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İnsansız hava aracı tasarımı

Design of an unmanned air vehicle

  1. Tez No: 46354
  2. Yazar: GÜRHAN ÇETİN
  3. Danışmanlar: DOÇ.DR. SÜLEYMAN TOLUN
  4. Tez Türü: Yüksek Lisans
  5. Konular: Uçak Mühendisliği, Aircraft Engineering
  6. Anahtar Kelimeler: Belirtilmemiş.
  7. Yıl: 1995
  8. Dil: Türkçe
  9. Üniversite: İstanbul Teknik Üniversitesi
  10. Enstitü: Fen Bilimleri Enstitüsü
  11. Ana Bilim Dalı: Belirtilmemiş.
  12. Bilim Dalı: Belirtilmemiş.
  13. Sayfa Sayısı: 89

Özet

ÖZET Bu çalışmada görev gereksinimleri ve performans değerleri verilen bir insansız hava aracının ön tasarımı yapılmıştır. İzlenen yöntemde öncelikle tasanm ihtiyaçlarına bağlı olarak aracın başlangıç boyutlandırması yapılmaktadır. Daha sonra başlangıç boyutlandırmadan elde edilen değerler kullanılarak boyutlandırma iterasyonu gerçekleştirilmekte ve boyutlandırılan uçağın performansı ve stabilitesi hesaplan maktadır. Değişen parametreler için tasarımın kolaylıkla tekrarlanabilmesi amacı ile kavramsal uçak tasarımını bilgisayar ortamında yapabilecek bir program yazılmıştır. Yazılan program ve sayısal optimizasyon rutinleri kullanılarak aracın farklı amaçlar için optimizasyonunun gerçekleştirilebilmesi sağlanmıştır. Sayısal optimizasyon ile kağıt üzerinde yapılan tasanm çalışmalarında uygulanabilenden çok daha fazla sayıda parametre değiştirilerek uçağın optimum konfıgürasyonu bulunabilmektedir. BÖLÜM 8' de 8 adet tasanm parametresi kullanılarak gerçekleştirilen bir optimizasyonun sonuçlan yer almaktadır.

Özet (Çeviri)

SUMMARY DESIGN OF AN UNMANNED AIR VEHICLE In this thesis, conceptual design of an unmanned air vehicle was done according to given mission profile and performance requirements. Definition of the vehicle mission profile is briefly carrying the payload (including camera, computer, receiver and transmitter) to reconnaissance field and after the surveillance is performed turning back to the takeoff point. Mission definition and performance requirements are given below, Takeoff altitude Takeoff ground run Max. rate of climb Cruise altitude Flight distance to the target Flight speed to the target Surveillance period Flight speed at surveillance Max. level speed Stall speed 3500 ft 400 ft 800 ft/min. 11500ft 40 km 65 knots 5.5 hours 55 knots 70 knots 35 knots UAV's is varying widely in design objects, fight principles, general configuration, range and endurance, takeoff and recovery types. Selected configuration for the vehicle to be designed is shown in Figure 1. -J fl u TV u Figure 1. General configuration of the vehicle. XIThe design process which is applied can be shown in Figure 2. In order to repeat the design easily with changed variables, a software was written to perform the complete conceptual design process in computer environment. MISSION PROFILE AND PERFORMANCE PARAMETERS INITIAL SIZING '-SIZING ITERATION1PERFORMANCE CALCULATIONS''STABILITY AND CONTROL1'OPTIMIZATION ROUTINES1'RESULTSFigure 2. Design process INITIAL SIZING In initial sizing, selections of wing loading (W/S) and power loading (W/hp ) are performed. Wing loading selection is based on stall speed requirements. For power loading, the required power loading is calculated for each mission segment and then the minimum one is selected. Then mission segment weight fractions and fuel fraction are calculated using the selected (W/S) and (W/hp). Using a statistical equation along with an iterative solution; takeoff weight, fuel weight and empty weight are calculated. All aerodynamic and performance values used in initial sizing are based on statistical data and previous similar designs. xnSIZING ITERATION For calculating dimensions and weight of aircraft, engine horsepower and total fuel weight, it is necessary to perform a sizing iteration. Inputs to sizing iteration are the values obtained from initial sizing. Outputs of it includes mean dimensions of the aircraft, component weights, total empty weight, engine horsepower, mission segment weight fractions and fuel weight. The mean steps performed in sizing iteration includes:. calculation of dimensions,. aerodynamic calculations,. determination of component weights, empty weight and center of gravity position,. propulsion system calculation,. mission segment weight fractions The above items are calculated at every iteration step and at the end of each iteration takeoff weight is calculated by summing component weights, fuel weight and payload. Calculated takeoff weight is compared with the value at previous iteration step to check convergence. Calculating of aircraft dimensions is based on takeoff weight, wing loading and power loading ( these values are obtained from initial sizing or previous iteration step). Selected configuration parameters (such as aspect ratios of wing and tails, horizontal and vertical tail volume coefficients) and some statistical equations are used to determine the aircraft component dimensions. In aerodynamic calculations wing lift curve slope, wing Oswald efficiency factor and parasite drag of the complete aircraft are found. For parasite drag calculation, component buildup method is used. The component buildup method estimates the subsonic parasite drag of each component of the aircraft using a calculated flat-plate skin-friction drag coefficient and a component form factor that estimates the pressure drag due to viscous separation. Parasite drag contributions of the engine and landing gears are also included in the calculations. The estimation of aircraft empty weight is a critical part of conceptual design process. For component weights, detailed statistical equations are used. These equations compute the component weights depending on components dimensions, takeoff weight and ultimate load factor. Once the component weights are calculated, the center of gravity position of the aircraft is determined based on the locations of the components, payload and fuel tank. In propulsion system calculations, variation of engine thrust is found by using engine horsepower and propeller efficiency. The variation is found at sea level condition and scaled for different altitudes using an equation that estimates the engine power loss for different altitudes. xmMission segment weight fractions are calculated by using Breguet equations along with detailed information available for aerodynamics and propulsion system from previous steps. Then, total fuel weight including reserve fuel is found. PERFORMANCE CALCULATIONS In performance calculations, takeoff ground run distance and takeoff distance for 35 ft obstacle clearance are calculated. In this calculation takeoff phase of the mission are analyzed by dividing it into different segments. The segments of takeoff analysis consists of the level ground roll, the ground roll during rotation to the angle of attack to liftoff, transition to climb and finally the climb to 35 ft altitude. Maximum rate of climb and maximum level speed are found for the aircraft. In these calculations aerodynamic characteristics and thrust values are used from the previous calculations along with numerical methods. STABILITY AND CONTROL Longitudinal and lateral stability analysis are performed for the aircraft. In the longitudinal stability analysis, the pitching stability of the aircraft is estimated. Lift curve slope of the wing including fuselage effect is found using empirical equations. Wing and horizontal tail incidence angles according to fuselage are calculated based on a specified trim condition. In addition; to determine the size of elevator and booms, the control power required to stall the aircraft and the control power required for the takeoff rotation at the ground are calculated by including the ground effect on wing and tail. In the lateral stability analysis, the yaw stability of the aircraft are estimated and rudder is sized according to the trim requirement at crosswind-landing condition. OPTIMIZATION Conceptual design process of an aircraft includes the synthesis of some major disciplines such as geometry, aerodynamics, propulsion, stability and control, etc. The design process requires the estimation of the shape and main dimensions of the aircraft and the fuel requirement to fly a prescribed mission and this is an iterative process. Here a computer code was developed to perform all conceptual design process in the computer environment and it is a common practice to run this code repeatedly, changing the design parameters: But, a great deal of effort is required to size the unmanned air vehicle so that all geometric limits, performance and stability requirements are consistent with the design specifications. While only one or two design parameters can be changed, the problem is relatively simple and can be solved by using graphical methods. But the number of design parameters which can be changed is more than two, this is a very time-consuming process for which optimization techniques can be used most effectively. xivIn this study, a numerical optimization code named ADS is used along with the developed computer code performing the conceptual design of the unmanned air vehicle. ADS is a general purpose numerical optimization programs which is written in FORTRAN, and it offers many different algorithms for different design problems. sizing iteration design obtained from the initial sizing ADS numerical optimization routines RESULTS Figure 3. Optimization process of the vehicle SUMT (sequential unconstrained minimization techniques) algorithms of ADS program are used to optimize the unmanned aircraft. Optimization was performed to maximize the payload for a specified engine. In the optimization; geometrical constraints and performance, stability and control constraints were considered. By using numerical optimization, it is possible to optimize the aircraft by changing numbers of design parameters, while the all design constraints are satisfied. xv

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