The aim of this work has been to identify hidden parameter value patterns during evasive maneuvering for a typical jet fighter. The work has created a performance model for a fighter aircraft and this model has then been combined with a missile model to simulate an enemy attack. By doing different kinds of simulations with a certain amount of predetermined scenarios, different outcomes could be evaluated when making small changes in the maneuvers during each specific scenario. The span of parameters that conducts a flying airplane’s trajectory is vast and the evaluation of different decisions that is up on the table for a pilot in a given situation might give new insights when optimizing tactical air fighting scenarios.After evaluating different scenarios with different input values in form of different turn and climb angles etc, it was clear that small changes resulted in vast differences regarding the outcome, when being chased by the missile. By analyzing the results, it can be concluded that there are underlying patterns regarding controllable parameter values when the airplane tries to get rid of the chasing missile. For example; one section in this work describes that by keeping a straight flight path for a certain amount of seconds after a specified value of turn angle - results in survival of the attack. Keeping level flight for too many seconds however has a lethal outcome. The results seem also to follow a continuous - non-randomized - pattern. This type of detailed analysis could be used to help a pilot to optimizise the performance of the maneuver.

To reduce the cost of launch, several aerospace companies are exploring the possibility of partly reusable launchers. In order to be reusable with minimum refurbishment cost the reusable part of the launcher has to suffer little damage and land in optimal conditions. In this paper, a guidance algorithm to achieve the return of the reusable vehicle on ground through a vertical landing is described. Different mission scenarios are taken into account and the performance of the guidance algorithm is assessed using a 6 Degrees Of Freedom simulator.

Operations on contaminated runways are a permanent challenge for the aviation community. Among the stakeholders, Airbus has decided to improve its knowledge by developing a fast post-processing software aiming at quickly identifying the runway state on operational recordings. First and foremost, the context of operations on contaminated runways is presented, with a particular emphasis on runway condition assessment methods. In a second part, the models embedded in this function are validated against simulation and more elaborated flight test analysis software. Then, the identification algorithm is validated by comparing its outputs to actual operational conditions. Finally, the robustness of the identification is assessed.

The thesis work content is to evaluate the use of more advanced turbulence models available in the ANSYS CFX software for aerodynamic calculations. In particular for flows over airfoils with thick trailing edges, the turbulence modeling is challenging to traditional methods, as both thin boundary layers as well as an unsteady wake needs to be well represented. This is done by using the standard SST and then performing unsteady computations using the more advanced unsteady SAS-SST model to get the relevant CFD results. By comparing to tests performed at GKN and results from literature the improvement could be assessed in terms of modeling quality and computational cost. The results presented give a good contribution to how the modeling of unsteady wakes can be improved and used for design purposes.

The analysis of the reliability of CREO simulate will be done step by step. First of all, the precisionof the simulation has to be measured thus a comparison with simple theoretical computations willbe done. Then, the scope of the capacity of the software will be analyzed and if elements deemednecessary to thermal simulations are missing, back up solutions are to be found. Also, the influenceof the meshing will be studied and measured to ensure that the software guarantee convergence evenin the hand of persons unfamiliar with simulations.In parallel, one will experiment with realistic hardware that could be used to compare reality withthe simulations. Those experiments will be handmade using regular materials from the company.

The 1/3 disc fragment model for fan disc burst shall allow aircraft manufacturers to minimize the hazards to the aeroplane without unnecessarily constraining its design. This paper presents a finite energy approach, along with its benefit assessment on the electrical segregation height and rapid decompression rotor burst design constraints. The fan case energy absorption is estimated through the AGARD impact mechanics formula, considering its capability to withstand fan blade-off events. The results for a standardized single-aisle engine suggest 34% energy absorption, 95mm minimum electrical segregation height decrease and 83% rapid decompression maximum opening decrease. The first estimation of energy absorption shows that it is unlikely that the fan retention belt would retain the 1/3 disc fragment. Assumptions regarding material failure, ring shape effect, and blade deformation energy absorption could be refined.

Nowadays, as the demand for fuel efficiency is increasing and the manufacturing techniques improved to allow the design of laminar wings, the laminar concepts are being revived and studied, and seen as one of the most promising techniques to reduce skin friction and hence the overall drag of the airframe. In order to predict the kind of benefits that could be obtained from such concepts, computational fluid dynamics (CFD) computations can now be carried out but it is necessary to be able to predict flow transition from laminar to turbulent around the lifting surfaces of the aircraft, such as the wing and the horizontal tail plane. Such capability has been developed in the last years by the French Aerospace Laboratory (ONERA) with their CFD code elsA, and by the German Aerospace Center (DLR) with their CFD code Tau. These two codes are widely used in Airbus in their CFD campaigns, and this study aims to assess the capability of the different approaches to predict transition and the benefits from laminar flow in the overall drag. The methods used, the limitations and the suggestions for further improvement of the codes will be detailed in this paper.

Hazards associated to Wake Vortex Encounters (WVE) are increasing with the growing traffic and the appearance of super heavy aircrafts such as the A380. The project comes with the need from airports and airlines to clearly analyze the phenomenon by detecting WVE and quantifying their strength from Flight Data Recorder data. A tool is developed with this purpose using Roll Control Ratio as severity metric. Flight data from flight simulator are used for first analyses. The obtained results are encouraging and enable proposed improvements and consideration to further work.

In this work the so-called μ-k method for robust flutter analysis is generalized to theLaplace domain. Although the generalization as such is a straightforward matter, the resultingμ-p method is far more versatile than the previous formulation. With the new method, a singlestructured singular value (or μ) evaluation can be used to determine if a certain eigenvalue p is asolution to the uncertain flutter equation or not. This result makes it possible to compute regionsof feasible eigenvalues in the complex plane, as well as extreme eigenvalues that can be used topredict damping bounds and perform robust flutter analysis. The capability to predict dampingbounds at any flight condition is a very attractive feature of the μ-p method, as flight testing israrely taken to the flutter point. This feature also makes it possible to estimate the magnitude ofthe uncertainty based on the measured frequency and/or damping of a particular mode, whichcan reduce the potential conservativeness of the robust analysis. Finally, the capabilities of thenew method is demonstrated by applying it to a low-speed wind tunnel model.

This paper introduces a new technique for robust aeroelastic analysis that extends standard linear flutter analysis to take deterministic uncertainty and variation into account. The basic principle of the proposed mu-p method is to exploit structured-singular-value (or mu) analysis to investigate if the system uncertainties can make the flutter determinant zero for a given flutter eigenvalue p. This makes it possible to compute regions of feasible eigenvalues in the complex plane as well as extreme eigenvalues that can be used to predict damping bounds and perform robust flutter analysis. The capability to predict damping bounds at subcritical flight conditions is a very attractive feature of the new method, as flight testing is rarely taken to the flutter point. The A-p formulation also opens up new possibilities to bound the magnitude of the system uncertainties based on frequency and/or damping estimates from flight testing. In the final part of the paper, the mu-p framework is successfully applied to perform robust aeroelastic analysis of a low-speed wind-tunnel model.

The use of upper-bound μ-k estimation method in the development of robust flutter analysis and flutter testing, was described. Since only the frequency-domain aerodynamic forces are required to compute μ(k), established aerodynamic methods can be used for the robust flutter analysis. A robust flutter analysis considering wing-tip aerodynamic uncertainty was developed in MATLAB® for a wind-tunnel model in low-speed airflow. The results show that the extended procedure for robust flutter analysis was successfully applied to a wind-tunnel model in low-speed airflow.

A simple and efficient algorithm for robust flutter analysis is presented. First, a general linear fractional transformation formulation of the mu-k method is provided, making it straightforward to pose the uncertain flutter equation in a form suitable for structured singular value analysis. The new formulation establishes a close connection between mu-k flutter analysis and traditional frequency-domain flutter analysis, which is used to formulate an efficient algorithm for computation of robust flutter boundaries. The proposed method is successfully applied to an F-16 sample test case with uncertain external stores aerodynamics, showing that standard tools for structural dynamics and unsteady aerodynamics can be used to perform robust flutter analysis with only modest additional modeling.

The solution of the flutter eigenvalue problem with mixed structural/aerodynamic uncertainty was examined. A delta-wing wind-tunnel model was used as a test case. The model had a simple structural design and was made of glass-fiber and carbon-fiber composite materials. It had a semispan of 0.88 m and mean chord of 0.70m and was mounted vertically in the low-speed wind tunnel L2000. The numerical analysis was based on a NASTRAN model with shell elements for the wing, mass elements for the missile, and aerodynamic panels for doublet-lattice aerodynamic loads. The relatively stiff missile was modeled as a rigid body attached to the wing tip. The nominal eigenvalue was always an interior point of the corresponding eigenvalue set. A parameter sweep taking only aerodynamic uncertainty into account was performed by solving the eigenvalue problem for a set of parameter values. The eigenvalue sets based on one patch and seven patches, respectively, showed only slight difference.

Extensive research has been done on the aerodynamics of supersonic aircrafts, especially in the military and commercial airplanes’ field. Regarding supersonic business jets (SSBJs), two major problems have been addressed in past investigations: reducing the sonic boom and decreasing the NOx emissions. This report focuses on a di˙erent aspect, the controls and stability of this type of aircrafts. This field has not been addressed thoroughly by the di˙erent companies and universities investigating SSBJs, as most of the existing concepts are preliminary designs that have not been developed extensively. With this report I try to put my two cents in analyzing the longitudinal stability and control surfaces of three di˙erent SSBJs designs.

This paper deals with the adaptation of a flight-test tool on A350 simulators in order to support a flight profile. This adaptation has been performed by doing first, an anal- ysis of the current situation of the existing tool on old programs like the A320 or the A330. Then, several modifications on the tool have been performed in order to adapt this tool on simulators. Finally, the validation of this tool has been done by performing several tests on A350 simulators. The difficulty of this adaptation was due to the appari- tion of new state-of-the-art technologies. The main improvement concerned the signals technology that appeared from the A380. Therefore, a modification on the parameters acquisition process has been done to take into account these improvements. Finally, the implementation of this mock-up on A350 simulators consists in a feasibility study that enables simulator test engineers to automate the tests performed on simulators and to do a traceability of these tests. The realisation of this mock-up is necessary because the simulator test engineers need this tool for the end of 2012 in order to develop the A350 whereas the real evolution of this tool will be done when the real aircraft A350 will be produced in 2014.

Laminar separation bubbles evolve on laminar wing at low Reynolds number.These bubbles have especially been observed on winglets or on wings with flapsor slats. They can stretch up to 30% chord, taking them into account is crucialsince their presence greatly modifies the forces acting on a wing. However, theestablishment of a model is extremely complex and many researches are stillundertaken on this subject. The aim of this work is to achieve automation ofdetecting and taking into account long bubbles in the Navier -Stokes code ofDassault Aviation named Aether. The key issue is the proper detection of thetransition position inside the bubble. This detection is possible by performing alinear stability calculation, but this method is not very robust in this case andvery costly in computation time, as conventional methods are not practicable inthe presence of separation. Finally, the chosen method uses a boundary layercode to predict the position of separation, then a criterion involving local integralquantities to determine the intermittent transition area. This method has beentested for various 2D airfoils and provided quite satisfactory results.

This paper gives an overview of the mathematical and practical tools that can be used in turbo-machinery RANS simulation to realize unstructured mesh adaptation. It first presents the concept of metric and recalls that the hessian of the physical flow properties can become, thanks to small modifications, both a metric and a upper bound of the P1 projection error. The resulting metric is then studied on a simple 2D case. In a second part, the industrial application of this concept is addressed and the tools used to overcome the turbo-machinery specificities are explained. Finally, some 2D and 3D results are presented.

For designing and evaluating new concepts for HALE platforms, the energy provided by solar cells is a key factor. The purpose of this thesis is to model the electrical power which can be harnessed by such a platform along any flight trajectory for different aircraft designs. At first, a model of the solar irradiance received at high altitude will be performed using the solar irradiance models already existing for ground level applications as a basis. A calculation of the efficiency of the energy generation will be performed taking into account each solar panel’s position as well as shadows casted by the aircraft’s structure. The evaluated set of trajectories allows a stationary positioning of a hale platform with varying wind conditions, time of day and latitude for an exemplary aircraft configuration. The qualitative effects of specific parameter changes on the harnessed solar energy is discussed as well as the fidelity of the energy generation model results.

A hazardous and not fully understood aeroelastic phenomenon is the transonic dip,the decrease in flutter dynamic pressure that occurs for most aircraft configurationsin transonic flows. The difficulty of predicting this phenomenon forces aircraft manufacturersto run long and costly flight test campaigns to demonstrate flutter-free behaviourof their aircraft at transonic Mach numbers.In this project, subsonic and transonic flutter calculations for the KTH-NASA genericfighter research model have been performed and compared to existing experimentalflutter data from wind tunnel tests performed at NASA Langley in 2016. For the fluttercalculations, industry-standard linear panel methods have been used together with afinite element model from NASTRAN.Further, an alternative approach for more accurate transonic flutter predictions usingthe full-potential solver Phi has been investigated. To predict flutter using this newmethodology a simplified structural model has been used together with aerodynamicmeshes of the main wing. The purpose of the approach was to see if it was possibleto find a method that was more accurate than panel methods in the transonic regimewhilst still being suitable for use during iterative design processes.The results of this project demonstrated that industry-standard linear panel methodssignificantly over-predict the flutter boundary in the transonic regime. It was alsoseen that the flutter predictions using Phi showed potential, being close to the linearresults for the same configuration as tested in Phi. For improved transonic accuracy inPhi, an improved transonic flow finite element formulation could possibly help .Another challenge with Phi is the requirement of an explicit wake from all liftingsurfaces in the aerodynamic mesh. Therefore, a method for meshing external storeswith blunt trailing edges needs to be developed. One concept suggested in this projectis to model external stores in "2.5D", representing external stores using airfoils withsharp trailing edges.

This master thesis gives a brief sum up of the implementation of a large strain formulation in an existing industrial finite element code. The logarithmic strain tensor additively decomposed in an elastic and a plastic contribution was selected as a basis of this formulation. The main advantage of this formalism is that the existing plastic resolution algorithms written in the small strain formalism can be re-used. Only a pre-processing phase (construction of the logarithmic strain tensor) and a post-processing of the stress tensor is needed. This formulation, detailed in this paper was implemented in FORTRAN and integrated to the existing code. The final implementation was then validated thanks to some benchmarks with Abaqus.

Dubai International Airport has set as an aim to become the leading connecting hub between Europa and Asia. As to be cope with the stresses being put on the airspace in terms of holding times, fuel efficiency and environmental impacts, an analysis has been made concerning the future technology in air traffic management: Time based Operations. This project has revealed that it is possible to maintain the current mean capacity of 29 aircraft/ hour while flying an optimal vertical descent profile. However, exceeding that capacity is not possible with the methods used in this project. The conclusion is that the only forward is Dynamic Controlled of Arrival, together with the reorganization of the entire airspace of Dubai.

In the following thesis the overset method, also called chimera or overlapping meshes, is discussed and applied toa formula race car, in order to calculate its aerodynamic map. The proposed method would allow reducing set-uptime through automation and avoided re-meshing process. a A theoretical background is presented before thediscussion of the way this kind of approach has been set-up in Star-CCM+. Results are obtained and discussedfor various car positions. Further investigations are finally suggested to further assess the viability of the method.

The report hereby presents the work performed within the Acoustic design and R&T service of Safran Aircraft Engines in Vi11arnche through a six month master thesis. The service is in charge of the development of noise prediction methods, analysis of tests results, design supervision, definition of specifications. The primary purpose of the present work was to develop a method of assessing the far-field pressure of the turbofan engine without far-field measurements. The foundation of the method relying on a particular signal processing method called beamforming, a complete explanation of the so-called method was realized according to the bibliography, which detects with precision unique modes. The numerical process of the method has been entirely developed, and a law to find the relation between the in-duct and farfield directivity angles was conceived and explained. The validation of the transfer function method was started, using the measurements of a test bench in Ecole Centrale de Lyon.

This paper deals with my master thesis done at Airbus Helicopters from the 2nd of February to the 31st of July 2015 in order to complete my studies for the Ecole Centrale de Lille and the Royal Institute of Technology as a double degree student. Airbus Helicopters is the leader on the civil helicopter market and one of the main actors on the military market. Therefore, innovative concepts are perpetually being tested to improve actual rotorcraft. During the development process of a new helicopter, wind tunnel tests are of prime interest to validate the design of different helicopter parts. In that perspective, the goal of my work is to improve the post-processing of wind tunnel data in order to statute about the efficiency of different shapes tested. This is done first by finding relevant handling qualities criteria and values for the aerodynamic coefficients, then by looking at the coefficients computed during wind tunnel tests and comparing them to the criteria to get the optimum configurations.

The field of computational fluid dynamics (CFD) is exponentially growing in terms of performance, robustness, and applications. The expansion of CFD also means more users and more simulations, which translates into more human errors and mistakes in the simulation set up. Because the simulation set up should be the correct in order to accurately reproduce the desired phenomenon, such errors must be mitigated in order to increase the reliability and robustness of the simulations. In this project a tool has been developed to tackle this issue, within the CFD software SIMULIA PowerFLOW. The tool extracts and analyzes the data of the cases before simulation, reporting the results to the user for error detection. The present work aims to present the implementation, the application and the benefits of the designed tool.

A person falling of a ship can be difficult to locate since the wake behind the ship forms achaotic field, making it extremely difficult to predict the location of the victim even if thetime when they fell overboard is known. Survivability for humans immersed at sea is verydependent on the time spent in the water, and varies significantly with sea temperature; thismakes it imperative that the victim is retrieved rapidly. Our current research is aimed atreducing this time using several UAV's searching the ships wake simultaneously, as a swarm.Since the wake is chaotic, a simulation was developed to model different random motions of avictim based on a chaotic equation. Our current research is making use of an establishedsimulator environment and developing it further to investigate how different platforms mayaffect rescue time, varying on the size of the ship, the weather conditions and whether thesearch is operated during day or night. Two different search strategies were implemented inthe developed simulator; these are Expanding Square search and Parallel search. An overallconclusion based on the results obtained is that the expanding square search tends to be amore rigid and reliable search strategy. Also the results show that for any scenario, the soughtperson is detected within minutes.

A variable-order finite-element method for the solution of the steady nonlinear potential flow equations is presented. To achieve robustness and computational efficiency, the formulation is restricted to purely subsonic flow by means of a density Modification in sonic flow regions. A test case that triggers the activation of this modification is presented to show that the method yields pressure results that are very close to those obtained with a mature Euler solver while reducing computational cost by an order of magnitude. Linear and quadratic elements are implemented, and the substantial benefit of using higher-order elements is demonstrated by means of a mesh-convergence study, showing how the convergence of induced drag and neutral point location is improved by the use or quadratic elements. For large surface meshes, the computational cost is found to be competitive with a linearized-potential boundary-element code accelerated by panel clustering.

The effect of nonlinear friction and freeplay in the control system on flut- ter is investigated for the case of light aircraft with manually operated control surfaces. A standard linear modal subspace model of the aeroelastic system is extended with nonlinear terms, so that time-domain simulations can be per- formed. Furthermore, the harmonic balance method is employed to obtain a frequency-domain formulation, which allows a convenient, though approxi- mate, computation of the stability boundary. Comparison with time-domain analyses for the case of a full aircraft configuration show that typical magni- tudes of hinge friction can stabilize oscillations with small displacement am- plitudes, while large amplitude motion remains unstable. Harmonic balance methods can be used to determine if such a behavior is present, but time- domain computations are necessary to evaluate the response of the system to a given excitation.

An efficient method for the non-iterative solution of the nonlinear flutter eigenvalue problem is presented. The properties of the piecewise quadratic decomposition employed make it particularly suitable for the parallel solution of aeroelastic stability problems where interaction of rigid-body and elastic motion is of interest or where the flutter damping must be obtained with accuracy, e.g. when comparison to flutter flight testing is intended. Ad- ditionally, the method allows the use of Laplace-domain aerodynamics if available. The paper presents a realistically complex test case which exposes some of the advantages and current shortcoming of the proposed solution procedure.

An approach to minimize the induced drag of an aeroelastic configuration by means of multiple leading- and trailing-edge control surfaces is investigated. A computational model based on a boundary-element method is constructed and drag-reducing flap settings are found by means of numerical optimization. Further, experiments with an elastic wind-tunnel model are performed in order to evaluate the numerically obtained results. Induced-drag results are obtained by analyzing lift distributions computed from optically measured local angles of attack because standard techniques proved insufficient. Results show that significant reductions of induced drag of flexible wings can be achieved by using optimal control surface settings.

An existing finite-element solver for steady, non-linear subsonic flow is extended in order to treat unsteady potential flow problems in the Laplace domain. The numerical formulation makes use of unstructured meshes including explicitly modelled, triangulated wake surfaces. Meshes can contain both linear and quadratic volume and surface elements, thus allowing to favor either very fast solution times or high spatial resolution. Aeroelastic problems are dealt with by modelling moving or deforming bodies by means of transpiration boundary conditions. Deformations of the aerodynamic mesh are computed either by projection of aerodynamic mesh nodes onto the finite elements of a structural shell model, or by radial basis function interpolation suitable for beam-type structural models. In addition to simple validation cases, an application of the solver for the evaluation of gust loads on a commuter aircraft is presented. In order to evaluate the use of the method in the context of a relevant, industrial-scale load analysis, typical geometrical and structural models for a twin-turboprop aircraft in the 15t-class (e.g. Saab 340, CN-235, Dash 8, Do 328) are employed.

An open-source implementation of an efficient mesh generation procedure for hybrid prismatic-tetrahedral meshes intended for use in Reynolds-averaged Navier-Stokes solutions is presented. The method employed combines the established, and very fast, Delaunay-based tetrahedral mesh generator TetGen with a novel technique for the creation of a prismatic layer, where constrained global optimization of the envelope is employed. Once a well-shaped envelope is thus obtained, a semi-structured layer of pentahedral elements is extruded between wall and envelope surface. Satisfactory mesh quality is demonstrated by comparing solutions obtained using the new meshes with reference data computed on high-quality advancing-front grids. Mesh generation time is shown to be substantially smaller than with many other methods. Overall, the presented implementation is deemed a valuable tool for cases where many meshes need to be generated for routine analyses and turnaround time is critical. 

The objectives of this study were to prepare the investment for a new facility in order to mechanise the flow and workstations on the Falcon fuselage equipped skin panels production line. This work took part of the in-production reactivity improvement (equivalent to LEAN) project within the company. 

One of the major challenges for helicopter pilots are low level flights and landings in degraded visual environments (DVE). Without proper assistance systems, the pilots are prone to lose their situational awareness (SA) when fog, heavy precipitation, limited sunlight and stirred-up sand or snow degrades their view. In recent years, various synthetic and enhanced vision systems were de-veloped so as to assist the pilots in these demanding situations. This work enhances the existing sys-tems by proposing a concept for the visualization of traffic information in helmet-/head-mounted displays (HMD). The intuitive representation provides additional cues about the environment and decreases the pilots’ workload, especially during flights in offshore windparks or while search and rescue operations with many other vehicles operating within a small range.The thesis, which was created at Airbus Defence & Space Friedrichshafen, analyzes the strengths and weaknesses of conventional head-down cockpit displays of traffic information (CDTIs) and of a very basic head-up traffic cueing system. Based on the results of this assessment, a prototype implementation of an integrated traffic visualization concept is developed where both display types – HDD and HMD – complement each other in their roles. All traffic information is retrieved from the on-board ADS-B system, which provides data such as the position and velocity vector of ADS-B equipped vehicles in the vicinity.The main focus of the work is placed on the development of methods for de-cluttering the HMD and increasing the information content of the head-up symbology by coding important parame-ters visually. In order to create a de-cluttered representation, the high-dimensional traffic data is clustered with a derivation of DBSCAN so as to identify groups of traffic that can be visualized by a single symbol. Ellipsoidal shapes are applied to airborne traffic while rectangular forms high-light vehicles on ground. To avoid distraction to the pilot due to sudden symbology changes when clusters are merged or split up, a smooth transition visualization method is developed. For the graphical representation of important information on the HMD, a measure indicating the threat potential of an intruder is proposed. This parameter, which is derived from the classification ap-proach of the TCAS II, is visualized by the color of the traffic symbol. In analogy to the TCAS, it is faded from green via yellow to red when the time to a predicted conflict decreases.The realized functions are integrated into a flight simulator and tested with synthetic ADS-B data. Furthermore, the proposed algorithms are evaluated in terms of their effectiveness and run-time efficiency, and finally the advantages and limitations of the approach are discussed. Now, the application must be integrated into an operational synthetic vision system so as to conduct further tests.

This project outlines a cost-effective numerical simulation method for the analysis of the longitudinal mechanics of paraglider systems. It is built on static stability methods for the analyses of low subsonic aircraft, non-linear lifting line methods for aerodynamic parameterization, and frequency domain analysis methods derived from system theory. Paragliders possess a glide polar in the range of ≈ 25-60 km.h−1 and display underdamped dynamic responses dominated by a long-period mode. The simulation results for performance and dynamic response are qualitatively valid relative to experimental data and in the same order of magnitude.

The continuous progress made in computer science and mathematics has allowed simulation to take a growing place in modern engineering. It is today a strategic stake for research as well as industry. Indeed, it allows a better apprehension of many fields of study, going from physics to biology through economics to mention a few. In aeronautics, simulation might come before production to avoid any fatal error in aircraft safety, or after production as a maintenance tool. It has many purposes, but it is always a compromise between the degree of subtlety of modeling and the calculation capacity of a computer.

This paper presents and compares two simulation codes in rotorcraft engineering. Both are used professionally by private companies. The main purpose is to enter one complete model of helicopter in both simulation software and to compare their outputs, understanding the models used in each case. The helicopter model is entered gradually, from main rotor to stabilizing back elements. Final results display the outputs of trim simulations performed on both programs.

The tip leakage vortex is responsible for the generation of stagnation pressure losses inside the compressor along with the outbreak of rotating stall and surge. The current paper analytically proved that a part of the losses is proportional to the vortex circulation squared. The evolution of this circulation has been investigated as part of a parametric study which tested several clearance heights.

The work consists in adopting a simplified single blade configuration to study the physics of the flow by means of wind tunnel experiments and numerical calculations. Upon visualising the main features of the flow, a model based on the study of jet in crossflows was implemented to describe the tip clearance flow for small gap sizes. For big gaps, the flow is assumed to behave as an isolated wing tip vortex which circulation is easily computed by the so called lifting line theory. The main vortical structures highlighted by the topology of the flow justified the use of the model of a jet in crossflow for small gap sizes. This model was challenged by experimental and numerical data and proved to well predict the evolution of the clearance vortex circulation for an increasing clearance height although some numerical results remain further away from the model.

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

This paper deals with the strategy for developing a very specific mechanism simulator. This mechanism is currently designed by Safran Group. A previous version of the same mechanism does exist but the scale is not on a like-for-like basis. Therefore, physical phenomena involved in this mechanism and their magnitude are not comparable to the previous version of the mechanism and this is why a new simulator has been developed (specified, coded and validated) from scratch. The paper addresses the strategy adopted for modelling a mechanism laying on a wide set of parameters as well as its use.

This thesis was performed with the purpose of investigate what influence different control system characteristics have on pilots' perception of handling qualities. Two gliders were investigated and compared, the JS1 [1] and the ASH26 [2]. Both are high performance gliders with similar type of configuration and aerodynamic performance, but with slightly different handling qualities. The stick force transfer functions in roll axis were calculated, the rate of angular change for a certain control input was measured and the subjective experience of the pilots were acquired through a pilot survey from flight tests. The pilots' experiences were then linked to how the gliders reacted on control input and control system design. It was found that the most important feature of the control system was control harmony, meaning that the responses in each control axis of the glider are in correspondence with each other. Also, the relationship between the stick force and response was much more important than the actual force itself.

The levelized cost of hydrogen for municipal fuel cell buses has been determined using the DOE H2A model for steam methane reforming (SMR), molten carbonate fuel cell reforming (MCFC), and wood gasification using wastewater biogas and willow wood chips as energy feedstocks. 300 kg H2/day was chosen as the design capacity.

Greenhouse gas emissions were calculated for each for the three processes and compared to diesel bus emissions in order to assess environmental impact. The levelized cost for SMR, MCFC, and gasification is $5.12, $8.59, and $10.62, respectively. SMR provided the lowest sensitivity to feedstock price, and lowest levelized cost at various scales, with competitive cost to diesel on a cost/km basis. All three technologies provide a reduction in total greenhouse gases compared to diesel bus emissions, with MCFC providing the largest reduction. These results provide preliminary evidence that small scale distributed hydrogen production for public transportation can be relatively cost-effective and have minimal environmental impact.

An approach to assess critical fuel configurations using robust flutter analysis is presented. A realistic aircraft model is considered to demonstrate how an available finite element model can be adapted to easily apply robust flutter analysis with respect to structural variations, such as fuel level variations. The study shows that standard analysis tools can be used to efficiently generate the system data that is required to perform robust flutter analysis. The μ-k method is used to compute the worst-case flutter speed and the corresponding worst-case fuel configuration is found. The main advantage of the proposed approach is that μ analysis guarantees robustness with respect to all possible fuel configurations represented by the tank model.

A study of the viabliness of a UAV for humanitarian relief missions has been done. With an aeronautical model fulfilling humanitarian logistics has then the manufacture and operational life-cycle costs as well as the present air jurisdiction. In contrast to this has also an evaluation of the empty weight conceptual estimation equation been reviewed and what conditions fairly describes the sizing of a UAV.

The study finds there’s a technical possibility, economic plausibility but a need for legal development.

The coefficients which are sufficient to estimate the empty weight ratio are the Home built aircrafts and powered sailplane. The study concludes what present off-the-shelf systems that might be suitable to sustain humanitarian relief missions.

Aircraft design is an act of art requiring dedication and careful work to ensure good results. An essential tool in that work is a flight mechanics simulator. Such simulators are often built up of modules/models that are executed in a sequential order in each time iteration. This project aims to analyze potential improvements to the model execution order based on the dependency structure of one such simulator. The analysis method Design Structure Matrix (DSM), was used to define/map the dependencies and then Binary Linear Programming (BLP) was utilized to find five new potentially improved model orders to minimize the number of feedbacks from one iteration to the next one. Those five proposed execution orders were next compared and evaluated. The result is a model order that reduce the number of models receiving feedbacks from the previous iteration from 13 to 6, with insignificant changes in the precision of the simulator.