One of the major challenges of scramjet propulsion remains the generation of sufficient thrust to overcome the large drag of hypersonic vehicles. Since the viscous drag constitutes a large portion of the overall drag, its mitigation offers potential for performance improvement. Viscous drag is generated on all wetted surfaces of the vehicle but is largest in the scramjet combustion chamber, where the fluid not only has a high flow speed, but also a high density. Reduction of the skin friction drag in the combustor hence promises large improvements to the efficiency of the propulsive system. Stalker (2005) proposed a novel approach to skin friction reduction that is based on the combustion of hydrogen within the turbulent boundary layer of supersonic or hypersonic flow. An extension to the theory of van Driest II was developed that suggests that the effectiveness of this method is significantly superior to that of film cooling without combustion effects. In essence, the combustion heat release reduces the velocity gradient at the wall and the density in the boundary layer so that the momentum transfer to the wall is decreased. This work investigates the applicability of this skin friction reduction method to scramjet combustors that would operate at flight Mach numbers between 8 and 13 at altitudes between 34 and 39 km. The corresponding combustor Mach number is approximately 4.5 and the total enthalpies are between 3.6 and 7.8 MJ/kg. Experiments that directly measured the skin friction drag on the internal scramjet combustor surface were conducted in the T4 Stalker tube at The University of Queensland. A constant area, axisymmetric combustor was tested with a matching constant area, axisymmetric inlet that did not compress the oncoming flow. Therefore, the experiments were of a quasi-direct-connect nature where the inlet was used to condition the wall layer of the flow that enters the combustion chamber. The start of the combustor was formed by a step at the end of the inlet which contained an annular slot for the injection of the gaseous hydrogen fuel. Fuel was injected tangentially to the main stream flow into the circular combustor as a uniform layer underneath the established boundary layer from this annular slot. Combustion was monitored via the measurement of the axial pressure distribution in the combustor and viscous forces on the combustor were measured with a stress wave force balance. Two different inlet lengths were tested to assess the effect of the boundary layer state and thickness on the ignition and combustion of the injected hydrogen. The leading edge of the inlet was either sharp or blunt to investigate the effect of the hot gas that is contained in an entropy layer that is generated by a blunt leading edge. Finally, the diameter of the duct was varied to ensure that the experimental data was not subject to duct scaling effects. The effect on skin friction of the combustion of fuel in the boundary layer was assessed directly by measurement as well as analytically with several prediction methods. The experimental data show reductions of skin friction drag of up to 77% when stable combustion was established. A thick, turbulent boundary layer results in ignition for lower enthalpy conditions than a thin, laminar layer. The blunted leading edge configuration creates conditions that results in ignition of the injected fuel at all tested flow enthalpies and when a sharp leading edge configuration does not. Analytical predictions of the skin friction drag are in close agreement with the experimental data for fuel-off, film cooling and boundary layer combustion cases. It is demonstrated that the characteristics of boundary layer combustion do not change when the duct diameter is increased and the hydrogen mass flow rate per unit circumferential length is kept constant.

Orientador: Prof. Dr. João Felipe de Araújo Martos / Dissertação (mestrado) - Universidade Federal do ABC, Programa de Pós-Graduação em Engenharia Mecânica, 2014. / O veículo hipersônico aeroespacial brasileiro 14-X B (VHA 14-X B) é um demonstrador tecnológico do sistema de propulsão hipersônica aspirada com base em combustão supersônica (scramjet) projetado para voar na atmosfera da Terra a 30 quilômetros de altitude e número de Mach 7. O VHA 14-X B encontra-se em desenvolvimento no Laboratório de Aerotermodinâmica e Hipersônica Professor Henry T. Nagamatsu do Instituto de Estudos Avançados (IEAv). Uma das metodologias mais importantes na análise e desenvolvimento de um veículo aeroespacial hipersônico são os túneis de vento hipersônicos pulsados, os quais são instalações experimentais em terra capazes de simular as condições de voo, fornecendo dados experimentais para o projeto dos veículos aeroespaciais hipersônicos. Neste trabalho, foi utilizado o túnel de vento hipersônico pulsado T3 que possui 61 cm de diâmetro no bocal de saída e é operado com base na técnica de onda de choque refletida. O T3 foi financiado pela Fundação de Amparo a Pesquisa de São Paulo (FAPESP) e foi projetado para Pesquisa e Desenvolvimento na área de combustão supersônica. Trata-se de um tubo de choque equipado com um bocal convergente-divergente utilizado para produzir elevado número Mach e escoamentos de alta entalpia na seção de teste próximos aos encontrados durante o voo de um veículo aeroespacial na atmosfera da Terra a hipervelocidade. O modelo de 1 metro de comprimento em aço inoxidável do 14-X B foi instrumentado, com vinte e oito transdutores de pressão piezelétricos PCB nas superfícies de compressão, câmara de combustão e de expansão. Utilizando o túnel T3 no modo de operação de equilíbrio de interface para atingir escoamento livre com número de Mach entre 7 e 8 foi realizada a investigação experimental. Medidas da pressão estática no intradorso do modelo 14-X B, bem como fotografias schlieren, feitas a partir do bordo de ataque do modelo forneceram dados experimentais, que foram comparados com as análises teórica-analítica e simulações computacionais de dinâmica de fluidos, ambos utilizadas no projeto do modelo do VHA 14-X B. / The Brazilian VHA 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet) intended to fly into the Earth¿s atmosphere at 30 km altitude and Mach number 7. The 14-X B was designed o the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics of the Institute for Advanced Studies (IEAv). Hypersonic wind tunnels are one of the most important ground-based experimental facilities intended to simulate the flight conditions providing experimental data to design hypersonic aerospace vehicles and engines. The IEAv 0.60-m nozzle exit diameter Hypersonic Reflected Shock Tunnel named T3 and funded by São Paulo Research Foundation (FAPESP), was designed as a research & development facility for basic investigations in supersonic combustion. The T3 Hypersonic Shock Tunnel is a shock tube equipped with a convergent-divergent nozzle to produce high Mach number and high enthalpy flows in the test section close to those encountered during the flight of a aerospace vehicle into the Earth's atmosphere at hypersonic flight speeds. A 1-m long stainless steel VHA 14-X B model was instrumented with twenty-eight piezoelectric pressure transducers along its compression surface, combustion chamber and nozzle. It was experimentally investigated on the equilibrium interface operational mode of the T3 Hypersonic Shock Tunnel, yielding a freestream Mach number from 7 to 8. Static pressure measurements at the lower surface of the VHA 14-X B as well as high speed schlieren photographs taken from the 5.5° leading edge and the 14.5° deflection compression ramp provide experimental data that was compared to the analytic theoretical analysis and computational fluid dynamics simulation, both applied to the design of the VHA 14-X B.

A Supersonic Combustion Ramjet (scramjet) is, at least in theory, an efficient air-breathing propulsion system for sustained hypersonic flight at Mach numbers above approximately M=5. Important design issues for such hypersonic propulsion systems, are the lack of ground based facilities capable of testing a full-sized engine at cruise flight conditions and the absence of general scaling laws for the extrapolation of wind tunnel data to flight configurations. Therefore, there is a strong need for the development and validation of CFD tools to support the design process of scramjet-powered vehicles. The aims of this thesis are, in this context, to assess the applicability of, to further develop, and to validate the DLR TAU flow solver for the CFD analysis of the complete flow-path of a scramjet vehicle. The basis of this validation and of the identification of critical modelling assumptions is the recalculation of a series of wind tunnel tests of the HyShot II generic scramjet configuration that were performed in the High Enthalpy Shock Tunnel Göttingen (HEG) at the German Aerospace Center, DLR. / Staustrahlantriebe, bei denen sich die Strömung im gesamten Triebwerksbereich im Überschall befindet (supersonic combustion ramjets, Scramjets), stellen ein - zumindest theoretisch - effektives Antriebessystem für den Hyperschallflug im Machzahlbereich von M > 5 dar. Die Auslegung und der Entwurf von luftatmenden Hyperschallantrieben sind in der Praxis mit Schwierigkeiten verbunden. Der Einsatz von Bodenversuchsanlagen ist auf kleinskalige Konfigurationen oder einzelne Triebwerkskomponenten begrenzt. Die Ergebnisse von numerischen Strömungssimulationsverfahren sind mit hohen Unsicherheiten behaftet, die ihren Ursprung in der Modellbildung für die komplexen Strömungsphänomene in chemisch reagierenden, kompressiblen und turbulenten Über- und Hyperschallströmungen haben. Weiterhin existieren keine allgemein gültigen Skalierungsgesetze um Aussagen aus Windkanalexperimenten auf Flugkonfigurationen zu übertragen.Die vorliegende Arbeit beschäftigt sich in diesem Zusammenhang mit der Erweiterung des DLRStrömungslösers TAU für die Berechnung von Überschallverbrennungsphänomenen in Scramjets sowie mit der Anwendung des Verfahrens für die numerische Analyse von Windkanalexperimenten, die im Hochenthalpiekanal Göttingen (HEG) des Deutschen Zentrums für Luft- und Raumfahrt (DLR) zur Untersuchung der generischen HyShot II Scramjet-Konfiguration durchgeführt wurden. Die wichtigsten Ziele waren die genaue Charakterisierung der freien Anströmung im Windkanal, der Nachweis der Anwendbarkeit des verwendeten Rechenverfahrens und die Analyse des Einflusses verschiedener numerischer Modellierungsansätze für die Strömungssimulation in Scramjets sowie die Nutzung der numerischen Daten für eine verbesserte Interpretation der experimentellen Ergebnisse.

La propulsion aérobie à grande vitesse est depuis longtemps identifiée comme l'un des prochains sauts technologiques à franchir dans le domaine des lanceurs spatiaux. Cependant, les véhicules hypersoniques (HSV) fonctionnant dans des domaines de vitesse extrêmement élevées, de nombreuses contraintes et incertitudes entravent les garanties des propriétés des contrôleurs. L'objet de cette thèse est d'étudier la synthèse de commande d'un tel véhicule.Pour commencer, il s'agit de définir un modèle représentatif d'un HSV exploitable pour la commande. Dans ce travail, nous construisons deux modèles de HSV. Un pour la simulation en boucle fermée, et le second afin de poser précisément le problème de commande.Nous proposons ensuite une synthèse de commande de la dynamique longitudinale dans le plan vertical de symétrie. Celle-ci est robuste aux incertitudes de modélisation, tolérante à des saturations, et n'excite pas les dynamiques rapides négligées. Ses propriétés sont évaluées sur différents cas de simulation. Puis, une extension est proposée afin de résoudre le problème de commande simultanée des dynamiques longitudinale et latérale, sous les mêmes contraintes.Ce résultat est obtenu par une assignation de fonction de Lyapunov, suite à une étude des dynamiques longitudinale et latérale. Par ailleurs, pour traiter les erreurs de poursuite dues aux incertitudes de modélisation, nous nous intéressons au problème de régulation asymptotique robuste par retour d'état. Nous montrons que cette régulation peut être accomplie en stabilisant le système augmenté d'un intégrateur de la sortie. Ceci constitue une extension de la structure de contrôle proportionnel-intégral au cas des systèmes non linéaires. / High speed airbreathing thrust has been known for a long time as one of the next technological step to be overcome in space launchers domain. However, HyperSonic Vehicles (HSV) speed operating ranges being extremely high, numerous constraints and uncertainties restrict the ensuring of control properties. The purpose of this thesis is to study control synthesis for such a vehicle.First, it concern the definition of a HSV model for controlling purpose. In this work is constructed two HSV models. One in order to effect closed loop simulation, and the other in order to precisely establish the control problem.Then, is proposed a control synthesis for the longitudinal dynamics restricted to the symmetric vertical plane. It is robust to modelling uncertainties, allows saturation, and does not excite neglected fast dynamics. Its properties are evaluated on different cases of simulation. Next, an extension is proposed in order to solve the problem of controlling simultaneously longitudinal and lateral dynamics, under the same constraints.This result is obtained by the use of control Lyapunov functions, following the study of longitudinal and lateral dynamics. Furthermore, in order to solve tracking errors due to modelling uncertainties, the problem of robust asymptotic regulation by state feedback has been addressed. It is shown that such a regulation can be achieved by stabilizing the system augmented by an output integrator. This constitutes an extension for nonlinear systems of the proportional-integral control structure.

A Supersonic Combustion Ramjet (scramjet) is, at least in theory, an efficient air-breathing propulsion system for sustained hypersonic flight at Mach numbers above approximately M=5. Important design issues for such hypersonic propulsion systems, are the lack of ground based facilities capable of testing a full-sized engine at cruise flight conditions and the absence of general scaling laws for the extrapolation of wind tunnel data to flight configurations. Therefore, there is a strong need for the development and validation of CFD tools to support the design process of scramjet-powered vehicles. The aims of this thesis are, in this context, to assess the applicability of, to further develop, and to validate the DLR TAU flow solver for the CFD analysis of the complete flow-path of a scramjet vehicle. The basis of this validation and of the identification of critical modelling assumptions is the recalculation of a series of wind tunnel tests of the HyShot II generic scramjet configuration that were performed in the High Enthalpy Shock Tunnel Göttingen (HEG) at the German Aerospace Center, DLR. / Staustrahlantriebe, bei denen sich die Strömung im gesamten Triebwerksbereich im Überschall befindet (supersonic combustion ramjets, Scramjets), stellen ein - zumindest theoretisch - effektives Antriebessystem für den Hyperschallflug im Machzahlbereich von M > 5 dar. Die Auslegung und der Entwurf von luftatmenden Hyperschallantrieben sind in der Praxis mit Schwierigkeiten verbunden. Der Einsatz von Bodenversuchsanlagen ist auf kleinskalige Konfigurationen oder einzelne Triebwerkskomponenten begrenzt. Die Ergebnisse von numerischen Strömungssimulationsverfahren sind mit hohen Unsicherheiten behaftet, die ihren Ursprung in der Modellbildung für die komplexen Strömungsphänomene in chemisch reagierenden, kompressiblen und turbulenten Über- und Hyperschallströmungen haben. Weiterhin existieren keine allgemein gültigen Skalierungsgesetze um Aussagen aus Windkanalexperimenten auf Flugkonfigurationen zu übertragen.Die vorliegende Arbeit beschäftigt sich in diesem Zusammenhang mit der Erweiterung des DLRStrömungslösers TAU für die Berechnung von Überschallverbrennungsphänomenen in Scramjets sowie mit der Anwendung des Verfahrens für die numerische Analyse von Windkanalexperimenten, die im Hochenthalpiekanal Göttingen (HEG) des Deutschen Zentrums für Luft- und Raumfahrt (DLR) zur Untersuchung der generischen HyShot II Scramjet-Konfiguration durchgeführt wurden. Die wichtigsten Ziele waren die genaue Charakterisierung der freien Anströmung im Windkanal, der Nachweis der Anwendbarkeit des verwendeten Rechenverfahrens und die Analyse des Einflusses verschiedener numerischer Modellierungsansätze für die Strömungssimulation in Scramjets sowie die Nutzung der numerischen Daten für eine verbesserte Interpretation der experimentellen Ergebnisse.

No description available.