purpose or mission profile

 The type(s) and amount of payload  The cruise and maximum speeds  The normal cruise altitude  The range or radius with normal payload  The endurance  The take-off distance at the maximum weight  The purchase cost 1.1.1 Aircraft Purpose The starting point of any new aircraft is to clearly identify its purpose. With this, it is often possible to place a design into a general category. Such categories include combat aircraft, passenger or cargo transports, and general aviation aircraft. These may also be further refined into subcategories based on particular design objectives such as range (short or long), take-off or landing distances, maximum speed, etc. The process of categorizing is useful in identifying any existing aircraft that might be used in making comparisons to a proposed design. With modern military aircraft, the purpose for a new aircraft generally comes from a military program office. For example, the mission specifications for the X-29 pictured in figure 1.1 came from a 1977 request for proposals from the U.S. Air Force Flight Dynamics Laboratory in which they were seeking a research aircraft that would explore the forward swept wing 7 Concept and validate studies that indicated such a design could provide better control and lift qualities in extreme maneuvers. With modern commercial aircraft, a proposal for a new design usually comes as the response to internal studies that aim to project future market needs. For example, the specifications for the Boeing

commercial aircraft (B-777) were based on the interest of commercial airlines to have a twin-engine aircraft with a payload and range in between those of the existing B-767 and B-747 aircraft. Since it is not usually possible to optimize all of the performance aspects in an aircraft, defining the purpose leads the way in setting which of these aspects will be the “design drivers.” For example, with the B-777, two of the prominent design drivers were range and payload. 1.2 Design Motivation Fundamentally, an aircraft is a structure. Aircraft designers design structures. The structures are shaped to give them desired aerodynamic characteristics, and the materials and structures of their engines are chosen and shaped so they can provide needed thrust. Even seats, control sticks, and windows are structures, all of which must be designed for optimum performance. Designing aircraft structures is particularly challenging, because their weight must be kept to a minimum. There is always a tradeoff between structural strength and weight. A good aircraft structure is one which provides all the strength and rigidity to 8 Allow the aircraft to meet all its design requirements, but which weighs no more than necessary. Any excess structural weight often makes the aircraft cost more to build and almost always makes it cost more to operate. As with small excesses of aircraft drag, a small percentage of total aircraft weight used for structure instead of payload can make the difference between a profitable airliner or successful tactical fighter and a failure. Designing aircraft structures involves determining the loads on the structure, planning the general shape and layout, choosing materials, and then shaping, sizing and optimizing its many components to give every part just enough strength without excess weight. Since aircraft structures have relatively low densities, much of their interiors are typically empty space which in the complete aircraft is filled with equipment, payload, and fuel. Careful layout of the aircraft structure ensures structural components are placed within the interior of the structure so they carry the required loads efficiently and do not interfere with placement of other components and payload within the space. Choice of materials for the structure can profoundly influence weight, cost, and manufacturing difficulty. The extreme complexity of modern aircraft structures makes optimal sizing of individual components particularly challenging. An understanding of basic structural concepts and techniques for designing efficient structures is essential to every aircraft designer.

1.3 Design Process The process of designing an aircraft and taking it to the point of a flight test article consists of a sequence of steps, as illustrated in the figure. It starts by identifying a need or capability for a new aircraft that is brought about by (1) a perceived market potential and (2) technological advances made through research and development. The former will include a market-share forecast, which attempts to examine factors that might impact future sales of a new design. These factors include the need for a new design of a specific size and performance, the number of competing designs, and the commonality of features with existing aircraft. As a rule, a new design with competitive performance and cost will have an equal share of new sales with existing competitors. The needs and capabilities of a new aircraft that are determined in a market survey go to define the mission requirements for a conceptual aircraft. These are compiled in the form of a design proposal that includes (1) the motivation for initiating a new design and (2) the “technology readiness” of new technology for incorporation into a new design. It is essential that the mission requirements be defined before the design can be started. Based on these, the most important performance aspects or “design drivers” can be identified and optimized above all others. Following the design proposal, the next step is to produce a conceptual design. The conceptual design develops the first general size and configuration for a new aircraft. It involves the estimates of the weights and the choice of aerodynamics characteristics that will best suited to the mission requirements stated in the design proposal.

The conceptual design is driven by the mission requirements, which are set in the design proposal. In some cases, these may not be attainable so that the requirement may need to be relaxed in one or more areas. This is shown in the iterative loop in the flow chart. When the mission requirements are satisfied, the design moves to the next phase, which is the preliminary design. 1.4 Conceptual Design This article deals with the steps involved in the conceptual design of an aircraft. It is broken down in to several elements, which are followed in order. These consist of: 1. Literature survey

2. Preliminary data acquisition 3. Estimation of aircraft weight a. Maximum take-off weight b. Empty weight of the aircraft c. Weight of the fuel d. Fuel tank capacity 4. Estimation of critical performance parameters a. Wing area b. Lift and drag coefficients c. Wing loading d. Power loading e. Thrust to weight ratio 5. Engine 6. Performance curves 7. 3 View diagrams

1.5 Design Process Breakdown

 Conceptual Design  Competing concepts What Drives the design? evaluated.  Performance goals established. Will it work/meet requirement?  Preferred concept selected. What does it look like?

 Preliminary Design  Refined sizing of preferred concept tests.

Do serious wind tunnel tests.

 Designs examined data/establish parameters.

Make actual cost estimate.

 Some changes allowed.

 Detail Design  Final detail design Certification process  Drawings Released Components /systems tests  Detailed performance Manufacturing  Only “tweaking” of design allowed.

Flight control system design

2. INTRODUCTION TO AIRLINER

An airliner is a type of aircraft for transporting passengers and air cargo. Such aircraft are most often operated by airliners. Although the definition of an airliner can vary from country to country, an airliner is typically defined as an aircraft intended for carrying multiple passengers or cargo in commercial service. The largest airliners are wide body jets. These aircraft are frequently called twin aisle aircraft because they generally have two separate aisles running from the front to the back of the passenger cabin. These aircraft are usually used for long haul flights between airline hub and major cities with many passengers. A smaller, more common class of airliners is the narrow body or single aisle aircraft. These smaller airliners are generally used for short to medium-distance flights with fewer passengers than their wide-body counterparts. Regional airliners typically seat fewer than 100 passengers and may be powered by turboprop or turbofan. These airliners are the non-mainline counterparts to the larger aircraft operated by the major carriers, legacy carriers, and flag carriers and are used to feed traffic into the large airline hubs. These regional routes then form the spokes of a hub-and-spoke air transport model. The lightest (light aircraft) of short haul regional feeder airliner type aircraft that carry 19 or fewer passenger seats are called commuter aircraft, commuterliners,

feederliners, and air taxis, depending on their size, engines, how they are marketed, region of the world, and seating configurations. The Beechcraft 1900, for example, has only 19 seats.

2.1 CLASSIFICATION OF AIRCRAFTS 1. AIRLINER AIRCRAFT General characteristics and performance of SAAB 2000 1. Crew = two 2. Capacity = 50-58 Passengers 3. Payload = 5900Kg (13,010lb) 4. Length = 28.27m(89ft 6 inch) 5. Wingspan = 27.76m(81ft 3 inch) 6. Height = 7.73m(25ft 4 inch) 7. Wing Area = 55.7m2(600sq. ft) 8. Airfoil = NASA MS(1)013 9. Aspect Ratio = 11:7 10.Empty Weight = 13,800Kg(30,424lb) 11.Max.takeoff weight = 22,800Kg(50,265lb) 12.Powerplant = 2xallison AE 2100A turboprop,3096Kw(4,152shp) each 13.Propellers = Six blades constant speed dowty propellers , 1 per engine PERFORMANCE OF SAAB 2000 1 Cruise Speed = 665km/hr. (370Knots, 424mph) 2 Range = 2869Km (1,549nmi;1,782mi) at Long cruise 3 Service Ceiling = 9450m (3100ft)

4 Rate of climb = 11.4m/s (2,250ft/min) 2. AMPHIBIAN AIRCRAFT: An Amphibian aircraft or Amphibious aircraft is an aircraft that can take off and land both on water and land. General Charactertics of RAAF Wackett Widgeon: 1 Crew = One 2. Capacity =Three 3. Length = 9.010m(29ft 3 inch) 4. Upper wingspan = 28.9345m(29ft 3.75inch) 5. Lower wingspan = 8.9345m(29 ft 3.75inch) 6. Height = 4.2863m (14ft 0.75inch) to tip of propeller 7. Wing Area = 39.24m2(424.3sq ft) 8. Empty weight = 1315Kg(1310lb) 9. Gross weight = 1796Kg(3960lb) 10 Fuel capacity = 44 Gallons 11 Propellers = 4 bladed wooden, 8ft 6 inch (2.59) diameter PERFORMANCE: 1 Maximum Speed = 103mph(166km/hr,90km) at sea level 2 Endurance = 3 hours 3 Service ceiling = 3400m (11000ft)

3 MILITARY AIRCRAFT: A military aircraft is any fixed wing or rotary wing aircraft that is operated by a legal or insurrectionary armed source of any type. General Characteristics of Republic Thunderbolt-II 1. Crew = One 2. Length = 16.26m(53ft 4 inch) 3. Wingspan = 17.53m(57ft 6 inch) 4. Height = 4.47m(14 ft. 8 inch) 5. Wing Area = 47m2(506ft2) 6. Airfoil = NACA6716 root, NACA 6713 tip 7. Empty weight = 11321Kg(24959lb) 8. Loaded weight = 13782Kg(30384lb) CAS mission = 21361Kg (47094kb) Anti-armor mission = 19083Kg (42017lb) 9. Max.takeoff weight = 23000Kg(50,000lb) 10.Internal Fuel capacity = 4990Kg(11000lb)\ 11.Powerplant = 2xgeneral electric TF34-GE-100A turbofans,9065 lbf(40.32KN)each PERFORMANCE; 1. Never exceed speed = 518mph,833km/hr(450Knots) at 5000 ft 2. Maximum speed = 706mp/h,439km/h(381knots)at sea level 3. Cruise speed = 340mph, 560km/h(300knots) 4. Stall speed = 138mph,220 km/h(120knots) 5. Service ceiling = 13700m(45000ft) 6. Rate of climb = 30m/s(6000 ft/min) 7. Wing loading = 482kg/m2(99lb/ft2) 8. Thrust/weight = 0.36

4. CARGO AIRCRAFT: Cargo aircraft is a fixed wing aircraft that is assigned or connected for the carrier of cargo rather than passengers. General Characteristics of Lockheed C-5 Galaxy 

         

Crew: 7 typical (aircraft commander, pilot, two flight engineers, three loadmasters) 4 minimum (Pilot, copilot, two flight engineers) Payload: 270,000 lb. (122,470 kg) Length: 247 ft. 1 in (75.31 m) Wingspan : 222 ft. 9 in (67.89 m) Height: 65 ft. 1 in (19.84 m) Wing area: 6,200 ft2 (576 m2) Empty Weight: 380,000 lb. (172,371 kg) Useful load: 389,000 lb. (176,450 kg) Loaded weight: 769,000 lb. (348,800 kg) Max. Takeoff Weight : 840,000 lb. (381,000 kg) ; Powerplant: 4 × General electric TF 39 -GE-1C high-bypass turbofan, 43,000 lbf (190 kN) each

PERFORMANCE:

         

Maximum Speed : Mach 0.79 (462 kn, 531 mph, 855 km/h) Cruise speed : Mach 0.77 (450 kn, 518 mph, 833 km/h) Range: 2,400 nmi (2,760 mi, 4,440 km) with a 263,200 lb (119,400 kg) payload Service ceiling : 35,700 ft. (10,600 m) at 615,000 lb (279,000 kg) gross weight Rate of climb : 1,800 ft./min (9.14 m/s) Wing Loading : 120 lb./ft2 (610 kg/m2) Thrust/weight : 0.22 Takeoff roll: 8,400 ft. (2,600 m) Landing roll: 3,600 ft. (1,100 m) Fuel capacity: 51,150 US gal (193,600 L)

5. AIRAMBULANCE AIRCRAFT: It is a specially outfitted fixed wing aircraft that transports injured or sick people in a medical emergency or over distance or terrain impractical for a conventional ground ambulance. General Characteristics of SAI KZIII Aircraft:            

Crew: one Capacity: one Length: 6.6 m (21 ft. 8 in) Wingspan: 9.6 m (31 ft. 6 in) Height: 2.1 m (6 ft. 11 in) Wing area: 13 m2 (140 sq. ft.) Airfoil : NACA 120213 Empty weight: 368 kg (811 lb.) Max takeoff weight: 706 kg (1,556 lb.) with auxiliary tank full, 34 kg (75 lb.) and extra 21 kg (46 lb.) luggage Fuel capacity: 41 l (11 US gal; 9.0 imp gal) Powerplant: 1 × Blackburn Cirrus Minor II 4-cyl. inverted air-cooled in-line piston engine, 75 kW (100 hp) Propellers: 2-bladed Weybridge, 1.9 m (6 ft. 3 in) diameter wooden fixed pitch propeller. Performance:

        

Maximum speed: 185 km/h (115 mph; 100 kn) at sea level Cruise speed: 170 km/h (106 mph; 92 kn) at 2,200 rpm Range: 500 km (311 mi; 270 nmi) Ferry range: 805 km (500 mi; 435 nmi) Service ceiling: 4,115 m (13,501 ft.) Maximum glide ratio: 1:8 (flaps up); 1:5 flaps down Rate of climb: 3.6 m/s (700 ft./min) Wing loading: 48.83 kg/m2 (10.00 lb./sq. ft.) Power/mass : 9.72 kg/kW (16 lb./hp)

6. PASSENGER AIRCRAFT: It is a type of aircraft for transporting passenger and air cargo. Such aircraft are mostly operated by airliners.

General Characteristics of Boeing 747 SP: 1. Cockpit crew : three 2. Seating layout ; 276 3. Seating for exit layout : 400 4. Length : 56.3m(184ft. 9inch) 5. Wingspan ; 59.6m(195ft.8inch) 6. Wing area = 511m2(5500ft2) 7. Wing sweep = 37.5 degree 8. Aspect Ratio = 7 9. Tail Height = 19.9m(65ft.5 inch) 10.Cargo capacity = 110m3(3900cubic feet) 11.Empty weight = 152,900Kg (337,100lb.) 12.Max.takeoff weight = 320,000Kg (700,000lb.) 13.Cruise speed = 907Km/hr. 14.Takeoff = 320,00Kg (700,000lb.) 15.Range = 10,800Km;6701mi) 16.Fuel Capacity = 190,360Litre(50,359 US gal) 17.Turbofan engines = Pratt&whittney JT9 D-7 18.Thrust per engine = 206-253KN(46,300-59,900lbf)

7. EXPERIMENTAL AIRCRAFT: It is an aircraft that has not yet been fully proven in flight. General Characteristics of FMA I. Ae.37:       

Crew: 1 Length: 11.78 m (38 ft. 8 in) Wingspan: 10 m (32 ft. 10 in) Height: 4.92 m (16 ft. 2 in) Wing area: 48 m2 (520 sq. ft.) Empty weight: 3,300 kg (7,275 lb.) Gross weight: 4,800 kg (10,582 lb.)

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Powerplant: 1 × Rolls –Royce Derwent V Turbojet , 16.02 kN (3,600 lbf) thrust

Performance   

Maximum speed: 800 km/h (497 mph; 432 kn) Range: 2,000 km (1,243 mi; 1,080 nmi) Service ceiling: 11,000 m (36,000 ft.)

9. SAILPLANE AIRCRAFT: It is a type of glider aircraft used in the sport of fighting. General Charactertics of DG Flugzeugban DDG-800:          

Crew: One pilot Capacity: 156 kg (343 lb.) water ballast Length: 7.06 m (23 ft. 2 in) Wingspan: 18.00 m (59 ft. 1 in) Height: ca. 1.35 m (4 ft. 5 in) Wing area: 11.8 m2 (127 ft2) Aspect ratio: 27.4 Empty weight: ca. 344 kg (757 lb.) Gross weight: 600 kg (1,320 lb.) Powerplant: 1 × Solo, 40 kW (ca. 53 hp)

Performance   

Maximum speed: 270 km/h (170 mph) Maximum glide ratio: 50 (51.5 with winglets) Rate of climb: 5.2 m/s (1,000 ft./min)

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Rate of sink: 0.5 m/s (98 ft./min)

10. AEROBATICS AIRCRAFT: It is the aircraft which is used for practice of flying maneuvers involving aircraft attitudes that are not used in normal flight. General Characteristics of Renard R.34:  Crew = Two  Wingspan = 9.20m(30ft. 2 inch)  Powerplant = 1xRenard 200 radial,180Kw PERFORMANCE:  Maximum Speed = 215km/hr. (134mph;116Kn) 11.FIRE FIGHTERS AIRCRAFT: It is the use of aircraft and other aerial resources to combat wildfires. General Characteristics of Lockheed P-2 Neptune: 

Crew: 9-11  Length: 77 ft. 10 in (23.72 m)  Wingspan : 100 ft. 0 in (30.48 m)  Height: 28 ft. 4 in (8.56 m)  Wing area: 1,000 ft² (92.9 m²)  Empty Weight : 34,875 lb (15,819 kg)  Max. Takeoff Weight : 64,100 lb. (29,076 kg)  Powerplant : 2 ×Wright R-3350 -26W Cyclone-18 Radial engine, 3,200 hp (2,386 kW) wet each  Propellers: 4 bladed propeller, 1 per engine

Performance   

Maximum Speed : 278 kn (313 mph) (515 km/h) Cruise speed : 155 kn (174 mph) (286 km/h) (max) Range : 3,458 nmi (3,903 mi) (6,406 km)

12. AGRICULTURAL AIRCRAFT: It is the aircraft that has been built or converted for agriculture use-usually aerial application of pesticides or fertilizers. General Characteristics of Cessna 188:               

Crew: one pilot Capacity: Hopper: 280 US gal (230 imp gal; 1,100 L) Length: 26 ft. 3 in (8.00 m) Wingspan : 41 ft. 8 in (12.70 m) Height: 7 ft. 8 1⁄2 in (2.35 m) Wing area: 205 ft² (19.05 m²) Airfoil : NACA 2412 (modified) Empty Weight : With no dispersal equipment installed: 2059 lb. (934 kg) With liquid dispersal system: 2214 lb. (1004 kg) Loaded weight: 3300 lb. (1497 kg) Max. Takeoff weight : Normal category: 3300 lb. (1496 kg) Restricted category: 4200 lb. (1905 kg) Powerplant : 1 × Continental 10-520 -D air-cooled Flat six, 300 hp (224 kW)

Performance       

Maximum speed : 121 mph (105 knots, 195 km/h) Cruise speed : 113 mph (98 knots, 182 km/h) (at 75% power) Stall speed : Clean: 61 mph (53 knots, 99 km/h) With full flaps: 57 mph (50 knots, 92 km/h) Range : 370 mi (321 nmi, 595 km) Service ceiling : 11,100 ft. (3,383 m)

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Length Capacity Wingspan Height Wing area Empty weight Max. Takeoff weight Powerplant

Thrust Max.speed Cruise speed Range Service ceiling Airfoil

Rate of climb : 690 ft./min (3.5 m/s) 3. COMMON COMPARATIVE STUDY

Saab 90 Scandia

Aerospatiale N262

Short SC7 Skyvan

21.30m 24-32 Pass 28m 7.40m 85.70m2 9960Kg 15900Kg

19.28m 29 Pass 21.90m 6.21m 55.00m2 6654Kg 10300Kg

12.21m 19 Pass 19.78m 4.6m 35.12m2 3331Kg 5670Kg

2xpratt whitney R2180-E-twin wasp E14 Cylinder radial engine 1361Kw each 450Km/h at 2600m 340Km/h 2650km 7500m

2xturbomeca bastan V1 C turboprop 794kw each

Rate of climb 7.5m/s Stall speed Wing loading Crew 4

De Havilland Canda BMC 6 twin otter 19.8m 19-20 Pass 19.8m 5.9m 39m2 2653Kg 5246Kg

Embraer

2xgarrett Ai research TPE-331-201 ,turboprop 533kw

Pratt&whittn ey PT 6A-20 550 Shp each

Pratt&whittney Canada PT 6A34 turboprop

794 Kw each

533Kw each

680Kw each

559 kw each

385Km/h

325Km/h

459 km/hr

360Km/hr 1110km 7300m

317Km/hr 1117km 6858m

297km/h at cruise altitude 278Km/h 1427km 7620m

NACA 23016 at root, NACA 23012 at tip 6.3m/s 128Km/hr 1

15.10 18 pass 15.33m 4.92 29.10m2 3393kg 5900kg

341 km/hr 1964km 6550m NACA 63A516

500m/min 136.6Kg/m2 1-2

8.1m/s 107Km/hr 1-2

803m/s 85.25km/hr 53.60kg/m2 2

Aspect Ratio

91:1

8.7:1

11.1:1

10:1

8:1

TABLE 1

Let L-410 Turbojet

Short 330

CASAC212 Aviocar

Saab 2000

Capacity

19 Pass

30Pass

26Pass

50-58Pass

Length

14.47m

17.69m

16.20m

27.28m

Wingspan

19.48m

22.76m

20.28m

24.76m

Height

5.83m

4.95m

6.30m

7.73m

Wing Area

34.86m2

42.1m2

41m2

55.7m2

Empty Weight

4200Kg

6680Kg

3780Kg

13800Kg

Max.Takeof f Weight

6600Kg

10387Kg

8100Kg

22800Kg

Powerplant

Prattxwhitne y 2xCanada PT-6A TUrbpprop

2xPratt7whitne y Canada PT6A-45 R Turboprop

Garett Ai 2xResearc hR 33/Turbo

2xAllison AE 2100A turbopro p

Thrust

597Kw

833Kw

690Kw

3096Kw

Maximum Speed

405Km/h

445Km/h

370Km/h

Cruise speed 405Km/h

352Km/h

360Km/

665Km/h

Range

1500km

1695Km

1960Km

2869Km

Service ceiling

8832m

6400m

7925m

9450m

NACA 63A

NACA

NASA

Airfoil

653-218

MS(1) 013 11.4m/s

Rate of climb

6m/s

8.3m/s

Stall speed

136Km/h

145Km/h

Wing Loading

247kg/m3

Crew

3

2

2

Aspect Ratio 11:45:1

12:1

10:1

11:1

4. COMPARATIVE GRAPHS 1.

Range vs Cruise speed

2.

Range vs empty weight

3. Aspect Ratio vs Range

4. Aspect Ratio vs Cruise speed

5. Rate of climb Vs Aspect Ratio

5 Weight Estimation

1. 2. 3. 4. 5.

Number of passengers = 20 V(cruise) = 340km/hr Safe Range = 1960kms Service Ceiling = 6500m or 6.5km Balanced field length for take off = 0.5m Gross Weight, Wo 1. Weight of payload plus Weight of crew 2. Fuel Fraction 3. We/Wo(Empty weight fraction) 4. Solve : W0 = Wcrew + Wpayload / [1 – (Wf/W0) – (We/W0)]

I

Wp + We = Wp + Wbaggage + Wcarry-on +Wcrew

Weight of 1 person = 100Kgs Weight of 20 Persons = 2000Kgs Crew Member = 100*4 = 400Kgs

Wp + We = 2000 + 400 = 2400Kgs

II Fuel Fraction a) Warm and take off Wo = Take off weight W1 = End take off weight W1/Wo = 0.98 b) Climb W2 = Weight at end of climb W2/W1 = 0.99 c) Cruise W3 = End of cruise weight W3/W2 = exp { -R x BSFC / (3600 x ͷp x L/D)max. } Safe Range = 1960 Maximum Head Wind = 15m/s or 54km/hr T = 1960/340 = 5.6Hrs Extra Drag = 5.76 x 54 = 311Kms Alternate Airport = 300Kms Total Range (R) = 1960 + 311 +300 = 2571Kms ͷp = 0.8 BSFC = 2.7N/Kw.hr (L/D)Max. Calculation Cd = Cdo + K𝐶𝑙 2 𝐶𝐷 = 𝐶𝐷0 + 𝐶𝐷𝑜 = 𝐶𝐹𝑒 +

1 𝜋𝑒𝐴

𝐶𝑙 2

𝑆𝑤𝑒𝑡 𝐵

𝐶𝐹𝑒 = 0.003 2

1

𝑆𝑤𝑒𝑡 = π𝑑𝑓 𝑙𝑓 (1 − ) 2⁄3 (1 + ) 3 𝜆 𝑓

𝑑𝑓 = 1.8m 𝑙𝑓 = 9.55m

𝜆𝑓 =

9.55 1.6

= 5.968≈6

𝑆𝑤𝑒𝑡 = π x 1.6 x 9.55x 0.768x1.027 = 37.61𝑚2 𝐶𝑑𝑜 = 𝐶𝑓𝑒 +

37.61 23.10

= 0.0038 e=

1

; Q = Inviscid Drag (Wake) , P = Viscid drag(skin friction &

𝑄+𝑃𝜋𝐴

pressure) e= =

1 1.05+0.007×𝜋×23.10 1

1.6899

= 0.59≈0.60 K= =

1 𝜋𝑒AR 1

𝜋×0.6×8.1

= 0.065

𝐶𝐷 = 𝐶𝑑𝑜 + K𝑐𝑙2 = 0.0038 + 0.065𝑐𝑙2 𝐿

1

𝐷

2√𝐾𝐶𝐷𝑜

( ) Max. = = 31.8

W3/W2 = exp {-R x BSFC / (3600 x ͷp x L/D)max. } = 0.97 (D) Loiter W4/W3 = exp {-E x BSFC x V / (1000 x ͷp x L/D)}

𝐶𝑙 = √ V=√

3𝐶𝑑𝑜 𝐾

= 0.418

2𝑊3 𝜌𝑆𝐶𝑙

W3 = (

𝑊1 𝑊2 𝑊3

)(

)(

𝑊𝑜 𝑊1 𝑊2

)Wo

= 0.98× 0.99 × 1.07 × 𝑊𝑜 = 1.04Wo V = 104.885m/s ≈ 105m/s (From above V equation) ͷ𝑝 = 0.75 , BSFC = 2.85N/Kw-Hr L

𝐿

(D) loiter. = 0.866 × (𝐷)Max. = 0.866 ×31.8 = 27.5

W4/W3 = exp {-E x BSFC x V / (1000 x ͷp x L/D)} = 1.0019 ≈ 1.002 10Mins before Landing (E) Ascent, Descent and Taxing 𝑊5 𝑊4 𝑊5 𝑊𝑜

= 0.98

=

𝑊1 𝑊𝑜

×

𝑊2 𝑊1

×

𝑊3 𝑊2

×

𝑊4 𝑊3

×

𝑊5 𝑊4

= 0.98 × 0.99 × 0.97 × 1.002 × 0.98 =0.9173≈0.92 Allowing 6 % trapped Fuel

𝑊𝑓 𝑊𝑜

= 𝐾𝑇𝐹 (1 −

𝑊5 𝑊𝑜

)

= 1.06(1 – 0.92) = 0.0848 𝑊𝑒

III)

𝑊𝑜

𝑊𝑒 𝑊𝑜

=A

= A𝑊𝑜𝑐 = 0.92 𝑊𝑜−0.05

IV) 𝑊𝑜 = Wp + Wc/1-Wf/Wo-We/Wo = 2400 /1-0.0848-0.92𝑊𝑜−0.05

𝑊𝑜 (Guessed

𝑊𝑒 𝑊𝑜

(III)

8000 0.58 7312 0.58 7321 0.58 7366 0.58 Wo (Gross Weight) = 7366Kgf

WING LOADING (W/S)

Stall:

Vstall = Vapproach / 1.3

= 30.78/ 1.3

𝑊𝑜 (IV) 7312 7371 7366 7366

Vstall = 23.68 m/s

CLmax = CLʌ=0 cos˄1/4

= 1.722 cos 250 = 1.56

W/S = ( ρVstall2 CLmax )/2

= (1.225 x 23.682 x 1.56) / (2 x 10)

W/S = 53.60 kg/m2

Landing:

Ground roll distance, S = 80 (W/S) / ( CLmax )

(W/S) = S (σCLmax )/ 80

= (259.70x 0.82 x 1.56) / 80

(W/S)landing = 4.15 kg/m2 Cruise:

Skin friction coefficient, Cfe = 0.003 (subsonic)

Assuming Swet/S = 1.29

Parasite drag CDo = Cfe (Swet / Sref)

CDo = 0.0038

e= .6

Vcruise= 94.72 m/s

ρ = .660 kg/m2

q0 = ℓ V2 / 2

= 296.07 kg/m2

(W/S)optimum cruise = q0 (πeARCD0 /3)0.5

= 41.173 kg/m2

(W/S)takeoff = (W/S)opt cruise x (W1/W0)-1 x (W2/W1)-1

= 41.173 x (0.98)-1 x (0.99)-1 (W/S)takeoff = 42.416 kg/m2

Wing Loading 60 50 40 30

Wing Loading

20 10 0 sea level

take off

cruise

landing

6. AIRFOIL SELECTION AIRFOIL GEOMETRY An airfoil is a surface designed to obtain a desirable reaction from the air through which it moves.

Chord line: Straight line connecting leading edge and trailing edge. Thickness: Measured perpendicular to chord line as a % of it. Camber: Curvature of section – perpendicular distance of section mid-points from chord line as a % of it. ANGLE OF ATTACK ( )

Angle of attack ( ) is the angle between the free stream and the chord line. 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 airfoil categories: Early on, airfoil selection was based on trial & error. NACA 4 digit was introduced during the 1930’s. NACA 5-digit is aimed at pushing position of max camber forwards for increased Clmax. NACA 6-digit is designed for lower drag by increasing region of laminar flow. The modern airfoil 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) It becomes necessary to use high speed airfoils, i.e., the 6x series, which have been designed to suit high subsonic cruise Mach numbers.

NACA 23016 (Root Airfoil) Design 𝐶𝑙 = 0.3 Maximum Chamber = 15% of airfoil Maximum Thickness = 16% NACA 23012 (Tip Airfoil) Design 𝐶𝑙 = 0.3 Maximum Camber = 15% Maximum Thickness = 12%

Maximum Thickness of airfoil desired to producec 𝐶𝑙 max. is 16% 2 𝐶𝑙 max. = (2*W/s)/ρ𝑉𝑆𝑡𝑎𝑙𝑙

𝐶𝑙 max. = 2 × 53.60/1.225 × 23.682 𝐶𝑙 max. = 1.56

2𝑤/𝑠

𝐶𝐿 cruise =

2 𝜌𝑉𝑐𝑟𝑢𝑖𝑠𝑒

= 0.31 S = 29.10𝑚2 b = 15.33m 2𝑆 𝑏(1+𝜆)

=

= 𝑐𝑅 (Root chord)

2 ×29.10 15.33 ×(1+0.25)

= 3.04m λ × 𝑐𝑅 = 𝑐𝑇 (Tip Chord) = 0.25 × 3.07 = 0.76m Mean aerodynamic Chord 2

1+𝜆+𝜆2

3

1+𝜆

Mac = ⌊( 2

1+0.25+0.252

3

1+0.25

= ⌊(

)⌋c

)⌋ = 2.128

Sweep angle LE = Sweep angle c/4 + [(1-Sweep Angle ) / AR (1+Sweep Angle)] Sweep Angle (LE) = 25 + 0.75 / (8.1 x 1.25) = 25.07 𝑀𝑒𝑓𝑓 = 𝑀∞ cos(sweep angle) LE = 0.37 cos25.07 = 0.335 Wing sweep reduces effective Mach number over the wing. 2 𝛽 = √1 − 𝑀𝑒𝑓𝑓 = 0.3 dCL/d𝛼 = 2.π.AR / [2+ √{4+(AR𝛽 2 . (1+tan2Sweep angle t/c / 𝛽 2 )}] = 2 x π x 8.1/ [2+ {4+(8.1 x 0.32 x (1+𝑡𝑎𝑛2 25 /0.32 )}] = 4.56

Wing Design Parameters W/S S AR B 𝜆 𝑐𝑟

Values 53.6Kg/𝑚2 26.10𝑚2 8.1 15.33 0.25 3.04 0.76 2.128 25 0.31

𝐶𝑚 Sweep Angle 𝐶𝑙

7 Powerplant Selection Camparision of turboprop engine Engine

Length

Allison 3 m or AE 2100A 118 inch Garett 1.16m Research TPE 331 Turboprop Pratt & 1.84m Whittney Canada PT-6A Turboprop General 1.7 Electric H30-200 Turbomec 1.4m

Diamet Thrust er 0.73m 3458K w

Weight 783

Thrust/Weight (Kw/Kg) 4.53

Pressure Ratio 16.6:1

0.53m

429Kw

153

2.8

10.55:1

0.48m

431Kw

122

3.52

6.3:1

0.56m

569Kw

180

3.21

6.7:1

0.46m

440Kw

160

2.75

8:1

a Astozou XVIC2

Engine Selection

Power Loading P/w = 3.5 P req = 3.5 x 8000 x 9.81 = 272.54Kw Pmax = 274.65Kw Single Engine = 274.68/2 = 137.340Kw PT -6A – Twin Shaft Turboprop engine Engine Power Rating 500 – 2000 Shp Strengths : 1. 2. 3. 4. 5.

Dependability Versatility4 Reduced green house gas emission Increased maintenance intervals Ease of Opeation

Turbine – 2PT( 2-stage axial Power turbine) Selected = Pratt & whittney Canda PT6A Pratt & Whittney Canada PT6A

The Pratt & Whitney Canada PTA6 is a family of twin- shaft turboprop engines and the world’s most popular engines in its class. As many as 69 different versions of the PT6A have been produced – with the engine power range from to 2000 shp. In service,PT6A- Powered aircraft are used for many different purposes such as transport of people ; (Both private and commercial), dropping cargo in adverse weather condiions ; military pilot training,surveillance and other special missions ; and supporting various environmental efforts.. The strength pf the PT6A famiy is its dependability and versatility. The PT6A is in service with both the U.S>Air force(MC -12W) and theU.S.Army (C-12) . The MC-12W and C-12 are military special mission aircraft and mainly provide intelligence , surveillance , and reconnaissance (ISR). The C-12 is powered by the two PT6A-6 engine with 1050shp each.

Also , the PT6A is used on the T-6 Texan II military trainer aircraft , which provides undergraduate pilot training for the U.S Air force and U.S.Navy . The T-6 Texan II is powered by a single PT6A – 68 engine woth 1100 shp. The PT6A engine is one of the most popular and proven power plants in its class. Over the years, new aerodynamics technologies have enabled the PT6A engine to gain more power without significantly increasing its size, or weight. Other innovations have reduced greenhouse gas emissions, increased maintenance intervals , and enhanced the ease of operation.

PT6A engines are currently in service with more then 6500 aircraft operators in over 170 countries. Since the PT6A family entered service in the 1960s , more than 41,000 engines have been produced. To date, the PT6A has accumulated 335 million flying hours and is therefore a highly proven and durable engine. Engine Specifications: Manufacturer : Pratt & Whittney Canada(United Technologies) Power: PT6A -60A : 1,050 shp ;PT6A-68 : 1,250 shp Overall Pressure Ratio at maximum power ; Unknown Compressor: Axial flow/centrifugal Compressor Stage: 4-stage axial/1-stage centrifugal Turbine: 2 PT (2 Stage axial power turbine) FADEC: No Length: 1.84m (72.5inch) Diameter: 48cm (19inch) Dry weight: Unknown Platforms: PT6A -60A : C – 12 Huron / MC-12W PT6A -68 : T-6 Texan II Price/cost: PT6A – 60A : $954,869(in 2015) PT6A-68 : $969,000(in 2015) Introduced : 1961 (First PT6 engine) First Run : Unknown First Flight : May 30,1961

Specifications

Multi – Stage axial and single-Stage centrifugal compressor 1. Reverse flow , radial inlet with screen for FOD(Foreign Object Damage ) Protection. 2. Large High Protection PT6A models incorporate 4-Stage axial and 1-stage centrifugal. 3. Small and medium PT6A models incorporate 3- stage axial and 1staeg compressor. Reverse Axial combustor 1. Low emissions, High Stability , Easy starting , Durable

Single-Stage compressor turbine 1. Cooled vanes in some models to maintain high durability. Independent ‘free’ power turbine with shrouded blades. 1. Large and medium PT6A models incorporate 2-Stage axial power turbine. 2. Small PT6A models incorporate 1-Stage axial power turbine. 3. Forward facing output for fast hot section refurbishment. Epicyclic Speed reduction gearbox 1. Enables compact installation. 2. Output speed optimized for highest power and low propeller noise 3. 1,700 to 2,200 rpm output speed Electronic engine controls on multiple PT6A models 1. Other models incorporate various control modules and over – ride features to promote ease of operation and safety of flight.

Drag Estimation

Drag Polar 2 𝐶𝑑 = 𝐶𝑑𝑜 + K𝐾𝐶𝑙

Where K = 1 / πAe e = Oswald efficiency factor Parasite drag CDo = Cfe (Swet / Sref) Where, Cfe = equivalent skin friction drag coefficient ; Swet = Wetted area of the airplane. Swet/Sref = 1.29 K = 0.065 (L/D)max. = 31.8

Cdo = 0.0038 ESTIMATION OF K: Oswald efficiency factor, 1/e = 1/ewing + 1/efuselage + 0.05 ewing = 0.6 ewing(Sweept Angle) = ewing(Sweep Angle) =0 cos (Sweep-5) = 0.563 1/efuselage = 0.1 1/e = 1.776 + 0.1+ 0.05 Therefore, e = 1.9276 e = 0.518 K = 1 / πAe = 1 / 13.815 = 0.0758 (L/D)max = 1 / 2Sweep angle (CD0 K) CD0 = 1 / 4K (L/D)2max = 1 / 4 x 0.0476 x 18.562 = 0.03 Cfe = 0.003 / 1.29 = 0.002 The drag polar is: CD = 0.0165 + 0.0476 CL2 Drag, D = (1/2)

V2SCD

Takeoff: 𝜌 = 1.225 kg/m3 V = 1.15 Vstall = 1.15 (23.57) = 27.22 m/s S =29.10 m2 Drag, D = 0.5 x 1.225 x 27.222 x 29.10𝑐𝑑 Dtakeoff = 99.04N

Landing: 𝜌 = 1.225 kg/m3 V = 1.3 Vstall = 0.771m/s Drag, D = 0.5 x 1.225 x 30.7712 x 7.5x10−3 x29.10 DLanding = 126.29N Cruise: 𝜌= 0.466 (at 10 km altitude)

V = 94.72 m/s S = 29.10 m2 Drag, D = 0.5 x 0.46 x 94.722 x 29.10 x 7.5 × 10−3 Dcruise = 45.62 kN also, (T/W)cruise = 1 / (L/D)cruise T/W = 0.0622 T = 0.0622 x (8000 x 9.81 ) Tcruise = 48.81 kN In straight and level flight, D ~ T 8 PERFROMANCE CHARACTERITCS TAKEOFF 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 of 50ft lift-off speed (VLO = about 1.2 x Vstall) = 28.416M/S SLO = 1.21W / g𝜌SCLmax(T/WD ) = 112.26 CLIMBING: Consider aircraft in a steady unaccelerated climb with vertical climb speed of Vc 𝐿=𝑊 𝑐𝑜𝑠𝛾C 𝑇=𝐷+𝑊 𝑠𝑖𝑛𝛾C

VC = (T - D)Vstall / W Sin 𝛾 = T –D /W R/Cmax = Sin 𝛾 VC = 3.19 *94.7/15 R/Cmax = 20.13 m/s LEVEL TURN: 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 maneuvers. In steady condition: T = D Force balance gives: W = Lcos 𝜃 Fr = mV2 / r = Lsin𝜃 tan𝜃 = V2 / Rg So for given speed and turn radius there is only one correct bank angle for a coordinate (no sideslip) turn. In the turn, n = L/W

W = Turn Rate = V/R = g√𝑛2−1 /v R = 𝑉 2 / g√𝑛2−1

Let 𝜃 = 600 n = 𝐿/𝑊 = 2 R = 94.72/ 9.81√4 − 1 R = 5.57 m 𝜔 = V/R 𝜔 = 0.058 rad/s

GLIDING: The thrust can be assumed to be zero while the aircraft is gliding.

∅= tan-1 [1/ (L/D)] ∅= tan-1 [ 1 / 31.8] ∅ = 1.80 is the glide angle. LANDING PERFORMANCE: APPROACH & LANDING: - Airborne approach at constant glide angle (around 30 ) and at constant speed. - Flare - transitional maneuver with airspeed reduced from about 1.3Vstall down to touch speed.

Ground roll - From touch-down to rest. Ground roll distance (STD): STD = 1.69 W2 / 𝑔𝜌𝑆𝐶𝐿𝑚𝑎𝑥 [D + 𝜇𝑟 (W-L)] Ground roll (STD) = 259.70m 9 Centre Of grvatiy The weight of an airplane changes in the flight due to consumption of fuel and dropping off / release of armament or supplies. Further, the payload and the amount of fuel carried by the airplane may vary from flight to flight. These factors lead to change in the location of the Centre of gravity (c.g.) of the airplane. The shift in the c.g location affects the stability and controllability of the airplane. The weight of entire airplane can be sub divided into empty weight and useful Load. The empty weight can be further subdivided into: (i) structures group (ii) Propulsion group and (iii) Equipment group. The structures group consists of the following components: - Wing

- Horizontal tail /canard - Vertical tail - Fuselage - Landing gear - main and nose/tail wheel - Nacelle, engine pod and air intake The propulsion group consists of the following components: - Engine as installed - Reduction gear - Propeller for piston and turboprop engines - cooling provisions - Engine controls - Fuel system and tanks The equipment group consists of the following items: - Flight controls - Auxiliary power unit (APU) - Instruments - Hydraulic, pneumatic, electrical, armament, air conditioning, anti-icing - Avionics - Furnishings in passenger airplanes The useful load consists of: (i) Crew (ii) Fuel - usable and trapped (iii) Oil (iv) Payload - passengers, cargo and baggage in transport airplane; ammunition, expendable weapons and other items in military airplanes.

The aim of estimating the weights of individual components and their c.g. is to obtain the location of the c.g. of the airplane. Then, the shift in the airplane c.g. is examined under various conditions. At this stage of preliminary design, the weights of individual components are estimated using simpler method like using the table above. The gross weight of the airplane estimated is 80000 kg The weights and c.g. locations of various components are estimated below:

Wing: S = 29.10 m2 b = 15.33 m bsemi = (15.53/2) – Wf/2) = 6.865 m cR = 3.04m cT = 0.76 m Fuselage width = 1.6 m mac 2.128 m Location of L.E of mac from L.E of wing = 0.85 m S (exposed) wing = s – (Cr x Wf) = 24.236 m2 From the table, the weight of the wing is, Wwing = 24.236 x 12 = 290.832 kg Wwing / W0 = 3.94% Hence, the location of the c.g. of wing from the leading edge of the root chord is, 0.85 + 0.4 x 2.128 = 1.70 m