Dissertationen zum Thema „Flying-wing aircraft“

Flying-wing aircraft are considered to have great advantages and potentials in aerodynamic performance and weight saving. However, they also have many challenges in design. One of the biggest challenges is the structural design of the inner wing (fuselage). Unlike the conventional fuselage of a tube configuration, the flying-wing aircraft inner wing cross section is limited to a noncircular shape, which is not structurally efficient to resist the internal pressure load. In order to solve this problem, a number of configurations have been proposed by other designers such as Multi Bubble Fuselage (MBF), Vaulted Ribbed Shell (VLRS), Flat Ribbed Shell (FRS), Vaulted Shell Honeycomb Core (VLHC), Flat Sandwich Shell Honeycomb Core (FLHC), Y Braced Box Fuselage and the modified fuselage designed with Y brace replaced by vaulted shell configurations. However all these configurations still inevitably have structural weight penalty compared with optimal tube fuselage layout. This current study intends to focus on finding an optimal configuration with minimum structural weight penalty for a flying-wing concept in a preliminary design stage. A new possible inner wing configuration, in terms of aerodynamic shape and structural layout, was proposed by the author, and it might be referred as ‘Wave-Section Configuration’. The methodologies of how to obtain a structurally efficient curvature of the shape, as well as how to conduct the initial sizing were incorporated. A theoretical analysis of load transmission indicated that the Wave-Section Configuration is feasible, and this was further proved as being practical by FE analysis. Moreover, initial FE analysis and comparison of the Wave-Section Configuration with two other typical configurations, Multi Bubble Fuselage and Conventional Wing, suggested that the Wave-Section Configuration is an optimal design in terms of weight saving. However, due to limitations of the author’s research area, influences on aerodynamic performances have not yet been taken into account.

The laminar-flying-wing aircraft appears to be an attractive long-term prospect for reducing the environmental impact of commercial aviation. In assessing its potential, a relatively straightforward initial step is the conceptual design of a version with restricted sweep angle. Such a design is the topic of this thesis. In addition to boundary layer laminarisation (utilising distributed suction) and limited sweep, a standing-height passenger cabin and subcritical aerofoil flow are imposed as requirements. Subject to these constraints, this research aims to: provide insight into the parameters affecting practical laminar-flow-control suction power requirements; identify a viable basic design specification; and, on the basis of this, an assessment of the fuel efficiency through a detailed conceptual design study. It is shown that there is a minimum power requirement independent of the suction system design, associated with the stagnation pressure loss in the boundary layer. This requirement increases with aerofoil section thickness, but depends only weakly on Mach number and (for a thick, lightly-loaded laminar flying wing) lift coefficient. Deviation from the optimal suction distribution, due to a practical chamber-based architecture, is found to have very little effect on the overall suction coefficient. In the spanwise direction, through suitable choice of chamber depth, the pressure drop due to frictional and inertial effects may be rendered negligible. Finally, it is found that the pressure drop from the aerofoil surface to the pump collector ducts determines the power penalty; suggesting there is little benefit in trying to maintain an optimal suction distribution through increased subsurface-chamber complexity. For representative parameter values, the minimum power associated with boundary-layer losses alone contributes some 80% - 90% of the total power requirement. To identify the viable basic design specification, a high-level exploration of the laminar-flying-wing design space is performed, with an emphasis above all on aerodynamic efficiency. The characteristics of the design are assessed as a function of three parameters: thickness-to-chord ratio, wingspan, and unit Reynolds number. A feasible specification, with 20% thickness-to-chord, 80 m span and a unit Reynolds number of 8 x 10[superscript 6] m[superscript -1], is identified; it corresponds to a 187 tonne aircraft which cruises at Mach 0.67 and altitude 22,500 ft, with lift coefficient 0.14. The benefit of laminarisation is manifested in a high lift-to-drag ratio, but the wing loading is low, and the structural efficiency and gust response are thus likely to be relatively poor. On the basis of this specification, a detailed conceptual design is undertaken. A 220-passenger laminar-flying-wing concept, propelled by three turboprop engines, with a cruise range of 9000 km is developed. The estimated fuel burn is 13.9 g/pax.km. For comparison, a conventional aircraft, propelled by four turboprop engines, with a high-mounted, unswept, wing is designed for the same mission specification and propulsion characteristics, and is shown to have a fuel burn of 15.0 g/pax.km. Despite significant aerodynamic efficiency gains, the fuel burn of the laminar flying wing is only marginally better as it suffers from a poor cruise engine efficiency, due to extreme differences between takeoff and cruising requirements, and is much heavier. The laminar flying wing proposed in this thesis falls short of the performance improvements expected of the concept, and is not worth the development effort. It is therefore proposed that research efforts either be focussed on improving the engine efficiency, or switching to a low aspect ratio, high sweep, design configuration.

The increasing demand of more economical and environmentally friendly aero engines leads to the proposal of a new concept – geared turbofan. In this thesis, the characteristics of this kind of engine and relevant considerations of integration on a flying wing aircraft were studied. The studies can be divided into four levels: GTF-11 engine modelling and performance simulation; aircraft performance calculation; nacelle design and aerodynamic performance evaluation; preliminary engine installation. Firstly, a geared concept engine model was constructed using TURBOMATCH software. Based on parametric analysis and SFC target, the main cycle parameters were selected. Then, the maximum take-off thrust was verified and corrected from 195.56kN to 212kN to meet the requirements of take-off field length and second segment climb. Besides, the engine performance at offdesign points was simulated for aircraft performance calculation. Secondly, an aircraft performance model was developed and the performance of FW-11 was calculated on the basis of GTF-11 simulation results. Then, the effect of GTF-11 characteristics performance on aircraft performance was evaluated. A comparison between GTF-11 and conventional turbofan, RB211- 524B4, indicated that the aircraft can achieve a 13.1% improvement in fuel efficiency by using the new concept engine. Thirdly, a nacelle was designed for GTF-11 based on NACA 1-series and empirical methods while the nacelle dimensions of conventional turbofan RB211-525B4 were obtained by measure approach. Then, the installation thrust losses caused by nacelle drags of the two engines were evaluated using ESDU 81024a. The results showed that the nacelle drags account for about 4.08% and 3.09% of net thrust for GTF-11 and RB211-525B4, respectively. Finally, the considerations of engine installation on a flying wing aircraft were discussed and a preliminary disposition of GTF-11 on FW-11 was presented.

This research is concerned with the flight dynamic, pitch flight control and flying qualities assessment for the reference BWB aircraft. It aims to develop the longitudinal control laws which could satisfy the flying and handing qualities over the whole flight envelope with added consideration of centre of gravity (CG) variation. In order to achieve this goal, both the longitudinal stability augmentation system (SAS) and autopilot control laws are studied in this thesis. Using the pole placement method, two sets of local Linear-Time-Invariant (LTI) SAS controllers are designed from the viewpoints of flying and handing qualities assessment and wind disturbance checking. The global gain schedule is developed with the scheduling variable of dynamic pressure to transfer gains smoothly between these two trim points. In addition, the poles movement of short period mode with the varying CG position are analysed, and some approaches of control system design to address the problem of reduced stability induced by CG variation are discussed as well. To achieve the command control for the aircraft, outer loop autopilot both pitch attitude hold and altitude hold are implemented by using the root locus method. By the existing criteria in MIL-F-8785C specifications being employed to assess the augmented aircraft response, the SAS linear controller with automatic changing gains effectively improve the stability characteristic for the reference BWB aircraft over the whole envelope. Hence, the augmented aircraft equals to a good characteristic controlled object for the outer loop or command path design, which guarantee the satisfactory performance of command control for the BWB aircraft. The flight control law for the longitudinal was completed with the SAS controller and autopilot design. In particular, the SAS was achieved with Level 1 flying and handing qualities, meanwhile the autopilot system was applied to obtain a satisfactory pitch attitude and altitude tracking performance.

In recent years unconventional aircraft configurations, such as Blended-Wing-Body (BWB) aircraft, are being investigated and researched with the aim to develop more efficient aircraft configurations, in particular for very large transport aircraft that are more efficient and environmentally-friendly. The BWB configuration designates an alternative aircraft configuration where the wing and fuselage are integrated which results essentially in a hybrid flying wing shape. The first example of a BWB design was researched at the Loughead Company in the United States of America in 1917. The Junkers G. 38, the largest land plane in the world at the time, was produced in 1929 for Luft Hansa (present day; Lufthansa). Since 1939 Northrop Aircraft Inc. (USA), currently Northrop Grumman Corporation and the Horten brothers (Germany) investigated and developed BWB aircraft for military purposes. At present, the major aircraft industries and several universities has been researching the BWB concept aircraft for civil and military activities, although the BWB design concept has not been adapted for civil transport yet. The B-2 Spirit, (produced by the Northrop Corporation) has been used in military service since the late 1980s. The BWB design seems to show greater potential for very large passenger transport aircraft. A NASA BWB research team found an 800 passenger BWB concept consumed 27 percent less fuel per passenger per flight operation than an equivalent conventional configuration (Leiebeck 2005). The purpose of this research is to assess the aerodynamic efficiency of a BWB aircraft with respect to a conventional configuration, and to identify design issues that determine the effectiveness of BWB performance as a function of aircraft payload capacity. The approach was undertaken to develop a new conceptual design of a BWB aircraft using Computational Aided Design (CAD) tools and Computational Fluid Dynamics (CFD) software. An existing high-capacity aircraft, the Airbus A380 Contents RMIT University, Australia was modelled, and its aerodynamic characteristics assessed using CFD to enable comparison with the BWB design. The BWB design had to be compatible with airports that took conventional aircraft, meaning a wingspan of not more than 80 meters for what the International Civil Aviation Organisation (ICAO) regulation calls class 7 airports (Amano 2001). From the literature review, five contentions were addressed; i. Is a BWB aircraft design more aerodynamically efficient than a conventional aircraft configuration? ii. How does the BWB compare overall with a conventional design configuration? iii. What is the trade-off between conventional designs and a BWB arrangement? iv. What mission requirements, such as payload and endurance, will a BWB design concept become attractive for? v. What are the practical issues associated with the BWB design that need to be addressed? In an aircraft multidisciplinary design environment, there are two major branches of engineering science; CFD analysis and structural analysis; which is required to commence producing an aircraft. In this research, conceptual BWB designs and CFD simulations were iterated to evaluate the aerodynamic performance of an optimal BWB design, and a theoretical calculation of structural analysis was done based on the CFD results. The following hypothesis was prompted; A BWB configuration has superior in flight performance due to a higher Lift-to-Drag (L/D) ratio, and could improve upon existing conventional aircraft, in the areas of noise emission, fuel consumption and Direct Operation Cost (DOC) on service. However, a BWB configuration needs to employ a new structural system for passenger safety procedures, such as passenger ingress/egress. The research confirmed that the BWB configuration achieves higher aerodynamic performance with an achievement of the current airport compatibility issue. The beneficial results of the BWB design were that the parasite drag was decreased and the spanwise body as a whole can generate lift. In a BWB design environment, several advanced computational techniques were required to compute a CFD simulation with the CAD model using pre-processing and CFD software.

The blended-wing-body (BWB) configuration appears as a promising contender for the next generation of large transport aircraft. The idea of blending the wing with the fuselage and eliminating the tail is not new, it has long been known that tailless aircraft can suffer from stability and control problems that must be addressed early in the design. This thesis is concerned with identifying and then evaluating the flight dynamics, stability, flight controls and handling qualities of a generic BWB large transport aircraft concept. Longitudinal and lateral-directional static and dynamic stability analysis using aerodynamic data representative of different BWB configurations enabled a better understanding of the BWB aircraft characteristics and identification of the mechanisms that influence its behaviour. The static stability studies revealed that there is limited control power both for the longitudinal and lateral-directional motion. The solution for the longitudinal problem is to limit the static margins to small values around the neutral point, and even to use negative static margins. However, for the directional control problem the solution is to investigate alternative ways of generating directional control power. Additional investigation uncovered dynamic instability due to the low and negative longitudinal and directional static stability. Furthermore, adverse roll and yaw responses were found to aileron inputs. The implementation of a pitch rate command/attitude hold flight control system (FCS) improved the longitudinal basic BWB characteristics to satisfactory levels, or Level 1, flying and handling qualities (FHQ). Although the lateral-directional command and stability FCS also improved the BWB flying and handling qualities it was demonstrated that Level 1 was not achieved for all flight conditions due to limited directional control power. The possibility to use the conventional FHQs criteria and requirements for FCS design and FHQs assessment on BWB configurations was also investigated. Hence, a limited set of simulation trials were undertaken using an augmented BWB configuration. The longitudinal Bandwidth/Phase delay/Gibson dropback criteria, as suggested by the military standards, together with the Generic Control Anticipation Parameter (GCAP) proved possible to use to assess flying and handling qualities of BWB aircraft. For the lateral-directional motion the MIL-F-8785C criteria were used. Although it is possible to assess the FHQ of BWB configuartions using these criteria, more research is recommended specifically on the lateral-directional FHQs criteria and requirements of highly augmented large transport aircraft.

This thesis presents a research project and results of design and optimization of a composite wing structure for a large aircraft in flying wing configuration. The design process started from conceptual design and preliminary design, which includes initial sizing and stressing followed by numerical modelling and analysis of the wing structure. The research was then focused on the minimum weight optimization of the /composite wing structure /subject to multiple design /constraints. The modelling, analysis and optimization process has been performed by using the NASTRAN code. The methodology and technique not only make the modelling in high accuracy, but also keep the whole process within one commercial package for practical application. The example aircraft, called FW-11, is a 250-seat commercial airliner of flying wing configuration designed through our MSc students Group Design Project (GDP) in Cranfield University. Started from conceptual design in the GDP, a high-aspect-ratio and large sweepback angle flying wing configuration has been adopted. During the GDP, the author was responsible for the structural layout design and material selection. Composite material has been chosen as the preferable material for both the inner and outer wing components. Based on the derivation of structural design data in the conceptual phase, the author continued with the preliminary design of the outer wing airframe and then focused on the optimization of the composite wing structure. Cont/d.

This thesis contains the specific description of Group Design Project (GDP) and Individual Research Project (IRP) that are undertaken by the author and form part of the degree of Master of Science. The target of GDP is to develop a novel and unique commercial flying wing aircraft titled FW-11. FW-11 is a three-year collaborative civil aircraft project between Aviation Industry Corporation of China (AVIC) and Cranfield University. According to the market analysis result conducted by the author, 250 seats capacity and 7500 nautical miles were chosen as the design targets. The IRP is the further study of GDP, which is to enhance the competitive capability by deploying prognostics and health management (PHM) technology to the fuel system of FW-11. As a novel and brand-new technology, PHM enables the real-time transformation of system status data into alert and maintenance information during all ground or flight operating phases to improve the aircraft reliability and operating costs. Aircraft fuel system has a great impact on flight safety. Therefore, the development of fuel system PHM concept is necessary. This thesis began with an investigation of PHM, then a safety and reliability analysis of fuel system was conducted by using FHA, FMEA and FTA. According to these analyses, fuel temperature diagnosis and prognosis were chosen as a case study to improve the reliability and safety of FW-11. The PHM architecture of fuel temperature had been established. A fuel temperature prediction model was also introduced in this thesis.

A method was developed to assist with the understanding of a unique configuration and investigate some of its stability and control attributes. Oblique wing aircraft concepts are a design option that is well understood, but has yet to be used in a production aircraft. Risk involved in choosing such a design can be averted through additional knowledge early in the concept evaluation phase. Analysis tools commonly used in early conceptual level analysis were evaluated for applicability to a non-standard aircraft design such as an oblique flying wing. Many tools used in early analyses make assumptions that are incompatible with the slewed wing configuration of the vehicle. Using a simplified set of tools, an investigation of a unique configuration was done as well as showing that the aircraft could be trimmed at given conditions. Wave drag was investigated to determine benefits for an oblique flying wing. This form of drag was reduced by the distribution of volume afforded by the slewing of the aircraftâ s wing. Once a reasonable concept was developed, aerodynamic conditions were investigated for static stability of the aircraft. Longitudinal and lateral trim were established simultaneously due to its asymmetric nature.
Master of Science

Effects of engine placement on flutter characteristics of a very flexible high-aspect-ratio wing are investigated using the code NATASHA (Nonlinear Aeroelastic Trim And Stability of HALE Aircraft). The analysis was validated against published results for divergence and flutter of swept wings and found to be in excellent agreement with the experimental results of the classical wing of Goland. Moreover, modal frequencies and damping obtained for the Goland wing were found in excellent agreement with published results based on a new continuum-based unsteady aerodynamic formulation. Gravity for this class of wings plays an important role in flutter characteristics. In the absence of aerodynamic and gravitational forces and without an engine, the kinetic energy of the first two modes are calculated. Maximum and minimum flutter speed locations coincide with the area of minimum and maximum kinetic energy of the second bending and torsion modes. Time-dependent dynamic behavior of a turboshaft engine (JetCat SP5) is simulated with a transient engine model and the nonlinear aeroelastic response of the wing to the engine's time-dependent thrust and dynamic excitation is presented. Below the flutter speed, at the wing tip and behind the elastic axis, the impulse engine excitation leads to a stable limit cycle oscillation; and for the ramp kind of excitation, beyond the flutter speed, at 75% span, behind the elastic axis, it produces chaotic oscillation of the wing. Both the excitations above the flutter speed are stabilized, on the inboard portion of the wing. Effects of engine placement and sweep on flutter characteristics of a backswept flying wing resembling the Horten IV are explored using NATASHA. This aircraft exhibits a non-oscillatory yawing instability, expected in aircraft with neither a vertical tail nor yaw control. More important, however, is the presence of a low frequency “body-freedom flutter” mode. The aircraft center of gravity was held fixed during the study, which allowed aircraft controls to trim similarly for each engine location, and minimized flutter speed variations along the inboard span. Maximum flutter speed occurred for engine placement just outboard of 60% span with engine center of gravity forward of the elastic axis. The body-freedom flutter mode was largely unaffected by the engine placement except for cases in which the engine is placed at the wing tip and near the elastic axis. In the absence of engines, aerodynamics, and gravity, a region of minimum kinetic energy density for the first symmetric free-free bending mode is also near the 60% span. A possible relationship between the favorable flutter characteristics obtained by placing the engines at that point and the region of minimum kinetic energy is briefly explored. Effects of multiple engine placement on a similar type of aircraft are studied. The results showed that multiple engine placement increases flutter speed particularly when the engines are placed in the outboard portion of the wing (60% to 70% span), forward of the elastic axis, while the lift to drag ratio is affected negligibly. The behavior of the sub- and supercritical eigenvalues is studied for two cases of engine placement. NATASHA captures a hump body-freedom flutter with low frequency for the clean wing case, which disappears as the engines are placed on the wings. In neither case is there any apparent coalescence between the unstable modes. NATASHA captures other non-oscillatory unstable roots with very small amplitude, apparently originating with flight dynamics. For the clean-wing case, in the absence of aerodynamic and gravitational forces, the regions of minimum kinetic energy density for the first and third bending modes are located around 60% span. For the second mode, this kinetic energy density has local minima around the 20% and 80% span. The regions of minimum kinetic energy of these modes are in agreement with calculations that show a noticeable increase in flutter speed at these regions if engines are placed forward of the elastic axis. High Altitude, Long Endurance (HALE) aircraft can achieve sustained, uninterrupted flight time if they use solar power. Wing morphing of solar powered HALE aircraft can significantly increase solar energy absorbency. An example of the kind of morphing considered in this thesis requires the wings to fold so as to orient a solar panel to be hit more directly by the sun's rays at specific times of the day. In this study solar powered HALE flying wing aircraft are modeled with three beams with lockable hinge connections. Such aircraft are shown to be capable of morphing passively, following the sun by means of aerodynamic forces and engine thrusts. The analysis underlying NATASHA was extended to include the ability to simulate morphing of the aircraft into a “Z” configuration. Because of the “long endurance” feature of HALE aircraft, such morphing needs to be done without relying on actuators and at as near zero energy cost as possible. The emphasis of this study is to substantially demonstrate the processes required to passively morph a flying wing into a Z-shaped configuration and back again.

This research programme explores, theoretically and experimentally, a new liftsystem for Vertical/Short Take-off and Landing (V/STOL) Aircraft. It is based upon an annular wing wrapped around a centrifugal flow generator, potentially creating a vehicle with no external moving parts, reduced vehicle aerodynamic losses compared to previous V/STOL technologies and substantially eliminating induced drag. It is shown that such a wing works best with a thick aerofoil section, and appears to offer greatest potential at a micro-aerial vehicle scale with regard to fundamental performance parameter “lift to weight ratio”. Certain efficiency losses are encountered mainly occurring from annular flow expansion and problems with achieving acceptable blower slot heights. Experimental methods are described along with results, and a comparison shows that the experimental values remain below theoretical values, partly due to flow asymmetry but possibly also other factors. Symmetrical blowing, as initially hypothesised, was found to be impracticable; this suggested use of pure upper surface blowing with Coanda effect. The modified approach was further explored and proved viable. The ultimate goal of this work was to develop an understanding and the facility to integrate the annular-wing into a vehicle to achieve controlled powered flight. To serve the purpose, issues encountered on current and past V/STOL aircraft are being investigated to set a path for further research/development and to validate/justify the design of future V/STOL aircraft. Also, presented is a feasibility study where different physical scales and propulsion systems are considered, and a turbofan has shown to achieve the best performance in terms of Range and Endurance. This privilege allows one to accurately study the V/STOL technologies around.

The cabin environment of a commercial aircraft, including cabin layout and the quality of air supply, is crucial to the airline operators. These aspects directly affect the passengers’ experience and willing to travel. This aim of this thesis is to design the cabin layout for flying wing aircraft as part of cabin environment work, followed by the air quality work, which is to understand what effect the ECS can have in terms of cabin air contamination. The project, initially, focuses on the cabin layout, including passenger cabin configuration, seat arrangement and its own size due to the top requirements, of a conventional aircraft and further into that of a flying wing aircraft. The cabin work in respect of aircraft conceptual design is discussed and conducted by comparing different design approaches. Before the evaluation of cabin air quality, an overall examination of the main ECS components involved in the contaminants access will be carried on and, therefore, attempt to discover how these components influence the property of the concerned contaminants. By case study in the B767 ECS, there are some comments and discussions regarding the relationship between the cabin air contaminations and the passing by ambient environment. The thesis ends up with a conclusion explaining whether or not the contaminated air enters the occupants’ compartments on aircraft and proposing some approaches and engineering solutions to the continue research.

The Multiple Simultaneous Specification controller design method is an elegant means of designing a single controller to satisfy multiple convex closed loop performance specifications. In this thesis, the method is used to design pitch and roll attitude controllers for a Zagi flying-wing unmanned aerial vehicle from Procerus Technologies. A linear model of the aircraft is developed, in which the lateral and longitudinal motions of the aircraft are decoupled. The controllers are designed for this decoupled state space model. Linear simulations are performed in Simulink, and all performance specifications are satisfied by the closed loop system. Nonlinear, hardware-in-the-loop simulations are carried out using the aircraft, on-board computer, and ground station software. Flight tests are also executed to test the performance of the designed controllers. The closed loop aircraft behaviour is generally as expected, however the desired performance specifications are not strictly met in the nonlinear simulations or in the flight tests.

Applications such as urban surveillance, search and rescue, agricultural applications, military applications, etc., require miniature air vehicles (MAVs) to fly for a long time. But they have restricted flight duration due to their dependence on battery life, which necessitates optimal path planning. The generated optimal path should obey the curvature limits prescribed by the minimum turn radius/ maximum turn rate of the MAV. Further, in a dynamically changing environment, the final configuration that the MAV has to achieve may change en route, which demands the path to be replanned by an airborne processor in real-time. As MAVs are small in size and light in weight, wind has a very significant effect on the flight of MAVs and the computation of the minimum-time path in the presence of wind plays an important role. The thesis develops feasible trajectory generation algorithms which are fast, efficient, optimal and implementable in an onboard computer for rectilinear and circular path convergence problems and waypoint following problems both in the absence and in the presence of wind. The first part of the thesis addresses the problem of computation of optimal trajectories when MAVs fly on a two-dimensional (2D) plane maintaining a constant altitude. The shortest path is computed for MAVs from a given initial position and orientation to a given final path with a specified direction as required for a given mission. Unlike the classical Dubins problem where the shortest path was computed between two given configurations (position and orientation), the final point in this case is not specified. However, the final path, which can either be a rectilinear path or a circular path, and the direction to which the MAV should converge, is specified. The time-optimal path of MAVs is developed in the presence of wind mainly using the geometric approach although a few important properties are also obtained using optimal control theory, specifically, Pontryagin’s minimum principle (which provides only the necessary condition for optimality) for control-constrained systems. The complete optima l solution to this problem in all its generality is a major contribution of this thesis as existing methods in the literature that address this problem are either not optimal or do not give a complete solution. Further, the time-optimal path for specified initial and final configurations is generated in reasonably short time without computing all the path lengths of possible candidate paths, which is the method that exists in the literature for similar problems. Simulation results illustrate path generation for various cases, including the presence of steady and time-varying wind. Another problem in MAV path planning in 2D addressed in this thesis computes an extremal path that transitions between two consecutive waypoint segments (obtained by joining two way points in sequence) in a time-optimal fashion. This designed trajectory, named as γ-trajectory, is also used to track the maximum portion of waypoint segments in minimum time and the shortest distance between this trajectory and the associated waypoint can be set to a desired value. Another optimal path, called the loop trajectory, that goes through the way points as well as through the entire waypoint segments, is also proposed. Subsequently, the thesis proposes algorithms to generate trajectories in the presence of steady wind and compares these with the optimal trajectory generated using nonlinear programming based multiple shooting method to show that the generated paths are optimal in most cases. In three-dimensional (3D) space, if the initial and final configurations – in terms of (X,Y,Z) position, heading angle and flight path angle- of the vehicle are specified then shortest path computation is an interesting problem in literature. The proposed method in this thesis is based on 3D geometry and, unlike the existing iterative methods which yield suboptimal paths and are computationally more intensive, this method generates the shortest path in much less time. Due to its simplicity and low computational requirements, this approach can be implemented on a MAV in real-time. But, If the path demands very high pitch angle (as in the case of steep climbs), the generated path may not be flyable for an aerial vehicle with limited range of flight path angles. In such cases numerical methods, such as multiple shooting, coupled with nonlinear programming, are used to obtain the optimal solution. The time-optimal 3D path is also developed in the presence of wind which has a magnitude comparable to the speed of MAVs. The simulation results show path generation for a few sample cases to show the efficacy of the proposed approach as compared to the available approach in the literature. Next, the path convergence problem is studied in 3D for MAVs. The shortest path is generated to converge to a rectilinear path and a circular path starting from a known initial position and orientation. The method is also extended to compute the time-optimal path in the presence of wind. In simulation, optimal paths are generated for a variety of cases to show the efficacy of the algorithm. The other problem discussed in this thesis considers curvature-constrained trajectory generation technique for following a series of way points in 3D space. Extending the idea used in 2D, a γ-trajectory in 3D is generated to track the maximum portion of waypoint segments with a desired shortest distance between the trajectory and the associated waypoint. Considering the flyability issue of the plane a loop-trajectory is generated which is flyable by a MAV with constrained flight path angle. Simulation results are given for illustrative purposes. The path generation algorithms are all based on a kinematic model, considering the vehicle as a point in space. Implementing these results in a real MAV will require the dynamics of the MAV to be considered. So, a 6-DOF SIMULINK model of a MAV is used to demonstrate the tracking of the computed paths both in 2D plane and in 3D space using autopilots consisting of proportional-integral-derivative (PID )controllers .Achieving terminal condition accurately in real-time, if there is noisy measurement of wind data, is also addressed.