Tesis: "Aircraft Wing"

1

McKechnie, Gregor. "Wing Design for the ECO1 Aircraft". Thesis, KTH, Flygdynamik, 2015. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-180460.

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An analysis of a new wing for the ECO1 general aviation aircraft. A new wing for this aircraft is hoped to provide better high-lift performance and a higher maximum angle of attack. To this end, using computational fluid dynamics with the SST k-ω turbulence model, this study explores modifying the wing profile and augmenting the wings with winglets, a feature not commonly used in general aviation. Using the CFD tool Fluent, two-dimensional simulations on variations of the Osquavia aerofoil indicate that both the slimmer and the elongated variants will provide for a greater maximum coefficient of lift and a higher stall angle. The three-dimensional CFD analysis predict that the winglets will decrease the overall drag in high lift-coefficient flight conditions without greatly penalising performance in the cruise condition. The addition of winglets are also shown to provide a higher angle of stall and improve flow across the control surfaces at high angles of attack.

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2

Go, Tiauw Hiong. "Aircraft wing rock dynamics and control". Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/50081.

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Thesis (Sc.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1999.
Includes bibliographical references (p. 232-236).
The dynamics of wing rock on rigid aircraft having single, two, and three rotational degrees-of-freedom are analyzed. For the purpose of the analysis, nonlinear mathematical models of the aircraft are developed. The aerodynamic expressions contained in the models can be built by fitting the appropriate aerodynamic data into the model. The dynamic analysis is performed analytically using a technique combining the Multiple Time Scales method, Center Manifold Reduction principle, and bifurcation theory. The technique yields solutions in parametric forms and leads to the separation of fast and slow dynamics, and a great insight into the system behavior. Further, a unified framework for the investigation of wing rock dynamics and control of aircraft is developed. Good agreement between the analytical results and the numerical simulations is demonstrated. Based on the results of the dynamic analysis, appropriate control strategies for the wing rock alleviation are developed. The control power limitation of the conventional aerodynamics control surfaces is considered and its effects on the alleviation of wing rock are investigated. Finally, the potential use of advanced controls to overcome the conventional controls limitation is discussed.
by Tiauw Hiong Go.
Sc.D.

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3

Huang, Haidong. "Optimal design of a flying-wing aircraft inner wing structure configuration". Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7439.

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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.

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4

Andersson, Daniel. "The performance of an iced aircraft wing". Thesis, Högskolan Väst, Institutionen för ingenjörsvetenskap, 2011. http://urn.kb.se/resolve?urn=urn:nbn:se:hv:diva-4098.

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The goal of this thesis work has been to develop and manufacture an ice layer which was to be mounted on the tip of a scaled down wing model. The iced wing should be tested in a wind tunnel and aerodynamic comparisons should be made to the same wing without ice.The development of the ice was carried out as a modified product development process. The main differences are that there is no costumer and that the actual shape and functions of the product are more or less predetermined. The challenge was to find the best way to create the ice layer and how to mount it to the wing without damaging it or covering any pressure sensors. Product development methods such as pros and cons lists and prototypes were used to solve problems before printing the plastic ice layer in a rapid prototyping machine.Wind tunnel experiments were then conducted on the wing with and without the manufactured ice. Raw data from the wind tunnel were processed and lift and drag coefficients were calculated using mathematical equations. Finally, conclusions were drawn by comparing the results from the wind tunnel tests with theory, other works as well as CFD simulations.The ice layer was successfully manufactured and it met the target specifications. The aerodynamic performance of an iced aircraft wing proved to be considerably worse compared to a blank wing. The maximum achievable lift force decreased by 22% and an increased drag force will require more thrust from the airplane.

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5

Xia, YuXin M. B. A. Sloan School of Management. "M28 Fixed wing transport aircraft cost reduction". Thesis, Massachusetts Institute of Technology, 2011. http://hdl.handle.net/1721.1/66038.

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Thesis (M.B.A.)--Massachusetts Institute of Technology, Sloan School of Management; and, (S.M.)--Massachusetts Institute of Technology, Engineering Systems Division; in conjunction with the Leaders for Global Operations Program at MIT, 2011.
Cataloged from PDF version of thesis.
Includes bibliographical references (p. 146-148).
The M28 is a Polish short-takeoff-and-landing (STOL) light cargo aircraft developed in 1984 and currently built by PZL Mielec, a subsidiary of United Technology Corporation (UTC). There has been renewed interest in the product from military and commercial markets due to its impressive STOL capabilities. However, in order to become price-competitive, its cost would need to be reduced significantly. Multiple cost-reduction concepts have been proposed by the manufacturing and procurement groups. An Optimization Team was also formed to lead the cost-reduction effort. However, a more systematic approach is required in order to achieve the ambitious reduction goals. The proposed solution is to create a top-down systematic cost-reduction framework used to coordinate and prioritize the team's current bottom-up approach. A top-down cost reduction strategy was developed based on UTC Otis' Octopus Fishing concept. Such methodology, heavily finance driven, systematically breaks M28 into sub-systems, and prioritizes improvement recommendations based on cost-reduction potentials. It also leverages on the wealth of knowledge from global cross-functional teams to generate explosive amount of improvement recommendations. The sub-systems were benchmarked against competitors cost structures. The framework will be linked to concepts generated from the database to create a process that combine top-down and bottom-up approaches. After tasks were prioritized using the outlined framework, a three-prong approach was implemented to enhance cost reduction capability. Manufacturing of labor intensive parts such as nacelle deflection cover was automated using CNC machines. A set of commodity purchasing strategies were formulated for forgings, avionics, raw materials, interior and composite materials. Lastly, a discrete Kaizen event was described to aid redesign-for-manufacturing.
by Yuxin Xia.
S.M.
M.B.A.

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6

Miao, Zhisong. "Aircraft engine performance and integration in a flying wing aircraft conceptual design". Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7249.

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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.

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7

Mesrobian, Chris Eden. "Concept Study of a High-Speed, Vertical Take-Off and Landing Aircraft". Thesis, Virginia Tech, 2009. http://hdl.handle.net/10919/35574.

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The purpose of the study was to evaluate the merits of the DiscRotor concept that combine the features of a retractable rotor system for vertical take-off and landing (VTOL) with an integral, circular wing for high-speed flight. Tests were conducted to generate basic aerodynamic characteristics of the DiscRotor in hover and in fixed-wing flight.

To assess the DiscRotor during hover, small scale tests were conducted on a 3ft diameter rotor without the presence of a fuselage. A â hover rigâ was constructed capable of rotating the model rotor at speeds up to 3,500 RPM to reach tip speeds of 500fps. Thrust and torque generated by the rotating model were measured via a two-component load cell, and time averaged values were obtained for various speeds and pitch angles. It has been shown that the DiscRotor will perform well in hover. Ground Effects in hover were examined by simulating the ground with a movable, solid wall. The thrust was found to increase by 50% compared to the ground-independent case. Pressure distributions were measured on the ground and disc surfaces. Velocity measurements examined the flow field downstream of the rotor by traversing a seven hole velocity probe. A wake behind the rotor was shown to contract due to a low pressure region that develops downstream of the disc.

Wind tunnel experimentation was also performed to examine the fixed wing flight of the DiscRotor. These experiments were performed in the VA Tech 6â X6â Stability Tunnel. A model of the fuselage and a circular wing was fabricated based upon an initial sizing study completed by our partners at Boeing. Forces were directly measured via a six degree of freedom load cell, or balance, for free stream velocities up to 200fps. Reynolds numbers of 2 and 0.5 million have been investigated for multiple angles of attack. Low lift-to-drag ratios were found placing high power requirements for the DiscRotor during fixed-wing flight. By traversing a seven-hole velocity probe, velocities in a 2-D grid perpendicular to the flow were measured on the model. The strengths of shed vortices from the model were calculated. A method to improve fixed-wing performance was considered where two blades were extended from the disc. An increase of 0.17 in the CL was measured due to the interaction between the disc and blades.

This research utilized a wide range of experiments, with the aim of generating basic aerodynamic characteristics of the DiscRotor. A substantial amount of quantitative data was collected that could not be included in this document. Results aided in the initial designs of this aircraft for the purpose of evaluating the merit of the DiscRotor concept.
Master of Science

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8

Ikeda, Toshihiro y toshi ikeda@gmail com. "Aerodynamic Analysis of a Blended-Wing-Body Aircraft Configuration". RMIT University. Aerospace, Mechanical and Manufacturing Engineering, 2006. http://adt.lib.rmit.edu.au/adt/public/adt-VIT20070122.163030.

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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.

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9

Zan, Steven James. "An investigation of low-speed wing buffet". Thesis, University of Cambridge, 1990. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.358845.

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10

Qiao, Yuqing. "Effect of wing flexibility on aircraft flight dynamics". Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7280.

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The purpose of this thesis is to give a preliminary investigation into the effect of wing deformation on flight dynamics. The candidate vehicle is FW-11 which is a flying wing configuration aircraft with high altitude and long endurance characteristics. The aeroelastic effect may be significant for this type of configuration. Two cases, the effect of flexible wing on lift distribution and on roll effectiveness during the cruise condition with different inertial parameters are investigated. For the first case, as the wing bending and twisting depend on the interaction between the wing structural deflections and the aerodynamic loads, the equilibrium condition should be calculated. In order to get that condition, mass, structure characteristics and aerodynamic characteristics are estimated first. Then load model and aerodynamic model are built. Next the interaction calculation program is applied and the equilibrium condition of the aircraft is calculated. After that, effect of wing flexibility on lift parameters is investigated. The influence of CG, location of lift and location of flexural axis are investigated. The other case is to calculate the transient roll rate response and estimate the rolling effectiveness of flexible aircraft, and compared with the rigid aircraft’s. A pure roll model is built and derivatives both for the rigid wing and the flexible wing are estimated. It has been found that flexible wing leads to the loss of control effectiveness, even cause reversal when reduces the structure natural frequency. The influence of inertia data for flexible roll is also investigated.

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11

Theos, Athanasios. "Design and analysis of welded aircraft wing panels". Thesis, Cranfield University, 2005. http://hdl.handle.net/1826/3901.

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Nowadays, increasing manufacturing cost effectiveness becomes a vital condition for the commercial success of the next generation of large wide body aircrafts. Welding is a very strong candidate process to be used in manufacturing, allowing both sensible cost reductions and structural efficiency. The main aim of the work is to study the fatigue crack propagation in welded structures. The study is focused on the effect of welding residual stresses to the damage tolerance behaviour of the structure. The welding technique under investigation is the Variable Polarity Plasma Arc (VPPA). Two stringer panels were designed, one tension panel to simulate the lower wing skin cover and one compression panel to mimic the upper wing skin cover. The main design driving force for the upper stiffened panel is buckling since it is under compression. Damage tolerance is the main design criterion for the lower stiffened panel due tensile fatigue loading. Design of the end- fittings for the tension stiffened panel was also carried out using finite element modelling in order to ensure uniform stress distribution at the cross section of the test area of the structure. A fatigue analysis at the various locations of the bolts and at the weld line has been performed. This is necessary in order to ensure that the crack initiation site comes from the weld line rather than from the fastener holes at the end-fittings during the fatigue testing. The research was focused on fatigue crack growth behaviour of welded aluminium panels. The FE model of the CCT coupon is the main tool for the comparison of the fatigue crack behaviour between the parent and the welded coupons. Furthermore AFGROW software is used in conjuction with the output of the FE model to compare the experimental and numerical results in terms of fatigue crack growth lives of welded coupons. In welded coupons a faster crack propagation growth was demonstrated at the region of the weld line and the heat affected zone (HAZ) due to the tensile welding residual stresses. Away from this region, a decrease in crack growth took place due to the compressive welding residual stresses in this area. Finally, a calculation effort in large-scale stiffened panels was made in terms of the stress intensity factor for both welded and non-welded cases. Possible future work was also addressed in such large-scale structures.

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12

Daniels, Charles L. "Comparison of fixed wing aircraft algorithms for JANUS". Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1994. http://handle.dtic.mil/100.2/ADA288503.

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13

Sweeney, Joseph Woods III y Edward M. Wu. "Computer aided deflection measurements of an aircraft wing". Thesis, Monterey, California. Naval Postgraduate School, 1987. http://hdl.handle.net/10945/22217.

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14

Gilmore, Christopher K. (Christopher Kenneth). "Electro-aerodynamic thrust for fixed-wing aircraft propulsion". Thesis, Massachusetts Institute of Technology, 2017. http://hdl.handle.net/1721.1/112452.

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Thesis: Ph. D., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2017.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 301-314).
Aviation operations negatively impact global climate, degrade surface air quality, and create noise. Towards mitigating these effects, this thesis considers electro-aerodynamic (EAD) propulsion, a form of in-atmosphere electrostatic propulsion, which requires no on-board propellant and has zero primary gaseous emissions. In addition, thrust generation has the potential to be nearly silent and requires no moving parts. Despite these advantages, however, EAD propulsion has yet to be implemented in fixed-wing aircraft, in part due to the limited understanding of EAD thruster performance. The objective of this thesis is to determine the feasibility and viability of EAD propulsion in fixed-wing aircraft applications. This thesis begins with a theoretical assessment of EAD thruster performance. This includes quantification of fundamental thrust density limits and the effect of interacting electric fields on thrust-to-power performance due to closely spaced electrode pairs. Additionally, performance as a function of altitude and vehicle flight speed is quantified, where thrust-to-power ratio is estimated to decrease as both increase. Next, this thesis experimentally assesses the achievable thrust density of EAD propulsion. Current and thrust generated from arrays of electrode pairs are observed to be a function of non-dimensional pair spacings for both parallel and staged operation. A thrust per unit area of 2 - 3 N/m² and per unit volume of 5 - 15 N/m³ are estimated, achieving approximately 50 and 10% of the corresponding one-dimensional space-charge limits, respectively. Results suggest that EAD propulsion is most readily viable at the small unmanned aerial vehicle (UAV) scale. Finally, based on the conclusions of the thrust density assessment, this thesis presents the development of a first-of-its-kind EAD-propelled, small-UAV prototype with the goal of achieving steady-level flight. A design space analysis is performed, determining that designs capable of steady-level flight potentially exist. The prototype development effort concludes with at-scale performance quantification of the primary EAD UAV subsystems. Results indicate that the achievable weight-to-thrust ratio is comparable to the vehicle lift-to-drag ratio. This thesis concludes that at the selected scale of the UAV prototype, EAD propulsion is potentially viable, and steady-level flight is, at worst, "nearly" feasible with the current design.
by Christopher Kenneth Gilmore.
Ph. D.

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15

Weed, Philip Andrew. "Hybrid wing-body aircraft noise and performance assessment". Thesis, Massachusetts Institute of Technology, 2010. http://hdl.handle.net/1721.1/62320.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2010.
This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from student-submitted PDF version of thesis.
Includes bibliographical references (p. 89-92).
Hybrid wing-body aircraft noise generation and boundary layer ingestion (BLI) performance trends with increased fan face Mach number inlet designs are investigated. The presented topics are in support of the NASA subsonic fixed wing project, which seeks to lower noise and increase performance by improving prediction methods and technologies. The aircraft configurations used for study are the N2A, using conventional podded engines, and the N2B, using an embedded propulsion system. Preliminary FAR Part 36 noise certification assessments are completed using the NASA Aircraft Noise Prediction Program (ANOPP). The limitations of applying current ANOPP noise prediction methods to hybrid wing-body aircraft are investigated. Improvements are made to the landing gear and airfoil self-noise modules, while a diffraction integral method is implemented in a companion thesis to enhance noise shielding estimates. The N2A overall takeoff and landing noise estimate is found to be 5.3 EPNdB higher than the N+2 goal. The dominant noise sources are the fan rearward and jet on takeoff and the main landing gear and elevons on approach. A lower fan pressure ratio and advanced landing gear fairings are recommended to decrease N2A overall noise levels. The available engine noise estimation tools were inadequate to model the N2B distributed propulsion system and rectangular exhaust nozzle; therefore, overall N2B aircraft noise results are presented for reference only. A simplified embedded propulsion system integration study is carried out to explore the N2B fan design space. A 2-D computational domain with contoured slip boundaries around the centerbody is used to replicate the effects of 3-D relief on the airframe and inlet aerodynamics. The domain includes the S-shaped inlet duct and is extended far downstream for a Trefftz plane power balance analysis to determine the propulsive power required for steady level flight. A fan actuator volume is included to couple the airframe external and the engine internal flows. Aircraft power savings, fan efficiency, and boundary layer thickness trends are examined to determine if increasing fan face Mach number improves system performance while mitigating the total pressure distortion risk of boundary layer ingestion. A fan face Mach number near 0.7 is found to increase aircraft power savings 12% relative to the baseline design and to reduce centerbody boundary layer kinetic energy thickness by 4.7%. In addition, power balances at lower fan pressure ratios as fan face Mach number increases suggesting that high-flow low pressure ratio fans are desirable for BLI.
by Philip Andrew Weed.
S.M.

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16

Quinn, Wilma W. "Multivariable control of a forward swept wing aircraft". Thesis, Massachusetts Institute of Technology, 1986. http://hdl.handle.net/1721.1/15015.

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Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Electrical Engineering and Computer Science, 1986.
MICROFICHE COPY AVAILABLE IN ARCHIVES AND ENGINEERING
Bibliography: leaves 121-122.
by Wilma W. Quinn.
M.S.

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17

Dababneh, Odeh. "Multidisciplinary design optimisation for aircraft wing mass estimation". Thesis, Cranfield University, 2016. http://dspace.lib.cranfield.ac.uk/handle/1826/10172.

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The implementation of key technologies in the initial stages of the aircraft wing design process has always represented a substantial challenge for aircraft designers. The lack of reliable and accessible wing mass prediction methods ¬which allow assessment of the relative benefits of new technologies for reducing structural wing weight - is of significant importance. This necessitates the development of new and generally applicable wing mass estimation methods. This thesis aims to create a new framework for estimating the mass of metallic and composite transport aircraft wings via finite element multidisciplinary analysis, and design optimisation techniques. To this end, the multidisciplinary static strength and stiffness, dynamic aeroelastic stability, and manufacturing constraints are simultaneously addressed within an optimisation environment through a gradient-based search algorithm. A practical optimisation procedure is presented as part of the sizing optimisation process, with enhanced features in solving large-scale nonlinear structural optimisation problems, incorporating an effective initial design variable value generation scheme based on the concept of the fully stressed design. The applicability and accuracy of the proposed approaches is accomplished by conducting a number of case studies in which the wingbox structure of the public domain NASA wing - commonly referred to as the Common Research Model (CRM) - is optimised to produce a minimum mass design. The results of a case study examining minimisation of the mass of the CRM wingbox structures designed using four different models of increasing structural fidelity prove that the multidisciplinary design optimisation framework can successfully calculate the mass of realistic real-world aircraft wing designs. This provides an insight into the competence of certain wingbox models in predicting the mass of the metallic and composite primary wing structures to an acceptable level of accuracy, and in demonstrating the relative merits of the wingbox structural complexity models under consideration and the computational resources necessary to achieving the required degree of accuracy ... [cont.].

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18

Saeed, Tariq Issam. "Conceptual design for a laminar-flying-wing aircraft". Thesis, University of Cambridge, 2012. https://www.repository.cam.ac.uk/handle/1810/243926.

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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.

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19

Okonkwo, Paulinus Peter Chukwuemeka. "Conceptual design methodology for blended wing body aircraft". Thesis, Cranfield University, 2016. http://dspace.lib.cranfield.ac.uk/handle/1826/10132.

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The desire to create an environmentally friendly aircraft that is aerodynamically efficient and capable of conveying large number of passengers over long ranges at reduced direct operating cost led aircraft designers to develop the Blended Wing Body(BWB) aircraft concept. The BWB aircraft represents a paradigm shift in the design of aircraft. The design offers immense aerodynamics and environmental benefits and is suitable for the integration of advanced systems and concepts like laminar flow technology, jet flaps and distributed propulsion. However, despite these benefits, the BWB is yet to be developed for commercial air transport. This is due to several challenges resulting from the highly integrated nature of the configuration and the attendant disciplinary couplings. This study describes the development of a physics based, deterministic, multivariate design synthesis optimisation for the conceptual design and exploration of the design space of a BWB aircraft. The tool integrates a physics based Athena Vortex Lattice aerodynamic analysis tool with deterministic geometry sizing and mass breakdown models to permit a realistic conceptual design synthesis and enables the exploration of the design space of this novel class of aircraft. The developed tool was eventually applied to the conceptual design synthesis and sensitivity analysis of BWB aircraft to demonstrate its capability, flexibility and potential applications. The results obtained conforms to the pattern established from a Cranfield University study on the BlendedWing Body Aircraft and could thus be applied in conceptual design with a reasonable level of confidence in its accuracy.

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20

Grima, Alexander. "Aerodynamic characterisation of an experimental tilt-wing aircraft". Thesis, KTH, Flygdynamik, 2016. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-198526.

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Computational analysis of experimental aircraft prior to test ights can be a valuable tool to estimate ight characteristics and determine areas of elevated caution. It can also provide feedback to software and model developers as to the accuracy of models used when the aircraft is ultimately own. This paper describes the aerodynamic analysis and characterisation of an experimental tilt-wing aircraft with a unique design. The paper covers what analysis is performed as well as results of these aircraft characterisations. Through this analysis a database le is created for use with NASA Design and Analysis of Rotorcraft (NDARC) tool.

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21

McCann, W. J. "Investigation of an over-wing propeller in conjunction with a flap". Thesis, Queen's University Belfast, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.356897.

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22

Smith, Kenneth Wayne. "Fighter Aircraft Synthesis/Design Optimization". Thesis, Virginia Tech, 2009. http://hdl.handle.net/10919/32821.

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This thesis presents results of the application of energy-based large-scale optimization of a two-subsystem (propulsion subsystem (PS) and airframe subsystem-aerodynamics (AFS-A)) air-to-air fighter (AAF) with two types of AFS-A models: a fixed-wing AFS-A and a morphing-wing AFS-A. The AAF flies 19 mission segments of a supersonic fighter aircraft mission and the results of the study show that for very large structural weight penalties and fuel penalties applied to account for the morphing technology, the morphing-wing aircraft can significantly outperform a fixed-wing AAF counterpart in terms of fuel burned over the mission. The optimization drives the fixed-wing AAF wing-geometry design to be at its best flying the supersonic mission segment, while the morphing-wing AFS-A wing design is able to effectively adapt to different flight conditions, cruising at subsonic speeds much more efficiently than the fixed-wing AAF and, thus yielding significant fuel savings. Also presented in this thesis are partially optimized results of the application of a decomposition strategy for large-scale optimization applied to a nine-subsystem AAF consisting of a morphing-wing AFS-A, turbofan propulsion subsystem (PS), environmental controls subsystem (ECS), fuel loop subsystem (FLS), vapor compression/polyalphaolefin loop subsystem (VC/PAOS), electrical subsystem (ES), central hydraulics subsystem (CHS), oil loop subsystem (OLS), and flight controls subsystem (FCS). The decomposition strategy called Iterative Local-Global Optimization (ILGO) is incorporated into a new engineering aircraft simulation and optimization software called iSCRIPTâ ¢ which also incorporates the models developed as part of this thesis work for the nine-subsystem AAF. The AAF flies 21 mission segments of a supersonic fighter aircraft mission with a payload drop simulating a combat situation. The partially optimized results are extrapolated to a synthesis/design which is believed to be close to the system-level optimum using previously published results of the application of ILGO to a five-subsystem AAF to which the partially optimized results of the nine-subsystem AAF compare relatively well. In addition to the optimization results, a parametric study of the morphing AFS-A geometry is conducted. Three mission segments are studied: subsonic climb, subsonic cruise, and supersonic cruise. Four wing geometry parameters are studied: leading-edge wing sweep angle, wing aspect ratio, wing thickness-to-chord ratio, and wing taper ratio. The partially optimized AAF is used as the baseline, and the values for these geometric parameters are increased or decreased up to 20% relative to an established baseline to see the effect, if any, on AAF fuel consumption for these mission segments. The only significant effects seen in any of the mission segments arise from changes in the leading-edge sweep angle and wing aspect ratio. The wing thickness-to-chord ratio shows some effect during the subsonic climb segment, but otherwise shows no effect along with the taper ratio in any of the three mission segments studied. It should be emphasized, however, that these changes are made about a point (i.e. synthesis/design), which is already optimal or nearly so. Thus, the conclusions drawn cannot be generalized to syntheses/designs, which may be far from optimal. Also note that the results upon which these conclusions are based may very likely highlight a weakness in the conceptual-level drag-buildup method used in this thesis work. Further optimization studies using this drag-buildup method may warrant setting the thickness-to-chord ratios and taper ratios rather than having them participate in the optimization as degrees of freedom (DOF). The final set of results is a parametric study conducted to highlight the correlation between the fuel consumption and the total exergy destruction in the AFS-A. The results for the subsonic cruise and supersonic cruise mission segments show that at least for the case when the AFS-A is optimized by itself for a fixed specific fuel consumption that there is a direct correlation between the fuel burned and total exergy destruction. However, as shown in earlier work where a three-subsystem AAF with AFS-A, PS, and ECS is optimized, this may not always be the case. Furthermore, based on the results presented in this thesis, there is a smoothing effect observed in the exergy response curves compared to the fuel-burned response curves to changes in AFS-A geometry. This indicates that the exergy destruction is slightly less sensitive to such changes.
Master of Science

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23

Leifsson, Leifur Thor. "Multidisciplinary Design Optimization of Low-Noise Transport Aircraft". Diss., Virginia Tech, 2005. http://hdl.handle.net/10919/26327.

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The objective of this research is to examine how to design low-noise transport aircraft using Multidisciplinary Design Optimization (MDO). The subject is approached by designing for low-noise both implicitly and explicitly. The explicit design approach involves optimizing an aircraft while explicitly constraining the noise level. An MDO framework capable of optimizing both a cantilever wing and a Strut-Braced-Wing (SBW) aircraft was developed. The framework employs aircraft analysis codes previously developed at the Multidisciplinary Design and Analysis (MAD) Center at Virginia Tech (VT). These codes have been improved here to provide more detailed and realistic analysis. The Aircraft Noise Prediction Program (ANOPP) is used for airframe noise analysis. The objective is to use the MDO framework to design aircraft for low-airframe-noise at the approach conditions and quantify the change in weight and performance with respect to a traditionally designed aircraft. The results show that reducing airframe noise by reducing approach speed alone, will not provide significant noise reduction without a large performance and weight penalty. Therefore, more dramatic changes to the aircraft design are needed to achieve a significant airframe noise reduction. Another study showed that the trailing-edge (TE) flap can be eliminated, as well as all the noise associated with that device, without incurring a significant weight and performance penalty. To achieve approximately 10 EPNdB TE flap noise reduction the flap area was reduced by 82% while the wing reference area was increased by 12.4% and the angle of attack increased from 7.6 degrees to 12.1 degrees to meet the required lift at approach. The wing span increased by approximately 2.2%. Since the flap area is being minimized, the wing weight suffers only about a 2,000 lb penalty. The increase in wing span provides a reduction in induced drag to balance the increased parasite drag due to a lower wing aspect ratio. As a result, the aircraft has been designed to have minimal TE flaps without any significant performance penalty. If noise due to the leading-edge (LE) slats and landing gear are reduced, which is currently being pursued, the elimination of the flap will be very significant as the clean wing noise will be the next 'noise barrier'. Lastly, a comparison showed that SBW aircraft can be designed to be 10% lighter and require 15% less fuel than cantilever wing aircraft. Furthermore, an airframe noise analysis showed that SBW aircraft with short fuselage-mounted landing gear could have similar or potentially a lower airframe noise level than comparable cantilever wing aircraft. The implicit design approach involves selecting a configuration that supports a low-noise operation, and optimizing for performance. A Blended-Wing-Body (BWB) transport aircraft has the potential for significant reduction in environmental emissions and noise compared to a conventional transport aircraft. A BWB with distributed propulsion was selected as the configuration for the implicit low-noise design in this research. An MDO framework previously developed at the MAD Center at Virginia Tech has been refined to give more accurate and realistic aircraft designs. To study the effects of distributed propulsion, two different BWB configurations were optimized. A conventional propulsion BWB with four pylon mounted engines and two versions of a distributed propulsion BWB with eight boundary layer ingestion inlet engines. A 'conservative' distributed propulsion BWB design with a 20% duct weight factor and a 95% duct efficiency, and an 'optimistic' distributed propulsion BWB design with a 10% duct weight factor and a 97% duct efficiency were studied. The results show that 65% of the possible savings due to 'filling in' the wake are required for the 'optimistic' distributed propulsion BWB design to have comparable $TOGW$ as the conventional propulsion BWB, and 100% savings are required for the 'conservative' design. Therefore, considering weight alone, this may not be an attractive concept. Although a significant weight penalty is associated with the distributed propulsion system presented in this study, other characteristics need to be considered when evaluating the overall effects. Potential benefits of distributed propulsion are, for example, reduced propulsion system noise, improved safety due to engine redundancy, a less critical engine-out condition, gust load/flutter alleviation, and increased affordability due to smaller, easily-interchangeable engines.
Ph. D.

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24

Zhu, Yan. "Longitudinal control laws design for a flying wing aircraft". Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7423.

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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.

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25

Jemitola, Paul Olugbeji. "Conceptual design and optimization methodology for box wing aircraft". Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7938.

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A conceptual design optimization methodology was developed for a medium range box wing aircraft. A baseline conventional cantilever wing aircraft designed for the same mis- sion and payload was also optimized alongside a baseline box wing aircraft. An empirical formula for the mass estimation of the fore and aft wings of the box wing aircraft was derived by relating conventional cantilever wings to box wing aircraft wings. The results indicate that the fore and aft wings would use the same correction coe cient and that the aft wing would be lighter than the fore wing on the medium range box wing aircraft because of reduced sweep. As part of the methodology, a computational study was performed to analyze di erent wing/tip n xities using a statically loaded idealized box wing con guration. The analy- ses determined the best joint xity by comparing the stress distributions in nite element torsion box models in addition to aerodynamic requirements. The analyses indicates that the rigid joint is the most suitable. Studies were also performed to investigate the structural implications of changing only the tip n inclinations on the box wing aircraft. Tip n inclination refers to the angle the tip n makes to the vertical body axis of the aircraft. No signi cant variations in wing structural design drivers as a function of tip n inclination were observed. Stochastic and deterministic optimization routines were performed on the baseline box wing aircraft using the methodology developed where the variables were wing area, av- erage thickness to chord ratio and sweep angle. The conventional aircraft design showed similar performance and characteristics to the equivalent in-service aircraft thereby pro- viding some validation to the methodology and the results for the box wing aircraft. Longitudinal stability investigations showed that the extra fuel capacity of the box wing in the ns could be used to reduce trim drag. The short period oscillation of the conventional cantilever wing aircraft was found to be satisfactory but the box wing aircraft was found to be unacceptable hence requiring stability augmentation systems. The eld and ight performance of the box wing showed to be better than the conventional cantilever wing aircraft. Overall, the economic advantages of the box wing aircraft over the conventional cantilever wing aircraft improve with increase in fuel price making the box wing a worthy replacement for the conventional cantilever wing aircraft.

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26

Kim, Kiun. "Nonlinear aeroelastic analysis of aircraft wing-with-store configurations". Diss., Texas A&M University, 2004. http://hdl.handle.net/1969.1/361.

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The author examines nonlinear aeroelastic responses of air vehicle systems. Herein, the governing equations for a cantilevered configuration are developed and the methods of analysis are explored. Based on the developed nonlinear bending-bending-torsion equations, internal resonance, which is possible in future air vehicles, and the possible cause of limit cycle oscillations of aircraft wings with stores are investigated. The nonlinear equations have three types of nonlinearities caused by wing flexibility, store geometry and aerodynamic stall, and retain up to third-order nonlinear terms. The internal resonance conditions are examined by the Method of Multiple Scales and demonstrated by time simulations. The effect of velocity change for various physical parameters and stiffness ratio is investigated through bifurcation diagrams derived from Poinar´e maps. The dominant factor causing limit cycle oscillations is the stiffness ratio between in-plane and out-of-plane motion.

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27

Moodie, Simon James. "The hydraukic shock analysis of aircraft wing box structures". Thesis, Imperial College London, 2010. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.530455.

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28

Ericsson, Max. "Simulating Bird Strike on Aircraft Composite Wing Leading Edge". Thesis, KTH, Hållfasthetslära (Inst.), 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-103783.

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In this master thesis project the possibility to model the response of a wing when subjected to bird strike using finite elements is analyzed. Since this transient event lasts only a few milliseconds the used solution method is explicit time integration. The wing is manufactured using carbon fiber laminate. Carbon fiber laminates have orthotropic material properties with different stiffness in different directions. Accordingly, there are damage mechanisms not considered when using metal that have to be modeled when using composites. One of these damage mechanisms is delamination which occurs when cured layers inside a component become separated. To simulate this phenomenon, multiple layers of shell elements with contact in between are used as a representation of the interface where a component is likely to delaminate. By comparing experimental and simulated results the model of delamination is verified and the influence of different parameters on the results is investigated. Furthermore, studies show that modeling delamination layers in each possible layer of a composite stack is not optimal due to the fact that the global stiffness of the laminate is decreased as more layers are modeled. However, multiple layers are needed in order to mitigate the spreading of delamination and obtain realistic delaminated zones. As the laminates are comprised of carbon fiber and epoxy sheets it is of importance to include damage mechanisms inside each individual sheet. Accordingly, a composite material model built into the software is used which considers tensile and compressive stress in fiber and epoxy. The strength limits are then set according to experimental test data. The bird is modeled using a mesh free technique called Smooth Particle Hydrodynamics using a material model with properties similar to a fluid. The internal pressure of the bird model is linked to the change in volume with an Equation of State. By examining the bird models behavior compared to experimental results it is determined to have a realistic impact on structures. A model of the leading edge is then subjected to bird strike according to European standards. The wing skin is penetrated indicating that reinforcements might be needed in order to protect valuable components inside the wing structure such as the fuel tank. However, the results are not completely accurate due to the fact that there is little experimental data available regarding soft body penetration of composite laminates. As a consequence, the simulation cannot be confirmed against real experimental results and further investigations are required in order to have confidence in modeling such events. Furthermore, the delamination due to the bird strike essentially spreads across the whole model. Since only one layer of delamination is included the spread is most likely overestimated.

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29

Wright, Steven Roland. "A wing rock model for the F-14A aircraft". Thesis, Monterey, California. Naval Postgraduate School, 1992. http://hdl.handle.net/10945/38532.

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"Submitted in partial fulfillment of the requirements for the degree of Aeronautical and Astronautical Engineer from the Naval Postgraduate School June 1992"
Approved for public release; distribution is unlimited.
An investigation of inertial coupling and its contribution to wing rock in the F-14A aircraft has been conducted. Wind tunnel data was used to obtain the stability parameters for angles of attack from zero to 25 degrees, after which linear and nonlinear analyses of the equations of motion were completed. The linearized analysis of the uncoupled longitudinal and lateral-directiondl equations was included to provide a baseline for comparison with the fully coupled, nonlinear equations. In both cases, the equations of motion were solved numerically and time history traces produced to illustrate aircraft response. Results indicate that a stable short period mode can feed damping energy into an unstable dutch roll mode via the coupling of the equations to produce a stable limit cycle very similar to those experienced in the aircraft. Numerous suggestions for follow on research are presented.

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30

Meo, Michele. "Application of welding to a large civil aircraft wing". Thesis, Cranfield University, 1999. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.323959.

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31

Agenbag, Daniël Sarel. "Longitudinal handling characteristics of a tailless gull-wing aircraft". Pretoria : [s.n.], 2008. http://upetd.up.ac.za/thesis/available/etd-09182008-132941.

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32

Herrera, Jason (Jason Richard). "Evaluation of control systems for automated aircraft wing manufacturing". Thesis, Massachusetts Institute of Technology, 2013. http://hdl.handle.net/1721.1/82484.

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Thesis (M.B.A.)--Massachusetts Institute of Technology, Sloan School of Management; and, (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics; in conjunction with the Leaders for Global Operations Program at MIT, 2013.
This electronic version was submitted and approved by the author's academic department as part of an electronic thesis pilot project. The certified thesis is available in the Institute Archives and Special Collections.
Cataloged from department-submitted PDF version of thesis
Includes bibliographical references (p. 62-64).
The Boeing Company is looking to bring aircraft manufacturing technology into the 21st century. As part of this process, several projects have been started to develop the technologies required to achieve Boeing's vision for the future of aircraft manufacturing. To date, much of this work has focused on hardware, including robotic and other automation technologies. However, in order to use this hardware, a significant effort must also be made in the area of factory control and coordination. This thesis advances knowledge in this area by evaluating the suitability of different control system approaches for aircraft wing box assembly. First, general classes of control systems are discussed and several criteria are proposed for evaluating their performance in an aircraft manufacturing environment. The current wing box assembly process is then examined in order to develop simplified but representative task networks to which various algorithms can be applied. The Tercio algorithm, developed at MIT, is used to generate schedules for several problem structures of interest in order to characterize the algorithm's performance in this context. The Tercio algorithm is then benchmarked against the Aurora scheduling tool, showing that Tercio can generate more efficient schedules than Aurora, but at the cost of increased computation time. Next, management considerations with respect to product design, manufacturing technology development, and implementation associated with advanced manufacturing technologies are discussed. Finally, recommendations are provided for how Boeing can accelerate the development of useful and practical advanced, automated manufacturing systems.
by Jason Herrera.
S.M.
M.B.A.

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33

Sobolic, Frantisek Michal. "Agile flight control techniques for a fixed-wing aircraft". Thesis, Massachusetts Institute of Technology, 2009. http://hdl.handle.net/1721.1/51640.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2009.
Includes bibliographical references (p. 91-94).
As unmanned aerial vehicles (UAVs) become more involved in challenging mission objectives, the need for agility controlled flight becomes more of a necessity. The ability to navigate through constrained environments as well as quickly maneuver to each mission target is essential. Currently, individual vehicles are developed with a particular mission objective, whether it be persistent surveillance or fly-by reconnaissance. Fixed-wing vehicles with a high thrust-to-weight ratio are capable of performing maneuvers such as take-off or perch style landing and switch between hover and conventional flight modes. Agile flight controllers enable a single vehicle to achieve multiple mission objectives. By utilizing the knowledge of the flight dynamics through all flight regimes, nonlinear controllers can be developed that control the aircraft in a single design. This thesis develops a full six-degree-of-freedom model for a fixed-wing propeller-driven aircraft along with methods of control through non conventional flight regimes. In particular, these controllers focus on transitioning into and out of hover to level flight modes. This maneuver poses hardships for conventional linear control architectures because these flights involve regions of the post-stall regime, which is highly nonlinear due to separation of flow over the lifting surfaces. Using Lyapunov back stepping control stability theory as well as quaternion-based control methods, control strategies are developed that stabilize the aircraft through these flight regimes without the need to switch control schemes. The effectiveness of each control strategy is demonstrated in both simulation and flight experiments.
by Frantisek Michal Sobolic.
S.M.

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34

Vavalle, Armando. "Response surface aerodynamic optimisation for blended wing body aircraft". Thesis, Cranfield University, 2005. http://dspace.lib.cranfield.ac.uk/handle/1826/11015.

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This study is concerned with a methodology for the aerodynamic analysis and preliminary design of a novel configuration for high subsonic civil transport, based on the flying wing concept, known a Blended Wing Body (BWB). A response surface based optimisation method is developed, enabling the designer to monitor the effect of shape modification on the controllability of the aircraft in both longitudinal and lateral/directional motion and on the Wing structural weight, while maximising the aerodynamic efficiency. The design aspects considered included high- speed aerodynamics, flight static-stability and trim characteristics. The response surface Scheme employs a space filling design of experiment technique to build least square fitting quadratic polynomials, used in place of the original computational modules in a gradient based search. A optimisation test indicated that the present method is more effective in leading the design near to the global optimum as opposed to a conventional gradient method with direct search, despite that the constructed approximation may not represent accurately the actual surface. With this system, multiple constrained optimisation problems are successfully solved in the favourable case of smooth objective/constraint function. Where these functions may exhibit high non-linear trends, an iterative response surface method refining both approximation and bounds of the design space is proposed. The capabilities of such a technique are shown for transonic aerofoil optimisation problems, demonstrating that the proposed method is more efficient and more effective than some other state-of- the-art methods. As a result of these studies, the aerodynamic efficiency of a large capacity BWB configuration has been considerably improved by re-designing the external shape to generate a spanwise loading intermediate between triangular and elliptic. The longitudinal static stability analysis revealed that the aircraft is stable except at low- weights with zero-payload. The lateral/directional analyses showed that the aircraft is stable in roll, but unstable in yaw. Despite that the winglets are found to stabilise the aircraft, it is directionally unstable without additional vertical stabilisers. I

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35

Schoř, Pavel. "Load State of an Aircraft with an Elastic Wing". Doctoral thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2018. http://www.nusl.cz/ntk/nusl-383528.

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V této práci je navržena metoda výpočtu zatížení letadla s netuhým křídlem, založená na spojení panelové metody prvního řádu dle Katz and Plotkin, Low-Speed Aerodynamics, 2001 s metodou stukturální analýzy dle Píštěk et al., Pevnost a životnost letadel I, 1988 a Lebofsky,Numerically Generated Tangent Stiffness Matrices for Geometrically Non-Linear Struc- tures, 2013. Panelová metoda poskytuje přasná data pro výpočet zatížení křídla od vzdušných sil za předpokladu že lze dané proudění aproximovat po- mocí potenciálního proudění, Narozdíl metod založených na interakci s CFD metodami lze navrženou metodu používat i na bežném počítači.

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36

Locatelli, Davide. "Optimization of Supersonic Aircraft Wing-Box using Curvilinear SpaRibs". Diss., Virginia Tech, 2012. http://hdl.handle.net/10919/26345.

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This dissertation investigates the advantages of using curvilinear spars and ribs, termed SpaRibs, to design supersonic aircraft wing-box in comparison to the use of classic design concepts that employ straight spars and ribs. The intent is to achieve a more efficient load-bearing mechanism and to passively control aeorelastic behavior of the structure under the flight loads. The use of SpaRibs broadens the design space and allows for the natural frequencies and natural mode shape tailoring. The SpaRibs concept is implemented in a new MATLAB-based optimization framework referred to as EBF3SSWingOpt. This framework interfaces different analysis software to perform the tasks required. VisualDOC is used as optimizer; the generation of the SpaRibs geometry and of the structure Finite Element Model (FEM) is performed by MD.PATRAN; MD.NASTRAN is utilized to compute the weight of the structure, the linear static stress analysis and the linear buckling analysis required for the calculation of the response functions. EBF3SSWingOpt optimization scheme performs both the sizing and the shaping of the internal structural elements. Two methods are compared while optimizing the wing-box; a One-Step method in which sizing and topology optimization are carried out simultaneously and a Two-Step method, in which the sizing and topology optimization are carried out separately but in an iterative way. The optimization problem statements for the One-Step and the Two-Step methodologies are presented. Three methods to define the shape of the SpaRibs parametrically are described: (1) the Bounding Box and Base Curves method defines the shape of the SpaRibs based on the shape of two curves called Base Curves which are positioned into the Bounding Box, a rectangular region defined on the plane z=0 and containing the projection of the wing plan-form onto the same plane; (2) the Linked Shape method defines the shape of a set of SpaRibs in a one by one square domain of the natural space. The set of curves is subsequently transformed in the physical space for creating the wing structure geometry layout. The shape of each curve of each set is unique however, mathematical relations link their curvature in an effort to reduce the number of design variables; and (3) the Independent Shape parameterization is similar to the Linked Shape parameterization however, the shape of each curve is unique. The framework and parameterization methods described are applied to optimize different types of wing structures. Following results are presented and discussed: (1) a rectangular wing-box subjected to a chord-wise linearly varying load, optimized using SpaRibs parameterized with Bounding-Box and Base Curves method; (2) a rectangular wing-box subjected to a chord-wise linearly varying load, optimized using SpaRibs parameterized with Linked Shape method; (3) a generic fighter wing subjected to uniform distributed pressure load, optimized using SpaRibs parameterized with Bounding-Box and Base Curves method; (4) a general business jet wing subjected to pull-up maneuver loads computed using ZESt (ZONA Technology Inc. Steady Euler equations solver), optimized using SpaRibs parameterized with Independent Shape method; (5) a preliminary application of the Linked Shape parameterization to place SpaRibs into a high speed commercial transport aircraft wing-box characterized by high geometry layout complexity; and (6) an optimization of panels subjected to axial and shear loads using curvilinear stiffeners and grids of curvilinear stiffeners. The results for the optimization of the rectangular wing-box show 36.8% weight reduction from the baseline, when the Bounding Box and Base Curves parameterization is applied and the Two-Step framework is implemented. For the same structure the weight reduction amounts to 46.7% when the Linked Shape parameterization and the Two-Step framework are used. Similar results are obtained for the generic fighter wing-box structure. In this case, the weight saving is about 20%. Bounding Box and Base Curves parameterization and Two-Step framework are used. Finally, the weight reduction for the general business jet wing-box structure amounts to 17% of the baseline weight. In this case, the computation is carried out using the Independent Shape parameterization and the Two-Step framework. In general, the Two-Step optimization framework finds better optimal structure configurations as compared to the One-Step optimization framework. However, the computational time required to find to optimum with the Two-Step optimization is larger when a small number of particles are used in the particle swarm optimization method. For larger number of particles, the computational time for the two methods is comparable. Finally for very large number of particles the Two-Step optimization requires less computational time. It is also important to notice how the Two-Step framework consistently leads to a better optimum than the One-Step framework, for the same number of particles.
Ph. D.

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37

Grasmeyer, Joel M. III. "Multidisciplinary Design Optimization of a Strut-Braced Wing Aircraft". Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/36729.

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The objective of this study is to use Multidisciplinary Design Optimization (MDO) to investigate the use of truss-braced wing concepts in concert with other advanced technologies to obtain a significant improvement in the performance of transonic transport aircraft. The truss topology introduces several opportunities. A higher aspect ratio and decreased wing thickness can be achieved without an increase in wing weight relative to a cantilever wing. The reduction in thickness allows the wing sweep to be reduced without incurring a transonic wave drag penalty. The reduced wing sweep allows a larger percentage of the wing area to achieve natural laminar flow. Additionally, tip-mounted engines can be used to reduce the induced drag. The MDO approach helps the designer achieve the best technology integration by making optimum trades between competing physical effects in the design space. To perform this study, a suite of approximate analysis tools was assembled into a complete, conceptual-level MDO code. A typical mission profile of the Boeing 777-200IGW was chosen as the design mission profile. This transport carries 305 passengers in mixed class seating at a cruise Mach number of 0.85 over a range of 7,380 nmi. Several single-strut configurations were optimized for minimum takeoff gross weight, using eighteen design variables and seven constraints. The best single-strut configuration shows a 15% savings in takeoff gross weight, 29% savings in fuel weight, 28% increase in L/D, and a 41% increase in seat-miles per gallon relative to a comparable cantilever wing configuration. In addition to the MDO work, we have proposed some innovative, unconventional arch-braced and ellipse-braced concepts. A plastic solid model of one of the novel configurations was created using the I-DEAS solid modeling software and rapid prototyping hardware.
Master of Science

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38

Türkoglu, Ercüment. "Application of robust control to a rotary-wing aircraft". Thesis, University of Leicester, 2007. http://hdl.handle.net/2381/30248.

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The thesis is concerned with the application of robust controller synthesis and analysis tools to a rotary-wing aircraft: the Bell 205 teetering-rotor helicopter. The Tioo loop-shaping approach is central to the work and two main issues concerned with its application will be considered. Firstly, the construction of diagonal (structured) and non- diagonal (unstructured) weighting functions will be considered. Secondly, the analysis of the implications of different weighting function structures in the controller implementation. A two stage cross-comparative analysis of a series of 1 Dof (Degree of Freedom) and 2 Dof controllers synthesized with both diagonal and non-diagonal weights using the Hqo loop- shaping technique will be presented for square and non-square multi input multi output, unstable, non-minimum phase and ill-conditioned models of the helicopter. Handling qualities of each control law augmented system will be assessed quantitatively and qualitatively. A quantitative analysis, in view of the specifications in ADS-33E, will be given based on a combination of flight data from in-flight tested controllers and, desk-top simula tions run on a fully augmented 12 Dof nonlinear helicopter model provided by QinetiQ, UK. A qualitative analysis will be given based on the pilot comments compiled (in view of the Cooper-Harper handling qualities rating scale) from the evaluated in-flight control laws.

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39

Dippold, Vance Fredrick III. "Numerical Assessment of the Performance of Jet-Wing Distributed Propulsion on Blended-Wing-Body Aircraft". Thesis, Virginia Tech, 2003. http://hdl.handle.net/10919/34878.

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Conventional airliners use two to four engines in a Cayley-type arrangement to provide thrust, and the thrust from these engines is typically concentrated right behind the engine. Distributed propulsion is the idea of redistributing the thrust across most, or all, of the wingspan of an aircraft. This can be accomplished by using several large engines and using a duct to spread out the exhaust flow to form a jet-wing or by using many small engines spaced along the span of the wing. Jet-wing distributed propulsion was originally suggested by Kuchemann as a way to improve propulsive efficiency. In addition, one can envision a jet-wing with deflected jets replacing flaps and slats and the associated noise.

The purpose of this study was to assess the performance benefits of jet-wing distributed propulsion. The Reynolds-averaged, finite-volume, Navier-Stokes code GASP was used to perform parametric computational fluid dynamics (CFD) analyses on two-dimensional jet-wing models. The jet-wing was modeled by applying velocity and density boundary conditions on the trailing edges of blunt trailing edge airfoils such that the vehicle was self-propelled. As this work was part of a Blended-Wing-Body (BWB) distributed propulsion multidisciplinary optimization (MDO) study, two airfoils of different thickness were modeled at BWB cruise conditions. One airfoil, representative of an outboard BWB wing section, was 11% thick. The other airfoil, representative of an inboard BWB wing section, was 18% thick. Furthermore, in an attempt to increase the propulsive efficiency, the trailing edge thickness of the 11% thick airfoil was doubled in size. The studies show that jet-wing distributed propulsion can be used to obtain propulsive efficiencies on the order of turbofan engine aircraft. If the trailing edge thickness is expanded, then jet-wing distributed propulsion can give improved propulsive efficiency. However, expanding the trailing edge must be done with care, as there is a drag penalty. Jet-wing studies were also performed at lower Reynolds numbers, typical of UAV-sized aircraft, and they showed reduced propulsive efficiency performance. At the lower Reynolds number, it was found that the lift, drag, and pitching moment coefficients varied nearly linearly for small jet-flap deflection angles.

Master of Science

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40

Adegbindin, Moustaine Kolawole Agnide. "Control Power Optimization using Artificial Intelligence for Forward Swept Wing and Hybrid Wing Body Aircraft". Thesis, Virginia Tech, 2017. http://hdl.handle.net/10919/74950.

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Many futuristic aircraft such as the Hybrid Wing Body have numerous control surfaces that can result in large hinge moments, high actuation power demands, and large actuator forces/moments. Also, there is no unique relationship between control inputs and the aircraft response. Distinct sets of control surface deflections may result in the same aircraft response, but with large differences in actuation power. An Artificial Neural Network and a Genetic Algorithm were used here for the control allocation optimization problem of a Hybrid Wing Body to minimize the Sum of Absolute Values of Hinge Moments for a 2.5-G pull-up maneuver. To test the versatility of the same optimization process for different aircraft configurations, the present work also investigates its application on the Forward Swept Wing aircraft. A method to improve the robustness of the process is also presented. Constraints on the load factor and longitudinal pitch rate were added to the optimization to preserve the trim constraints on the control deflections. Another method was developed using stability derivatives. This new method provided better results, and the computational time was reduced by two orders of magnitude. A hybrid scheme combining both methods was also developed to provide a real-time estimate of the optimum control deflection schedules to trim the airplane and minimize the actuation power for changing flight conditions (Mach number, altitude and load factor) in a pull-up maneuver. Finally, the stability derivatives method and the hybrid scheme were applied for an antisymmetric, steady roll maneuver.
Master of Science

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41

Venter, Jacob. "Development of an experimental tilt-wing VTOL unmanned aerial vehicle". Thesis, Link to the online version, 2006. http://hdl.handle.net/10019/225.

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42

Macci, S. H. M. "Structural and mechanical feasibility study of a variable camber wing (VCW) for a transport aircraft". Thesis, Cranfield University, 1992. http://hdl.handle.net/1826/4168.

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Aerodynamic investigations have shown' that variable camber wings (VCW) for transport aircraft have considerable potential in terms of improving aircraft performance and enhancing their operational flexibility. In order to justify these benefits it is essential that the camber varying system is structurally and mechanically feasible. This research examined the feasibility of providing variable camber to two supercritical aerofoil sections of different'characteristics. The unique method of camber vaTiation was applied by rotating the forward and aft regions of the aerofoil on a circular arc and keeping the surface continuous and matching at their attachment to the main wing box. The change in camber thus increased the chord due to translational motion of the aforementioned regions. The geometries required for varying the forward camber by this method presented formidable design difficulties and no immediate solutions could be found. As a result, an alternative geometry was devised which accepts camber by simply drooping the nose region. A novel idea was developed for aft camber variation, which is considered to be universal for all supercritical aerofoil sections. The system utilises a tracking mechanism which guides a trailing edge element on a continuous arc. Surface continuity is provided by a flexible skin on the upper side and a spring loaded hinged panel on the under side. The flexible skin remains attached to the trailing edge element through a series of roller link arrangement which locate the skin in a separate guide rail. The large moment arm and therefore the increased torsional loads created due to the translational motion of the trailing edge element necessitated investigation of alternative deployment geometries. As a result two additional geometries were schemed. One had reduced radius of rotation and therefore reduced extension, while the other changed camber by drooping the aft region without any chordal extension. Since there was no aerodynamic evidence on the possible benefits offered by these geometries it was decide to postpone them until such information was available. Some detailed aspects of the proposed concept for aft camber variation were considered by applying the system to a modem transport aircraft wing. This resulted in a design which is practically feasible. Justification of this concept was made by designing and testing a half scale structural model of one trailing edge segment. Three dimensional (3-D) geometric investigation showed that the camber-varying elements ride on a frustum of a cone and therefore their deployment is skewed to the line of flight. The 3-D geometric implications of variable camber clearly suggested that the camber variation by rotation on a circular arc, on a tapered wing can be possible if the rotating element is made to flex and twist or it utilises a pin jointed arrangement. To provide the necessary flexibility to the trailing edge element, its structural box best be made from fibre reinforced plastic material. The deployment of the trailing edge element on the structural m(; del was made possible by designing it in laminated wood. Comparison of the proposed variable camber system with a conventional single slotted flap arrangement suggests that the two systems could be equally complex but the variable camber could be slightly heavier., Further systems investigations are required to quantify overall aerodynamic, mass, and cost implications of the use of VCW on transport aircraft.

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43

Matos, Catherine Anne Moseley. "Download reduction on a wing-rotor configuation". Diss., Georgia Institute of Technology, 2001. http://hdl.handle.net/1853/12058.

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44

Naghshineh-Pour, Amir H. "Structural Optimization and Design of a Strut-Braced Wing Aircraft". Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/36142.

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A significant improvement can be achieved in the performance of transonic transport aircraft using Multidisciplinary Design Optimization (MDO) by implementing truss-braced wing concepts in combination with other advanced technologies and novel design innovations. A considerable reduction in drag can be obtained by using a high aspect ratio wing with thin airfoil sections and tip-mounted engines. However, such wing structures could suffer from a significant weight penalty. Thus, the use of an external strut or a truss bracing is promising for weight reduction. Due to the unconventional nature of the proposed concept, commonly available wing weight equations for transport aircraft will not be sufficiently accurate. Hence, a bending material weight calculation procedure was developed to take into account the influence of the strut upon the wing weight, and this was coupled to the Flight Optimization System (FLOPS) for total wing weight estimation. The wing bending material weight for single-strut configurations is estimated by modeling the wing structure as an idealized double-plate model using a piecewise linear load method. Two maneuver load conditions 2.5g and -1.0g factor of safety of 1.5 and a 2.0g taxi bump are considered as the critical load conditions to determine the wing bending material weight. From preliminary analyses, the buckling of the strut under the -1.0g load condition proved to be the critical structural challenge. To address this issue, an innovative design strategy introduces a telescoping sleeve mechanism to allow the strut to be inactive during negative g maneuvers and active during positive g maneuvers. Also, more wing weight reduction is obtained by optimizing the strut force, a strut offset length, and the wing-strut junction location. The best configuration shows a 9.2% savings in takeoff gross weight, an 18.2% savings in wing weight and a 15.4% savings in fuel weight compared to a cantilever wing counterpart.
Master of Science

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45

de, Castro Helena V. "Flying and handling qualities of a fly-by-wire blended-wing-body civil transport aircraft". Thesis, Cranfield University, 2003. http://hdl.handle.net/1826/119.

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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.

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46

Landolfo, Giuseppe. "Aerodynamic and Structural Design of a Small Nonplanar Wing UAV". University of Dayton / OhioLINK, 2008. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1262089704.

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47

Harris, Turner John. "CONSTRAINED VOLUME PACKING OF DEPLOYABLE WINGS FOR UNMANNED AIRCRAFT". UKnowledge, 2011. http://uknowledge.uky.edu/gradschool_theses/129.

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UAVs are becoming an accepted tool for sensing. The benefits of deployable wings allow smaller transportation enclosures such as soldier back packs up to large rocket launched extraterrestrial UAVs. The packing of soft inflatable wings and Hybrid inflatable with rigid section wings is being studied at the University of Kentucky. Rigid wings are volume limited while inflatable wings are mass limited. The expected optimal wing design is a hybrid approach. Previous wing designs have been packed into different configurations in an attempt to determine the optimal stowed configurations. A comparison of rigid, hybrid, and inflatable wings will be presented. Also a method for simulating optimally packed wings with respect to geometric constraints will be presented. A code has been written to study soft wing packing and verified the soft wing packing results. This code can be used during initial wing design to help predict wing size and packing configurations. In this thesis, an over view of the packing configurations as well as packing observations will be covered such , packing inefficiencies, wing mounting limits, long term storage, and scaling of packing.

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48

Hu, Xiaowei. "Analysis and Experiment of an Ultra-light Flapping Wing Aircraft". Thesis, Cranfield University, 2013. http://dspace.lib.cranfield.ac.uk/handle/1826/8466.

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II Inspired by flying animals in nature especially birds, human has designed and attempted to achieve man-powered flapping wing aircraft in very early aviation history. Limited by the understanding of the aerodynamic theory and materials in practise, the bird-like aircraft remains as a dream and ambition for over a contrary. As the relevant knowledge and technology are fast developing in the last decade, the research topic becomes attractive again with encouraging results from a few full scale aircraft flight tests. Although it is suspected that a manned scale flapping wing may not be as efficient as fixed wing, the unique advantages of high manoeuvrability and short take-off and landing capability will keep flapping wing as one of the most potential type of personal and aerobatic aircraft in the future market. The aim of this project is to investigate into the feasibility and development of a bio-inspired bird-like man-powered ultra-light flapping wing aircraft (ULFWA). The project is based on analytical and experimental study of a scaled model taking an existing hang glider as the baseline airframe. Based on the characteristics of flying animals in nature and manmade hang glider properties, this thesis focuses its study on evaluating the feasibility and analysis of primarily a human powered aircraft. For this purpose, there are four main features as guidance in the ULFWA design. Firstly the flapping frequency was limited to below 2Hz. Secondly the hang glider airframe was adapted with a simple flapping mechanism design. Thirdly the flapping wing stroke and kinematics has been kept with the simplest and resonant movement to achieve high mechanical efficiency. Finally the wing structure has flexible rib of chord wise unsymmetrical bending stiffness to offset the aerodynamic lift loss in upstroke. An engine powered mechanism design was also studied as additional option of the ULFWA. The initial design and aerodynamic calculation of the ULFWA was based on the hang glider data including dimensions, MTOW (226 kg) and cruising speed. The unsteady aerodynamic lift and thrust forces were calculated based on Theodorsen’s theory and unsteady panel method in 2D and extended to 3D using strip theory. A set of optimal flapping kinematic parameters such as amplitude and combination of the heaving and pitching motion of the 2D wing section were determined by calculation and comparison in the limited range. Considering the maximum power and lag motion that human could achieve, the flapping frequency in the ULFWA design is limited to 1Hz. This slow motion leads to a much lower propulsive efficiency in terms of the optimum Strouhal Number (St=0.2-0.4), which was used as the design reference. Mechanism and structure design with inertia force calculation was then completed based on the kinematics. This led to the evaluation of power requirement, which was divided into two components, drag and inertia forces. The results show that the ULFWA needs minimum 2452.25W (equals to 3.29Bhp) to maintain sustainable cruise flight. In order to demonstrate the ULFWA flapping mechanism and structure design, a 1:10 scaled model with two pairs of wings of different stiffness were built for testing and measurement. Two servomotors were used as to simulate human power actuation. With this model, simplified structure and one of mechanism designs was shown. Four experiments were carried out to measure the model’s lift and thrust force. Because of the limited response of the servo motors, the maximum flapping frequency achieved is only 0.75 Hz in the specified flapping amplitude which is close to reality and has improvement margin. By reducing the flapping amplitude, the frequency can be increased to gain higher thrust. Although it is fund that the result from scaled model test is a little lower than theoretical result, it has demonstrated the feasibility and potential of human powered flapping wings aircraft.

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49

Chaffee, Neil T. "Expanding fixed-wing aircraft capability in US Army aviation operations /". Fort Leavenworth, Kan. : Ft. Belvior, VA : Alexandria, Va. : U.S. Army Command and General Staff College ; Available to the public through the Defense Technical Information Center ; National Technical Information Service [distributor], 2009. http://www.dtic.mil/dtic/.

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Thesis (M.S. of Military Art and Science General Studies)--U.S. Army Command and General Staff College, June 2009.
"June 2009." Thesis advisor: David W. Christie. Performed by the U.S. Army Command and General Staff College, Fort Leavenworth, Kansas. "Presented to the faculty of the U.S. Army Command and General Staff College in partial fulfillment of the requirements for the degree Master of Military Art and Science General Studies from the U.S. Army Command and General Staff College, June 2009."--P. [i]. Includes bibliographical references. Also available online from the Combined Arms Research Library (CARL) at the U.S. Army Command and General Staff College and the DTIC Online Web site.

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50

Ng, Leo Wai-Tsun. "Design and acoustic shielding prediction of hybrid wing-body aircraft". Thesis, Massachusetts Institute of Technology, 2009. http://hdl.handle.net/1721.1/51635.

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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2009.
Includes bibliographical references (p. 99-101).
Recent research and developmental efforts in aircraft design have focused on the growing concerns about the environment impact of aviation and the rising costs of fuel. Under NASA's N+2 subsonics fixed-wing project, hybrid-wing-body (HWB) aircraft are investigated with the goal to meet the N+2 noise, fuel burn, and emissions requirements. As part of the N+2 program, this thesis is focused on the design and assessment of an HWB aircraft and the development of a prediction method for turbomachinery noise shielding. Based on MIT's previous experience in the Silent Aircraft Initiative, the SAX-40 aircraft concept was further developed into the N+2 HWB aircraft. The design effort resulted in two aircraft configurations: the N2A aircraft with conventional podded engines, and the N2B aircraft with a distributed propulsion system embedded in the airframe. The initial performance assessment shows that the N2A and the N2B aircraft can both meet the N+2 fuel burn goal and that the N2A aircraft is 5.7 EPNdB short of the noise goal. Also, the assessment revealed that current noise prediction methods cannot model the advanced propulsion system of the N2B aircraft, requiring the development of noise assessment tools for advanced engine-airframe configurations. NASA's Aircraft Noise Prediction Program employs the barrier shielding method to predict the airframe shielding of engine noise. However, it is an empirical formulation for straight edges and thus it is not appropriate for the planform shape of an HWB aircraft.
(cont.) At the same time, high fidelity methods such as boundary element methods and ray tracing methods are too computationally expensive if used in the early aircraft design and assessment stage. A compromise is the previously formulated diffraction integral concept based on the Maggi-Rubinowicz representation of Kirchhoff's diffraction theory. The diffraction integral method was implemented and applied to the N2A and the N2B aircraft. A noise reduction of over 20 dB in OASPL due to airframe shielding was predicted, demonstrating the shielding benefit of the HWB configuration. This shielding method is shown to be applicable to any aircraft configuration and planform geometry. The contributions of this thesis are the design of an HWB aircraft to be used as a platform for the development and evaluation of advanced analysis methods. In addition, a fast and improved-fidelity method for noise shielding prediction was developed, applicable to conventional and advanced airframe configurations such as, for example, the N2A and the N2B HWB aircraft.
by Leo Wai-Tsun Ng.
S.M.

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