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 purpose of this thesis is to design and study an aircraft which implements the Co-Flow Jet (CFJ) airfoil concept, as well as to study the CAARC standard highrise building. The design concept is verified mainly by the use of a Computational Fluid Dynamics (CFD) package. A thorough methodology for geometry and mesh generation is developed, and subsequently applied to the two cases. The first case studied is that of the CFJ Airplane (CFJA). It consists of a threedimensional, highly blended, ying wing geometry implementing the Co-Flow Jet airfoil concept. Though a thorough comparison to a baseline geometry, it is shown that usage of the CFJ airfoil cross-section greatly improves aircraft performance by increasing lift, reducing drag, and providing a source of thrust over the operational range of angles of attack. A steady state CFD simulation is used for this case, as the air ow around an airfoil cross-section is inherently steady for attached ows. CFD results are used to support the Engineless Aircraft" concept, where the CFJ airfoil is used as the sole form of propulsion. The second case studied consists of a rectangular high-rise building undergoing a wind condition with Mach number of 0:1 and a Reynolds number of 160000. Due to the non-streamlined geometry of the building cross-section, aerodynamic instabilities due to uid separation are present, and therefore an unsteady CFD analysis is necessary to fully resolve all of the ow phenomena. Preliminary steady state results are presented, and a plan is laid down for the future study of this highly complex case. Results are presented for a variety of angles of attack in the case of the CFJA, and for the main ow direction in the case of the CAARC building. Results are compared with baseline geometry in the case of the CFJ Airplane. The CFJ Airplane case is simulated using a 3rd order steady state scheme, which is sufficient to achieve valid results for the ow regime. The CAARC building, which has inherent ow separation, requires the use of high order schemes.

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.

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.

Unmanned aerial systems have been widely used for variety of civilian applications over the past few years. Some of these applications require accurate guidance and control. Consequently, Unmanned Aerial Vehicle (UAV) guidance and control attracted many researchers in both control theory and aerospace engineering. Flying wings, as a particular type of UAV, are considered to have one of the most efficient aerodynamic structures. It is however difficult to design robust controller for such systems. This is due to the fact that flying wings are highly sensitive to control inputs. The focus of this thesis is on modeling and control design for a UAV system. The platform understudy is a flying wing developed by SmartPlanes Co. located in Skellefteå, Sweden. This UAV is particularly used for topological mapping and aerial photography. The novel approach suggested in this thesis is to use two controllers in sequence. More precisely, Linear Quadratic Regulator (LQR) is suggested to provide robust stability, and Proportional, Integral, Derivative (PID) controller is suggested to provide reference signal regulation. The idea behind this approach is that with LQR in the loop, the system becomes more stable and less sensitive to control signals. Thus, PID controller has an easier task to do, and is only used to provide the required transient response. The closed-loop system containing the developed controller and a UAV non-linear dynamic model was simulated in Simulink. Simulated controller was then tested for stability and robustness with respect to some parametric uncertainty. Obtained results revealed that the LQR successfully managed to provide robust stability, and PID provided reference signal regulation.

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 aircraft'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.

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 flying-wing is a type of configuration which is a tailless airplane accommodating all of its parts within the outline of a single airfoil. Theoretically, it has the most aerodynamic efficiency. The fuel consumption can be more efficient than the existed conventional airliner. It seems that this configuration can achieve the above mentioned requirements. According to these outstanding advantages, many aircraft companies did a great deal of projects on the flying-wing concept. However, the application was only for sport and military use; for airliner, none of them entered production. FW-11 is a flying-wing configuration airliner which is a design cooperation between Cranfield University and Aviation Industry Corporation of China (AVIC). Aiming the spatial economic and environmental needs, this 200-seat airliner would attract attention from airline companies for cost saving and environmental protection. Before start, this program is designated for a new generation commercial aircraft to compete with the existing same capability airliner, such as Airbus A320 and Boeing 767. As the first team of this program, the aim is to finish the conceptual design and prepare the relevant document for next two teams that will perform preliminary and detail design. As a member of FW-11 program and as part of the GDP, the author has been through the four conceptual design stages: engine manufacturers, aircraft family issues, structure design and the establishment of 3-D CAD model. The aim of IRP study is to focus on the initial fuselage design.

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.