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To improve the autonomous flight capability of endo-atmospheric flight vehicles, such as cruise missiles, drones, and other small, low-cost unmanned aerial vehicles (UAVs), a novel minimum-effort waypoint-following differential geometric guidance law (MEWFDGGL) is proposed in this paper. Using the classical differential geometry curve theory, the optimal guidance problem of endo-atmospheric flight vehicles is transformed into an optimal space curve design problem, where the guidance command is the curvature. On the one hand, the change in speed of the flight vehicle is decoupled from the guidance problem. In this way, the widely adopted constant speed hypothesis in the process of designing the guidance law is eliminated, and, hence, the performance of the proposed MEWFDGGL is not influenced by the varying speed of the flight vehicle. On the other hand, considering the onboard computational burden, a suboptimal form of the MEWFDGGL is proposed to solve the problem, where both the complexity and the computational burden of the guidance law dramatically increase as the number of waypoints increases. The theoretical analysis demonstrates that both the original MEWFDGGL and its suboptimal form can be applied to general waypoint-following tasks with an arbitrary number of waypoints. Finally, the superiority and effectiveness of the proposed MEWFDGGL are verified by a numerical simulation and flight experiments.

Purpose Although windy weather conditions have a significant effect on the flight safety and stability of any aircraft, the fact that quadrotors are lighter than other aircraft makes them more sensitive to the wind. This study aims to examine the extent to which quadrotors and their sensors, which are used in many fields and whose use is expected to increase significantly in the future, are affected by wind. Design/methodology/approach Flight experiments were carried out on different routes assigned by using Pixhawk Holybro 4 and Radiolink flight controllers. In these flight experiments, quadrotors were exposed to winds at different speeds and directions. Findings In the flight experiments, the deviation amounts in the quadrotor’s route at which wind speed was determined, and it was seen that these deviations were very serious and affected the safe flight at high wind speeds. According to the sensor information obtained from both different quadrotors’ flight experiments at different wind speeds, it was determined that the wind decreased the sensors’ accuracy. Practical implications It is foreseen that the data obtained in this study will be a source to be used in the design of quadrotors to be used in public areas in the future and to take the wind into account for safe flight. Originality/value In this study, numerous flight tests were carried out experimentally at various speeds from low speeds to high speeds on different routes using different flight controllers. The deviation data on the obtained routes and the effect of the wind on the sensors are experienced in real atmospheric conditions.

Having an exciting array of applications, the scope of unmanned aerial vehicle (UAV) application could be far wider one if its flight endurance can be prolonged. Solar-powered UAV, promising notable prolongation in flight endurance, is drawing increasing attention in the industries’ recent research and development. This work arose from a Bachelor’s degree capstone project at Hong Kong Polytechnic University. The project aims to modify a 2-metre wingspan remote-controlled (RC) UAV available in the consumer market to be powered by a combination of solar and battery-stored power. The major objective is to greatly increase the flight endurance of the UAV by the power generated from the solar panels. The power system is first designed by selecting the suitable system architecture and then by selecting suitable components related to solar power. The flight control system is configured to conduct flight tests and validate the power system performance. Under fair experimental conditions with desirable weather conditions, the solar power system on the aircraft results in 22.5% savings in the use of battery-stored capacity. The decrease rate of battery voltage during the stable level flight of the solar-powered UAV built is also much slower than the same configuration without a solar-power system.

We present analyses of Global Navigation Satellite System (GNSS) carrier phase observations in multiple kinematic scenarios for different receiver types. Multi-GNSS observations are recorded on high sensitivity and geodetic-grade receivers operating on a moving zero-baseline by conducting terrestrial urban and aerial flight experiments. The captured data is post-processed, carrier phase residuals are computed using the double difference (DD) concept. The estimated noise levels of carrier phases are analysed with respect to different parameters. We find DD noise levels for L1 carrier phase observations in the range of 1.4&ndash, 2 mm (GPS, Global Positioning System), 2.8&ndash, 4.6 mm (GLONASS, Global Navigation Satellite System), and 1.5&ndash, 1.7 mm (Galileo) for geodetic receiver pairs. The noise level for high sensitivity receivers is at least higher by a factor of 2. For satellites elevating above 30 ∘ , the dominant noise process is white phase noise. For the flight experiment, the elevation dependency of the noise is well described by the exponential model, while for the terrestrial urban experiment, multipath and diffraction effects overlay, hence no elevation dependency is found. For both experiments, a carrier-to-noise density ratio (C/N 0 ) dependency for carrier phase DDs of GPS and Galileo is clearly visible with geodetic-grade receivers. In addition, C/N 0 dependency is also visible for carrier phase DDs of GLONASS with geodetic-grade receivers for the terrestrial urban experiment.

We hypothesized that an incorrect expectation due to spatial disorientation may induce roll reversal errors. To test this, an in-flight experiment was performed, in which forty non-pilots rolled wings level after receiving motion cues. A No-leans condition (subthreshold motion to a bank angle) was included, as well as a Leans-opposite condition (leans cues, opposite to the bank angle) and a Leans-level condition (leans cues, but level flight). The presence of leans cues led to an increase of the roll reversal error (RRE) rate by a factor of 2.6. There was no significant difference between the Leans-opposite and Leans-level condition. This suggests that the expectation strongly affects the occurrence of an RRE, and that people tend to base their responses on motion cues instead of on information on the AI. We conclude that expectation and spatial disorientation have a large effect on piloting errors and may cause hazardous aircraft upsets. © 2019 Elsevier Ltd

We hypothesized that an incorrect expectation due to spatial disorientation may induce roll reversal errors. To test this, an in-flight experiment was performed, in which forty non-pilots rolled wings level after receiving motion cues. A No-leans condition (subthreshold motion to a bank angle) was included, as well as a Leans-opposite condition (leans cues, opposite to the bank angle) and a Leans-level condition (leans cues, but level flight). The presence of leans cues led to an increase of the roll reversal error (RRE) rate by a factor of 2.6. There was no significant difference between the Leans-opposite and Leans-level condition. This suggests that the expectation strongly affects the occurrence of an RRE, and that people tend to base their responses on motion cues instead of on information on the AI. We conclude that expectation and spatial disorientation have a large effect on piloting errors and may cause hazardous aircraft upsets. © 2019 Elsevier Ltd

The present thesis develops a new control approach for scale-model airplanes. The proposed control solutions exploit a simple but pertinent nonlinear model of aerodynamic forces acting on the aircraft. Nonlinear controllers are based on a hierarchical structure, and are derived on the basis of theoretical stability and convergence analyses. They are designed to operate on a large spectrum of operating conditions. In particular, they avoid the singularities associated with the parameterization of the attitude and the heading of the vehicle, and do not rely on a decoupling between longitudinal and lateral dynamics. First, the trajectory tracking problem is addressed by extending the thrust vectoring method used for small rotor vehicles to the case of fixed wing vehicles. In the case of airplanes, the main challenge is to take into account the aerodynamic forces in the design of control systems. In order to solve this problem, the proposed control is designed and analyzed on the basis of the proposed aerodynamic forces model. The flight envelope is thus broadened beyond trim trajectories which are classically used in the literature. This solution is then adapted to the path following problem, and kinematic guidance and dynamic control laws are developed within a single coherent framework that applies to almost any regular 3D path. The proposed control laws incorporate integral terms that robustify the control with respect to unmodelled dynamics. Several practical issues are addressed and the proposed control laws are validated via hardware-in-the-loop simulations. Finally, successful flight test results illustrate the soundness and performance of the proposed control laws.; Cette thèse développe une nouvelle approche de contrôle pour les avions à échelle réduite. Les lois de commande proposées exploitent un modèle non linéaire simple mais pertinent des forces aérodynamiques appliquées à l’aéronef. Ils reposent sur une structure hiérarchique de contrôle non linéaires, et sont synthétisées sur la base d’analyse de stabilité et de convergence théoriques. Ils sont conçus pour fonctionner sur un large domaine de vol. En particulier, ils évitent les singularités associées à la paramétrisation de l'attitude et la direction de la vitesse. Dans un premier temps, le problème de stabilisation de trajectoires de référence est résolu en étendant la méthode du "thrust vectoring", utilisée pour les véhicules à voilure tournante, au cas des aéronefs à voilure fixe. Dans le cas des avions, le principal défi est de prendre en compte les forces aérodynamiques dans la conception des systèmes de commande. Afin de résoudre ce problème, le contrôle proposé est conçu et analysé sur la base du modèle de forces aérodynamique proposé. Le domaine d'utilisation de cette loi de commande est élargi et englobe les trajectoires d'équilibre (trim trajectories) qui sont classiquement utilisées dans la littérature. Cette solution est ensuite adaptée au problème de suivi de chemin, afin de concevoir des lois de guidage cinématique et de contrôle dynamique applicables à presque tout chemin 3D régulier. Les lois de contrôle proposées contiennent des termes intégraux qui robustifient le contrôle vis-à-vis de dynamiques non modélisées. Plusieurs problèmes pratiques sont adressés et les lois de commande proposées sont validées par des simulations du type "hardware-in-the-loop". Enfin, des résultats d'essais en vol illustrent la performance des lois de contrôle proposées.

The present thesis develops a new control approach for scale-model airplanes. The proposed control solutions exploit a simple but pertinent nonlinear model of aerodynamic forces acting on the aircraft. Nonlinear controllers are based on a hierarchical structure, and are derived on the basis of theoretical stability and convergence analyses. They are designed to operate on a large spectrum of operating conditions. In particular, they avoid the singularities associated with the parameterization of the attitude and the heading of the vehicle, and do not rely on a decoupling between longitudinal and lateral dynamics. First, the trajectory tracking problem is addressed by extending the thrust vectoring method used for small rotor vehicles to the case of fixed wing vehicles. In the case of airplanes, the main challenge is to take into account the aerodynamic forces in the design of control systems. In order to solve this problem, the proposed control is designed and analyzed on the basis of the proposed aerodynamic forces model. The flight envelope is thus broadened beyond trim trajectories which are classically used in the literature. This solution is then adapted to the path following problem, and kinematic guidance and dynamic control laws are developed within a single coherent framework that applies to almost any regular 3D path. The proposed control laws incorporate integral terms that robustify the control with respect to unmodelled dynamics. Several practical issues are addressed and the proposed control laws are validated via hardware-in-the-loop simulations. Finally, successful flight test results illustrate the soundness and performance of the proposed control laws.; Cette thèse développe une nouvelle approche de contrôle pour les avions à échelle réduite. Les lois de commande proposées exploitent un modèle non linéaire simple mais pertinent des forces aérodynamiques appliquées à l’aéronef. Ils reposent sur une structure hiérarchique de contrôle non linéaires, et sont synthétisées sur la base d’analyse de stabilité et de convergence théoriques. Ils sont conçus pour fonctionner sur un large domaine de vol. En particulier, ils évitent les singularités associées à la paramétrisation de l'attitude et la direction de la vitesse. Dans un premier temps, le problème de stabilisation de trajectoires de référence est résolu en étendant la méthode du "thrust vectoring", utilisée pour les véhicules à voilure tournante, au cas des aéronefs à voilure fixe. Dans le cas des avions, le principal défi est de prendre en compte les forces aérodynamiques dans la conception des systèmes de commande. Afin de résoudre ce problème, le contrôle proposé est conçu et analysé sur la base du modèle de forces aérodynamique proposé. Le domaine d'utilisation de cette loi de commande est élargi et englobe les trajectoires d'équilibre (trim trajectories) qui sont classiquement utilisées dans la littérature. Cette solution est ensuite adaptée au problème de suivi de chemin, afin de concevoir des lois de guidage cinématique et de contrôle dynamique applicables à presque tout chemin 3D régulier. Les lois de contrôle proposées contiennent des termes intégraux qui robustifient le contrôle vis-à-vis de dynamiques non modélisées. Plusieurs problèmes pratiques sont adressés et les lois de commande proposées sont validées par des simulations du type "hardware-in-the-loop". Enfin, des résultats d'essais en vol illustrent la performance des lois de contrôle proposées.

Identification process of dynamic systems, such as the case of an aircraft for control tunning, is a common practice in engineering. Nowadays several flight, wind tunnel or CFD databases are used to carry it out. However, even involving a great deal of e ort both economically and temporally, the validation of the model and therefore the aplicability of the model is restricted to the area of the envelope where the flight tests have been gathered. This fi nal bachelor thesis demonstrates how decoupled simulations of flight data from complete maneuvers, both linear and nonlinear, are able to support the local identi cation of the system, a great help in local system identi cation (in the data collection area), and how they can greatly reduce the amount of required flight data. On the other hand, a new system identi cation scheme called Global System Identi cation will be tested. Aerodynamic simulations based on the Vortex Lattice Method are proposed as the fi rst approach in areas where flight data collection is not possible, allowing the extrapolation of the model to that area building correction factors. The validation of the simulations in the local area is carried out using the flight data from two di erent aircraft, and veri cation of the proposed scheme for global identi cation is performed by extrapolation in the known zone (ie, the model identi ed at a point of the envelope is corrected for another by the simulations (which in this case also exists flight data) and are veri ed with the flight data. A first approach for modeling by extrapolation at high altitude is proposed, in the development phases of HALE platforms., La identificación de sistemas dinámicos, como el caso de una aeronave para el modelado del controlador, es una práctica habitual en ingeniería. Hoy en día se usan grandes bases de datos de vuelo para llevarlo a cabo, datos de túnel de viento o CFD. Sin embargo incluso implicando una gran cantidad de esfuerzo tanto en lo económico como en lo temporal, la validación del modelo y por tanto el mismo está restringido a la zona de la envolvente la cual se han realizado los test de vuelo. En este Trabajo de Fin de Grado se demuestra cómo simulaciones desacopladas de los datos de vuelo de maniobras completas, tanto lineales como no lineales, son de gran ayuda en la identi cación local del sistema (en la zona de recolección de datos) y cómo pueden reducir en gran medida los datos de vuelo necesarios. Por otra parte se pondría a prueba un nuevo esquema para la identi cación de sistemas llamado Identi cación Global del Sistema. En éste se proponen simulaciones aerodinámicas, basadas en Vortex Lattice Method como primera aproximación, en las zonas donde no es posible la recolección de datos de vuelo permitiendo mediante el proceso de identi cación de dichas maniobras sintéticas, la extrapolación del modelo a dicha área. La validación de las simulaciones en la zona local es realizada mediante los datos de vuelo para dos aeronaves distintas, y la veri cación del esquema propuesto para la identi cación global es realizada mediante extrapolación en la zona conocida (Es decir, el modelo identifi cado en un punto de la envolvente se corrige para otro mediante las simulaciones) y son veri ficadas con los datos de vuelo. Se propone entonces una primera aproximación al modelado por extrapolación a gran altitud en fases de desarrollo de plataformas HALE., DLR, German Aerospace Center, Institut für Robotik und Mechatronik, 178 páginas

High Altitude Long Endurance (HALE) platforms are the aerial platforms capable of flying in the stratosphere for long periods of time. This master thesis presents aircraft system identification procedures geared towards such fixed wing platforms where aerodynamic forces and moments are parametrically modelled with so-called stability and control derivatives. The first part of the thesis addresses local System identification procedures intended for controller synthesis at low altitude flights whereas the second part of the thesis deals with a preliminary study on a new global system identification method. The local system identification procedure is based on the two step method, which offers flexibility regarding the aerodynamic structure. Therefore, it is suitable for the development of a system identification tool chain for various fixed wing platforms. Various system identification experiments have been conducted to collect flight test data. The parameters for the estimation of aerodynamic forces and moments are then found through an optimization procedure. Such parameters have been validated using a validation set from flight test data and their applicability for controller synthesis has been demonstrated. Global system identification typically requires the collection of flight test data at multiple points in the flight envelope and often, is combined with extensive Computational Fluid Dynamics (CFD) solutions as well as wind-tunnel experiments. Such an approach is time consuming and costly. This thesis presents a new method to overcome the limitations of the current methodology by applying a Parameter search on VLM-based (Vortex Lattice Method) dynamic simulations of aircraft System identification manoeuvres and correcting the estimated models with available flight test data. The current study shows improvements in fidelity with decrease in Root Mean Squared Error (RMSE) by factor 0.2 and 0.5 for x-axis and z-axis forces in body frame respectively, while reducing the effort for obtaining a model with similar fidelity.

We report the first results of the LISA Pathfinder in-flight experiment. The results demonstrate that two free-falling reference test masses, such as those needed for a space-based gravitational wave observatory like LISA, can be put in free fall with a relative acceleration noise with a square root of the power spectral density of 5.2±0.1 fm s−2/Hz, or (0.54±0.01)×10−15 g/Hz, with g the standard gravity, for frequencies between 0.7 and 20 mHz. This value is lower than the LISA Pathfinder requirement by more than a factor 5 and within a factor 1.25 of the requirement for the LISA mission, and is compatible with Brownian noise from viscous damping due to the residual gas surrounding the test masses. Above 60 mHz the acceleration noise is dominated by interferometer displacement readout noise at a level of (34.8±0.3) fm/Hz, about 2 orders of magnitude better than requirements. At f≤0.5 mHz we observe a low-frequency tail that stays below 12 fm s−2/Hz down to 0.1 mHz. This performance would allow for a space-based gravitational wave observatory with a sensitivity close to what was originally foreseen for LISA. CNES 1316634/CNRS 103747 UnivEarthS Labex program/ANR-10-LABX-0023 UnivEarthS Labex program/ANR-11-IDEX-0005-02 DLR Federal Ministry for Economic Affairs and Energy/FKZ 50OQ0501 Federal Ministry for Economic Affairs and Energy/FKZ 50OQ1601 Agenzia Spaziale Italiana Instituto Nazionale di Fisica Nucleare AYA2010-15709 (MICINN) ESP2013-47637-P (MINECO) ESP2015-67234-P (MINECO) Fundacion General CSIC Swiss Space Office (SSO) Swiss National Science Foundation United Kingdom Space Agency (UKSA) University of Glasgow University of Birmingham Imperial College Scottish Universities Physics Alliance (SUPA) U.S. National Aeronautics and Space Administration (NASA)

The LISA Pathfinder mission will demonstrate the technology of drag-free test masses for use as inertial references in future space-based gravitational wave detectors. To accomplish this, the Pathfinder spacecraft will perform drag-free flight about a test mass while measuring the acceleration of this primary test mass relative to a second reference test mass. Because the reference test mass is contained within the same spacecraft, it is necessary to apply forces on it to maintain its position and attitude relative to the spacecraft. These forces are a potential source of acceleration noise in the LISA Pathfinder system that are not present in the full LISA configuration. While LISA Pathfinder has been designed to meet it's primary mission requirements in the presence of this noise, recent estimates suggest that the on-orbit performance may be limited by this `suspension noise'. The drift-mode or free-flight experiments provide an opportunity to mitigate this noise source and further characterize the underlying disturbances that are of interest to the designers of LISA-like instruments. This article provides a high-level overview of these experiments and the methods under development to analyze the resulting data., 13 pages, 5 figures. Accepted to Journal Of Physics, Conference Series. Presented at 10th International LISA Symposium, May 2014, Gainesville, FL, USA

資料番号: PA0910106000

Real-time monitoring allows new possibilities in applications like disaster management or traffic observation and guidance. The German Aerospace Center is currently developing an aircraft based observation system. Among other sensors a high resolution camera platform together with an optical downlink terminal is an integral part of the system. The optical terminal was tested in the first stage of expansion in November and December 2008. At distances up to 85 km the achieved mean tracking offset with pure CPA tracking was 266 μrad. Initial communication tests have been successfully performed up to a distance of 40 km.

The increasing role of the computational fluid dynamics in the design process of space transportation vehicls is here emphasized. The coupling of CFD with many disciplines is being possible due to the use of modern computational cluster resources and due to the validation of the applied numerical tools by means of carefully designed experiments for ground based facilities and for in-flight demonstration. But CFD, wind tunnel test and flight experiments have also shortcomings. While high fidelity CFD results are today based on the solution of the Navier-Stokes equations, they still require validated physical models. Improvements of these models require considerable progress in novel measurement techniques and combination of systematic, efficient and more demanding ground tests and in-flight experiments. Thus, in order to advance hypersonic space transportation CFD, ground based testing and flight experiments, need to be continuously improved and linked in an optimal manner.

Here is presented an overview of the hypersonic experiment on sharp edge concepts, SHEFEX-I. The project, being performed under responsibility of the German Aerospace Center (DLR), is aimed to investigate the behavior and the possibilities of an improved shape for aerospace vehicles considering sharp edges and facetted surfaces. It is a basic in-flight experimentation research on hypersonic technologies for future launcher vehicles but not a re-entry experiment. Additionally, the SHEFEX-I project is the starting point for a series of experiments which enable the acquisition of important knowledge in hypersonic free flight experimentation and which are an excellent test bed for new technological concepts. The experiment which successfully flew on top of a two-stage solid propellant sounding rocket on October 27th, 2005 from Andøya Rocket Range in northern Norway, enabled time accurate investigation of the flow effects and their structural answer during a hypersonic flight Ma~6 from 90 km down to an altitude of 20 km. The present paper gives an overview about the aerothermodynamic philosophy and introduces some main outcomes of the post-flight analysis.