Hafif siklet bir uçağın kanat yüklerinin analizi
The Analysis of the wing loads of a light airplane
- Tez No: 14342
- Danışmanlar: PROF.DR. AHMET NURİ YÜKSEL
- Tez Türü: Yüksek Lisans
- Konular: Uçak Mühendisliği, Aircraft Engineering
- Anahtar Kelimeler: Belirtilmemiş.
- Yıl: 1991
- Dil: Türkçe
- Üniversite: İstanbul Teknik Üniversitesi
- Enstitü: Fen Bilimleri Enstitüsü
- Ana Bilim Dalı: Belirtilmemiş.
- Bilim Dalı: Belirtilmemiş.
- Sayfa Sayısı: 91
Özet
ÖZET Bu yüksek lisans çalışmasında, hafif, pervaneli bir uçağın kanadına çeşitli uçuş hallerinde gelen aerodinamik yükler ve bu yüklerin dağılımlarının tesbiti yapılmıştır. Hesaplarda, pazmany'nin hafif, pervaneli uçaklar için ge liştirmiş olduğu dizayn metodu takip edilmiştir. Dış boyutlar ve ağırlık, kalemleri belirlenmiş olan uçağın, hizmet ömrü boyunca çeşitli uçuş hallerinde maruz kalabileceği sınır yükleri ihtiva eden V-n manevra zarfı çizilmiştir. V-n manevra zarfının çeşitli noktalarına te kabül eden yükleme halleri göz önüne alınarak, nervür, ka natçık gibi kanat elemanları üzerindeki yük dağılımı ve zorlamalar incelenmiştir. Hesaplarda kompleks denklemler yerine, dizaynı yapı lan uçağın özelliklerine yakın performanslara sahip uçak ların istatistiki bilgilerinden ve havacılık nizamnamele rinden faydanılmıştır. vıı
Özet (Çeviri)
THE ANALYSIS OF THE WING LOADS OF A LIGHT AIRPLANE SUMMARY The most difficult problem in designing an airplane is to find a guide that shows step by step the sequence in which the different problems have to be approached and solved. In this thesis, the analysis of the load on the wing of a lignt airplane has been done by using pazmany's Design Method. In every phase of the calculations, modern aerodynamic and structural data were applied. No higher mathematics were used only the four basic algebra opera tions, along with many graphs and diyagrams. 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 Definition”. The dimensions and the weights of the airlane have had to be determined before being started to the aerodynamic calculations. The general character istics of the airplane are as follows? -Specific Use : Trainer, two place -Fuselage : Type of construction; Semimo- nocogue, Basic structural material; aluminum alloy, Skin Material; aluminum alloy -Wing : Type of construction; cantilever Shape; rectangular, Location; low, airfoil; NACA 632615, Aspect ratio; 7.592, Area; 105.4 Sq.ft; High lift device; Plaxn flaps -Power Plant : Type; Opposed air-cooled 90-95 Hp, tank location; wing tips. -Landing Gear : Type; Tricycle, fixed -Weight : Empty; 674 lb., Gross; 1200 lb -Desired Performance: Range; 400 miles, service Ceiling; 14000 ft The airplane to be designed is in the utility category. The airplanes in this category are intended 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, spins, stalls (except whip stalls) viiilazy eights, and chandelles. Relatively few components are affected by the higher load factor due to the acrobatic manuevers, while a great part of the structure is dimensioned by minumum practical gauges. Thus, with a small weight penalty, used to“beef -up”some critical parts such as the wing spar, the airplane can be designed to meet the requirements imposed by the“Acrobatic”category instead of the“Utility”. The additional strength also covers future possibilities of increasing the engine powers. The semi-monocoque construction has been selected, as the type of construction of the fuselage. The semi- monocoque 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 stress 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 aluminum alloy has been used for all the structural parts The uniformity in quality is better than plywood or spruce. With a metallic skin, the expensive fabric finishing is eliminated, not only as an inital investment but also at the periodical overhaul. Some aerodynamic and manufacturing advantages can be obtained, when the aluminum alloy is used for all the structural parts. A relatively thick skin on the leading edge allows the use 6f countersunk rivets, and also reduces the wrinkles. Both conditions are very desirable to obtain some laminarity in the airflow, at least up to the maximum thickness of the airfoil. This is an ideal condition diffucult to reach but if obtained, will result in a general improvement in the performance. The construc tion of an all-rnetal small airplane does not require a big investment in tooling or machinery. Double curved parts much and can be avoided, but single curvature instead of flat panels is panels is desired to reduce oil canning and improve stiffness. The most dangerous parts of every flight probably are the take-off, landing and flying the pattern. Visibility in a turn !s greatly desired during these maneuvers. In a high wing aircraft the visibility during these critical moments is reduced mostly toward the ixinside of the turn. These considerations alone will decide the choice between high wing and low wing, but there are many others that can be enumerated. From aerodynamic viewpoints, the fuselage cross- sectional area of a low wing airplane could be made smaller than of a high wing; the ocupants could be seated over the wing. The main landing gear struts in a low wing airplane can be very short if they are attached to the wing spar. This represents a minimum weight and parasite drag. The wing spar does not need to be strengthened to take the landing gear loads because the air loads are critical. In view of these considerations, it was decided to use low wing for the airplane to be designed. In view of structural and aerodynamic considerations the airfoil selection has been done. The weight of a beam which has to carry a certain bending moment is inversely proportional to the square of the depth of the beam. Therefore, the thickest air foil will house the deepest beam, and this in turn will result in the lightest construction. The lift force in an airfoil is approximately located at 30 per cent of the cord. If the maximum thickness is also at 30 per cent, obviusly this will be the ideal location for the main spar. As the high lift device, flaps have been used. From a constructional view point, it is easier to make the wing with flaps than without them. An advantage of wing flaps is the increase in drag, resulting in a steeper glidepath. The newest“laminar type”airfoils have been used in high performance airplanes and gliders all over the world. The characteristics of low drag airfoils are no worse than those of conventional airfoils and, if sufficient care is taken with the surface condition, definitive advantages are associated with their use. 632615 airfoil has been chosen by comparing to others. At higher section lift coefficient, the 63 2615 has low drag. Another reason in the selection is that the 632615 got high section lift coefficient; therefore, landing speed without flap is slightly reduced. The 63 615 has got a great moment coefficient which results in a propotionally large trim drag. This is the main disadven- tege using this airfoil.In chapter 2, the aerodynamic loads which act on the airplane in the different flight condition have been calculated and their distributions have been drawn. The V-n diagram which represents the limit fligh load factors and limit flight conditions was drawn. Symetrical flight conditions have been investigated for two different conditions: flap retracted and flap extended. In flap retracted condition, all loads were calculated for one side of the airplane. Normal loads and torsional moments distribution have been assigned. In flap extended conditions, normal loads and torsional moments distribution have been assigned, too. In unsymetrical flight conditions, air loads act on the wing unsymetrical ly. The wing and wing carry-through structure have been designed for 100 % of condition point“A”in V-n diagram on one side of the plane of syraetry and 60 % on the other opposite side, normal loads and torsional moments spanwise distribution have been assigned. In the other unsymmetrical flight conditions, the wing and wing carry-through structure have been designed for the loads resulting from a combination of 75 % of positive maneuver wing load on both sides of the plane of symmetry combined with the maximum wing torsion resulting from aileron displacement. The determination of critical aileron displacement has been done. The loads on the ribs have been investigated for different conditions. Chordwise load distrubutions have been assigned. In flap retracted condition, the ribs have been investigated for loads at points A-D-E and G in the V-n diagram. The pressure distribution on the airfoil has been drawn. The areas under the pressure distribution diagrams represent the total load on the rib. The same calculations have been done for the flap extended condi tions and the aileron deflected conditions. In the flap extended condition, the loads on the ribs have been investigated for two different conditions; VFmin =90 mph, nfiap= 0 and vFmin= 9° mPn nflap= 2.2. In the aileron deflected condi tion, the ribs have been investigated for loads at points“A”in the V-n diagram. In the analysis of the pressure distribution at these conditions, the theory of wing sec tion has been used. The theory of thin wing section shows that the load xidistribution of a thin section may be considered to consist of a basic distribution at the ideal angle of attack and an additional distribution proportional to the angle of attack. The first load distribution is a function only of the shape of the thin wing section or of the mean-line geometry. The second load distribu tion results from changing the angle of attack. The velocity distribution about the wing section is considered to be composed of theree seperate and independent components as follows - The distribution corresponding to the velecity distribution over the basic thickness form at zero angle of attack. - The distribution corresponding to the load distribu tion of the mean line at its ideal angle of attack. - The distribution corresponding to the additional load distribution associated with angle of attack. The local load at any chordwise position is caused by a difference of velocity between the upper and the lower surfaces. It is assumed the velocity increment on one surface is equal to the velocity decrement on the other surface. XII
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