This report focuses on the utilisation of flight mechanics in the context of aircraftconceptual design to assess stability, control, and motion characteristics. The pri-mary objective is to acquire the equations of motion and implement longitudinalstability and control criteria using Pacelab Aircraft Preliminary Design 8.1, a com-mercial software tool. The equations and criteria employed in this study are derivedfrom an extensive review of relevant literature.By incorporating a dedicated Flight Mechanics chapter within the software, it be-comes possible to evaluate aircraft concepts under varying conditions. To ensureaccuracy and validity, DATCOM+ and OpenVSP were employed for testing andverification purposes.The key aspects covered in this report include flight mechanics, its implementationin Pacelab APD 8.1, determination of aerodynamic derivatives, formulation of equa-tions of motion, and their application to the B747 aircraft model. The emphasis liesin assessing longitudinal stability and control, including specific characteristics suchas the phugoid and short period modes.This report provides valuable insights into the integration of flight mechanics withinthe Pacelab APD 8.1 software for aircraft conceptual design. The results contributeto a better understanding of stability and control parameters and their impact onaircraft performance.

Ce mémoire de thèse traite du développement des dynamiques et des lois de commande de vol d’un projectile d’artillerie gyrostabilisé guidé par une tête découplée. Un modèle non linéaire du projectile est proposé, et sert à calculer un modèle linéarisé de la dynamique de roulis du nez et un modèle q-LPV des chaînes de tangage et de lacet à paramètres fortement variants. Les incertitudes de modélisation sont prises en compte pour concevoir l’autopilote. Des propriétés importantes des projectiles gyrostabilisés, qui sont liées au couplage dynamique tangage/lacet, aux modes internes et à la stabilité, sont mises en valeur grâce au modèle q-LPV. En vue de l’utiliser pour calculer une loi de commande, la dimension de son vecteur des paramètres est réduite et la position des capteurs intégrés dans le nez est considérée. Un seul correcteur linéaire est suffisant pour la dynamique de l’angle de roulis du nez alors qu’une stratégie systématique de commande par séquencement de gains basée sur une linéarisation est élaborée séparément pour générer un correcteur séquencé des facteurs de charge de tangage et de lacet. Des structures de commande fixées d’ordre réduit sont conçues en appliquant la même approche de synthèse linéaire H∞ par façonnage de gain de boucle pour les axes de roulis et de tangage/lacet. De très bonnes propriétés de performance et de robustesse en boucle fermée, comparables à celles fournies par des correcteurs d’ordre plein, sont obtenues. Finalement, l’efficacité de l’autopilote augmenté d’une loi de guidage par navigation proportionnelle pure est vérifiée via de nombreuses simulations non linéaires de trajectoires. Ces dernières correspondent à des scénarios de vol nominaux d’interception de cibles balistiques, non balistiques immobiles, ou manœuvrantes, ainsi qu’à des scénarios considérant des perturbations sur les conditions de tir ou sur les dynamiques du projectile guidé / This thesis addresses the development of the flight dynamics and control laws for an artillery spin-stabilized projectile equipped with a decoupled guidance nose. A projectile nonlinear model is discussed, and it is used for computing a linearized model of the nose roll dynamics along with a q-LPV model of the highly parameter-varying pitch/yaw-dynamics. Modeling uncertainty is taken into account for autopilot design. Important properties specific to spin-stabilized projectiles, which are relevant to pitch/yaw-channel cross-coupling, internal modes and stability, are highlighted using the q-LPV model. In order to use the latter for calculating a control law, the dimension of its parameter vector is reduced and the position of the nose-embedded sensors is considered. A single linear controller is sufficient for the nose roll angle dynamics whereas a systematic linearization-based gain-scheduled control strategy is separately devised to provide a pitch/yaw-axis load factor gain-scheduled controller. Controllers of reduced-order fixed structures are computed by applying the same H∞ linear design loop-shaping approach for the roll and pitch/yaw-axes. Very good closed-loop performance and robustness properties, which are similar to those provided by full order controllers, are obtained. Finally, the effectiveness of the autopilot augmented by a pure proportional navigation guidance law is verified through a variety of nonlinear trajectory simulations. The latter correspond to nominal flight scenarios with ballistic, non-ballistic stationary, and maneuvering interception points, and to scenarios with perturbed launch conditions or guided projectile dynamics

Cette thèse présente une méthodologie de conception et d’évaluation de lois de commande pour projectiles guidés, au moyen d’un prototype placé dans une soufflerie via un support autorisant plusieurs degrés de liberté en rotation. Ce dispositif procure un environnement permettant à la fois de caractériser expérimentalement le comportement de la munition et d’évaluer les performances des lois de commande dans des conditions réalistes, et est mis en œuvre pour l’étude d’autopilotes de tangage et de lacet, à vitesse fixe et à vitesse variable, pour un prototype de projectile empenné piloté par canards. La modélisation d’un tel système aboutit à un modèle non-linéaire dépendant de nombreuses conditions de vol telles que la vitesse et des angles d’incidence. Les méthodes de séquencement de gain basées sur des linéarisations d’un modèle non-linéaire sont couramment employées dans l’industrie pour la commande de ce type de systèmes. A cette fin, le système est représenté au moyen d’une famille de modèles linéaires dont les paramètres sont directement estimés à partir de données recueillies sur le dispositif expérimental. L’observation du comportement à différents points de vol permet de considérer la vitesse de l’air comme unique variable de séquencement. La synthèse des différents contrôleurs est réalisée au moyen d’une méthode H∞ multi-objectifs à ordre et structure fixes, afin de garantir la stabilité et la robustesse du système vis-à-vis d’incertitudes liées à la variation du point de fonctionnement. Ces lois de commande sont alors validées au moyen d’analyses de robustesse, puis par leur implémentation sur le dispositif expérimental. Les résultats obtenus lors d’essais en soufflerie correspondent aux simulations numériques et sont conformes aux spécifications attendues. / This work presents a novel methodology for flight control law design and evaluation, using a functional prototype installed in a wind tunnel by the means of a support structure allowing multiple rotational degrees of freedom. This setup provides an environment allowing experimental characterization of the munition’s behavior, as well as for flight control law evaluation in realistic conditions. The design and validation of pitch and yaw autopilots for a fin-stabilized, canard-guided projectile is investigated, at fixed and variable airspeeds. Modeling such a system leads to a nonlinear model depending on numerous flight conditions such as the airspeed and incidence angles. Linearization-based gain scheduling techniques are widely employed in the industry for controlling this class of systems. To this end, the system is represented with a family of linear models whose parameters are directly estimated from experimentally collected data. Observation of the projectile’s behavior for different operating points indicates the airspeed can be considered as the only scheduling variable. Controller synthesis is performed using a multi-objective, fixed-order, fixed-structure H∞ technique in order to guarantee the stability and robustness of the closed-loop against operating point uncertainty. The obtained control laws are validated with robustness analysis techniques and are then implemented on the experimental setup, where wind-tunnel tests results correlate with numerical simulations and conform to the design specifications.

Military helicopter operations require precise maneuverability characteristics for performance to be determined for the entire helicopter flight envelope. Historically, these maneuverability analyses are combinatorial in nature and involve human-interaction, which hinders their integration into conceptual design. A model formulation that includes the necessary quantitative measures and captures the impact of changing requirements real-time is presented. The formulation is shown to offer a more conservative estimate of maneuverability than traditional energy-based formulations through quantitative analysis of a typical pop-up maneuver. Although the control system design is not directly integrated, two control constraint measures are deemed essential in this work: control deflection rate and trajectory divergence rate. Both of these measures are general enough to be applied to any control architecture, while at the same time enable quantitative trades that relate overall vehicle maneuverability to control system requirements. The dimensionality issues stemming from the immense maneuver space are mitigated through systematic development of a maneuver taxonomy that enables the operational envelope to be decomposed into a minimal set of fundamental maneuvers. The taxonomy approach is applied to a helicopter canonical example that requires maneuverability and design to be assessed simultaneously. The end result is a methodology that enables the impact of design choices on maneuverability to be assessed for the entire helicopter operational envelope, while enabling constraints from control system design to be assessed real-time.

The present work focuses on the system identification method of aerodynamic parameter estimation which is used to calculate the stability and control derivatives required for aircraft flight mechanics. A new rotorcraft parameter estimation technique is proposed which uses a type of artificial neural network (ANN) called radial basis function network (RBFN). Rotorcraft parameter estimation using ANN is an unexplored research topic and the earlier works in this area have used the output error, equation error and filter error methods which are conventional parameter estimation methods. However, the conventional methods require an accurate non-linear rotorcraft simulation model which is not required by the ANN based method. The application of RBFN overcomes the drawbacks of multilayer perceptron (MLP) based delta method of parameter estimation and gives satisfactory results at either end of the ordered set of estimates. This makes the RBFN based delta method for parameter estimation suitable for rotorcraft studies, as both transition and high speed flight regime characteristics can be studied. The RBFN based delta method for parameter estimation is used for computation of aerodynamic parameters from both simulated and real time flight data. The simulated data is generated from an 8-DoF non-linear simulation model based on the Level-1 criteria of rotorcraft simulation modeling. The generated simulated data is used for computation of the quasi-steady and the time-variant stability and control parameters for different flight conditions using the RBFN based delta method. The performance of RBFN based delta method is also analyzed in the presence of state and measurement noise as well as outliers. The established methodology is then applied to compute parameters directly from real time flight test data for a BO 105 S123 helicopter obtained from DLR (German Aerospace Center). The parameters identified using the RBFN based delta method are compared with the identified values for the BO 105 helicopter from published literature which have used conventional parameter estimation techniques for parameter estimation using a 6-DoF and a 9-DoF rotorcraft simulation model. Finally, the estimated parameters are verified from the flight data generated by a frequency sweep pilot control input for assessing the predictive capability of the RBFN based delta method. Since the approach directly computes the parameters from flight data, it can be used for a reliable description of the higher frequency range, which is needed for high bandwidth flight control and in-flight simulation.

Les récentes recherches sur les rotors d'hélicoptère conduisent au développement de pales de nouvelle génération présentant des géométries courbes. La double flèche de la pale BlueEdgeTM proposée par Eurocopter impose de reconsidérer les outils de calcul des charges rotors pour déterminer le torseur des efforts appliqués aux pales et aux éléments constitutifs du moyeu rotor afin de satisfaire aux exigences de conception et de certification. Les charges rotors se décomposent en contributions aéro- et élasto-dynamiques prises en compte par des modélisations distinctes. La thèse vise à définir une méthodologie de calcul de charges applicable aux pales à double flèche. Ainsi sont présentés les modèles aérodynamiques bi-dimensionnels pour calculer les vitesses induites du rotor et déterminer la répartition des efforts aérodynamiques sur le rotor. Le calcul des charges rotor nécessite de recourir à des modèles élasto-dynamiques. En résolvant les équations de la dynamique des solides pour un système mécanique, le code de mécanique du vol HOST considère une modélisation élastique de pale pour déterminer le torseur des efforts, les efforts de commande étant fournis par l'ensemble bielle de pas et plateaux cycliques. Le comportement non linéaire des adaptateurs de traînée interpales est décrit par des modèles de force de restitution. Ces travaux ont utilisé des caractérisations expérimentales sur des machines de traction de laboratoire ainsi que des essais en vol afin d'évaluer le niveau de représentativité des outils et méthodes proposés. La mise en oeuvre de l'ensemble de ces modèles détermine avec satisfaction les charges dynamiques du rotor pour des vols stabilisés. / New generation blades have led to new load computation problems due to the evolution of the general shape, with forward and backward sweep. The BlueEdgeTM blade pattented by Eurocopter imposes to reconsider the development methodology and thus it is no longer possible to speak of straight blades and the models used for load computation have to be evaluated. The objective of this thesis is to determine what has to be modified and improved in current load computation methodology in order to reach an acceptable predictive level. This work considers both aerodynamic and dynamic models implemented in the HOST multi-body computer code. The aerodynamics models are based on the hypothesis of a two dimensional flow. The use of the CFD software \emph{elsA} is evaluated. Attention is given to rotor dynamics models that have an impact on loads, such as lead-lag damper models, blade element models and hub models. This thesis presents the different models and gives orientations relating to efficient load computation methodology. The aerodynamics models are compared to windtunnels experiments from the literature. This study leads also to perform flight tests and to investigate the dampers behavior on test benches in order to confront the computed loads to the reality of the helicopter operation. The proposed methodology is able to compute with a good accuracy rotor loads for stabilized flight cases.

Ziel - Die Zeit die ein Flugzeug benötigt, um auf eine bestimmte Höhe zu steigen (die Steigzeit) kann mit einer Formel berechnet werden, die vereinfachend annimmt, dass die Steiggeschwindigkeit über dem gesamten Steigflug mit zunehmender Höhe linear abnimmt. Ziel der Untersuchung ist, zu ermitteln, ob die Annahme einer linear abnehmenden Steiggeschwindigkeit realistisch ist bzw. welche Fehler sich aus der Annahme ergeben. ----- Methode - Mit der Höhe ändern sich Parameter wie Luftdichte, Widerstand, Schub und damit auch die optimale Fluggeschwindigkeit für den Steigflug. Die Parameter beeinflussen sich dabei gegenseitig. Der Schub wird dabei nach drei unterschiedlichen Methoden berechnet, gegeben von Bräunling, Scholz und Howe. Analysiert wird der Verlauf des Schubes mit der Höhe und der Verlauf der Steiggeschwindigkeit mit der Höhe für jede der drei Schubberechnungen. Abschließend wird für jede Schubberechnung die Steigzeit verglichen wie sie sich ergibt a) aus der einfachen Formel und b) aus einer Integrationsberechnung, bei der der Verlauf der Steiggeschwindigkeit durch eine Funktion beschrieben wird. ----- Ergebnisse - Die drei Schubberechnungen liefern ausgehend vom gleichen Startschub unterschiedliche Schübe in der Höhe. In die Methode nach Bräunling gehen mehr Parameter ein als in die anderen beiden Methoden. Es kann angenommen werden, dass die Methode nach Bräunling genauer ist, der Beweis kann aber nicht geführt werden. Der Schub nach Scholz und Howe fällt nahezu linear mit der Höhe ab. Der Schubverlauf nach Bräunling zeigt eine deutliche Nichtlinearität. Es wird die Steigzeit von 0 km auf 11 km Höhe berechnet nach a) und b), mit jeder der drei Schubberechnungen. Dabei wird jeweils der Unterschied in der Steigzeit ermittelt. Aufgrund der Nichtlinearität im Schubverlauf zeigt die Methode nach Bräunling dann auch den größten Unterschied zwischen den Berechnungsmethoden von 7,1 %. Bei einer Schubberechnung nach Scholz ergeben sich 1,7 % und nach Howe 1,4 %. Wenn bereits zu Beginn Vereinfachungen, z.B. bezüglich des Triebwerksschubes, vorgenommen wurden, ist es in Hinblick auf den Aufwand und die zu erreicheneden Ergebnisse möglich, und zum Teil sinnvoll, die Berechnungen der Steigzeit mittels linearer Abnahme der vertikalen Geschwindigkeit durchzuführen. Es wird ausdrücklich darauf hingewiesen, dass es hier um den Vergleich von zwei Methoden zur Berechnung der Steigzeit geht und nicht um die Bewertung von Methoden zur Schubberechnung (für die keine Vergleichswerte vorlagen). ----- Praktischer Nutzen - Es konnte festgestellt werden, dass eine einfache Formel zur Berechnung der Steigzeit mit geringem Fehler angewandt werden kann - insbesondere wenn Methoden zur Schubberechnung vorliegen, bei denen der Schub annähernd linear mit der Höhe abnimmt. Bei großem Aufwand und realitätsnaher Betrachtung, z.B. nach Bräunling, führt der lineare Ansatz jedoch zu einem zu großen Fehler. Hierfür sollte die Berechnung der Steigzeit mittels Integration durchgeführt werden.

Um método para avaliação do uso de ganho adequado em ensaios de verificação de PIO (Pilot Induced Oscillations) é apresentado. As tarefas de manobra sintética (Synthetic Tracking Task) são utilizadas para demonstração. A teoria é baseada no conceito de entropia estatística proveniente da teoria da informação e no modelo estrutural do piloto humano. O método é apresentado para manobras executadas no eixo lateral e oferece uma medida do ganho humano utilizado durante a sua execução em malha-fechada. Para a modelagem da planta é utilizado modelo black-box com equacionamento de espaço de estados e identificação de parâmetros. Dados de ensaios em voo provenientes de uma aeronave de transporte certificada FAR-25 são utilizados para medir a razão entre o tempo gasto pelo piloto humano em uma malha fechada em posição versus o tempo em uma malha de derivada da posição (roll vs. roll rate). Esta medida é proposta como validadora da execução correta do ensaio. / A method is proposed to verify losed-loop adequate flight test piloting gain in PIO aircraft certification. The synthetic tracking task PIO flight test is used. The theory is based on the entropy concept from information theory and on the structural pilot model of the human pilot. The method is presented for single axis pilot tracking maneuvers and offers a measure of the human pilot gain employed during its execution. A black-box, state-space, parameter-identified model is used for the plant. Flight test data from a FAR-25 transport aircraft is used to verify the theory of how to determine a measure of the ratio between time spent by the human pilot in the error loop versus in the error rate loop to control the aircraft. This measure is proposed as a test point validation method for PIO flight testing.

Um método para avaliação do uso de ganho adequado em ensaios de verificação de PIO (Pilot Induced Oscillations) é apresentado. As tarefas de manobra sintética (Synthetic Tracking Task) são utilizadas para demonstração. A teoria é baseada no conceito de entropia estatística proveniente da teoria da informação e no modelo estrutural do piloto humano. O método é apresentado para manobras executadas no eixo lateral e oferece uma medida do ganho humano utilizado durante a sua execução em malha-fechada. Para a modelagem da planta é utilizado modelo black-box com equacionamento de espaço de estados e identificação de parâmetros. Dados de ensaios em voo provenientes de uma aeronave de transporte certificada FAR-25 são utilizados para medir a razão entre o tempo gasto pelo piloto humano em uma malha fechada em posição versus o tempo em uma malha de derivada da posição (roll vs. roll rate). Esta medida é proposta como validadora da execução correta do ensaio. / A method is proposed to verify losed-loop adequate flight test piloting gain in PIO aircraft certification. The synthetic tracking task PIO flight test is used. The theory is based on the entropy concept from information theory and on the structural pilot model of the human pilot. The method is presented for single axis pilot tracking maneuvers and offers a measure of the human pilot gain employed during its execution. A black-box, state-space, parameter-identified model is used for the plant. Flight test data from a FAR-25 transport aircraft is used to verify the theory of how to determine a measure of the ratio between time spent by the human pilot in the error loop versus in the error rate loop to control the aircraft. This measure is proposed as a test point validation method for PIO flight testing.

In order to uncover the best kept secret in today's commercial aviation, this project deals with the calculation of fuel consumption of aircraft. With only the reference of the aircraft manufacturer's information, given within the airport planning documents, a method is established that allows computing values for the fuel consumption of every aircraft in question. The aircraft's fuel consumption per passenger and 100 flown kilometers decreases rapidly with range, until a near constant level is reached around the aircraft's average range. At longer range, where payload reduction becomes necessary, fuel consumption increases significantly. Numerical results are visualized, explained, and discussed. With regard to today's increasing number of long-haul flights, the results are investigated in terms of efficiency and viability. The environmental impact of burning fuel is not considered in this report. The presented method allows calculating aircraft type specific fuel consumption based on publicly available information. In this way, the fuel consumption of every aircraft can be investigated and can be discussed openly.