In this thesis, Initial Value Problem of Level-set Method is applied to solid propellant combustion to find the grain burnback. For the performance prediction of the rocket motor, 0-D, 1-D or 3-D flow models are used depending on the type of thre grain configuration.

Solid propellant rocket motor is the primary propulsion technology used for short and medium range missiles. It is also commonly used as boost motor in many di_erent applications. Its wide spread usage gives rise to diversity of environments in which it is handled and stored. Ability to predict the storage life of solid propellants plays an important role in the design and selection of correct protective environments. In this study a methodology for the prediction of solid propellant storage life using cumulative damage concepts is introduced. Finite element mesh of the solid propellant grain is created with the developed parametric grain geometry generator. Finite element analyses are carried out to obtain the temperature and stress response of the propellant to the environmental thermal loads. Daily thermal cycles are assumed to be sinusoidal cycles represented by their means and amplitudes. With the cumulative damage analyses, daily damage accumulated in the critical locations of the solid propellant grain are investigated. Meta-models relating the daily damage amount with the daily temperature cycles are constructed in order to compute probability of failure. The results obtained in this study imply that it is possible to make numerical predictions for the storage life of solid propellants even in the early design phases. The methodology presented in this study provides a basis for storage life predictions.

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.

This dissertation documents the building of computational propellant/ingredient models toward predicting AP/HTPB/Al cookoff events. Two computer codes were used to complete this work; a steady-state code and a transient ignition code Numerous levels of verification resulted in a robust set of codes to which several propellant/ingredient models were applied. To validate the final cookoff predictions, several levels of validation were completed, including the comparison of model predictions to experimental data for: AP steady-state combustion, fine-AP/HTPB steady-state combustion, AP laser ignition, fine-AP/HTPB laser ignition, AP/HTPB/Al ignition, and AP/HTPB/Al cookoff. A previous AP steady-state model was updated, and then a new AP steady-state model was developed, to predict steady-state combustion. Burning rate, temperature sensitivity, surface temperature, melt-layer thickness, surface species at low pressure and high initial temperature, final flame temperature, final species fractions, and laser-augmented burning rate were all predicted accurately by the new model. AP ignition predictions gave accurate times to ignition for the limited experimental data available. A previous fine-AP/HTPB steady-state model was improved to predict a melt layer consistent with observation and avoid numerical divergence in the ignition code. The current fine-AP/HTPB model predicts burning rate, surface temperature, final flame temperature, and final species fractions for several different propellant formulations with decent success. Results indicate that the modeled condensed-phase decomposition should be exothermic, instead of endothermic, as currently formulated. Changing the model in this way would allow for accurate predictions of temperature sensitivity, laser-augmented burning rate, and surface temperature trends. AP/HTPB ignition predictions bounded the data across a wide range of heat fluxes. The AP/HTPB/Al model was based upon the kinetics of the AP/HTPB model, with the inclusion of aluminum being inert in both the solid and gas phases. AP/HTPB/Al ignition predictions bound the data for all but one source. AP/HTPB/Al cookoff predictions were accurate when compared to the limited data, being slightly low (shorter time) in general. Comparisons of AP/HTPB/Al ignition and cookoff data showed that the experimental data might be igniting earlier than expected.

<div>Enhancing the burn rates of solid propellants and energetics is a crucial step towards improving the performance of several solid propellant based micro-propulsion systems. In addition to increasing thrust, high burn rates also help simplify the propellant grain geometry and increase the volumetric loading of the rocket motor, which in turn reduces the overall size and weight. <b><i>Thus, in this work, burn rate enhancement of solid propellants when coupled to highly conductive graphene-based micro-structures was studied using both experiments and molecular dynamic (MD) simulations.</i></b></div><div><b><i><br></i></b></div><div><div>The experiments were performed using three different types of graphene-structures i.e. graphite sheet (GS), graphene nano-pellets (GNPs) and graphene foam (GF), with nitrocellulose (NC) as the solid propellant.</div></div><div><br></div><div><div>For the NC-GS samples, propellant layers ranging from 25 µm to 170 µm were deposited on the top of a 20 µm thick graphite sheet. Self-propagating combustion waves were observed, with burn rate enhancements up to 3.3 times the bulk NC burn rate (0.7 cm/s). The burn rates were measured as a function of the ratio of fuel to graphite layer thickness and an optimum thickness ratio was found corresponding to the maximum enhancement. Moreover, the ratio of fuel to graphite layer thickness was also found to affect the period and amplitude of the combustion wave oscillations. Thus, to identify the important non-dimensional parameters that govern the burn rate enhancement and the oscillatory nature of the combustion waves, a numerical model using 1-D energy conservation equations along with simple first-order Arrhenius kinetics was also developed.</div><div><br></div><div><div>For the GNP-doped NC lms, propellant layers, 500 30 µm thick, were deposited on the top of a thermally insulating glass slide with the doping concentrations of GNPs being varied from 1-5% by mass. An optimum doping concentration of 3% was obtained for which the burn rate enhancement was 2.7 times. In addition, the effective thermal conductivities of GNP-doped NC lms were also measured experimentally using a steady state, controlled, heat flux method and a linear increase in the thermal conductivity value as a function of the doping concentration was obtained.</div></div><div><br></div><div><div>The third type of graphene structure used was the GF - synthesized using a chemical vapor deposition (CVD) technique. The effects of both the fuel loading ratio and GF density were studied. Similar to the GNPs, there existed an optimum fuel loading ratio that maximized the burn rates. However, as a function of the GF density, a monotonic decreasing trend in the burn rate was obtained. Overall, burn rate enhancement up to 7.6 times was observed, which was attributed to the GF's unique thermal properties resulting from its 3D interconnected network, high thermal conductivity, low thermal boundary resistance and low thermal mass. Moreover, the thermal conductivity of GF strut walls as a function of the GF density was also measured experimentally.</div></div><div><br></div><div><div>Then as a next step, the GF structures were functionalized with a transition metal oxide (MnO₂). The use of GF-supported catalyst combined the physical eect of enhanced thermal transport due to the GF structure with the chemical effect of increased chemical reactivity (decomposition) due to the MnO₂ catalyst, and thus, resulted in even further burn rate enhancements (up to 9 times). The burn rates as a function of both the NC-GF and MnO₂-NC loadings were studied. An optimum MnO₂-NC loading corresponding to the maximum burn rate was obtained for each NC-GF loading. In addition, thermogravimetric (TG) and differential scanning</div><div>calorimetry (DSC) analysis were also conducted to determine the effect of NC-GF and MnO₂-NC loadings on the activation energy (E) and peak thermal decomposition (PTD) temperatures of the propellant NC.</div></div><div><br></div><div><div>In addition to the experimental work, molecular dynamics simulations were also conducted to investigate the thermal transport and the reactivity of these coupled solidpropellant/graphene-structures. A solid monopropellant, Pentaerythritol Tetranitrate (PETN), when coupled to highly conductive multi-walled carbon nanotubes (MWCNTs) was considered. The thickness of the PETN layer and the diameter of the MWCNTs were varied to determine the effect of PETN-MWCNT loading on the burn rates obtained. Burn rate enhancement up to 3 times was observed and an optimal PETN-MWCNT loading of 45% was obtained. The enhancement was attributed to the faster heat conduction in CNTs and to the layering of PETN molecules around the MWCNTs surface. Moreover, the CNTs remained unburned after the combustion process, conrming that these graphene-structures do not take part in the chemical reactions but act only as thermal conduits, transferring heat from the burned to the unburned portions of the fuel.</div></div><div><br></div><div><div>A long-pursued goal, which is also a grand challenge, in nanoscience and nanotechnology is to create nanoscale devices, machines and motors that can do useful work. However, loyal to the scaling law, combustion would be impossible at nanoscale because the heat loss would profoundly dominate the chemical reactions. <b><i>Thus, in addition to the solid propellant work, a preliminary study was also conducted to understand as how does the heat transfer and combustion couple together at nano-scales.</i></b></div></div><div><b><i><br></i></b></div><div><div>First, an experimental study was performed to understand the feasibility of combustion at nano-scales for which a nano-scale combustion device called "nanobubbles" was designed. These nanobubbles were produced from short-time (< 2000 µs) water electrolysis by applying high-frequency alternating sign square voltage pulses (1-500 kHz), which resulted in H₂ and O₂ gas production above the same electrode. Moreover, a 10 nm thick Pt thermal sensor (based on resistance thermometry) was also fabricated underneath the combustion electrodes to measure the temperature changes obtained. A signicant amount of bubble production was seen up to 30 kHz but after that the bubble production decreased drastically, although the amount of faradaic current measured remained unchanged, signifying combustion. The temperature changes measured were also found to increase above this threshold frequency of 30 kHz.</div></div><div><br></div><div><div>Next, non-reactive molecular dynamic simulations were performed to determine as how does the surface tension of water surrounding the electrodes is affected by the presence of dissolved external gases, which would in turn help to predict the pressures inside nanobubbles. Knowing the bubble pressure is a perquisite towards understanding the combustion process. The surface tension of water was found to decrease with an increase in the supersaturation ratio (or an increase in the external gas concentration), thus, the internal pressure inside a nanobubble is much smaller than what would have been predicted using the planar-interface surface tension value of water. Once the pressure behavior as a function of external gas supersaturation was understood, then as a next step, reactive molecular dynamic simulations were performed to study the effects of surface-assisted dissociation of H₂ and O₂ gases and initial system pressure on the ignition and reaction kinetics of the H₂/O₂ system at nano-scales. A signicant amount of hydrogen peroxide (H₂O₂), 6-140 times water (H₂O), was observed in the combustion products. This was attributed to the low temperature(~300 K) and high pressure (2-80 atm) conditions at which the chemical reactions were taking place. Moreover, the rate at which heat was being lost from the combustion chamber (nanobubble) was also compared to the rate at which heat was being released from the chemical reactions and only a slight rise in the reaction temperature was observed (~68 K), signifying that, at such small-scales, heat losses dominate.</div></div><div><br></div></div>

<p>A variety of methods have been developed to enhance solid propellant burning rates, including adjusting oxidizer particle size, modifying metal additives, tailoring the propellant core geometry, and adding catalysts or wires. Fully consumable reactive wires embedded in propellant have been used to increase the burning rate by increasing the surface area; however, the manufacture of propellant grains and the observation of geometric effects with reactive components has been restricted by traditional manufacturing and viewing methods. In this work, a printable reactive filament was developed that is tailorable to a number of use cases spanning reactive fibers to photosensitive igniters. The filament employs aluminum fuel within a printable polyvinylidene fluoride matrix that can be tailored to a desired burning rate through stoichiometry or aluminum fuel configuration such as particle size and modified aluminum composites. The material is printable with fused filament fabrication, enabling access to more complex geometries such as spirals and branches that are inaccessible to traditionally cast reactive materials. However, additively manufacturing the reactive fluoropolymer and propellant together comes attendant with many challenges given the significantly different physical properties, particularly regarding adhesion. To circumvent the challenges posed by multiple printing techniques required for such dissimilar materials, the reactive fluoropolymer was included within a solid propellant carrier matrix as small fibers. The fibers were varied in aspect ratio (AR) and orientation, with aspect ratios greater than one exhibiting a self-alignment behavior in concordance with the prescribed extrusion direction. The effective burning rate of the propellant was improved nearly twofold with 10 wt.% reactive fibers with an AR of 7 and vertical orientation. </p> <p>The reactive wires and fibers in propellant proved difficult to image in realistic sample designs, given that traditional visible imaging techniques restrict the location and dimensions of the reactive wire due to the necessity of an intrusive window next to the wire, a single-view dynamic X-ray imaging technique was employed to analyze the evolution of the internal burning profile of propellant cast with embedded additively manufacture reactive components. To image complex branching geometries and propellant with multiple reactive components stacked within the same line of sight, the dynamic X-ray imaging technique was expanded to two views. Topographic reconstructions of propellants with multiple reactive fibers showed the evolution of the burning surface enhanced by the geometric effects caused by the faster burning fibers. These dual-view reconstructions provide a method for accurate quantitative analysis of volumetric burning rates that can improve the accessibility and viability of novel propellant grain designs.</p>

L’ajout de particules d’aluminium dans le chargement des moteurs à propergol solide améliore les performances propulsives, mais peut aussi entraîner différents phénomènes néfastes, dont des oscillations de pression. Des travaux de recherche sont réalisés depuis de nombreuses années afin d’améliorer la compréhension de ces phénomènes, notamment par l’utilisation de la simulation numérique. Cependant les données d’entrée de la simulation numérique, en particulier la taille et la vitesse initiale des particules d’aluminium dans l’écoulement, sont souvent difficiles à obtenir pour des propulseurs réels. L’ONERA développe depuis plusieurs années un montage d’ombroscopie permettant de visualiser les particules d’aluminium proches de la surface de petits échantillons en combustion. La présente étude porte sur le développement d’outils pour analyser les images expérimentales du montage d’ombroscopie et améliorer l’interaction avec la simulation numérique diphasique. Une première partie concerne des échantillons de propergol contenant des particules inertes, dont l’intérêt est de permettre de valider les méthodes de mesure sur des images relativement simple et avec des données de référence. Les outils mis en œuvre portent sur la détection et le suivi des particules dans des séquences d’image, ainsi que sur la localisation de la surface du propergol. Une bonne correspondance des distributions de taille a été obtenu avec les distributions de référence. La mise en vitesse des particules quittant la surface a été confrontée à un modèle simplifié de transport de particules dans un écoulement constant. L'utilisation de ce modèle a permis de souligner l'importance de la population de pistes détectées pour bien exploiter un profil de vitesse moyen, en particulier en termes de diamètre moyen. Une simulation numérique diphasique a ensuite été réalisée pour l’expérience d’ombroscopie. Différents paramètres ont été étudiées (type et taille de maillage, paramètres thermodynamiques...) afin d'obtenir un champ stationnaire simulé pour les gaz du propergol. Le mouvement des particules inertes simulées a pu être comparé aux profils expérimentaux pour différentes stratégies d'injection, soit en utilisant un diamètre moyen, soit à partir d’une distribution lognormale. L’autre partie de l'étude est consacrée à l’analyse des images expérimentales de la combustion de particules d’aluminium. La complexité des images dans ces conditions a conduit à utiliser une approche de segmentation sémantique par apprentissage profond, visant à classer tous les pixels de l'image en différentes classes, en particulier goutte d'aluminium et flamme d'aluminium. L’apprentissage a été mené avec une base restreinte d’images annotées en utilisant le réseau U-Net, diverses adaptations pour le traitement des images d’ombroscopie ont été étudiées. Les résultats sont comparés à une technique de référence basée sur une détection d’objets MSER. Ils montrent un net gain à l’utilisation de techniques neuronales pour la ségrégation des gouttes d'aluminium de la flamme. Cette première démonstration de l'utilisation de réseau de neurones convolutifs sur des images d'ombroscopie propergol est très prometteuse. Enfin nous traçons des perspectives côté analyse d’image expérimentales et simulation numériques pour améliorer l’utilisation conjointe de ces deux outils dans l’étude des propergols solides. / The addition of aluminum particles in the solid propellant loading improves propulsive performance, but can also lead to various adverse phenomena, including pressure oscillations. Research has been carried out for many years to improve the understanding of these phenomena, particularly through the use of numerical simulation. However, the input data of the numerical simulation, especially the size and the initial velocity of the aluminum particles in the flow, are often difficult to obtain for real rocket motors. ONERA has been developing a shadowgraphy set-up for several years to visualize aluminum particles near the surface of propellant samples in combustion. The present study deals with the development of tools to analyze the experimental images of the shadowgraphy set-up and to improve the interaction with the two-phase digital simulation. A first part concerns propellant samples containing inert particles, which interest is to make it possible to validate the measurement methods on relatively simple images and with reference data. The implemented tools concern the detection and the tracking of particles in image sequences, as well as the location of the surface of the propellant. Good correspondence of size distributions was obtained with reference distributions. The velocity of particles leaving the surface has been confronted with a simplified model of particle transport in a constant flow. The use of this model has made it possible to emphasize the importance of the population of detected tracks in order to make good use of an average velocity profile, particularly in terms of average diameter. A two-phase flow simulation was then carried out for the shadowgraphy experiment. Different parameters were studied (type and size of mesh, thermodynamic parameters ...) in order to obtain a simulated stationary field for propellant flow. The movement of the simulated inert particles could be compared to the experimental profiles for different injection strategies, either using a mean diameter or using a lognormal distribution. The other part of the study is devoted to the analysis of experimental images of the combustion of aluminum particles. The complexity of the images under these conditions has led to the use of a deep learning semantic segmentation approach, aiming to classify all the pixels of the image into different classes, in particular aluminum droplet and flame. The learning was conducted with a restricted base of annotated images using the U-Net neural network, with various adaptations on the processing of the experimental images were studied. The results are compared to a reference technique based on MSER object detection. They show a clear gain in the use of neural techniques for the segregation of aluminum drops of the flame. This first demonstration of the use of convolutional neuronal network on propellant shadowgraphy images is very promising. Finally, we draw perspectives on experimental image analysis and numerical simulation to improve the joint use of these two tools in the study of solid propellants.

<p>This work presents the development and/or application of several laser diagnostics for studying the flame structure of composite propellant flames. These studies include examining the flame structure of novel energetic materials with potential as propellant ingredients, the near-surface flame structure of basic composite propellants, and the global flame structure of propellants containing metal additives.<br></p><p><br></p><p>First, the characterization of the deflagration of various novel energetic cocrystals is presented. The synthesis and development of novel energetic materials is a costly and challenging process. Rather than synthesizing new materials, cocrystallization provides the potential opportunity to achieve improved properties of existing energetic materials. This work presents the characterization of the effect of cocrystallization on the deflagration of a 2:1 molar cocrystal of CL-20 and HMX as well as a 1:1 molar cocrystal of CL-20 and TNT. A hydrogen peroxide (HP) solvate of CL-20 as well as a polycrystalline composite of HMX and ammonium perchlorate (AP) were also studied. A physical mixture of each material was also tested for comparison. The burning rate of each material was measured as a function of pressure. Flame structure during self-deflagration was examined using planar laser-induced fluorescence (PLIF) of CN and OH. The burning rate of the HMX/CL-20 cocrystal and the CL-20/HP solvate closely matched that of CL-20, but the burning rate of the TNT/CL-20 cocrystal was between the burning rate of its coformers. All HMX/AP materials had a higher burning rate than either HMX or AP individually and the burning rate of a physical mixture was found to be a function of particle size. The differences in the burning rate of the physical mixtures and composite crystal of HMX/AP can be explained by changes in the flame structure observed using PLIF. Burning rates and flame structure of the cocrystals were found to closely match those of their respective physical mixtures when smaller particle sizes were used (approx. less than 100 um). The results obtained demonstrate that the deflagration behavior of the coformers is not indicative of the deflagration behavior of the resulting physical mixture or cocrystal. However, changes in the resulting flame structure greatly affect the burning rate.</p><p><br></p><p>Next, PLIF of nitric oxide (NO) was utilized to characterize the near surface flame structure of composite propellants of AP and hydroxyl-terminated polybutadiene (HTPB) containing varying particle sizes of AP burning at 1 atm in air. In all propellants, the NO PLIF signal was strongest close to the burning propellant surface and fell to a non-zero constant value within ~1 mm of the surface where it remained throughout the remainder of the flame. Distinct diffusion-flame-like structure was observed above large individual burning AP particles in the propellant containing a bimodal distribution of 400 and 40 um AP. In contrast, the flame of a propellant containing only fine AP (40 um) behaved like a homogeneous, premixed flame. The flame of the propellant containing a bimodal distribution of 200 and 40 um AP also showed similar behavior to a premixed flame with some heterogeneous structure indicating that, at this pressure, the propellant is approaching a limit where the particle sizing is small enough that the flame behaves like a homogeneous, premixed flame. Additionally, propellants containing aluminum were tested. No significant differences were observed in the NO PLIF behavior between the propellants with and without aluminum suggesting that, at these conditions, the aluminum does not have a significant effect on the AP/HTPB flame structure near the burning surface.</p><p><br></p><p>The effect of aluminum particle size on the temperature of aluminized-composite-propellant flames burning at 1 atm is also presented. In this work, measurements of 1) the temperature of CO (within the flame bath gas) and 2) the temperature of AlO (located primarily within regions surrounding the burning aluminum particles) within aluminized, AP-HTPB-propellant flames were performed as a function of height above the burning propellant surface. Three aluminized propellants with varying aluminum particle size (nominally 31 um, 4.5 um, or 80 nm) and one non-aluminized AP-HTPB propellant were studied while burning in air at 1 atm. A wavelength-modulation-spectroscopy (WMS) diagnostic was utilized to measure temperature and mole fraction of CO via mid-infrared wavelengths and a conventional AlO emission-spectroscopy technique was utilized to measure the temperature of AlO. The bath-gas temperature varied significantly between propellants, particularly within 2 cm of the burning surface. The propellant with the smallest particles (nano-scale aluminum) had the highest average temperatures and far less variation with measurement location. At all measurement locations, the average bath-gas temperature increased as the initial particle size of aluminum in the propellant decreased, likely due to increased aluminum combustion. The results support arguments that larger aluminum particles can act as a heat sink near the propellant surface and require more time and space to ignite and burn completely. On a time-averaged basis, the temperatures measured from AlO and CO agreed within uncertainty at near 2650 K in the nano-aluminum propellant flame, however, AlO temperatures often exceeded CO temperatures by ~250 to 800 K in the micron-aluminum propellant flames. This result suggests that in the flames studied here, and on a time-averaged basis, the micron-aluminum particles burn in the diffusion-controlled combustion regime, whereas the nano-aluminum particles burn within or very close to the kinetically controlled combustion regime.</p><p><br></p><p>The study of the effect of aluminum particle size on the temperature of aluminized, composite-propellant flames was then extended to characterize the same propellants burning at elevated pressures ranging from 1 to 10 atm. A novel mid-infrared scanned-wavelength direct absorption technique was developed to acquire measurements of temperature and CO in particle-laden propellant flames burning at up to 10 atm. The results from the application of this diagnostic are among the very first measurements of gas properties in aluminized composite propellant flames burning at pressures above atmospheric pressure. In all propellants, the flame temperature and combustion efficiency of the propellant flames increased with an increase in pressure. In addition, the propellants with smaller aluminum particle sizes achieved higher flame temperatures as the particles were able to ignite and react faster. However, the propellants containing nano-scale and the smallest micron-scale aluminum powders had similar global flame temperatures suggesting that at some point a decrease in particle size results in minimal gains in the overall flame temperature. The results demonstrate how well measurements of gas properties can be used to understand the behavior of the aluminum particle combustion in the flame.</p><p><br></p><p>Last, the design, development, and application of a laser-absorption-spectroscopy diagnostic capable of providing quantitative, time-resolved measurements of gas temperature and HCl concentration in flames of aluminized, composite propellant flames is presented. This diagnostic utilizes a quantum-well distributed-feedback tunable diode laser emitting near 3.27 um to measure the absorbance spectra of one or two adjacent HCl lines using a scanned-WMS technique which is insensitive to non-absorbing transmission losses caused by metal particulates in the flame. This diagnostic was applied to characterize the spatial and temporal evolution of temperature and/or HCl mole fraction in small-scale flames of AP-HTPB composite propellants containing either an aluminum-lithium alloy or micron-scale aluminum. Experiments were conducted at 1 and 10 atm. At both pressures, the flame temperature of the aluminum-lithium propellant, on a time-averaged basis, was 80 to 200 K higher than that of the aluminum-propellant (depending on location in the flame) indicating more complete combustion. In addition, the mole fraction of HCl in the aluminum-lithium propellant flame reached values 65-70% lower than the conventional aluminum-propellant flame at the highest measurement location in the flame. The measurements at both pressures showed similar trends in the reduction of HCl in the aluminum-lithium propellant flame but at 10 atm this occurred on a length scale an order of magnitude smaller than the flame at atmospheric pressure. The results presented further support that the use of an aluminum-lithium alloy is effective at reducing HCl produced by the propellant flame without compromising performance, thereby making it an attractive additive for solid rocket propellants.</p>

In the present thesis, the objective was to find the maximum amount of payload mass that can be put into a 500 km polar orbit by a 1400 kg air-launched multi-stage rocket launched from a fighter jet platform. To fulfill the objective an algorithm incorporating several modules was developed. The modules performed calculations based on theoretical models and literature values to arrive at optimal design variables. From the design the maximum payload mass was able to be derived and it was concluded that a three-stage launch vehicle was able to deliver a 22.0 kg payload to the desired orbit. / I den här avhandlingen var syftet att hitta den maximala mängden nyttolastmassa som kan transporteras av en 1400 kg flerstegsraket uppskjuten från luften till en 500 km polär bana. För att uppfylla målet utvecklades en algoritm med flera moduler. Modulerna utförde beräkningar baserade på teoretiska modeller och litteraturvärden för att komma fram till optimala designvariabler. Från konstruktionen kunde den maximala nyttolastmassan härledas och det konstaterades att en trestegsraket kunde leverera en nyttolast på 22.0 kg till den önskade omloppsbanan.

In the present thesis, the objective was to find the maximum amount of payload mass that can be put into a 500 km polar orbit by a 1400 kg air-launched multi-stage rocket launched from a fighter jet platform. To fulfill the objective an algorithm incorporating several modules was developed. The modules performed calculations based on theoretical models and literature values to arrive at optimal design variables. From the design the maximum payload mass was able to be derived and it was concluded that a three-stage launch vehicle was able to deliver a 22.0 kg payload to the desired orbit. / I den här avhandlingen var syftet att hitta den maximala mängden nyttolastmassa som kan transporteras av en 1400 kg flerstegsraket uppskjuten från luften till en 500 km polär bana. För att uppfylla målet utvecklades en algoritm med flera moduler. Modulerna utförde beräkningar baserade på teoretiska modeller och litteraturvärden för att komma fram till optimala designvariabler. Från konstruktionen kunde den maximala nyttolastmassan härledas och det konstaterades att en trestegsraket kunde leverera en nyttolast på 22.0 kg till den önskade omloppsbanan.