Hafif siklet bir uçağın yer yüklerine göre boyutlandırılması
Dimensioning of a light airplane depending on ground loads
- Tez No: 14343
- 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: 33
Özet
ÖZET Bu tez çalışmasında iki kişilik, İstanbul-Antalya arasındaki mesafeyi (400 mil) katedebilen 4 silindirli piston-pervaneli bir motora sahip, hafif siklet ev yapı mı bir uçağın gövde, kanopi, kontrol elemanları ve iniş takımlarına gelebilecek zorlamalar ile bunların sözkonu- su elemanlar boyunca dağılımları incelenmiştir. Bu yapı lirken NACA teknik raporları ile sivil havacılık nizamna mesinden faydanılmıştır. Başlangıçta motor sehpası ve destekleyici yapısı, motor dönme momenti etkisi ile birlikte belirli temel uçuş şartları için dizayn edilmiştir. Bundan sonra göv de üzerindeki elemanlara gelebilecek limit yanal yükler tespit edilmiştir. Uçağın yere çarpması sırasında pilot ve yolcuya gelebilecek nihai yükler bulunmuş, gövdenin arka kısmında münferit ve yayılı yüklerden oluşan atalet yükleri maksimum zorlanma hali için tespit edilmiştir. Gövde ile ilgili bu çalışmadan sonra, pilot mahali kapağı üzerindeki maksimum zorlanma hali için basınç dağılımı elde edilmiştir. Kontrol elemanlarına gelebilecek aero dinamik zorlamalar için sivil havacılık nizamnamesi FAR* dan istifade edilerek, sözkonusu elemanlar boyunca dağı lımları incelenmiştir. Son olarak, çeşitli iniş hallerin deki iniş takımlarına gelebilecek zorlamalar tespit edil miştir. Bu tez çalışması, belirli performans ve şartları ye rine getiren bir hafif siklet uçağın gövde ve diğer ana elemanlarına gelebilecek zorlamalar ve bu elemanlar boyun ca dağılımlarının incelenmesi ile sonuçlandırılmıştır. Bu elemanlara ait mukavemet hesap ve incelemelerine giril memiştir. vı
Özet (Çeviri)
DIMENSIONING OF A LIGHT AIRPLANE DEPENDING ON GROUND LOADS SUMMARY \ In this thesis, loads on fuselage, control loads, and ground loads of two seat light home-built airplane which can fly between İstanbul and Antalya (400 limes) have been examined. Air loads at different points of V-n diagram, and their distrubition on main parts of the airplane have been determined. It is not easy to predict all flying qualities of an airplane using calculations. This applies either to the simplest“home-built”or to a sophisticated supersonic fighter, mostly when they are of unconventional type, like the delta wing. There are too many variables in the game, especially when stability and control is the subject. The aerodynamicist needs the help of wind tunnel testing and simulators to rectify or ratify his calculations. The airplane to be examined is a conventional airplane for present standarts; neverheless, modern aerodynamic and structural data were applied in every phase of the study. The important part of an airplane is the fuselage. Engine mounts and their supporting structure have been designed for engine torque effects combined with certain basic flight conditions. By determining lateral limit load factor, the lateral loads on the parts on the fuselage have been determined. Crash loads have been determined by utilizing FAR. The ultimate accelerations to which occupants are assumed to be subjected are as follows: upward 3.0 g, forward 9.0 g, and sideward 1.5.g. The loads on the cano.py are maximum when the canopy is closed, when“q”is maximum, and at high angle of attack. Therefore point“D”in the V-n diagram has been investigated. And pressure distrubition on the buble canopy has been determined. Pilot efforst for controls have been given. For control surface loads, direction of loa ding, magnitude viiof loading, and chordwise distrubition on each component have been listed by utilizing FAR. After that stabilator spanwise load distrubition has been obtained. Therefore the following assumptions have been made: a) Rectangular air load distrubition b) No inertia relief. After that vertical tail spanwise loads distrubition has been obtained. So the following assumptions have been made: a) Air load distribition proportional to chord. b) No inertia relief. In the last parts of the thesis, the ground loads at different landing cases of the airplane have been determined. Landing cases to be examined are two types for the tricycle landing gear: 1) Level landing a) Nose wheel and main wheels contacting the ground simultaneously. b) Main wheels contacting the ground, nose wheel just clear the ground. 2) Tail down landing: Airplane in stalling attiude In addition, for the nose wheel aft, forward, and side loads have been determined. This study is a part of the airhane design. So, before this study, for the airplane to be designed, there are some steps to be done. The first step in the preliminary design of an airplane consists of defining the characteristics of the airplane and its use. In aeronautical engineering this is called“Mission Definition”. So, at the beginning, the general characteristics and desired performance of the airplane to be designed and its category are determined. The airplane to be examined is in utility category. 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 VlllAcrobatic Maneuvers”is interpreted to include steep turns, spins, stalls (except whip stalls), lazy eights, and chandelles. General characteristics of the airplane examined are as follows: Specific use : Utility Fuselage : Semi-monoc.oque of aliminum alloy structure. Wing : Cantilever, detachable, low Wing monoplane of aliminum alloy rectangular structure. As the high lift devices, flaps are used. Empennage : Conventional Power Plant : Opposed, air-cooled, 85-100 HP. Fuel tanks at wing tips. Desired Performance 50 mph 110 mph 127 mph 400 limes 14000 ft Stalling speed Cruising speed Max. Speed Range Service ceiling The semi-monocoque construction is widely used for airplane 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 by the four sides carries the torsional and shear loads.' As the basic structural and skin material, the aliminum alloy is used for all structural partes. The uniformity in quality is better than plywood or spruce. A relatively 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 lamina rity 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 impro vement 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 consi dered in the comparison, but also the extra fittings, bolts, turnbuckles etc. along with the added loss in aerodynamicefficiency due to the additional parasite and interference drag. From theoretical considerations and from pressure distrubition tests, it can be dem nstrated that the ideal wing form is the elliptical because it has the smallest induced drag. But using the same theory and tests, it was found that a rectangular wing of aspect ratio 6 has only about 5 per cent greater induced drag then than of ixan elliptical. Between these two wing plan forms, there is the tapered, which has roughly one per cent more induced drag then the elliptical. Both elliptical and tapered wings allow a lighter spar construction, but these advantages are of small importance when compared with the better stalling characteristics and simplified construction of a rectangular wing. The most dangeraus parts of every flight probably are the take off, landing and flying the pattern. Visibility in a turn is greatly desired during these maneuvers. In a high wing aircraft the visibility during these critical moments is reduced mostly toward the inside of the turn. These considera tion alone will decide the choice between high wing and low wing. After the determination of the general characteris tics the external dimensions of the airplane to be exa mined. Therefore at the beginning aerodynamics of the airplane must be determined. The first step of the external dimensions is the determination of the wing area. The wing area is a function of the landing speed. The second important step is the determination of the aspect ratio of wing. By utilizing aerodynamic equa tion, aspect ratio including wing tip fuel tanks is deter mined, depending on this, span and chord of wing can be determined easily. After the determination of wing characteristics empennage is designed. The tail surfaces of an airplane have to meet two basic reguirements: Stability and cont rol. The design of tail surfaces, the determination of their size, position, angle of incidence is not an easy problem. The effects of factors such as sliptream, downwash, interference, C.G. position, Reynolds number, and many others, complicate not only the problem, but also obscure the basic concepts for the amateur designer. It is always convenient to use non-dimensional coefficients for comperative purposes. Therefore, hori- zantel and vertical tail volume coefficients can be defined. In this manner, both tail surfaces are related to the wing chord which has a great importance on the airplane langitudinal stability and control, while the vertical tail is related to the wing span which has a great signi ficance on directional stability and control. For this purpose, the volume coefficients are selected by utilizing some well-known airplanes. So, horizantal and vertical tail areas and their chord and span can be determined.After the determination of the external dimension, the suitable power plant and landing gear of the airplane are selected. The choice of a tricyle landing gear is justified by the following reasons: 1- A Leveled position is more compfortable when entering or leaving the cockpit. 2- There is an improved forward vision from the cabin during ground runs. 3- The tricycle landing gear eliminates the ground loop; it gives better ground stability and permits full braking which in turn reduces the landing distance. After these calculations, the weight of the airplane to be designed, and weights of its components are estima ted. For this purpose, a table that prests data on two seat light airplanes is prepared. In this table, it can be seen that the relation useful load and gross weight varies very much depending on the airplane. It is diffucult to establish a trend because many factors are involved such as load factor, engine weight, aspect ratio, type of construction, etc. Therefore, instead of calcula ting an average value based on the whole table, it will be arvisable to select two or theree airplanes with similar characteristics to the proposed design and calculate the average value based on these few samples. So, for the airplane to be designed, gross weight can be estimated by this method. By estimating gross weight, the other weights of structure majir assemblies are estimated with the graphs and formulas. For some components of the airpane to be designed, the solution to find their weight is to prepare simple drawings and determine the weight analitically. One of the most important steps in the aiplane design is the location of the center of gravity. The aircraft designer should permanetly keep track of the weight and balance of the aiplane. The C.G. pobition is calculated simply by calculating the moments of each component with respect to reference lines. The maximum aft and forward C.G. positions are important. Because maximum aft. C.G. pasition is the most critical for stability. And maximum forward C.G. condition is critical for elevator dimensioning. And xifinally the C.G. for airplane Gross Weight is calculated. Obviously, it must fall between the two extremes explained before. After that, air loads on the airplane to be designed at different points of V-n diagram and their distrubition on main parts of the aiplane are determined. In general, there are two different flight canditions for an airplane: Symmetrical and unsymmetrical. In the symmetrical flight conditions, flap retracted and flap extended conditions are examined. For the flap retracted conditions, these flap retracted condition, these assump tion are made: The spanwise lift distrubition for any typical wing to consist of two parts. One part, called“the basic distrubition”, is the distrubition that depends principally on the twist of the wing and occurs when the total lift of the wing is zero. The second part of the lift distrubition, called the“additional distrubition”, is the lift due to change of the wing angle of attack. For the flap ewtended condition, in order to simplify the analysis conservative assumption are made: Rectangular air load distrubition and no inertia relief. So, for these conditions, normal loads and torsional moments spanwise distrubitions on wing are determined. In the unsymmetrical flight condition, there are two different types to be examined: 1- Unsymmetrical wing loads: The wing and wing carry-through structure are designed for 100 % of condi tion point“A”in V-n diagram on one side of the plane of symmetry and 60 % on the opposite side. 2- Symmetrical wing load + aileron displacement: The wing and wing carry-through structure are 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 resul ting from aileron displacement. For these two conditions air loads and their distrubition on each side of wing are determined. The other important component of the wing is the rib. The air loads on the rib at different points of V-n diagram and their distrubition are dekermined. xiiChordwise loads distrubitions with flap and aileron deflections are examined. So, these assumptions can be made: The total load distrubition is the sum of the incremantal basic distrubition, the incremental additional load distrubition, and the load distrubition of unf lapped section at the same angle of attack. So, by these assump tions air loads on the wing and their distrubition are determined. As mentioned before at the beginning enternal dimen sions, weights of the entire airplane and its parts, and air loads at different points of V-n diagram and their distrusition on the wing of the airplane are to be deter mined. So, in this study air loads and their distrubition on the other important parts have been determined. After this study the selection of material and strength calcula tions of these parts must be determined. Thus, entire design of the airplane will be finished. Xlll
Benzer Tezler
- Hafif siklet bir uçağın genel aerodinamik dizaynı
General aerodynamic design of a light airplane
CANER ŞENTÜRK
- Hafif siklet bir uçağın kanat yüklerinin analizi
The Analysis of the wing loads of a light airplane
OĞUZ TUBAY
- Hafif siklet bir uçağın kaplama perçin ve rib hesabı
The Calculations of skin rivets and rib design of light airplane
MEHMET SAİT SAFFET BAYSAL
- Hafif siklet bir uçağın kesme-eğilme ve burulmaya göre kanat hesabı
According to the shear loeds, bending and light airplane design torsional moments
ÖZDEN AKTAN
- A uml profile for role-based access control
Rol-tabanlı erişim denetimi için bir uml profili
ÇAĞDAŞ CİRİT
Yüksek Lisans
İngilizce
2009
Bilgisayar Mühendisliği Bilimleri-Bilgisayar ve Kontrolİstanbul Teknik Üniversitesiİleri Teknolojiler Ana Bilim Dalı
YRD. DOÇ. DR. FEZA BUZLUCA
