LEE- Uçak ve Uzay Mühendisliği Lisansüstü Programı

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http://hdl.handle.net/11527/19473

Gözat

Yazar "Nikbay, Melike" ile LEE- Uçak ve Uzay Mühendisliği Lisansüstü Programı'a göz atma
  • Öge
    Development of a nonlinear sonic boom propagation code
    (Graduate School, 2023-01-24) Demiroğlu, Yusuf ; Nikbay, Melike ; 511191144 ; Astronautics and Aeronautics Engineering
    Civil supersonic flight is still one of the most challenging research topics in the aerospace industry. Since Concorde's last flight in 2003, researchers tried to find efficient solutions to make supersonic flights more affordable and reliable. Meanwhile, with the advance of computational power, computational fluid dynamics (CFD) has been implemented in advanced optimization studies involved in elevating supersonic aircraft design processes with given operational criteria and requirements. However, reducing the cost of a supersonic flight by increasing aerodynamic efficiency is not the only concern in civil supersonic transport. The second most important factor for a supersonic aircraft is the noise produced on land due to the shock waves that propagate through the atmosphere to the ground. This phenomenon is called sonic boom which is addressed in this thesis study. A sonic boom generated by a supersonic aircraft can cause very loud noise on the ground that may exceed 100 decibels. This loudness value is not acceptable due to its effects on people's daily life. Therefore, to enable civil supersonic flight over land, sonic boom loudness must be eliminated or reduced below a certain level. This effort is called sonic boom minimization and there are several methodologies that are provided in this study. Lots of studies for sonic boom minimization utilize optimization algorithms that call sonic boom prediction tools along with the CFD solvers. Therefore, to reduce sonic boom loudness, a sonic boom propagation code that accurately predicts sonic boom loudness is essential for the multidisciplinary design optimization of civil supersonic aircraft. In this regard, a new nonlinear sonic boom prediction code, named ITUBOOM, is developed in-house to be incorporated into our design optimization studies to achieve a low-boom aircraft geometry. ITUBOOM is developed in Python programming language for ease of implementation for design studies. A sonic boom calculation process can be broken down into three main steps; a near-field solution with CFD to generate an initial acoustic signal, atmospheric propagation with acoustics methods, and loudness calculation. Unlike other sonic boom codes, ITUBOOM can also be used to generate a near-field pressure directly from CFD outputs by surface slicing or in-flow signature extraction. Then, it can be used to perform atmospheric propagation by taking into account nonlinear effects such as molecular relaxation and thermoviscous attenuation. Results of ITUBOOM are validated against NASA Langley Research Center's well-known sBOOM code for different conditions on benchmark problems and presented in this thesis in detail.
  • Öge
    Implementation of propulsion system integration losses to a supersonic military aircraft conceptual design
    ( 2021-10-07) Karaselvi, Emre ; Nikbay, Melike ; 511171151 ; Aeronautics and Astronautics Engineering ; Uçak ve Uzay Mühendisliği
    Military aircraft technologies play an essential role in ensuring combat superiority from the past to the present. That is why the air forces of many countries constantly require the development and procurement of advanced aircraft technologies. A fifth-generation fighter aircraft is expected to have significant technologies such as stealth, low-probability of radar interception, agility with supercruise performance, advanced avionics, and computer systems for command, control, and communications. As the propulsion system is a significant component of an aircraft platform, we focus on propulsion system and airframe integration concepts, especially in addressing integration losses during the early conceptual design phase. The approach is aimed to be appropriate for multidisciplinary design optimization practices. Aircraft with jet engines were first employed during the Second World War, and the technology made a significant change in aviation history. Jet engine aircraft, which replaced propeller aircraft, had better maneuverability and flight performance. However, substituting a propeller engine with a jet engine required a new design approach. At first, engineers suggested that removing the propellers could simplify the integration of the propulsion system. However, with jet engines for fighter aircraft, new problems arose due to the full integration of the propulsion system and the aircraft's fuselage. These problems can be divided into two parts: designing air inlet, air intake integration, nozzle/afterbody design, and jet interaction with the tail. The primary function of the air intake is to supply the necessary air to the engine with the least amount of loss. However, the vast flight envelope of the fighter jets complicates the air intake design. Spillage drag, boundary layer formation, bypass air drag, and air intake internal performance are primary considerations for intake system integration. The design and integration of the nozzle is a challenging engineering problem with the complex structure of the afterbody and the presence of jet and free-flow mix over control surfaces. The primary considerations for the nozzle system are afterbody integration, boat-tail drag, jet flow interaction, engine spacing for twin-engine configuration, and nozzle base drag. Each new generation of aircraft design has become a more challenging engineering problem to meet increasing military performances and operational capabilities. This increase is due to higher Mach speeds without afterburner, increased acceleration capability, high maneuverability, and low visibility. Tradeoff analysis of numerous intake nozzle designs should be carried out to meet all these needs. It is essential to calculate the losses caused by different intakes and nozzles at the conceptual design of aircraft. Since the changes made after the design maturation delay the design calendar or changes needed in a matured design cause high costs, it is crucial to accurately present intake and nozzle losses while constructing the conceptual design of a fighter aircraft. This design exploration process needs to be automated using numerical tools to investigate all possible alternative design solutions simultaneously and efficiently. Therefore, spillage drag, bypass drag, boundary layer losses due to intake design, boat-tail drag, nozzle base drag, and engine spacing losses due to nozzle integration are examined within the scope of this thesis. This study is divided into four main titles. The first section, "Introduction", summarizes previous studies on this topic and presents the classification of aircraft engines. Then the problems encountered while integrating the selected aircraft engine into the fighter aircraft are described under the "Problem Statement". In addition, the difficulties encountered in engine integration are divided into two zones. Problem areas are examined as inlet system and afterbody system. The second main topic, "Background on Propulsion," provides basic information about the propulsion system. Hence, the Brayton cycle is used in aviation engines. The working principle of aircraft engines is described under the Brayton Cycle subtitle. For the design of engines, numbers are used to standardize engine zone naming to present a common understanding. That is why the engine station numbers and the regions are shown before developing the methodology. The critical parameters used in engine performance comparisons are thrust, specific thrust and specific fuel consumption, and they are mathematically described. The Aerodynamics subtitle outlines the essential mathematical formulas to understand the additional drag forces caused by propulsion system integration. During the thesis, ideal gas and isentropic flow assumptions are made for the calculations. Definition of drag encountered in aircraft and engine integration are given because accurate definitions prevent double accounting in the calculation. Calculation results with developed algorithms and assumptions are compared with the previous studies of Boeing company in the validation subtitle. For comparison, a model is created to represent the J79 engine with NPSS. The engine's performance on the aircraft is calculated, and given definitions and algorithms add drag forces to the model. The results are converged to Boeing's data with a 5% error margin. After validation, developed algorithms are tested with 5th generation fighter aircraft F-22 Raptor to see how the validated approach would yield results in the design of next-generation fighter aircraft. Engine design parameters are selected, and the model is developed according to the intake, nozzle, and afterbody design of the F-22 aircraft. A model equivalent to the F-119-PW-100 turbofan engine is modeled with NPSS by using the design parameters of the engine. Additional drag forces calculated with the help of algorithms are included in the engine performance results because the model is produced uninstalled engine performance data. Thus, the net propulsive force is compared with the F-22 Raptor drag force Brandtl for 40000 ft. The results show that the F-22 can fly at an altitude of 40000 ft, with 1.6M, meeting the aircraft requirements. In the thesis, a 2D intake assumption is modeled for losses due to inlet geometry. The effects of the intake capture area, throat area, wedge angle, and duct losses on motor performance are included. However, the modeling does not include a bump intake structure similar to the intake of the F-35 aircraft losses due to 3D effects. CFD can model losses related to the 3D intake structure, and test results and thesis studies can be developed. The circular nozzle, nozzle outlet area, nozzle throat area, and nozzle maximum area are used for modeling. The movement of the nozzle blades is included in the model depending on the boattail angle and base area. The works of McDonald & P. Hughest are used as a reference to represent the 2D-sized nozzle. The method described in this thesis is one way of accounting for installation effects in supersonic aircraft. Additionally, the concept works for aircraft with conventional shock inlets or oblique shock inlets flying at speeds up to 2.5 Mach. The equation implementation in NPSS enables aircraft manufacturers to calculate the influence of installation effects on engine performance. The study reveals the methodology for calculating additional drag caused by an engine-aircraft integration in the conceptual design phase of next-generation fighter aircraft. In this way, the losses caused by the propulsion system can be calculated accurately by the developed approach in projects where aircraft and engine design have not yet matured. If presented, drag definitions are not included during conceptual design causing significant change needs at the design stage where aircraft design evolves. Making changes in the evolved design can bring enormous costs or extend the design calendar.