Control of vortex-driven instabilities was demonstrated via a scaled-down, cold-flow simulation that modeled closed-end acoustics. When vortex shedding frequencies couple with the natural acoustic modes of a choked chamber, potentially damaging low-frequency instabilities may arise. Although passive solutions can be effective, an active control solution is preferable. An experiment was performed to demonstrate an active control scheme for the reduction of vortex-driven oscillations. A non-reacting experiment using a primary flow of air, where both the duct exit and inlet are choked, simulated the closed-end acoustics. Two plates, separated by 1.27 cm, produced the vortex shedding phenomenon at the chamber's first longitudinal mode. Two active control schemes, closed-loop and open-loop, were studied via a cold-flow simulation for validating the effects of reducing vortex shedding instabilities in the system. Actuation for both control schemes was produced by using a secondary injection method. The actuation system consisted of pulsing compressed air from a modifed, 2-stroke model airplane engine, controlled and powered by a DC motor. The use of open-loop only active control was not highly effective in reducing the amplitude of the first longitudinal acoustic mode, near 93 Hz, when the secondary injection was pulsed at the same modal frequency. This was due to the uncontrolled phasing of the secondary injection system. A Pulse Width Modulated (PWM) signal was added to the open-loop control scheme to correct for improper phasing of the secondary injection flow relative to the primary flow. This addition allowed the motor speed to be intermittently increased to a higher RPM before returning to the desired open-loop control state. This proved to be effective in reducing the pressure disturbance by approximately 46%. A closed-loop control scheme was then test for its effectiveness in controlling the phase of the secondary injection. Feedback of the system's state was determined by placing a dynamic pressure transducer near the chamber exit. Closed-loop active control, using the designed secondary injection system, was proven as an effective means of reducing the problematic instabilities. A 50% reduction in the FFT RMS amplitude was realized by utilizing a Proportional-Derivative controller to modify the phase of the secondary injection.

<p>When characterizing a new solid propellant, one of the most important steps in determining its usefulness is discovering how the burning rate changes in response to changes in pressure. While there are many dynamic methods for directly measuring the regression rate of a burning propellant sample, few of them are capable of being used in typical harsh motor conditions: high pressures, high temperatures, and in an environment comprised of propellant exhaust products. This paper describes and evaluates the use of two custom-built microwave interferometers, one operating at 35 GHz and the other operating at 94 GHz, in several different configurations for the measurement of propellant regression rates. Four different configurations of interferometer and waveguide are presented and contrasted, with example results of experiments included. A polytetrafluoroethylene (PTFE) waveguide, utilized in previous works for explosives detonation velocity characterization, was used to directly couple interferometer signal with a burning propellant strand. This PTFE coupling is shown to be applicable to pressure vessel studies by simply using a cable feedthrough. In this configuration, signal quality is high but signal amplitude is low, especially when the waveguide is encased by support structures. A novel PTFE truncated cone waveguide expander is presented which performs three tasks: expanding the microwave signal such that an oversized (relative to signal wavelength) strand may be examined via microwave interferometry, functioning as a weak antenna that can observe phenomena through interstitial material without picking up significant amounts of environmental reflection, and acting as a sealing surface for pressure vessel experiments. Additionally, the use of a more-standard hollow-core waveguide and high-gain antenna is displayed, highlighting the increased signal strength but the larger number of spurious reflections in the signal. This study shows, through various experiments using the aforementioned configurations, the capability of microwave interferometry to quickly characterize a full propellant burning rate curve using a single dynamic-pressure test with 40g of propellant in a 2.5cm diameter propellant strand. Several novel combinations of mechanical configuration and propellant composition are shown that may guide future studies into the use microwave interferometry for solid propellant regression rate analysis.</p>

The payload environment of a rocket is random and dynamic during liftoff and flight, with acoustic noise, vibration, and acceleration falling under its definition. Characterization of this environment is important to finalize payload design requirements and insure mission completion. This report will focus on the study and measurement of acoustic noise using a low-cost microphone. Various spectral analysis techniques were utilized to characterize acoustic intensities and frequency content. Effects of vibration and acceleration, ground reflection, atmospheric absorption, and nonlinear propagation were investigated. Noise data were obtained from a rocket launch and several vertical, staticired hybrid motors. The propulsion system acoustic loads were compared to prediction methods from NASA SP-8072.

In this thesis, the design and particularities of a unique and revolution- ary scalable propulsion system are presented. A spacecraft mechanical design is included together with a mission definition, aiming to provide a context for a technology demonstration in space of an Hybrid Rocket Motor (HRM) as satellite thruster. Rocket motors have been around for many decades, with their use mainly focused on launch vehicles and large satellites, thus restricting the access to space to institutions with big budgets. To overcome this limitation, the application of a cost-effective type of rocket motor without a heritage of space utilization is explored. This is the implementation of an HRM as satellite thruster. In Chapter 2, the characteristics of this particular case of chemical rocket motor are presented in detail. The HRM applied for the present mission is a particular case of an in- house developed motor design method. As presented in Chapter 7, a scalable and versatile mechanical and propulsion design have been elab- orated following the maturation of a scalability software (Appendix A). The combination of these constitute a valuable tool allowing for a fast and accurate motor design for the desired scenario. Taking advantage of this straightforward tool, an attractive mission was defined to provide a meaningful context for the maiden use of an HRMin space. A micro satellite deep space mission, defined in Chapter 3, was chosen to validate the tool and prove Hybrid Rocket Motors (HRMs) capabilities, showing the benefits of its use over other propulsion systems already available, specifically in the small satellite family. The spacecraft design was tackled aiming to support the motor’s scalable concept while complying with the mission requirements and space standards. The out- come is an easily adaptable satellite design, justified in Chapter 8. The performed structural simulations are outlined in Appendix C to validate the developed design. Ultimately, this thesis work intends to provide the space community with a noteworthy product, opening the access to interplanetary missions provided the reduced mission costs of small satellites mounted with anHRM as propulsion system. Arising from the thesis content, research papers (Part v) have been published and presented in distinguished congresses, contributing to space development.

The present research project is carried out at the von Karman Institute for Fluid Dynamics (Rhode-Saint-Genèse, Belgium) with the financial support of the European Space Agency. The first stage of spacecrafts (e.g. Ariane 5, Vega, Shuttle) generally consists of large solid propellant rocket motors (SRM), which often consist of segmented structure and incorporate a submerged nozzle. During the combustion, the regression of the solid propellant surrounding the nozzle integration part leads to the formation of a cavity around the nozzle lip. The propellant combustion generates liquefied alumina droplets coming from chemical reaction of the aluminum composing the propellant grain. The alumina droplets being carried away by the hot burnt gases are flowing towards the nozzle. Meanwhile the droplets may interact with the internal flow. As a consequence, some of the droplets are entrapped in the cavity forming an alumina puddle (slag) instead of being exhausted through the throat. This slag reduces the performances. The aim of the present study is to characterize the slag accumulation process in a simplified model of the MPS P230 motor using primarily optical experimental techniques. Therefore, a 2D-like cold-gas model is designed, which represents the main geometrical features of the real motor (presence of an inhibitor, nozzle and cavity) and allows to approximate non-dimensional parameters of the internal two-phase flow (e.g. Stokes number, volume fraction). The model is attached to a wind-tunnel that provides quasi-axial flow (air) injection. A water spray device in the stagnation chamber realizes the models of the alumina droplets, which are accumulating in the aft-end cavity of the motor. To be able to carry out experimental investigation, at first the the VKI Level Detection and Recording(LeDaR) and Particle Image Velocimetry (PIV) measurement techniques had to be adapted to the two-phase flow condition of the facility. A parametric liquid accumulation assessment is performed experimentally using the LeDaR technique to identify the influence of various parameters on the liquid deposition rate. The obstacle tip to nozzle tip distance (OT2NT) is identified to be the most relevant, which indicates how much a droplet passing just at the inhibitor tip should deviate transversally to leave through the nozzle and not to be entrapped in the cavity. As LeDaR gives no indication of the driving mechanisms, the flow field is analysed experimentally, which is supported by numerical simulations to understand the main driving forces of the accumulation process. A single-phase PIV measurement campaign provides detailed information about the statistical and instantaneous flow structures. The flow quantities are successfully compared to an equivalent 3D unsteady LES numerical model. Two-phase flow CFD simulations suggest the importance of the droplet diameter on the accumulation rate. This observation is confirmed by two-phase flow PIV experiments as well. Accordingly, the droplet entrapment process is described by two mechanisms. The smaller droplets (representing a short characteristic time) appear to follow closely the air-phase. Thus, they may mix with the air-phase of the recirculation region downstream the inhibitor and can be carried into the cavity. On the other hand, the large droplets (representing a long characteristic time) are not able to follow the air-phase motion. Consequently, a large mean velocity difference is found between the droplets and the air-phase using the two-phase flow measurement data. Therefore, due to the inertia of the large droplets, they may fall into the cavity in function of the OT2NT and their velocity vector at the level of the inhibitor tip. Finally, a third mechanism, dripping is identified as a contributor to the accumulation process. In the current quasi axial 2D-like set-up large drops are dripping from the inhibitor. In this configuration they are the main source of the accumulation process. Therefore, additional numerical simulations are performed to estimate the importance of dripping in more realistic configurations. The preliminary results suggest that dripping is not the main mechanism in the real slag accumulation process. However, it may still lead to a considerable contribution to the final amount of slag.

Les moteurs à propergol solide sont parfois le siège d'instabilités aéroacoustiques résultant d'un couplage entre l'hydrodynamique des gaz brûlés et les modes acoustiques de la chambre de combustion. Ces instabilités se traduisent par de fortes Oscillations de Pression (ODP) dans la chambre de combustion du moteur. Ces ODP entrainent des vibrations de la structure, qui si elles venaient à dépasser certains niveaux pourraient nuire à la charge utile. Au vu du coût d'un essai, il est important de disposer d'outils permettant de prédire l'apparition de ces instabilités au moment de la conception. L'objectif de cette thèse est en premier lieu la mise au point d'une chaîne de couplage permettant d'évaluer l'impact des interactions fluide-structure sur l'amplitude des oscillations aéroacoustiques présentes au sein du propulseur. Une attention particulière est portée à l'algorithme de couplage entre les solveurs fluide et solide afin d'assurer une bonne conservation de l'énergie à l'interface fluide-structure, point clé dans l'étude d'instabilités. La chaîne numérique ainsi conçue est appliquée à une configuration réduite du moteur à propergol solide d'Ariane 5 dans le cadre de deux études. La première porte sur l'impact des vibrations de la structure sur les d'instabilités aéroacoustiques. L'effet d'un croisement de fréquences des modes propres longitudinaux de la structure et un des modes acoustiques de la chambre de combustion est traité. La seconde étude s'intéresse à l'effet des battements des protections thermiques du propulseur dans l'écoulement. Une structuration de l'écoulement et un net renforcement des ODP sont mis en évidence. / Large solid propellant rocket motors may be subjected to aero-acoustic instabilities arising from a coupling between the burnt gas flow and the acoustic eigenmodes of the combustion chamber. These instabilities lead to large pressure oscillations in the combustion chamber. These pressure oscillations cause vibrations which might jeopardize the payload if they happen to be larger than a certain threshold. Given the size and cost of any single firing test or launch, it is of first importance to rely on numerical tools able to predict these instabilities so that they can be avoided at the design level. The first purpose of this thesis is to build a numerical tool in order to evaluate how the coupling of the fluid flow and the whole structure of the motor influences the amplitude of the aeroacoustic oscillations living inside the rocket. A particular attention was paid to the coupling algorithm between the fluid and the solid solvers in order to ensure the best energy conservation through the interface.The numerical chain is applied to a sub-scaled configuration of Ariane 5 solid rocket motor in two studies. The first relates to the impact of vibration of the structure on aeroacoustic instabilities. The effect of a crossover frequency between the longitudinal modes of the structure and the acoustic modes of the combustion chamber is assessed. The second study examines the effect of thermal protection oscillations in the flow. An increased of the flow organisation and a significant strengthening of pressure oscillations are highlighted.

The focus of this study is the prediction of trajectories of solid particles injected into either a cylindrically- shaped solid rocket motor (SRM) or a bidirectional vortex chamber (BV). The Lagrangian particle trajectory is assumed to be governed by drag, virtual mass, Magnus, Saffman lift, and gravity forces in a Stokes flow regime. For the conditions in a solid rocket motor, it is determined that either the drag or gravity forces will dominate depending on whether the sidewall injection velocity is high (drag) or low (gravity). Using a one-way coupling paradigm in a solid rocket motor, the effects of particle size, sidewall injection velocity, and particle-to-gas density ratio are examined. The particle size and sidewall injection velocity are found to have a greater impact on particle trajectories than the density ratio. Similarly, for conditions associated with a bidirectional vortex engine, it is determined that the drag force dominates. Using a one-way particle tracking Lagrangian model, the effects of particle size, geometric inlet parameter, particle-to-gas density ratio, and initial particle velocity are examined. All but the initial particle velocity are found to have a significant impact on particle trajectories. The proposed models can assist in reducing slag retention and identifying fuel injection configurations that will ensure proper confinement of combusting droplets to the inner vortex in solid rocket motors and bidirectional vortex engines, respectively.

Within the scope of this thesis study, an optimization tool for the ballistic design of three-dimensional grains in solid propellant rocket motors is developed. The modeling of grain geometry and burnback analysis is performed analytically by using basic geometries like cylinder, cone, sphere, ellipsoid, prism and torus. For the internal ballistic analysis, a quasi-steady zero-dimensional flow solver is used. Genetic algorithms have been studied and implemented to the design process as an optimization algorithm. Lastly, the developed optimization tool is validated with the predesigned rocket motors.

The performance prediction of solid propellant rocket motor depends on the calculation of internal aerodynamics of the motor through its operational life. In order to obtain the control volume, in which the solutions will be carried out, a process called &ldquo / grain burnback calculation&rdquo / is required. During the operation of the motor, as the interface between the solid and gas phases moves towards the solid propellant in a direction normal to the surface, the combustion products are generated and added into the control volume. This phenomenon requires handling of moving boundaries as the solution proceeds. In this thesis, Fast Marching Method is implemented to the problem of grain burnback. This method uses the upwinding nature of the propellant interface motion and solves the Eikonal type equations on a fixed three-dimensional tetrahedron mesh. The control volume is coupled to a one-dimensional and a three-dimensional Euler aerodynamic solver in order to obtain the performance of the engine. The speed by which the interface moves depends on the static pressure on the surface of the propellant and comes from the solver. Therefore an iterative method has been proposed between the interface capturing algorithms and the flow solver. Both of the calculation results, which are obtained from one-dimensional and three-dimensional solvers are compared with actual rocket firing data and validated.

Dans un moteur à propergol solide, l’écoulement dépend fortement des gouttes d’alumine en suspension, dont la fraction massique est élevée. La distribution en taille des gouttes, qui s’élargit avec la coalescence, joue un rôle clef. Or résoudre des écoulements diphasiques polydisperses instationnaires avec une bonne précision sur la taille est un défi à la fois sur le plan de la modélisation et du calcul scientifique: (1) de très petites gouttes, par exemple résultant de la combustion de nanoparticules d’aluminium, subissent mouvement brownien et coalescence, (2) de petites gouttes ont leur vitesse conditionnée par leur taille de sorte qu’elles coalescent lorsqu’elles ont des tailles différentes, (3) des gouttes plus grosses peuvent se croiser par effet d’inertie et (4) toutes les gouttes interagissent de manière fortement couplée avec la phase porteuse. En complément des approches lagrangiennes, des modèles eulériens ont été développés pour décrire la phase dispersée à un coût raisonnable, et ils permettent un couplage aisé avec la phase porteuse ainsi que la parallélisation massive des codes: les approches eulériennes sont bien adaptées aux calculs industriels. Le modèle Multi-Fluide permet la description détaillée de la polydispersion, des coreélations taille/vitesse et de la coalescence, en résolvant séparément des “fluides” de gouttes triées par taille, appelés sections. Un ensemble de modèles est évalué dans cette thèse et une stratégie numérique est développée pour effectuer des calculs industriels de moteurs à propergol solide. (1) La physique des nanoparticules est évalué et incluse dans un modèle de coalescence complet. Des méthodes de moments d’ordre élevé sont ensuite développées: (2) une méthode à deux moments en taille est étendue à la coalescence pour traiter la physique de la polydispersion et les développements numériques connexes permettent d’effectuer des calculs applicatifs dans le code industriel CEDRE; (3) une méthode basée sur les moments en vitesse du deuxième ordre, un schéma de transport à l’ordre deux sur maillages structurés ainsi qu’un modèle de coalescence sont développés. Des validations académiques de la stratégie pour gouttes d’inertie modérée sont effectuées sur des écoulements complexes puis avec de la coalescence; (4) une stratégie d’intégration en temps est développée et mise en œuvre dans CEDRE pour traiter efficacement le couplage fort, dans des cas instationnaires et polydisperses incluant de très petites particules. L’ensemble des développements est soigneusement validé: soit par des formules analytiques ad hoc pour la coalescence et pour le couplage fort d’une onde acoustique; soit par des comparaisons numériques croisées avec une DPS pour la coalescence et avec des simulations lagrangiennes de cas applicatifs, coalescents et fortement couplés; soit par des résultats expérimentaux disponibles sur une configuration académique de coalescence et sur un tir de moteur à échelle réduite. La stratégie complète permet des calculs applicatifs à un coût raisonnable. En particulier, un cal- cul de moteur avec des nanoparticules permet d’évaluer la faisabilité de l’approche et d’orienter les efforts de recherche sur les propergols chargés de nanoparticules. / In solid rocket motors, the internal flow depends strongly on the alumina droplets, which have a high mass fraction. The droplet size distribution, which is wide and spreads up with coalescence, plays a key role. Solving for unsteady polydisperse two- phase flows with high accuracy on the droplet sizes is a challenge for both modeling and scientific computing: (1) very small droplets, e.g. resulting from the combustion of nanoparticles of aluminum fuel, encounter Brownian motion and coalescence, (2) small droplets have their velocity conditioned by size so they coalesce when having different sizes, (3) bigger droplets have an inertial behavior and may cross each other’s trajectory, and (4) all droplets interact in a two-way coupled manner with the carrier phase. As an alternative to Lagrangian approaches, some Eulerian models can describe the disperse phase at a moderate cost, with an easy coupling to the carrier phase and with massively parallel codes: they are well-suited for industrial computations. The Multi- Fluid model allows the detailed description of polydispersity, size/velocity correlations and coalescence by separately solving “fluids” of size-sorted droplets, the so-called sections. In the present work, we assess an ensemble of models and we develop a numerical strategy to perform industrial computations of solid rocket motor flows. (1) The physics of nanoparticles is assessed and included in a polydisperse coalescing model. High order moment methods are then developed: (2) a Two-Size moment method is ex- tended to coalescence to treat accurately the physics of polydispersity and coalescence and the related numerical developments allow to perform applicative computations in the industrial code CEDRE; (3) a second order velocity moment method is developed, together with a second order transport scheme, to evaluate a strategy for a moderately inertial disperse phase, and academic validations are performed on complex flow fields; (4) a time integration strategy is developed and implemented in CEDRE to treat efficiently two-way coupling, in unsteady polydisperse cases including very small particles. The developments are carefully validated, either through purposely derived analytical formulae (for coalescence and two-way acoustic coupling), through numerical cross-comparisons (for coalescence with a Point-Particle DNS, for applicative cases featuring coalescence and two-way coupling with a Lagrangian method), or through available experimental results (for coalescence with an academic experiment, for the overall physics with a sub-scale motor firing). The whole strategy allows to perform applicative computations in a cost effective way. In particular, a solid rocket motor with nanoparticles is computed as a feasibility case and to guide the research effort on motors with nanoparticle fuel propellants.