Hafif siklet bir uçağın genel aerodinamik dizaynı
General aerodynamic design of a light airplane
Yazar: CANER ŞENTÜRK
Danışman: PROF. DR. AHMET NURİ YÜKSEL
İstanbul Teknik Üniversitesi / Fen Bilimleri Enstitüsü
Derece: Yüksek Lisans, 1991
Konu: Uçak Mühendisliği = Aircraft Engineering
Dizin: Aerodinamik yük = Aerodynamic load ; Tasarım = Design ; Uçaklar = Airplanes

Özet:
ÖZET Bu çalışmada iki kişilik, piston-pervane motorlu, be lirli performans ve şartları yerine getirecek hafif sik let bir uçağın dizaynı yapılmıştır. Comple uçağın ve ele manlarının boyut ve alırlıkları tesbit edilerek, çeşitli elemanlar üzerindeki aerodinamik yüklerin hesaplanması ve yapısal mukavemet hesaplarına girilmesi için bir zemin teşkil edilmiştir. Birinci bülümde dizaynı yapılacak olan uçağın misyo nu ve diğer genel karakteristikleri belirlenmiş; başlan gıçta arzu edilen performansları tesbit edilmiştir. İkinci bölümde dış boyutların tesbit edilmesi ile uğraşılmış, kanat alanının tesbiti, açıklık oranının tes- biti, kuyruk yüzeylerinin boyutlandırılması dizaynı yapı lan uçağınkine benzer özelliklere sahip uçakların istatis tiki bilgilerinden de yararlanılarak yapılmıştır. üçüncü bölümde ise komple uçağın ve elemanlarının ağırlıkları tesbit edilmiştir. Bulunan ağırlıklardan fay dalanılarak, uçağın ağırlık merkezi bulunmuştur. Ağırlık merkezinin en önde ve en arkada olabileceği konumlar, çe şitli şartlar göz önüne alınarak bulunmuştur. Bu tez kapsamında yapılan hesaplardan yararlanılarak çeşitli uçuş şartlarında uçak elemanlarında oluşabilecek aerodinamik zorlamalar ve bu zorlamaların elemanlar boyun ca dağılımları incelenebilir. Böylece yapısal mukavemet hesaplarına geçilebilir. Bu tezin kapsamı, üstte sözü edilen çalışmalara geçilebilmesi için gerekli bütün hesap ları içermektedir. vıı

Abstract:
GENERAL AERODYNAMİC DESİGN OF A LİGHT AİRPLANE SUMMARY This thesis has been devoted to light airplane design according to the method used by Pazmany. Dimensions and weights of the entire airplane and its parts have been assigned by utilizing some well-know aircrafts with similar characteristics to the proposed design. The first step in the design of an airplane consists of defining the characteristics of the airplane and its use. In aeronautical engineering this is called 'mission defini tion'. The airplane to be designed is in the utility category. Airplanes in this category are intented for normal operations and limited acrobatic maneuvers. These airplanes are not suited for use in snap or inverted maneuvers. 'Limited acrobatic maneuvers' is interpreted to include steep turns, spin, stalls (except whip stalls), lazy eights and chandelles. In the first chapter, the general characteristics and the desired performance of the airplane as well as its category have been assigned as follow: - Specific use - Appearance - Fuselage - Wing - Empannage - Power plant - Cockpit - Landing gear Utility, sport, trainer. Functional Semimonocoque, basic structural and skin material is aluminum alloy. Cantilever, detachable, carry-through spar, rectangular shape, low-wing, high lift devices are flaps, spar and skin material is aluminum alloy. Airfoil : NACA 632615, Area : 100 sq. ft Aspect ratio : 7 Wing loading : 10 lbs/sq.ft. Fin-Rudder configuration : conventional stab-Elevator configuration : conven tional or slab tail. Type : opposed, air-colled, 85-100 Hp. Tank location : Wing tips Control type : stick Instruments : Nominal Canopy type : Sliding bubble Tricycle, fixed. vniWeight : Empty : 700 lbs. Gross : 1200 lbs, Desired performance : Stalling speed : 50 mph. Cruising speed : nomph. Maximun speed : l27mph. Range : 400 miles Service ceiling: 14000 ft. As the type of construction of the fuselage, the semimonocoque construction has been selected. This kind of construction is widely used for airplanes of this size and characteristics. The fuselage built around four longerons does not require complicated assembly jigs, and also, it is an efficient structure to transmit the loads. The analysis is simple; each side of the fuselage can be considered a beam, while the box formed by the four sides carries the torsional and shear loads. As the basic structural and skin material, the alumi num alloy was used. The uniformity in quality is better than plywood or spruce. A relativly thick skin on the leading edge allows the use of countersunk rivets, and also reduces the wrinkles. Both conditions are very desirable to obtain some laminarity in the flow, at last up to the maximum thickness of the airfoil. This is an ideal condition difficult to reach, but if obtained, will result in a gene ral improvement in the performance. With a 15% chord-thickness ratio airfoil, it is possible to build a cantilever wing with a weight comparable to a strut braced. It is not only the weight of the basic members that must be considered in the comparison, but also the extra fittings, bolts, turnbuckles etc., along with the added loss in aerodynamic efficiency due to the additional parasite and interference drag. So the cantilever wing has been selected for the airplane to be designed. As the location of the wing, low-wing has been selected. The most dangerous parts of every flight probably are take off, landing and flying the pattern. Visibility in a turn is greatly desired during these maneuvers. In a high wing aircaft, the visibility during these critical moments is reduced mostly toward the inside of the turn. These consi derations alone will decide the choise between high wing and low wing, but there are many others that can be enumerated. Aircraft accident investigations and simple reasoning indicate that the more structure between the occupants and the ground, in case of crash, the higher are dissipated in ixWing before starting with the passengers. From aerodynamic viewpoints, the fuselage cross-section area of a low wing airplane could be made smaller than of a high wing; the accupants could be seated over the wing. In chapter 2, dimensions of the entire airplane as well as of its parts have been assigned by utilizing the characteristics of some well know aircrafts. Airfoil selec tion has been made according to structural considerations and aerodynamic considerations. Wing area of the aircraft has been determined in a subchapter by utilizing landing and stalling speeds of several well known airplanes and the criterion that the wing area is a function of the landing speed. In this subchapter, the lift increment due to plain flaps has been calculated. The flap does not affect the whole wing, therefore, the lift increment calculated has been reduced accordingly to the flap span. Aspect ratio has been determined as 7 at the beginning of chapter 2. For light conventional airplanes a good compromise aspect ratio is about 7. A smaller aspect ratio will result in excessive induced Drag, penalizing mostly the climb and ceiling. On the other hand aspect ratios over 7 result in excessively reduced wing chords. From struc tural view point, the advantage of a small aspect ratio are double; first, because a larger chord will provide a propor tionally larger depth for the wing spars, second, a shorter wing represents smaller bending moments, which in turn requires lighter spars. Wing aerodynamic characteristics, power plant selec tion, empennage dimensioning and landing gear design have been made at the end of the chapter 2. The tail surfaces of an airplane to meet two basic requirements : stability and control. The design of tail surfaces, the determina tion of their size, position, angle of incidence is not an easy problem. The effects of factors such as slipstream, downwash, interference, center of gravity position, Reynolds Number complicate the problem of empennage dimensioning. In chapter 3, the weight estimation was made according to the statistics obtained from some existent airplanes. Determination of useful load, structural weight estimation have been made by utilizing some criterions. The wing weight, the fuselage weight, the horizontal tail and ver tical tail weight and the weight of landing gear, controls, major assemblies, spinner, engine cowling, fuel tanks, fuel lines, equipment, windshield and canopy, engine controls etc.. have been calculated.At the end chapter 3, in the subchapter " Airplane balance, the center of gravity position was calculated by calculating the moment of each components with respect to reference lines. The maximum aft. C.G. position and the most forward C.G. position were calculated, too. The calculation of the center of gravity position is one of the most important steps in the aircraft design. The aircraft designer should permanently keep track of the weight and balance of the airplane. This is so important that every aircraft factory has a "Weights Group" in its Engineering Department. "Weight Engineers" continuously check the weight of each component during the design. Each drawing should be signed by them before release. Sometimes the weight of parts or assemblies result higher than estimated, then a decision should be made to either redesign the part and tryto make it lighter even if it results more complicated, or to leave the design as it is and take weight penalty. Parts located away from the C.G. are more critical than parts close to the C.G. In the calculation of the C.G. position, the following procedure is applied: - Draw a side view of the airplane at a convenient scale (1/10 is adequate). Indicate the center gravity of each component by a small circle. It requires some practice to estimate by "eye ball" the position of the center of gravity of some components. As a general guide the C.G. of wings lies at 40 % of Mean Aerodynamic chord. The center of gravity of vertical and Horizontal Tails can be ocated 50 % of the respective mean chord. The center of gravity of fuselage could be estimated at 40 % of the fuselage length measured between the firewall and the tail cone. - Enter the weight of each component in the weight column - Draw a vertical reference line at the spinner vertice and a horizontal reference line at ground level. - Measure the horizontal and vertical distance of each component's center of gravity from the reference lines. Enter these values in the column of horizontal arm and vertical arm. - Multiply the weight of each component by its horizontal distance and enter the result in the horizontal moment column. XIMultiply the weight of each component by its vertical distance and enter the result in the vertical moment column. Add weight column to obtain the sum of weights. Add horizontal moment column to obtain the sum of horizontal moments. Add vertical moment column to obtain the sum of vertical moments. Divide the sum of horizontal moments by the sum of weights to obtain the horizontal location of the center of gravity, Divide the sun of vertical moments by the sum of weights to obtain the vertical location of the center of gravity For a new study, as a continue of this thesis aerody namic loads on different component of the aircraft may be analysed. For this study first the V-n diagram which determines limit flight load factors and limit flight conditions must be drawn. Then aerodynamic loads, torsional moments spanwise distribution, chordwise loads-spanwise distribution, normal loads spanwise distribution at any flight condition (Symmetrical ami unsymwetrical) may be calculated for each point of the V-n. diagram. Flap extended and flap retracted conditions, unsymmentrical wing loads, loads due to down aileron and their distribution may be analysed. Rib loads and chordwise pressure distribution may be analysed for flap deflected and aileron deflected conditions. For continue, loads on the fuselage, crash loads, inertia loads on aft. Fuselage, bubble canqpy loads, control loads, control surface loads, ground loads may be analysed, too. xn