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.

Laminar turbulent transition is known for a long time as a critical phenomenon influencing the thermal load encountered by hypersonic vehicle during their planetary re-entry trajectory. Despite the efforts made by several research laboratories all over the world, the prediction of transition remains inaccurate, leading to oversized thermal protection system and dramatic limitations of hypersonic vehicles performances. One of the reasons explaining the difficulties encountered in predicting transition is the wide variety of parameters playing a role in the phenomenon. Among these parameters, surface roughness is known to play a major role and has been investigated in the present thesis.<p><p>A wide bibliographic review describing the main parameters affecting transition and their coupling is proposed. The most popular roughness-induced transition predictions correlations are presented, insisting on the lack of physics included in these methods and the difficulties encountered in performing ground hypersonic transition experiments representative of real flight characteristics. This bibliographic review shows the importance of a better understanding of the physical phenomenon and of a wider experimental database, including real flight data, for the development of accurate prediction methods.<p><p>Based on the above conclusions, a hypersonic experimental test campaign is realized for the characterization of the flow field structure in the vicinity and in the wake of 3D roughness elements. This fundamental flat plate study is associated with numerical simulations for supporting the interpretation of experimental results and thus a better understanding of transition physics. Finally, a model is proposed in agreement with the wind tunnel observations and the bibliographic survey.<p><p>The second principal axis of the present study is the development of a hypersonic in-flight roughness-induced transition experiment in the frame of the European EXPERT program. These flight data, together with various wind tunnel measurements are very important for the development of a wide experimental database supporting the elaboration of future transition prediction methods. / Doctorat en Sciences de l'ingénieur / info:eu-repo/semantics/nonPublished

This thesis presents the derivation of the theory leading to the determination of the governing equations of chemically reacting flows under local thermodynamic equilibrium, which rigorously takes into account effects of elemental (de)mixing. As a result, new transport coefficients appear in the equations allowing a quantitative predictions and helping to gain deeper insight into the physics of chemically reacting flows at and near local equilibrium. These transport coefficients have been computed for both air and carbon dioxide mixtures allowing the application of this theory to both Earth and Mars entry problems in the framework of the methodology for the determination of the catalytic activity of Thermal Protections Systems (TPS) materials.<p>Firstly, we analyze the influence of elemental fraction variations on the computation of thermochemical equilibrium flows for both air and carbon dioxide mixtures. To this end, the equilibrium computations are compared with several chemical regimes to better analyze the influence of chemistry on wall heat flux and to observe the elemental fractions behavior along a stagnation line. The results of several computations are presented to highlight the effects of elemental demixing on the stagnation point heat flux and chemical equilibrium composition for air and carbon dioxide mixtures. Moreover, in the chemical nonequilibrium computations, the characteristic time of chemistry is artificially decreased and in the limit the chemical equilibrium regime, with variable elemental fractions, is achieved. Then, we apply the closed form of the equations governing the behavior of local thermodynamic equilibrium flows, accounting for the variation in local elemental concentrations in a rigorous manner, to simulate heat and mass transfer in CO2/N2 mixtures. This allows for the analysis of the boundary layer near the stagnation point of a hypersonic vehicle entering the true Martian atmosphere. The results obtained using this formulation are compared with those obtained using a previous form of the equations where the diffusive fluxes of elements are computed as a linear combination of the species diffusive fluxes. This not only validates the new formulation but also highlights its advantages with respect to the previous one :by using and analyzing the full set of equilibrium transport coefficients we arrive at a deep understanding of the mass and heat transfer for a CO2/N2 mixture.<p>Secondly, we present and analyze detailed numerical simulations of high-pressure inductively coupled air plasma flows both in the torch and in the test chamber using two different mathematical formulations: an extended chemical non-equilibrium formalism including finite rate chemistry and a form of the equations valid in the limit of local thermodynamic equilibrium and accounting for the demixing of chemical elements. Simulations at various operating pressures indicate that significant demixing of oxygen and nitrogen occurs, regardless of the degree of nonequilibrium in the plasma. As the operating pressure is increased, chemistry becomes increasingly fast and the nonequilibrium results correctly approach the results obtained assuming local thermodynamic equilibrium, supporting the validity of the proposed local equilibrium formulation. A similar analysis is conducted for CO2 plasma flows, showing the importance of elemental diffusion on the plasma behavior in the VKI plasmatron torch.<p>Thirdly, the extension of numerical tools developed at the von Karman Institute, required within the methodology for the determination of catalycity properties for thermal protection system materials, has been completed for CO2 flows. Non equilibrium stagnation line computations have been performed for several outer edge conditions in order to analyze the influence of the chemical models for bulk reactions. Moreover, wall surface reactions have been examined, and the importance of several recombination processes has been discussed. This analysis has revealed the limits of the model currently used, leading to the proposal of an alternative approach for the description of the flow-surface interaction. Finally the effects of outer edge elemental fractions on the heat flux map is analyzed, showing the need to add them to the list of parameters of the methodology currently used to determine catalycity properties of thermal protection materials. / Doctorat en sciences appliquées / info:eu-repo/semantics/nonPublished

An approach for designing a high-enthalpy wind tunnel driven by exothermic chemical reactions was developed. Nozzle contours were designed using CONTUR, a program implementing the method of characteristics, to design nozzle contours at various flow conditions. A reacting mixture including nitrous oxide has been identified as the best candidate for providing clean air at high temperatures. The nitrous oxide has a few performance factors that were considered, specifically the combustion of the gas. Initial CFD simulations were performed on the nozzle and test region to validate flow characteristics and possible issues. Initial results show a fairly uniform exit velocity and ability to perform testing. In a second phase of the work, two generic, high-speed missile configurations were explored using numerical simulation. The mean flow was computed on both geometries at 0 and 45 roll and 0, 1, and 10 angle of attack. The computations identified complex flow structures, including three-dimensional shock/boundary-layer interactions, that varied considerably with angle of attack.

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.

A series of experiments were carried out in Purdue University’s Boeing/AFOSR Mach6 Quiet Tunnel, to understand the amplification of streamwise vortices across a separated region in a quiet flow regime. Streamwise vortices were induced on the upstream end of an axisymmetric model consisting of a 7-degree half-angle cone, a cylinder, and a 10-degree flare. The instabilities were seeded using a pre-existing set of roughness inserts, with small, discrete roughness elements. The elements varied in spacing, height, and number of elements. The model was aligned to near 0.0 degree angle of attack. <div><br></div><div>The streamwise, Gortler-like instabilities travelled across the separated region onto the flare, where they were measured with pressure transducers and infrared thermography. The amplification of the instabilities was measured at a variety of Reynolds numbers, under both quiet and conventional noise flow. The results were compared to those of a smooth insert. Heat transfer results showed a streaking pattern, with a peak in heating visible in the streak. Heat flux increased linearly with Reynolds number. If transition was induced, the heat flux would begin to decrease. Power spectral density measurements of the pressure fluctuations indicated that the region within the streak contained two notable instabilities, one between 70 and 150 kHz, and one between 200 and 250 kHz. Transition was only measured in the spectral content in the region on the flare where a ”filling in” of streaks was visible in heat transfer results. Heat flux increased in an nonlinear manner with increasing roughness height. </div><div><br></div><div>The streak positioning and peak heat flux showed a high sensitivity to small, uncontrollable changes in run conditions throughout. Heat transfer results were largely repeatable for small angles of attack, less than 0.1 degrees. The streaks shifted slightly in width and position for angles of attack near 0.1 degrees. Small changes in the streak positioning and heat transfer magnitude were seen in repeatability runs; this is mostly attributable to small changes in initial run conditions. </div>

An approach for designing a high-enthalpy wind tunnel driven by exothermic chemical reactions was developed. Nozzle contours were designed using CONTUR, a program implementing the method of characteristics, to design nozzle contours at various flow conditions. A reacting mixture including nitrous oxide has been identified as the best candidate for providing clean air at high temperatures. The nitrous oxide has a few performance factors that were considered, specifically the combustion of the gas. Initial CFD simulations were performed on the nozzle and test region to validate flow characteristics and possible issues. Initial results show a fairly uniform exit velocity and ability to perform testing. In a second phase of the work, two generic, high-speed missile configurations were explored using numerical simulation. The mean flow was computed on both geometries at 0 and 45 roll and 0, 1, and 10 angle of attack. The computations identified complex flow structures, including three-dimensional shock/boundary-layer interactions, that varied considerably with angle of attack. <br>

This research focuses on measurements of a convective shear-layer instability seen naturally in quiet hypersonic flow. Experiments were carried out in the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. The BAM6QT provides low-disturbance hypersonic flow with freestream noise levels similar to what would be experienced by a flight vehicle. To obtain high-speed, off-the-surface measurements of the instability, a modified focused laser differential interferometer (FLDI) was first designed to work with the contoured Plexiglas windows available in the tunnel.<div><br>A cone-cylinder-flare geometry was then selected to study the instabilities related to an axisymmetric separation bubble at Mach 6. The sharp cone had a 5-degree half-angle, while flare angles of 10 degrees and 3.5 degrees were tested to compare axisymmetric compression with and without separation, respectively. Under quiet flow, laminar separation and reattachment was confirmed by schlieren and surface pressure-fluctuation measurements. Coherent traveling waves were observed. These were attributed to both the second-mode instability, as well as a shear-generated instability from the separation bubble. The symmetry of the bubble was found to be highly sensitive to angle of attack. Additionally, by introducing controlled disturbances on the cone upstream of the separation, larger-amplitude shear-generated waves were measured while the second-mode amplitudes remained unchanged. Therefore, the shear-generated waves were amplified moving through the shear layer, while the second mode remained neutrally stable. These appear to be the first measurements of traveling waves that are generated in the shear layer of a separation bubble in hypersonic flow. <br></div>

<p>Air intakes are a fundamental part of all high speed airbreathing propulsion concepts. The main purpose of an intake is to capture and compress freestream air for the engine. At hypersonic speeds, the intake’s surface and shock structure effectively slow the airflow through ram-air compression. In supersonic-combustion ramjets, the captured airflow remains supersonic and generates complicated shock structures. The design of these systems require careful evaluation of proposed operating conditions and relevant aerodynamic phenomena. The physics of these systems, such as the intake’s operability range, mass capture efficiency, back-pressure resiliency, and intake unstart margins are all open areas of research. </p> <p><br></p> <p>A high speed intake, dubbed the Indiana Intake Testbed, was developed for experimentation within the Boeing-AFOSR Mach 6 Quiet Tunnel at Purdue University. This inward-turning, mixed compression intake was developed from osculating axisymmetric theory and uses a streamtracing routine to create a shape-transitioned geometry. To account for boundary layer growth, a viscous correction was implemented on the intake’s compression surfaces. This comprehensive independent design code was pursued to generate an unrestricted geometry that satisfies academic inquiry into fluid dynamic interactions relevant to intakes. Additionally, the design code contains built-in analysis tools that are compared against CFD calculations and experimental data. </p> <p><br></p> <p>Two blockage models were constructed and outfitted with Kulite pressure transducers to detect possible intake start and unstart effects. Due to an error in the design code, the preliminary blockage models’ lower surfaces were oversized. The two intake models were tested over a freestream Reynolds number sweep, under noisy and quiet flow, at one non-zero angle of attack, and at a singular back-pressure condition. Back-pressure effects acted to unstart the intake and provide a comparison between forced-unstart and started states. The experimental campaign cataloged both tunnel starting and inlet starting conditions, which informed the design of the finalized model. The finalized model is presented herein. Future experiments to study isolator shock-trains, shock-wave boundary layer interactions, and possible instances of boundary layer transition on the intake’s compression surface are planned. </p>

<p>Accurate and reliable prediction of laminar-turbulent boundary layer transition at hypersonic velocities is important for the development of a variety of practical high-speed flight systems currently under development. Boundary layer transition can cause up to an order of magnitude increase in skin friction and heat flux on a flight vehicle, meaning that understanding boundary layer behavior is critical to the design of weight-efficient thermal protection systems. Despite the importance of the topic, significant gaps remain in the community's current understanding of boundary layer transition and control. </p> <p>One of the biggest areas of concern in the field of high-speed boundary layer transition is the effect of facility noise on wind tunnel measurements. Conventional hypersonic wind tunnels are contaminated by freestream fluctuations which can be as much as two orders of magnitude higher than free-flight atmospheric conditions. These disturbances are typically produced by turbulent boundary layers on the tunnel walls; they are acoustic in nature and consist of pressure waves which radiate into the test section. This facility noise plays a leading role in high-speed transition phenomena in conventional hypersonic tunnels.</p> <p><br></p> <p>The current work studies the effects of facility noise on hypersonic transition using both linear stability theory and direct numerical simulation. A model for the freestream disturbance environment of the von Karman Facility's Tunnel B based on experimental measurements of the disturbance spectra present in the tunnel is created and used to study a past experiment performed in the same wind tunnel using a sharp cone and hollow cylinder. The results show that while linear stability theory accurately captures the behavior of second-mode instability growth, it fails to predict the growth of low-frequency instabilities recorded in the experiments. The stability theory analysis also suggests that very fine scale variation in nose tip geometry can play an outsize role in the development of boundary layer instabilities significantly farther downstream.</p> <p><br></p> <p>The direct numerical simulation demonstrates that, using an artificial body forcing term to implement the constructed tunnel noise model, the experimental effects of facility noise can be adequately captured with a sufficiently dense computational grid. For the conical geometry used in the experiments, calculations of surface heat flux indicate good experimental agreement with in prediction of transition location, and total temperature spectra extracted from the flow compare favorably with the experimental data as well. Visualizations of the flowfield confirm the onset of turbulence as a result of the freestream forcing. The computations also suggest that nonlinear interactions may be present in the turbulent breakdown region, leading to the production of streamwise streaks along the cone's surface. Transition on the hollow cylinder was not achieved due to suspected resolution issues, so detailed physical comparison of the two cases was not possible.</p> <p><br></p> <p>Overall, the results of this work suggest that direct numerical simulation is a capable tool for studying the effects of facility noise on hypersonic transition for simple geometries, albeit one which can be difficult to practically realize considering the required computational cost. Computational results indicate that two phenomena may play a role in the development of boundary layer instabilities for a sharp cone --- the fine-scale shape of the tip, which may change the behavior of the entropy layer near the nose; and the interactions between low- and high-frequency waveforms, which seems to cause nonlinear breakdown in line with current experimental understanding.</p>