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1
Karlsson, Albin, and Anton Lomaeus. "Transport Aircraft Conceptual Design." Thesis, KTH, Skolan för teknikvetenskap (SCI), 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-210778.
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A conceptual design for a transport aircraft has been created, tailored for human-itarian missions along the equator with its home base in the European Union while optimizing for fuel eciency and speed. An initial estimate of the empty weight was made using historical data and Breguet equations, based on a required payload of 60 tonnes and range of 5 500 nautical miles. A constraint diagram consisting of require-ments for stall speed, takeo distance, climb rate and landing distance was used to determine wing loading and thrust to weight ratio, resulting in a main wing area of 387m 2 and thrust to weight ratio of 0:224, for which two Rolls Royce Trent 1000-H engines were selected. A high aspect ratio wing was designed with blended winglets to optimize against lift induced drag. Wing placement and tail volume were decided by iterative calculations, resulting in a centre of lift located aft of the centre of gravity during all stages of the mission. The resulting aircraft model has a high wing with a span of 62 m, length of 49m with a takeo gross weight of 221 tonnes, of which 83 tonnes are fuel.
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Carlson, Jesper, and Diyar Jazrawi. "Conceptual Design of a Transport Aircraft." Thesis, KTH, Skolan för teknikvetenskap (SCI), 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-211549.
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When flooding or earthquakes hits a country the population in the area could suffer a lot and are in a big need of help quickly. In these situations heavy transport aircraft are used around the world to help and support the victims in the area by delivering supplies. In our operational mission scenario the country Papa New guinea has been hit by an earthquake and approximately 5000 people have lost their home and are in need of help. The only problem is that there are no heavy transport airplanes available to fly from EU to this country and return without refuelling. The problem here is that the country is in a big need of help and if an airplane needs to land to refuel a lot of time is wasted. Therefore, in this task we have designed a conceptual transport aircraft that is able to fly from EU to Papa New Guinea to deliver supplies in form of food, aid, water etc. Due to the horrific accident the airplane does not have access to a runway and will have to deliver the supplies by airdrop in parachutes. In this report we will generate a requirements specification, which will state the requirements of the aircraft and be vital for the design. There will be precise estimations and calculations presented and it will include important parameters used in the Design of the aircraft.
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Rabizadeh, Nadja, and Bahar Kasbi. "Conceptual Design of a Transport Aircraft." Thesis, KTH, Skolan för teknikvetenskap (SCI), 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-211556.
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The mission of this project is to conceptually design a transport aircraft. A typical mission for a transport aircraft is to deliver supplies to countries in need of help. The given requirement is that the aircraft should be able to travel from a place within EU to a place along the equator in Africa, deliver the supplies, and return (to the takeoff location) without refueling. The operational mission will be to provide people in need with supplies such as food, water and tents. The aircraft will be able to carry necessities that will be able to provide 5000 persons during a week. Since a landing runway is not available at the destination, the payload will be airdropped in parachutes. First off, the desired requirements are defined, they are either already given or estimated. An analysis of´the mission and the desired performance of the aircraft is made by creating a mission profile. With the help of this a weight estimation is done, most importantly the takeoff weight of the aircraft is estimated. With the takeoff weight known and by the help of the desired performance requirements, a constraint diagram is made. By a constraint analysis the optimal wing loading and thrust-to-weight ratio is found. This makes it possible to choose an appropriate engine and to design the wings so that they are customized for the desired mission. Other parts of the aircraft such as the tail and fuselage are designed, and the center-of gravity of the aircraft is found. Throughout the project, different aerodynamic parameters are changed in order to optimize the aircraft and its performance to make it as adapted as possible to the desired mission.
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Jackson, David Wayne. "Robust aircraft subsystem conceptual architecting." Diss., Georgia Institute of Technology, 2013. http://hdl.handle.net/1853/50202.
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Aircraft subsystems are key components in modern aircraft, the impact and significance of which have been constantly increasing. Furthermore, the architecture selection of these subsystems has overall system-level effects. Despite the significant effects of architecture selections, existing methods for determining the architecture, especially early in design, are similar to the use of traditional point solutions.
Currently, aircraft subsystems are rarely examined during the conceptual design phase, despite the fact that this phase has a significant influence on aircraft cost and performance. For this reason, there is a critical need to examine subsystem architecture trades and investigate the design space during the conceptual design of an aircraft. Traditionally, after the aircraft conceptual design phase, subsystems are developed in a process that begins with the point selection of the architecture, then continues with its development and analysis, and concludes in the detailed development of the subsystems. The choice of the point design of the architecture to be developed can be made using simplified models to explore the design space. This method known as conceptual architecting is explored in this dissertation.
This dissertation also focuses on bringing actuation subsystem architecture trades into conceptual design because of the significant cost impact of this design phase and the interdependence of vehicle sizing with the subsystems impact on the aircraft. The extent of these interdependencies is examined and found to be significant. As a result, this coupling must be captured to enable better informed decision making. A methodology to examine the design space of aircraft subsystem architectures during the conceptual design of aircraft, while incorporating this coupling, is presented herein and applied specifically to actuation architectures.
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Gangadharan, Venkata Krishnan. "Conceptual Design Tool for Aircraft Electrical System." Thesis, Linköpings universitet, Fluida och mekatroniska system, 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-96162.
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The conceptual design stage of an aircraft involves many uncertainties with regard to prediction of weight of systems. The current trend is that electrical systems increasingly replace hydraulic and pneumatic systems in an aircraft. This leads to greater uncertainty in weight, size and power requirement prediction. This work is an attempt at developing a sizing tool that will allow users to estimate the power requirements and weight of electrical systems for a given size of an aircraft specified either by passenger capacity or by aircraft operating empty weight or by maximum take-off weight. As with all predictive tools, the results of this work are based on currently available data, i.e., the specification of existing aircraft. This collection of specification of existing aircrafts would constitute the data library. The accuracy of the result of this work depends greatly on the variety of aircrafts and the level of detail for which the data is available. The tool is made in Microsoft Excel with some codes made in VBA to perform Excel calculations.
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Raymer, Daniel. "Enhancing Aircraft Conceptual Design using Multidisciplinary Optimization." Doctoral thesis, KTH, Aeronautical Engineering, 2002. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3331.
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Research into the improvement of the Aircraft ConceptualDesign process by the application of MultidisciplinaryOptimization (MDO) is presented. Aircraft conceptual designanalysis codes were incorporated into a variety of optimizationmethods including Orthogonal Steepest Descent (full-factorialstepping search), Monte Carlo, a mutation-based EvolutionaryAlgorithm, and three variants of the Genetic Algorithm withnumerous options. These were compared in the optimization offour notional aircraft concepts, namely an advanced multiroleexport fighter, a commercial airliner, a flying-wing UAV, and ageneral aviation twin of novel asymmetric configuration. Tobetter stress the methods, the commercial airliner design wasdeliberately modified for certain case runs to reflect a verypoor initial choice of design parameters including wingloading, sweep, and aspect ratio.
MDO methods were evaluated in terms of their ability to findthe optimal aircraft, as well as total execution time,convergence history, tendencies to get caught in a localoptimum, sensitivity to the actual problem posed, and overallease of programming and operation. In all, more than a millionparametric variations of these aircraft designs were definedand analyzed in the course of this research.
Following this assessment of the optimization methods, theywere used to study the issue of how the computer optimizationroutine modifies the aircraft geometric inputs to the analysismodules as the design is parametrically changed. Since thiswill ultimately drive the final result obtained, this subjectdeserves serious attention. To investigate this subject,procedures for automated redesign which are suitable foraircraft conceptual design MDO were postulated, programmed, andevaluated as to their impact on optimization results for thesample aircraft and on the realism of the computer-defined"optimum" aircraft. (These are sometimes called vehicle scalinglaws, but should not be confused with aircraft sizing, alsocalled scaling in some circles.)
This study produced several key results with application toboth Aircraft Conceptual Design and MultidisciplinaryOptimization, namely:
MDO techniques truly can improve the weight and cost ofan aircraft design concept in the conceptual design phase.This is accomplished by a relatively small "tweaking" of thekey design variables, and with no additional downstreamcosts.In effect, we get a better airplane for free.
For a smaller number of variables (<6-8), adeterministic searching method (here represented by thefull-factorial Orthogonal Steepest Descent) provides aslightly better final result with about the same number ofcase evaluations
For more variables, evolutionary/genetic methods getclose to the best final result with far-fewer caseevaluations. The eight variables studied herein probablyrepresent the practical upper limit on deterministicsearching methods with todays computer speeds.
Of the evolutionary methods studied herein, the BreederPool approach (which was devised during this research andappears to be new) seems to provide convergence in the fewestnumber ofcase evaluations, and yields results very close tothe deterministic best result. However, all of the methodsstudied produced similar results and any of them is asuitable candidate for use.
Hybrid methods, with a stochastic initial optimizationfollowed by a deterministic final "fine tuning", proved lessdesirable than anticipated.
Not a single case was observed, in over a hundred caseruns totaling over a million parametric design evaluations,of a method returning a local rather than global optimum.Even the modified commercial airliner, with poorly selectedinitial design variables far away from the global solution,was easily "fixed" by all the MDO methods studied.
The postulated set of automated redesign procedures andgeometric constraints provide a more-realistic final result,preventing attainment of an unrealistic "better" finalresult. Especially useful is a new approach defined herein,Net Design Volume, which can prevent unrealisticallyhigh design densities with relatively little setup andcomputational overhead. Further work in this area issuggested, especially in the unexplored area of automatedredesign procedures for discrete variables.
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Munjulury, Venkata Raghu Chaitanya. "Knowledge Based Integrated Multidisciplinary Aircraft Conceptual Design." Licentiate thesis, Linköpings universitet, Fluida och mekatroniska system, 2014. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-106925.
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With the ever growing complexity of aircrafts, new tools and eventually methods to use these tools are needed in aircraft conceptual design. To reduce the development cost, an enhancement in the conceptual design is needed. This thesis presents a knowledge-based aircraft geometry design tool RAPID and the methodology applied in realizing the design. The parameters used to create a geometry need to be exchange between different tools. This is achieved by using a centralized database or onedata concept. One-database will enable creating a less number of cross connections between different tools to exchange data with one another. Different types of aircraft configurations can be obtained with less effort. As RAPID is developed based on relational design, any changes made to the geometric model will update automatically. The geometry model is carefully defined to carry over to the preliminary design. The validation of RAPID is done by implementing it in different aircraft design courses at Linköping University. In the aircraft project course, RAPID was effectively used and new features were added to the obtained desired design. Knowledge-base is used to realize the design performance for the geometry with an integrated database approach for a multidisciplinary aircraft conceptual design.
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Schäfer, Katharina [Verfasser]. "Conceptual Aircraft Design for Sustainability / Katharina Schäfer." Aachen : Shaker, 2018. http://d-nb.info/1161299424/34.
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Kay, Jacob. "Control authority assessment in aircraft conceptual design." Thesis, This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-03242009-040703/.
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Squire, Douglas J. "Afterbody drag prediction for conceptual aircraft design." Thesis, This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-09192009-040348/.
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Chai, Sonny T. "Landing gear integration in aircraft conceptual design." Thesis, This resource online, 1996. http://scholar.lib.vt.edu/theses/available/etd-09182008-063506/.
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Opgenoord, Max Maria Jacques. "Uncertainty budgeting methods for conceptual aircraft design." Thesis, Massachusetts Institute of Technology, 2016. http://hdl.handle.net/1721.1/103423.
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Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2016.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 (pages 107-112).
Quantification and management of uncertainty are critical in the design of engineering systems, especially in the early stages of conceptual design. This thesis presents an approach to defining budgets on the acceptable levels of uncertainty in design quantities of interest, such as the allowable risk in not meeting a critical design constraint and the allowable deviation in a system performance metric. A sensitivity-based method analyzes the effects of design decisions on satisfying those budgets, and a multiobjective optimization formulation permits the designer to explore the tradespace of uncertainty reduction activities while also accounting for a cost budget. For models that are computationally costly to evaluate, a surrogate modeling approach based on high dimensional model representation achieves efficient computation of the sensitivities. Example problems in aircraft conceptual design illustrate the approach. The first example investigates the influence of uncertainty in the propulsion technology on the overall aircraft design, whereas the second problem looks at the influence of six different uncertain design parameters from three different disciplines within the aircraft design. Secondly, the distributional sensitivity analysis (DSA) method is extended for better computational efficiency and wider applicability. Instead of assuming that all uncertainty in an input parameter can be reduced, DSA apportions output uncertainty as a function of the uncertainty reduction of a particular input parameter. This leads to more information on influences of uncertainty reduction, and to a more informative ranking of input parameters. In this thesis the ANOVA-HDMR framework is used for DSA to increase computational efficiency. Additionally, this approach allows for using DSA for more general distributions.
by Max Maria Jacques Opgenoord.
S.M.
13
Knöös, Franzén Ludvig, and Erik Magnusson. "Weight Penalty Methods for Conceptual Aircraft Design." Thesis, Linköpings universitet, Fluida och mekatroniska system, 2018. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-175012.
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This report addresses a project conducted at Saab Aeronautics during the spring of 2018. The goal of the project was to investigate aircraft weight estimations in the conceptual design phase. The work was divided into two major parts: finding new weight estimation techniques and implementing an existing technique called the Berry Weight Estimation in to the Pacelab APD software. Several weight estimation techniques were found during an extensive literature review but in the end, only one was chosen for further investigation. The chosen technique was the NASA Wing Weight Build-Up which proposed calculations for wing weights based on aircraft statistics. It contained material data tables for determining so called K-factors that were used to essentially scale the individual wing weight formulas. The data tables did not include K-factors up to a load factor of 9 which was a requirement from Saab. Extrapolations of the material data tables were done to approximate the missing values. The NASA wing weight build-up showed promising results with little deviation from the actual wing weight for a few chosen aircraft. This weight estimation technique was consequently chosen as a worthy candidate for a future implementation in the Pacelab APD software. The task of implementing the Berry Weight Estimation in Pacelab APD was divided into a fuselage- and a wing part. This was done to ease the implementation since it would resemble the original description of the method. The wing and fuselage weights were both calculated in two steps. The first step was to calculate a gross shell weight. This is the weight of an idealized structure without cut-outs or imperfections. The second step was to add so called weight penalties for various components within the wing or fuselage. Typical aircraft components had associating weight penalty functions described in the Berry Weight Estimation. Most of the implemented calculations used Pacelab APD to get involved parameters automatically. However, some of the needed parameters had to be user specified for the implemented Berry Weight Estimation to work. Once the implementation task was finished, several sensitivity studies were made to establish a perception about the involved parameters impact on the Berry Weight Estimation results. The new implementation gave benefits compared with the Berry Weight Estimation in Bex. One of these was the ability to perform extensive trade- and sensitivity studies. The sensitivity studies gave verdicts on the most influencing parameters of the implemented code and guide lines on future improvements of the calculations. These sensitivity studies show, among other things, that is recommended to increase the number of wing and fuselage stations significantly in order to get a converged result for the Berry Weight Estimation.
14
Malan, Paul. "Inlet drag prediction for aircraft conceptual design." Thesis, Virginia Polytechnic Institute and State University, 1989. http://hdl.handle.net/10919/53727.
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A research effort aimed at enhancing ACSYNT, a computer program for aircraft conceptual design, has necessitated the development of methods for predicting inlet drag. Originally, the drag of only one inlet type, the variable-geometry conical inlet, could be calculated within ACSYNT. This prompted the present research which resulted in the creation of a modular suite of subroutines that extend the capability of ACSYNT. Using this new source code, ACSYNT can now predict the drag of subsonic and supersonic pitot inlets, fixed- and variable-geometry conical inlets, and two-dimensional supersonic inlets.
Even though the requirement of computational efficiency has necessitated that many simplifications be made in the analysis, the drag calculations have a sound physical basis. The semi-empirical methods have been extracted from a number of sources based on an extensive literature survey, and these have been enhanced to encompass the full range of inlet operating conditions. The effectiveness of the methods has been demonstrated by comparing some results of the predictions to published data.Master of Science
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Harasani, W. "Aircraft conceptual design decision through operational modelling." Thesis, Cranfield University, 2005. http://dspace.lib.cranfield.ac.uk/handle/1826/11093.
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Aircraft manufacturing is not only a difficult business but also a very competitive one,
the consequences of any drop in sales would cost billions, loss of jobs, and maybe an
economical failure. Therefore, concentrating on just flight performance and adding new
technologies just because they exit is not enough to win the airlines attention, especially
the flow cost carriers.
Manufactures must be able to convince operators that the application of a new design or
technology will produce a favourable change in the bottom line of their balance sheets
and not just a reduction in fuel burn.
Aircraft designers must put more emphasis on what happens to the aircraft after it leaves
the assembly line, through the designed life operation cycle of the aircraft with the airline
customer, quality should be built in to the aircraft. Knowing what are the airline's
concerns, how the aircraft with a given design behaves, and the issues that the airline has,
is vital.
Firstly, it is important to know what are the issues that the airline has, the costumer
(airlines) needs are identified, and, since fleet planning is the top level decision making
department in the airline in which a decision is made to buy one aircraft over the other, it
is important to understand the process and the elements that are involved in fleet
planning. So fleet planning was studied.
Second different technologies for the design have been looked at and selected. Then the
aircraft, airline, airport, and air traffic control are studied, as well as the interaction
between them.
A key element of the research is a simulation program DEBOS that has been built to see
the impact of the different design technologies and concepts through the operation of a
simulation fleet size of 23 aircraft. The Boeing777 aircraft has been chosen to be the base
line of the study.
Finally, it was found that a given technology with improved performance, or a new
concept, would improve the aircraft attractiveness only if it has better life cycle behaviour
characteristics.
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Gleeson, David A. "A second law approach to aircraft conceptual design." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1998. http://handle.dtic.mil/100.2/ADA356093.
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Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, September 1998."September 1998". Thesis advisor(s): Conrad F. Newberry. Includes bibliographical references (p. 55-56). Also available online.
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Gavel, Hampus. "On aircraft fuel systems : conceptual design and modeling." Doctoral thesis, Linköping : Division of Machine Design, Department of Mechanical Engineering, Linköpings universitet, 2007. http://www.bibl.liu.se/liupubl/disp/disp2007/tek1067s.pdf.
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Zhao, Tienan. "Acquisition cost estimating methodology for aircraft conceptual design." Thesis, Cranfield University, 2008. http://dspace.lib.cranfield.ac.uk/handle/1826/9587.
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The research was conducted in the light of a training programme which will train a total
of 150 engineers of AVIC I in Cranfield University during a period of 3 years.
Cost has become an essential driver to aircraft design, as well as performances due to
either the limited defence budget or competitive airline market. Consequently, knowing
the possible cost prior to making actual expenditure will help managers to make proper
decisions and allocate resources efficiently, and designers to optimize their work.
Existing aircraft cost estimating models are outdated and mainly based on a database
including both military and civil aircraft with various missions. This research
concentrated on commercial jet aircraft and was to develop a suitable acquisition cost
estimating methodology for conceptual design from a commercial aircraft
manufacturer’s perspective.
The literature reviewing took a comprehensive overview of some widely-applied cost
estimating methods: Analogy, Parametric, Bottom-up, Feature-based costing, Activitybased
costing (ABC), Expert judgement, and etc. Some practical cost models were also
reviewed to learn the application of cost estimating in the aerospace industry. Then,
analogy and parametric approaches were selected to perform the methodology
development considering the limited data available at the conceptual design phase.
An investigation was deployed to identify the actual problems in practice. The results
helped to recognize the needs of industry. Also, the preparation works for development
are presented to understand the environment.
With subjective judgement and statistical techniques, a series of cost estimating
relationships (CERs) were achieved, in which some historic explanatory parameters
remained or were eliminated, and some new ones introduced. Size of aircraft became
another variable besides weight. As to engines, all developed explanatory variables have
been revealed in prior researches. The validation of CERs proves that they can provide
reliable cost estimates with high accuracy and can be applied to conceptual design. In
addition, a case study was conducted using a baseline aircraft defined in the group
design project (GDP) and presents cost forecasting for the proposed aircraft.
At last, discussion and conclusion presents an overview of the research. A framework
for cost estimating system can be educed. Also, the future work is proposed for in-depth
research.
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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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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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Chicken, S. H. "Conceptual design methodologies for waterborne and amphibious aircraft." Thesis, Cranfield University, 1999. http://dspace.lib.cranfield.ac.uk/handle/1826/9945.
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This study is laid out in 8 self-explanatory sections. The Introduction sets the scene for the thesis by describing the reasoning behind the study, defines terms and introduces the reader to the markets for amphibious aircraft which drive the design requirements. An overall floatplane design methodology is developed. The advantages and disadvantages of the 2 practical float configurations are identified, which result in a basic configuration choice methodology. A method of initially estimating float dimensions and mass for a required displacement is developed from existing references and the aircraft and float databases. Initial float support structure design solutions are proposed based, again, on the information from the databases. A method of positioning the resultant float and structure configuration relative to the existing land-based aircraft centre of gravity is then developed using existing guidance on lateral and longitudinal water-borne static stability and the aircraft database. Guidance on the initial purchase price of floats is gained from a study of commercially available items. The changes in performance due to fitting floats to a conventional aircraft are studied along with a drag comparison study of the main configurations. The work on flyingboats develops an overall flyingboat design methodology which identifies key areas where design methods are required. These methods are developed leading to initial configuration choice methodologies based on a series of generalised mass, configuration and role classifications. Having decided on the overall configuration, tools are developed to choose the method of providing on-water lateral stability and to complete the initial sizing of that choice. A method of estimating initial planing bottom dimensions is developed along with step position and configuration. Tools to estimate the mass of flyingboat-specific items are developed including planing bottom structure and the choice of lateral stability method. Knowing the mass and configuration of the flyingboat allows spray estimation and detailed on-water static stability calculations to be completed to check the acceptability of the initial configuration and dimensions. Performance estimation methods including take-off and landing, aerodynamic drag and on-water dynamic stability are proposed. Logistic support infrastructure, safety and water loading are common to both floatplanes and flyingboats and these are discussed in a separate section, along with a method of allocating values to amphibious aircraft design attributes to measure the success of the design. The methodologies are then used to design 5 floatplanes and 5 flyingboats based on a crosssection of relevant aircraft specification types. This use of the methodologies illustrates that the concept of a linked series of tools to complete the rapid conceptual design of an amphibious aircraft has been successfully achieved. A discussion chapter summarises the key discoveries in each of then former chapters and a conclusion details how the study's aim to develop integrated conceptual design methodologies for waterborne and amphibious aircraft has been successfully achieved. The study's contribution to knowledge, which includes mass, sizing, performance and cost equations for both floatplanes and flyingboats, are also detailed. A list of further work is included which concentrates on the need for further empirical information to increase confidence in the methodologies. A comprehensive bibliography of relevant texts is included.
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Fantini, Paolo. "Effective multiobjective MDO for conceptual design - An aircraft design perspective." Thesis, Cranfield University, 2007. http://hdl.handle.net/1826/2219.
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Once the requirements for a new aircraft have been defined, the Conceptual design phase is launched. During this phase one or more designers have the goal of defining and investigating a number of alternative solutions. Through discussion with industry it has become apparent that optimisation tools are seldom used, even though these could greatly enhance the work of the designers.
The objective of the work carried forward has been of identifying, comparing and where necessary improving the most suitable techniques, as well as schemes for their integration, in order to perform effectively Multidisciplinary Design and Optimisation (MDO) in the Conceptual phase of the aircraft design. The techniques that have been investigated include: multi-objective optimisation algorithms, MDO algorithms for treating non-hierarchically decomposable systems and Automatic Differentiation (AD).
As a result an algorithm for performing multiobjective MDO has been developed. Given a complete model for a complex non-hierarchically decomposable system and given a number of objectives and constraints, the algorithm is capable of determining a set of well distributed solutions, representative of both local and global Pareto frontiers.
A number of test cases have been used for evaluating the alternative methodologies and the proposed algorithm. These include a set of complex algebraic test cases typically used for evaluating global optimisation algorithms and a simplified aircraft conceptual design model, which was provided by industry.
The results demonstrate the unique capability of the algorithm of determining well distributed solutions on the global and local Pareto frontiers for global multiobjective continuous nonlinear constrained optimisation problems. The results also show this capability when the algorithm is applied to non-hierarchically decomposable systems, as typically encountered when performing MDO. Further work could extend the approach in order to handle mixed discrete/continuous variables.
23
Padulo, Mattia. "Computational engineering design under uncertainty : an aircraft conceptual design perspective." Thesis, Cranfield University, 2009. http://hdl.handle.net/1826/4462.
Full textAbstract:
Presented in this thesis is a novel methodology for aircraft design optimization in the presence of uncertainty, with emphasis on the conceptual design stage.
In the initial part of the thesis, the uncertainty typologies of interest for aircraft design are identied within a broader epistemological framework. The main implications for non-deterministic computational design are also outlined.
The focus is then restricted to uncertainties that can be modeled by probability theory. In this context, a methodology is developed to enhance robust design optimization (RDO). Firstly, the problem is formulated in order to relax, when required, the common RDO assumption about the normality of objectives and constraints. Secondly, starting from engineering considerations about the risk related with design unfeasibility, suitable estimates of tail conditional expectation are introduced in the set of robustness metrics.
The proposed formulation requires the estimation of mean and variance of objec¬tives and constraints. To calculate such moments, a novel uncertainty propaga¬tion technique is proposed, which achieves a favorable trade-obetween the ac-curacy of the estimates and the required computational cost. Peculiar features of the propagation technique are exploited to couple the propagation and the opti¬mization phases for the classes of gradient-based methods and the derivative-free pattern search methods. Also analyzed are the possible advantages achievable when the two types of algorithms are hybridized.
The usefulness of the proposed methodology for conceptual design optimization is demonstrated with the aid of two engineering design problems, concerning the sizing of passenger aircraft and the design of transonic airfoils.
24
Nunez, Marco. "Design exploration for engineering design optimisation : an aircraft conceptual perspective." Thesis, Cranfield University, 2010. http://dspace.lib.cranfield.ac.uk/handle/1826/6900.
Full textAbstract:
Most of the efforts in optimisation so far have been focused on the development of novel
or the improvement of existing numerical methods for an effective computation of
optimal solutions. Particular attention has been put on balancing multiple conflicting
objectives, handling the interaction between different disciplines, reducing computational
cost and managing uncertainty. Nonetheless, specific issues of this design methodology
still remain to be properly addressed. In this research, attention is concentrated on
advancing engineering optimisation as a tool for design exploration. The work is put in
the context of conceptual aircraft design.
The overall aim of the present research is to develop a methodology that allows the
designer to effectively conduct an exploration and analysis of alternative design solutions
via a set of methods that can be used separately or conjointly.
The initial part of the thesis introduces two novel methods for assisting the formulation of
an optimisation problem, which generally is assumed to be given a priori. Nonetheless,
the correctness of the optimisation statement, which is not addressed by established
optimisation methods, turns out to be decisive for the feasible design set determination.
The designer is thus provided with an adaptive formulation of functional and designvariable
constraints, which allows the exploration of further promising solutions initially
not contained in the feasible design set. Meaningless results or the loss of important
solutions can therefore be partially avoided.
In a second instance, attention is focused on the visualisation needs for design
exploration. A suitable visualisation methodology has been developed to make the large
multidimensional results of complex design optimisation procedures fully readable and
explanatory. This is achieved by integrating advanced visualisation techniques which provide the designer with diverse perspectives of the data under study and allow him/her
to conduct a number of analysis tasks on it, without the need to be an expert in numerical
optimisation methods.
Last, but not least, a methodology to address conceptual design change problems is
proposed. The decision-maker is enabled to formally state the new design requirements
and priorities introduced by the conceptual change via an adequate problem
reformulation. All the data previously collected can thus be re-used and exploited to drive
an effective exploration of alternative design solutions through design space regions of
interest.
The evaluation of the proposed methodologies has been carried out with a number of test
cases. Analytical examples have been used for the assessment of effectiveness, whereas
codes representative of aircraft sizing procedures have been adopted to evaluate the
methodologies functionality. A visualisation user interface prototype has also been
developed for demonstration and evaluation purposes.
25
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.
Full textAbstract:
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.
26
Al-Ahmed, S. M. "Integrating combat effectiveness disciplines into the aircraft conceptual/preliminary design phase." Thesis, Cranfield University, 1996. http://hdl.handle.net/1826/3631.
Full textAbstract:
An assessment methodology has been developed for - use during the
conceptual/preliminary design phase to quantify the effectiveness of newly designed
aircraft. The effectiveness is measured by a squadron Sortie Generation Rate (SGR). Key
elements of this methodology were the establishment of link parameters between
design synthesis and the main effectiveness disciplines. These were Reliability and
Maintainability (R&M), Survivability / Vulnerability and Acquisition Cost. A
programmable solid modeller was used to create a solid CAD assembly of the aircraft
critical components. A ray tracing technique has been used to develop an interactive
vulnerability assessment tool. A Mission Simulation Model (MSM) has been developed
which typically simulates the operation of a squadron of aircraft and gives the operational
activities such as flying sorties and maintenance actions. The methodology has been
validated based on real data from recent conflicts. The application aspects of the
methodology have been demonstrated by quantifying the effectiveness of two recent
combat aircraft.
27
Siegers, Frank. "Conceptual design synthesis and optimization for new generations of combat aircraft." Thesis, Cranfield University, 1996. http://dspace.lib.cranfield.ac.uk/handle/1826/11009.
Full textAbstract:
A numerical
design synthesis methodology for new generations of combat
aircraft has been
developed. It incorporates advanced technology in the form of
design for low observables. Aircraft capable of being modelled with this
methodology will have internal or external Weapons carriage, side mounted
intakes, a straight-tapered trapezoidal wing, aft-mounted tail with the option of
single or twin ns, and one or two engines with rectangular or axisymmetric
nozzles. The
design methodology incorporates sufficiently accurate and
realistic
algorithms for the calculation of the geometry and the estimation of the
aerodynamic, mass and performance properties of the aircraft. The inherent
flexibility of the design permits the examination of a wide range of
configurations whilst maintaining the accuracy required to examine minor
changes in the design requirements. A numerical optimization routine was
linked to the
synthesis, allowing the determination of optimum aircraft design
variables for a
given set of mission and performance requirements. Results
were obtained
showing the usefulness of this design tool for setting up
parametric trend studies. The numerical accuracy, flexibility of configuration
options and high level of advanced aircraft technology of this synthesis make a
significant contribution to the continuing development of automated design
tools.
28
Wilson, Joseph Scott. "Uncertainty quantifiation with mitigation actions for aircraft conceptual design." Diss., Georgia Institute of Technology, 2015. http://hdl.handle.net/1853/53586.
Full textAbstract:
There are always differences between conceptual design estimates and the performance of a final product. These differences may result in constraint violations, which can have severe financial impacts. Such violations may necessitate downstream changes to recover aircraft performance. The ability to estimate the likelihood and impact of late-stage changes is key to mitigating the overall risk of a design.
Reliability methods can treat design uncertainty; however, existing methods do not account for aspects of aircraft design such as sizing processes, the design freeze after conceptual design, and late-stage ``mitigation actions'' taken when a performance constraint is violated. By accounting for these elements, new reliability metrics can be developed. In addition to the probability of compliance, the designer can determine the probability of recovery through mitigation actions, which helps determine the true likelihood that a design can meet the requirements.
Hypotheses are developed to fill the identified gaps, resulting in Aircraft Recovery through Mitigation & Optimization under Uncertainty for Reliability. ARMOUR augments reliability methods by integrating aircraft sizing, uncertainty margins, and mitigation actions. ARMOUR is demonstrated on the conceptual design of a large civil transport and is exercised to explore previously obscured relationships.
The field of probabilistic aircraft design is enhanced by the concurrent quantification of three elements in one design environment: probability of compliance, probability of recovery after failure, and traditional design criteria. ARMOUR enables the identification of designs which both meets reliability goals and optimizes a traditional performance metric, selecting a design that efficiently meets reliability requirements.
29
de, Tenorio Cyril. "Methods for collaborative conceptual design of aircraft power architectures." Diss., Georgia Institute of Technology, 2010. http://hdl.handle.net/1853/34818.
Full textAbstract:
This thesis proposes an advanced architecting methodology. This methodology allows for the sizing and optimization of aircraft system architecture concepts and the establishment of subsystem development strategies. The process is implemented by an architecting team composed of subsystem experts and architects. The methodology organizes the architecture definition using the SysML language. Using meta-modeling techniques, this definition is translated into an analysis model which automatically integrates subsystem analyses in a fashion that represents the specific architecture concept described by the team. The resulting analysis automatically sizes the subsystems composing it, synthesizes their information to derive architecture-level performance and explores the architecture internal trade-offs. This process is facilitated using the Coordinated Optimization method proposed in this dissertation. This method proposes a multi-level optimization setup. An architecture-level optimizer orchestrates the subsystem sizing optimizations in order to optimize the aircraft as whole. The methodologies proposed in this thesis are tested and demonstrated on a proof of concept based on the exploration of turbo-electric propulsion aircraft concepts.
30
Vaziry-Zanjany, Mohammad Ali (F). "Aircraft conceptual design modelling incorporating reliability and maintainability predictions." Thesis, Cranfield University, 1996. http://hdl.handle.net/1826/3437.
Full textAbstract:
A computer assisted conceptual aircraft design program has been developed
(CACAD). It has an optimisation capability, with extensive break-down in
maintenance costs. CACAD's aim is to optimise the size, and configurations of
turbofan-powered transport aircraft.
A methodology was developed to enhance the reliability of current aircraft systems,
and was applied to avionics systems. R&M models of thermal management were
developed and linked with avionics failure rate and its maintenance cost prediction
methods. The impact of the environmental control system, and engine-provided bleed
flow was also modelled and incorporated into CACAD. The program showed the
ARINC 600 & 408A flow rates to the avionics bay, and to the deck instruments may
both profitably be increased by 50%. This keeps the direct operating cost (DOC)
increase at bay for long-range passenger aircraft, and offers a reduction of up to 1% in
DOC for the short to medium range passenger aircraft.
A methodology was developed to model all aspects of future high risk technologies,
with special consideration given to reliability, maintainability, and development cost
(R, M&D) predictions as applied to variable camber wings (VCW). Many aspects of
VCW were modelled. These included different types of drag saving due to chord-
wise, as well as span-wise camber variation. Models were also derived for mass,
maintenance cost, and extra development cost increments for wing trailing edge
devices, flight control, and hydraulic systems. On incorporation into CACAD, a
reduction in DOC of up to 3.5% was predicted. The VCW technology were evaluated
for DOC improvements, against a number of existing, future, and derivative aircraft,
under different sensitivity conditions. R, M&D predictions were shown to be decisive
in addressing the feasibility of a new technology.
The R&M predictions of the whole study shows that, long range, low to medium
capacity derivative transport aircraft are most appropriate for the VCW technology, and the short to medium range, low to medium capacity aircraft are most suitable for
reliability enhancement projects of aircraft advanced systems.
31
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.
Full textAbstract:
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.
32
Tan, Rendell Kheng Wah. "Quality functional deployment as a conceptual aircraft design tool." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2000. http://handle.dtic.mil/100.2/ADA378471.
Full textAbstract:
Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, March 2000.Thesis advisor(s): Newberry, Conrad F. "March 2000." Includes bibliographical references (p. 81-83). Also available online.
33
Peterson, Gustav. "Sizing and Balance Module Development for Aircraft Conceptual Design." Thesis, Linköping University, Department of Management and Engineering, 2007. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-9999.
Full textAbstract:
This thesis work was done in order to improve the capabilities in a preliminary aircraft analysis program, DIBA, at Saab Aerosystems. The areas that this was done are in the sizing and balance. One sizing tool was developed in order to make a performance analysis with the DIBA generated geometry and customer and/or regulation based criteria. A balance diagram, a neutral point estimation function, a landing gear plot and a trim program was created in order to extend the weight and balance analysis.
Results show that various aircraft both military and civil can be analyzed with good comparison to other analysis and reality. For example EXCEL implemented analysis and graphs over real aircrafts shown in the report.
34
Chakraborty, Imon. "Subsystem architecture sizing and analysis for aircraft conceptual design." Diss., Georgia Institute of Technology, 2015. http://hdl.handle.net/1853/54427.
Full textAbstract:
In traditional aircraft conceptual design, subsystems are largely accounted for implicitly based on available historical data and trends. Such an approach has limitations when novel subsystem architectures such as More Electric or All Electric aircraft are considered, since historical data regarding such architectures is either limited or non-existent. In such cases, the incorporation of more thorough and explicit consideration of the aircraft subsystems into the conceptual design phase is warranted.
The first objective of this dissertation is to integrate subsystem sizing and analysis methods that are suitable for the early design phases with the traditional aircraft sizing methodology. The goal is to facilitate the assessment subsystem architecture performance with respect to vehicle and mission level metrics. The second objective is to investigate how the performance of different subsystem architectures varies with aircraft size. The third and final objective is to assess the sensitivity of architecture performance to epistemic and technological uncertainty.
These objectives are pursued through the development of an integrated sizing and analysis environment where the subsystems are sized in parallel with the aircraft itself using subsystem models that are computationally inexpensive and do not require detailed aircraft definition. The effects of subsystem mass, secondary power requirements, and drag increments are propagated to the mission performance analysis following which the vehicle and subsystems are re-sized. A number of experiments are performed to first test the capabilities of the developed environment and subsequently assess the performance of numerous subsystem architectures and the sensitivity of select architectures to epistemic and technological uncertainty.
35
Sobron, Alejandro. "On Subscale Flight Testing : Applications in Aircraft Conceptual Design." Licentiate thesis, Linköpings universitet, Fluida och mekatroniska system, 2018. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-152488.
Full textAbstract:
Downscaled physical models, also referred to as subscale models, have played an essential role in the investigation of the complex physics of flight until the recent disruption of numerical simulation. Despite the fact that improvements in computational methods are slowly pushing experimental techniques towards a secondary role as verification or calibration tools, real-world testing of physical prototypes still provides an unmatched confidence. Physical models are very effective at revealing issues that are sometimes not correctly identified in the virtual domain, and hence can be a valuable complement to other design tools. But traditional wind-tunnel testing cannot always meet all of the requirements of modern aeronautical research and development. It is nowadays too expensive to use these scarce facilities to explore different design iterations during the initial stages of aircraft development, or to experiment with new and immature technologies. Testing of free-flight subscale models, referred to as Subscale Flight Testing (SFT), could offer an affordable and low-risk alternative for complementing conventional techniques with both qualitative and quantitative information. The miniaturisation of mechatronic systems, the advances in rapid-prototyping techniques and power storage, as well as new manufacturing methods, currently enable the development of sophisticated test objects at scales that were impractical some decades ago. Moreover, the recent boom in the commercial drone industry has driven a quick development of specialised electronics and sensors, which offer nowadays surprising capabilities at competitive prices. These recent technological disruptions have significantly altered the cost-benefit function of SFT and it is necessary to re-evaluate its potential in the contemporary aircraft development context. This thesis aims to increase the comprehension and knowledge of the SFT method in order to define a practical framework for its use in aircraft design; focusing on low-cost, short-time solutions that don’t require more than a small organization and few resources. This objective is approached from a theoretical point of view by means of an analysis of the physical and practical limitations of the scaling laws; and from an empirical point of view by means of field experiments aimed at identifying practical needs for equipment, methods, and tools. A low-cost data acquisition system is developed and tested; a novel method for semi-automated flight testing in small airspaces is proposed; a set of tools for analysis and visualisation of flight data is presented; and it is also demonstrated that it is possible to explore and demonstrate new technology using SFT with a very limited amount of economic and human resources. All these, together with a theoretical review and contextualisation, contribute to increasing the comprehension and knowledge of the SFT method in general, and its potential applications in aircraft conceptual design in particular.
36
McCormick, Daniel John. "An Analysis of Using CFD in Conceptual Aircraft Design." Thesis, Virginia Tech, 2002. http://hdl.handle.net/10919/33409.
Full textAbstract:
The evaluation of how Computational Fluid Dynamics (CFD) package may be incorporated into a conceptual design method is performed. The repeatability of the CFD solution as well as the accuracy of the calculated aerodynamic coefficients and pressure distributions was also evaluated on two different wing-body models. The overall run times of three different mesh densities was also evaluated to investigate if the mesh density could be reduced enough so that the computational stage of the CFD cycle may become affordable to use in the conceptual design stage. A farfield method was derived and used in this analysis to calculate the lift and drag coefficients. The CFD solutions were also compared with two methods currently used in conceptual design - the vortex lattice based program Vorview and ACSYNT. The unstructured Euler based CFD package FELISA was used in this study.Master of Science
37
Malakhoff, Lev A. "Combat aircraft mission tradeoff models for conceptual design evaluation." Diss., Virginia Polytechnic Institute and State University, 1988. http://hdl.handle.net/10919/53583.
Full textAbstract:
A methodology is developed to address the analyses of combat aircraft attrition. The operations of an aircraft carrier task force are modeled using the systems dynamics simulation language DYNAMO. The three mission-roles include: surface attack, lighter escort, and carrier defense. The level of analysis is performed over the entire campaign, going beyond the traditional single·sortie analysis level.
These analyses are performed by determining several measures of effectiveness (MOEs) for whatever constraints are applied to the model. The derived MOEs include: Campaign Survivability (CS), Fractlon of Force Lost (FFL), Exchange Ratio (ER), Relative Exchange Ratio (RER), Possible Crew Loss (PCL), and Replacement Cost (RC). RER is felt to be the most useful MOE since it considers the initial inventory levels of both friendly and enemy forces, and its magnitude is easy for the analyst to relate to (an RER greater than one is a prediction of a friendly force’s victory).
The simulation model developed in this research is run for several experiments. The effects of force size on the MOEs ls studied, as well as a hypothetical multimission aircraft deployed to perform any of the three missions (albeit at lower effectiveness than the speciallzed aircraft for their given roles but nonetheless with a higher availability).
Evaluation of specific technological improvements such as smaller radar cross section, higher thrust/weight, improved weapons ranges, is made using the MOEs. Also, a cost-effectiveness tradeoff methodology is developed by determining the acquisition cost ratio (ACR) for certain modified alternatives the baseline by determining the required initial inventory of modified aircraft to produce the same total effectiveness of the baseline aircraft.Ph. D.
38
Cabrera, Antonio Trani. "Combat aircraft scenario tradeoff models for conceptual design evaluation." Diss., Virginia Polytechnic Institute and State University, 1988. http://hdl.handle.net/10919/53920.
Full textAbstract:
The purpose of this research is to apply engineering-based knowledge to the field of combat aircraft survivability, and to create scenario-specific models in order to estimate the tradeoff between aircraft survivability and lethality metrics at the encounter and sortie levels. The development of scenario-specific models serves to identify and quantify technological changes that have Ieverage on the overall performance of the aircraft from a survivability point of view. Also, the models focus on the fighter aircraft susceptibility assessment and are capable of incorporating outputs from offline studies as inputs, such as in the area of vulnerability assessment where extensive databases are available.
The mission scenario models are microscopic in nature and relate important conceptual aircraft design parameters such as thrust-to-mass ratio, wing loading, empty mass, maneuverability, etc. and operational parameters (e.g., weapon payload, range, loiter time, flight profiles, etc.) to the aircraft sortie survivability and lethality under various threat scenarios.
This research proposes a methodology to estimate survivability and lethality aircraft performance at the sortie level where aircraft parameters can be implemented into scenario-specific models to assess their impact upon survivability-related metrics. While the project was conceived with naval aircraft in mind, the methodology, to the extent possible, is not to be aircraft-specific and thus could be applied to any particular design at the conceptual stage.Ph. D.
39
Sripawadkul, Vis. "Geometrical representations for efficient aircraft conceptual design and optimisation." Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/12164.
Full textAbstract:
Geometrical parameterisation has an important role in the aircraft design process due to its impact on the computational efficiency and accuracy in evaluating different configurations. In the early design stages, an aircraft geometrical model is normally described parametrically with a small number of design parameters which allows fast computation. However, this provides only a course approximation which is generally limited to conventional configurations, where the models have already been validated. An efficient parameterisation method is therefore required to allow rapid synthesis and analysis of novel configurations. Within this context, the main objectives of this research are: 1) Develop an economical geometrical parameterisation method which captures sufficient detail suitable for aerodynamic analysis and optimisation in early design stage, and2) Close the gap between conceptual and preliminary design stages by bringing more detailed information earlier in the design process.
Research efforts were initially focused on the parameterisation of two-dimensional curves by evaluating five widely-cited methods for airfoil against five desirable properties. Several metrics have been proposed to measure these properties, based on airfoil fitting tests. The comparison suggested that the Class-Shape Functions Transformation (CST) method is most suitable and therefore was chosen as the two-dimensional curve generation method. A set of blending functions have been introduced and combined with the two-dimensional curves to generate a three-dimensional surface. These surfaces form wing or body sections which are assembled together through a proposed joining algorithm. An object-oriented structure for aircraft components has also been proposed. This allows modelling of the main aircraft surfaces which contain sufficient level of accuracy while utilising a parsimonious number of intuitive design parameters ... [cont.].
40
Vaziry-Zanjany, Mohammad Ali. "Aircraft conceptual design modelling incorporating reliability and maintainability predictions." Thesis, Cranfield University, 1996. http://dspace.lib.cranfield.ac.uk/handle/1826/3437.
Full textAbstract:
A computer assisted conceptual aircraft design program has been developed (CACAD). It has an optimisation capability, with extensive break-down in maintenance costs. CACAD's aim is to optimise the size, and configurations of turbofan-powered transport aircraft. A methodology was developed to enhance the reliability of current aircraft systems, and was applied to avionics systems. R&M models of thermal management were developed and linked with avionics failure rate and its maintenance cost prediction methods. The impact of the environmental control system, and engine-provided bleed flow was also modelled and incorporated into CACAD. The program showed the ARINC 600 & 408A flow rates to the avionics bay, and to the deck instruments may both profitably be increased by 50%. This keeps the direct operating cost (DOC) increase at bay for long-range passenger aircraft, and offers a reduction of up to 1% in DOC for the short to medium range passenger aircraft. A methodology was developed to model all aspects of future high risk technologies, with special consideration given to reliability, maintainability, and development cost (R, M&D) predictions as applied to variable camber wings (VCW). Many aspects of VCW were modelled. These included different types of drag saving due to chord- wise, as well as span-wise camber variation. Models were also derived for mass, maintenance cost, and extra development cost increments for wing trailing edge devices, flight control, and hydraulic systems. On incorporation into CACAD, a reduction in DOC of up to 3.5% was predicted. The VCW technology were evaluated for DOC improvements, against a number of existing, future, and derivative aircraft, under different sensitivity conditions. R, M&D predictions were shown to be decisive in addressing the feasibility of a new technology. The R&M predictions of the whole study shows that, long range, low to medium capacity derivative transport aircraft are most appropriate for the VCW technology, and the short to medium range, low to medium capacity aircraft are most suitable for reliability enhancement projects of aircraft advanced systems.
41
Bérard, Adrien. "Method Development for Computer Aided Engineering for Aircraft Conceptual Design." Licentiate thesis, KTH, Aeronautical and Vehicle Engineering, 2008. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-9240.
Full textAbstract:
This thesis presents the work done to implement new computational tools and methods dedicated to aircraft conceptual design sizing and optimization. These tools have been exercised on different aircraft concepts in order to validate them and assess their relevance and applicability to practical cases.First, a geometry construction protocol has been developed. It is indeed essential to have a geometry description that supports the derivation of all discretizations and idealizations used by the different analysis modules (aerodynamics, weights and balance, stability and control, etc.) for which an aircraft concept is evaluated. The geometry should also be intuitive to the user, general enough to describe a wide array of morphologies and suitable for optimization. All these conditions are fulfilled by an appropriate parameterization of the geometry. In addition, a tool named CADac (Computer Aided Design aircraft) has been created in order to produce automatically a closed and consistent CAD solid model of the designs under study. The produced CAD model is easily meshable and therefore high-fidelity Computational Fluid Dynamics (CFD) computations can be performed effortlessly without need for tedious and time-consuming post-CAD geometry repair.Second, an unsteady vortex-lattice method based on TORNADO has been implemented in order to enlarge to scope of flight conditions that can be analyzed. It has been validated satisfactorily for the sudden acceleration of a flat plate as well as for the static and dynamic derivatives of the Saab 105/SK 60.Finally, a methodology has been developed to compute quickly in a semi-empirical way the buffet envelope of new aircraft geometries at the conceptual stage. The parameters that demonstrate functional sensitivity to buffet onset have been identified and their relative effect quantified. The method uses a combination of simple sweep theory and fractional change theory as well as the buffet onset of a seed aircraft or a provided generic buffet onset to estimate the buffet envelope of any target geometry. The method proves to be flexible and robust enough to predict within mainly 5% (and in any case 9%) the buffet onset for a wide variety of aircrafts, from regional turboprop to long-haul wide body or high-speed business jets.This work was done within the 6th European framework project SimSAC (Simulating Stability And Control) whose task is to create a multidisciplinary simulation environment named CEASIOM (Computerized Environment for Aircraft Synthesis and Integrated Optimization Methods), oriented toward stability and control and specially suited for aircraft conceptual design sizing and optimization.
SimSAC
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Lu, Zhijie. "Data Management in an Object-Oriented Distributed Aircraft Conceptual Design Environment." Diss., Georgia Institute of Technology, 2007. http://hdl.handle.net/1853/14548.
Full textAbstract:
Aircraft conceptual design, as the first design stage, provides major opportunity to compress design cycle time and is the cheapest place for making design changes. However, traditional aircraft conceptual design programs, which are monolithic programs, cannot provide satisfactory functionality to meet new design requirements due to the lack of domain flexibility and analysis scalability. Therefore, we are in need of the next generation aircraft conceptual design environment (NextADE). To build the NextADE, the framework and the data management problem are two major problems that need to be addressed at the forefront. Solving these two problems, particularly the data management problem, is the focus of this research.
In this dissertation, a distributed object-oriented framework is firstly formulated and tested for the NextADE. In order to improve interoperability and simplify the integration of heterogeneous application tools, data management is one of the major problems that need to be tackled. To solve this problem, taking into account the characteristics of aircraft conceptual design data, a robust, extensible object-oriented data model is then proposed according to the distributed object-oriented framework. By overcoming the shortcomings of the traditional approach of modeling aircraft conceptual design data, this data model makes it possible to capture specific detailed information of aircraft conceptual design without sacrificing generality. Based upon this data model, a prototype of the data management system, which is one of the fundamental building blocks of the NextADE, is implemented utilizing the state of the art information technologies.
Using a general-purpose integration software package to demonstrate the efficacy of the proposed framework and the data management system, the NextADE is initially implemented by integrating the prototype of the data management system with other building blocks of the design environment. As experiments, two case studies are conducted in the integrated design environments. One is based upon a simplified conceptual design of a notional conventional aircraft; the other is a simplified conceptual design of an unconventional aircraft. As a result of the experiments, the proposed framework and the data management approach are shown to be feasible solutions to the research problems.
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Lammering, Tim [Verfasser]. "Integration of aircraft systems into conceptual design synthesis / Tim Lammering." Aachen : Hochschulbibliothek der Rheinisch-Westfälischen Technischen Hochschule Aachen, 2014. http://d-nb.info/1056993960/34.
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Seeckt, Kolja. "Conceptual design and investigation of hydrogen-fueled regional freighter aircraft." Licentiate thesis, KTH, Farkost och flyg, 2010. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-26348.
Full textAbstract:
This thesis presents the conceptual design and comparison of five versions of regional freighter aircraft based on the ATR 72. The versions comprise four baseline designs differing in their propulsion systems (jet/turboprop) and the fuel they use (kerosene/hydrogen). The fifth version is an improved further development of the hydrogen-fueled turboprop aircraft. For aircraft modeling the aircraft design software PrADO is applied. The criteria for the overall assessment of the individual aircraft versions are energy use, climate impact in terms of global warming potential (GWP) and direct operating costs (DOC). The results indicate that, from an aircraft design perspective, hydrogen is feasible as fuel for regional freighter aircraft and environmentally promising: The hydrogen versions consume less energy to perform a reference mission of 926 km (500 NM) with a payload of 8.1 t of cargo. The climate impact caused by the emissions of hydrogen-fueled regional freighter aircraft is less than 1 % of that of kerosene-fueled aircraft. Given the circumstance that sustainably produced hydrogen can be purchased at a price that is equivalent to kerosene with respect to energy content, hydrogen-fueled regional freighter aircraft are also economically competitive to current kerosene-fueled freighters. In consequence, regional freighters appear especially favorable as first demonstrators of hydrogen as aviation fuel, and cargo airlines and logistics companies may act as technology drivers for more sustainable air traffic. The potential of regional freighter aircraft alone to mitigate climate change is marginal. The share of national and regional air cargo traffic in global manmade climate impact lies in the region of 0.016 % to 0.064 %, which also represents the maximum reduction potential. The presented work was to a large extend performed during the joint research project "The Green Freighter" under the lead of Hamburg University of Applied Sciences (HAW Hamburg).QC 20101123
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Bérard, Adrien. "Method development for computer aided engineering for aircraft conceptual design /." Stockholm : School of Engineering Sciences, Kungliga Tekniska högskolan, 2008. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-9240.
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Altman, Aaron. "A conceptual design methodology for low speed high altitude long endurance unmanned aerial vehicles." Thesis, Cranfield University, 2000. http://hdl.handle.net/1826/3998.
Full textAbstract:
A conceptual design methodology was produced and subsequently coded into a
Visual C++ (GUI) environment to facilitate the rapid comparison of several possible
configurations to satisfy High Altitude Long Endurance (FIALE) unmanned aircraft (UAV)
missions in the Low Speed (propeller driven aircraft) regime.
Several comparative studies were performed to verify the applicability of traditional
design methods. The traditional computational design methodologies fail in several areas
such as high aspect ratio wing weight estimation and design, low Reynolds number wing
design, high altitude engine performance, low Reynolds number drag estimation, unmanned
aircraft design, and the conceptual design of unconventional configurations. The
methodology developed for this thesis was robust enough to allow not only for
consideration of these areas of inadequacy in traditional methods, but also to allow for the
inclusion of advancements in the relevant technologies as they become more widely
available.
The following configurations were evaluated for suitability to the Low Speed HALE
UAV application: conventional, canard, twin boom, multiple fuselage (conventional or
canard), tandem wing, multiple fuselage tandem wing or flying wing configuration. The
configurations were compared on the basis of aircraft endurance for takeoff weights ranging
from 2,000 to 20,000 pounds and wing loadings ranging from 5 to 25 lbs1fe.
Initial drag estimates were made using traditional parabolic drag estimation
techniques. A more refined drag buildup was performed using a vortex lattice drag
estimation for the lift induced drag (for all lifting components) and calculated skin friction
coefficients for the parasite drag. Statistically based methods were used for other
components of drag having much smaller contributions. In addition, a statistical approach
was taken to the weight estimation of the major aircraft components. However, this
approach made comparison of alternative configurations more difficult. Thus wing bending
moments trends were evaluated and utilized in the development of weight saving values for
multiple fuselage wing weight estimation.
The comparative performance of each configuration is justified with direct reference
to the terms in the Breguet Endurance equation. Validation was performed where possible
on all modules and segments associated with the methodology, as well as for the
macroscopic results. In addition, parametric studies on endurance were performed for the
conventional configuration for geometric characteristics and operating conditions directly
and indirectly effecting the calculated endurance and generalized results presented. Finally, a
case study was performed to demonstrate this capability.
A new relation was developed for aircraft empty weight prediction, a low speed
airfoil figure of merit was proposed, and new constants were offered for UAV fuselage
length prediction. In addition, horizontal and vertical tail volume coefficients were proposed
for all of the Low Speed HALE UAV configurations considered. It was determined that the
multiple fuselage configurations showed comparatively superior endurance performance
across a range of takeoff weights, with several other configurations demonstrating marginal
endurance improvements. Finally, a highly flexible and robust computer based conceptual
design methodology was developed and validated enabling the quick comparison of a greater
number of possible configurations to satisfy a given mission for Low Speed HALE UAV's
and providing detailed drag and weight breakdown data.
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Amadori, Kristian. "On Aircraft Conceptual Design : A Framework for Knowledge Based Engineering and Design Optimization." Licentiate thesis, Linköping : Department of Management and Engineering, Linköpings universitet, 2008. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-11873.
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Koepke, John Allen. "Conceptual design of a stand-off weapon for maritime patrol aircraft." Thesis, Monterey, California. Naval Postgraduate School, 1988. http://hdl.handle.net/10945/22982.
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King, Daniel J. S. M. Massachusetts Institute of Technology. "Conceptual aircraft design for environmental impact : modeling operations for emissions assessment." Thesis, Massachusetts Institute of Technology, 2005. http://hdl.handle.net/1721.1/34138.
Full textAbstract:
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2005.Includes bibliographical references (p. 121-123).
Decisions that guide technology investment and policy-making for the future of air transportation will be based in part on the tradeoffs between environmental performance and economics. The Environmental Design Space (EDS) project explores the tradeoffs between noise, emissions, and economics for conceptual design of future aircraft. A key component to EDS is modeling the emissions of aircraft in operation. Traditional design tools need more detailed mission analysis to calculate operational emissions in the landing and takeoff cycle (LTO), and in cruise. This thesis presents a methodology for modeling an aircraft flight profile and the corresponding aircraft and engine states required to calculate emissions over that mission. The methodology was implemented as an operations model in EDS. The development of the methodology and demonstrations of the model are presented in this thesis. The model takes user-input flight procedures, including mission range, and uses aerodynamic models and engine models from EDS to calculate flight profiles. The operations model can be used for analyzing the relationships between flight procedures, and emissions and fuel burn for a fixed design. Alternatively, multidisciplinary design optimization (MDO) with EDS and the operations model can be used to optimize the aircraft design for minimized emissions in the flight profile.
(cont.) MDO with the new model enables exploration of a design space that includes operations along with design in evaluating tradeoffs between emissions, noise, and economics. In addition to the development and demonstration of the operations model, a detailed study of the effects of derated-thrust takeoffs on emissions and fuel burn for Boeing 777 flights is presented in this thesis. The emissions of airline flights are calculated from flight data and compared to International Civil Aviation Organization (ICAO) assumptions for the engines used. The results show that NOx emissions are significantly less than the ICAO assumed values for takeoff and climb-out. A second analysis compares the emissions of derated thrust; takeoffs to the emissions that would have resulted if the same aircraft had flown with full-power on the same day. The results show a relationship between percent derate and a change in the emissions produced in takeoff, and a tradeoff of increased fuel burn for a decrease in NOx production.
by Daniel J. King.
S.M.
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Henriquez, Acacio Alejandro Morales. "Flight control design for a flexible conceptual aircraft using backstepping technique." Instituto Tecnológico de Aeronáutica, 2011. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=2170.
Full textAbstract:
A nonlinear flight control system is proposed for a conceptual flexible aircraft using Backstepping technique to achieve global stability in the rigid and flexible dynamics. It is introduced a controller to lead the model to a rigid-body model approximation, minimizing structural dynamics effects using static Backstepping approach, that system can be called as flexible modes suppressor. Afterward, it is applied a controller with an internal loop involving the angular rates of the aircraft and an external loop which includes pitch angle, sideslip angle and bank angle without the two-timescale assumption to separate slow and fast dynamics and without consider aerodynamics forces and moments increments caused by structural dynamics. In addition, external looping are built using Backstepping for first order systems in order to control aircraft course and altitude, the results are reference inputs to be introduced in the previous loop developed for rigid body control. Also, it is implemented a separate controller to track velocity using Backstepping approach, as a result, aircraft autopilot system is completed. Nonlinear six degree of freedom simulation results for a conceptual model of a medium size jet, like Embraer 190/195 and Boeing 737-200/300, are presented to demonstrate the effectiveness of the proposed control law in several conditions. It is assumed that the aerodynamics coefficients are fixed and the model presents augmented flexible features.