A hot-flow axisymmetric Air Augmented Rocket (AAR) test apparatus was constructed to test various mixing duct configurations at static conditions. Primary flow for the AAR was provided through a liquid methanol-gaseous oxygen bipropellant rocket. Experimental thrust measurements were recorded and propellant mass flow rates and chamber conditions were calculated using an iterative solver dependant on recorded propellant line stagnation pressures. Primary rocket flow produced thrust ranging from 14 to 17.9lbf. Primary mass flow rate through testing ranged from 0.071 to 0.085lbm/s with calculated chamber pressures between 298-362psia. Calculated primary flow velocity ranged from 6,600ft/s to 8,000ft/s depending on propellant pressure inputs and calculated chamber conditions. The AAR test apparatus was capable of testing various mixing duct geometries and measuring the axial thrust of the mixing ducts separately from the total thrust of the system. Two mixing duct geometries, a straight wall mixing duct and diverging wall mixing duct, with identical exterior dimensions and inlet geometry were tested for a range of air/fuel mixture ratios from 0.82 to 2.2 spanning the stoichometric mixture ratio of 1.5. Mixing duct thrust did not vary greatly with primary flow characteristics. Straight mixing duct thrust averaged 0.97lbf and diverging mixing duct thrust averaged 0.18lbf. Total system thrust decreased by an average of 0.62lbf with a straight mixing duct and 0.74lbf with a diverging mixing duct. Decreases in total thrust are attributed to low pressure flow interaction between the mixing duct and the primary rocket assembly. Visual flow comparison between mixing duct configurations and fuel ratio cases were carried out using high definition video recording with a grid reference for comparison. The diverging mixing duct produced the greatest variation in visible flow when compared to a straight mixing duct and no mixing duct configuration. This indicated that the diverging mixing duct had a greater influence on primary and secondary flow field mixing than the straight mixing duct.

A 2-D Air Augmented Rocket, the Cal Poly Air Augmented Rocket (CPAAR) Test Apparatus operating as a mixer-ejector was tested to investigate high stagnation pressure ratio and transient flow fields of an ejector. The primary rocket ejector was supplied with high pressure nitrogen at a maximum chamber pressure of 1758 psia and a maximum mass flow rate of 1.4 lb/s. The secondary flow air was entrained from a fixed volume plenum chamber producing pressures as low as 3.3 psia. The maximum total pressure ratio achieved was 221. The original CPAAR apparatus was rebuilt re-instrumented and capability expanded. A fixed volume plenum was attached to the secondary ducts through a constant area square section to mimic the cross section of the secondary ducts with a bell mouth inlet. The mixing duct length was increased from 8 in. to 18 in. An investigation of the mixing duct flow-field was done with data from pressure and temperature instrumentation. A study of the transient operation of the rocket was compared with results from former research to qualify the quasi-steady assumption of the flow-field. The CPAAR produced Fabri-choked operation, the startup transient observed caused the secondary flow to become established during Fabri-choke mode operation. The supersonic saturated mode was not observed during quasi-steady operation. The quasi-steady operation was defined based on characteristics from previous quasi-steady models of transient operation of supersonic ejectors. The measurement of the data during testing resulted in a 2.96% experimental uncertainty in the entrainment ratio calculation. The smallest entrainment ratio observed was 0.05 at a total pressure ratio of 220. The location of the Fabri-choke point was shown through the interpretation of the primary and secondary flow as a result of the pressure and temperature measurements. The experimental evidence showed the location of the secondary choke point has a logarithmic relationship with the total pressure ratio. At a total pressure ratio of 220, the area of the aerodynamic throat of the secondary flow is 0.26 in2 and the location occurs 6 inches downstream from the nozzle exit. The secondary flow un-choke is related to the breakdown of the shock structure of the primary flow and produces a flow-field asymmetry which blocks the right duct flow. The CPSE simulation was unable to accurately predict AAR performance when the inputs are changed from the original CPAAR configuration. At high pressure ratios (PR=220), the error in the prediction is 90%.

<p dir="ltr">Rotating detonation engines (RDEs) are a promising technology by which to increase the efficiency of propulsion and power generation systems. Self-sustained, rotating detonation waves within the combustion chamber provide a means for combustion to occur at elevated local pressures, theoretically resulting in hotter temperature product gas than a constant pressure combustion process could provide at equivalent operating conditions. Despite theoretical advantages of RDEs, the thermodynamic benefit has yet to be achieved in experimental applications. Additionally, much of the experimental work to date has been conducted at mean operating pressures lower than industrial applications will require, especially for rocket or gas turbine combustion environments. The sensitivity of these devices to operating pressure has made clear the importance of chemical reaction rates on the successful operation of these combustors. This work addresses critical risks associated with implementing this technology at flight-relevant conditions by advancing the understanding of deflagrative loss mechanisms on delivered performance and by investigating the coupling between chemical kinetic timescales and operating modes produced by the combustor.</p><p dir="ltr">A novel pressure measurement technique was developed in which the stagnation pressure of exhausting gas produced by the RDC is measured through quantification of the under-expanded exhaust plume divergence angle at megahertz-rates. Time-averaged stagnation pressure measurements obtained with this technique are shown to be within 1.5% of the equivalent available pressure (EAP) measured. Time-resolved stagnation pressure measurements produced by this technique provide a means to quantify the detonation cycle pressure ratio. It was shown that increasing the total mass flow rate through the combustor, therefore increasing the mean operating pressure, results in a decrease in both detonation wave velocities and detonation cycle stagnation pressure ratios.</p><p dir="ltr">Numerical modeling of detonations was conducted to understand the coupling of stagnation pressure ratios and wave speeds to deflagrative modes of combustion within rotating detonation combustors. Using the experimental measurements, it is shown that significant amounts of propellant combusts as a result of deflagration prior to (i.e., preburning) and after (i.e., afterburning) the detonation wave. Increasing the RDC operating pressure by 4x is shown to increase the amount of preburned propellant by 4.5x. Relevant chemical kinetic reaction rates of the conditions tested are modeled to increase by 4.5x as well, indicating that the increase in reactant preburning is the result of faster chemical kinetic timescales associated with higher pressure combustion. Results from this testing suggest an operating pressure upper limit for this combustor exists around 20 bar. At these conditions, chemical kinetic rates would be fast enough that deflagration would be the primary mode of combustion and the detonation would not exist. It is suggested that different injector or combustor designs might be able to extend operating limits, however it is unclear if there is a chemical kinetic limit at which no design would be able to overcome.</p><p dir="ltr">Despite significant amounts of deflagrative combustion within the RDC, the vacuum specific impulse produced by the RDC was shown to be between 95.0% and 98.5% of what an ideal deflagrative combustor could produce for most conditions. Given conventional rocket combustors typically operate at specific impulse efficiencies in the range of 90%-99%, it is noted that the RDC tested in this work has demonstrated, at the very least, equal performance to the current state of the art for rocket propulsion combustors while utilizing an effective combustor length (L*) of only 63 mm (2.5 inches). A detailed RDC performance model was developed which considered losses associated with deflagration (both preburning and afterburning) and incomplete combustion. Using measurements obtained from the experiment it is determined that incomplete combustion contributes a larger performance loss than the deflagration which occurs within the combustor.</p><p dir="ltr">A total of 17 parametric studies were conducted experimentally to evaluate the response of the RDC specifically to changes in the propellant chemical reaction timescales. Detonation wave arrival times ranged between 10 microseconds and 178 microseconds as a result of testing at ranges of operating pressures, equivalence ratios, and utilizing nine unique propellant combinations. It was shown that the wave arrival time is primarily a function of chemical kinetic timescales and injector mixing processes. A model using the injector momentum ratio and modeled deflagrative preheat times is shown to be able to closely predict experimentally obtained detonation wave arrival times.</p>

A technical project alongside the University courses can deepen the understanding and increase the motivation for the subject of choice. As a student, there is often a hurdle to start such a project because of a lack of inspiration. And even after overcoming this, the costs associated with such a project may put students off. With my project I show how a 3rd semester Mechanical Engineering student can design and manufacture a rocket engine with all testing components on a student budget. Cost structure and resource planning are explained in detail. I launched the project in December 2020 and in September 2021 it was presented at the StuFoExpo21. A general curiosity for the topic and a basic understanding of mechanical engineering was sufficient for starting the project. Importantly, I gained the most valuable knowledge during the implementation of the project, through active failure-iteration and reading specialised literature. The project is focussed on the design and manufacturing of a rocket engine and its testing components. A special feature is the cooling jacket of the combustion chamber. It has been 3D printed in the SLUB Makerspace, a facility at TU Dresden. Further work packages of the project were the programming of sensors and control systems, first open-air combustion tests of the injector head, safety checks and a Risk & Safety analysis. The first testing and other preliminary work were performed in collaboration with fellow students. During the entire design and manufacturing process I was in continuous exchange with the research group “Space Transportation” of the Institute of Aerospace Engineering at TU Dresden. Special thanks go to Dipl.-Ing. Jan Sieder-Katzmann and Dipl.-Ing. Maximilian Buchholz for their help during this process. For 2022 I plan a test campaign of the rocket engine to collect sensor data and to perform engine thrust measurements. / Ein selbständiges, fachbezogenes Projekt neben dem eigentlichen Studium kann die Kenntnisse vertiefen und die Motivation für das Studienfach fördern. Oft jedoch fehlt Studenten die Inspiration für ein solches Projekt. Und selbst wenn diese Hürde genommen ist, schrecken die Kosten davon ab zu beginnen. Am Beispiel meines Projektes im Bereich Luft- und Raumfahrt zeige ich, wie man als Maschinenbaustudent im 3. Semester einen Raketenantrieb und alle Testkomponenten mit einem Studentenbudget entwerfen und herstellen kann. Kostenstruktur und Ressourcenplanung werden im Detail erläutert. Das Projekt startete ich im Dezember 2020. Im September 2021 wurde es auf der StuFoExpo21 vorgestellt. Für den Beginn reichten Neugier und ein Grundlagenverständnis im Maschinenbau aus. Die meisten und wichtigsten Wissensbausteine erlernte ich während des Projektes durch aktive Fehleriteration und aus der Fachliteratur. Das Projekt umfasst das Design und die Fertigung eines Raketenantriebs und der dazugehörigen Testkomponenten. Eine Besonderheit, die Kühlkammer-Ummantelung des Triebwerks, wurde unter Nutzung der Ressourcen an der TU Dresden, SLUB, mit einem 3D-Drucker hergestellt. Weitere wichtige Schritte waren die Programmierung der Sensorik und der Steuerungseinheiten für den Test, ein erster offener Injektor Test mit Treibstoffverbrennung, Sicherheitstests und eine „Risk & Safety Analysis“. Tests und Vorbereitungsarbeiten erfolgten in Zusammenarbeit mit Kommilitonen - der gesamte Entwicklungsprozess fand in ständigem Austausch mit der Forschungsgruppe „Raumtransportsysteme“ des Instituts für Luft- und Raumfahrt in Dresden statt. Besonderer Dank gilt Dipl.-Ing. Jan Sieder-Katzmann und Dipl.-Ing. Maximilian Buchholz für ihre wertvollen Hinweise. Für 2022 ist eine Testkampagne des Raketentriebwerks geplant, um Sensorwerte aufzuzeichnen und den Schub des Triebwerks zu messen.

<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>

<p> A proposal is made to extend the present capabilities of gun-launched rocket vehicles to include attitude control during flight. The problems involved are stated and design criteria for possible sensors are listed. A review of presently available sensing devices is made and rejection of unsuitable instruments is based on fundamentals of their design and operation. </p> <p> A report is made upon the sensors which most adequately fulfil the harsh environmental requirements of gun-launch. These sensors are infrared-horizon sensors and a tuning fork vibratory gyroscope. A preliminary design is given for the tuning fork gyroscope a well as a summary of fundamental design considerations. </p> / Thesis / Master of Engineering (ME)

A swirl injector for a rocket engine being developed by \emph{AESIR} (Association of EngineeringStudents in Rocketry) was simulated with different geometric parameters. The swirl injector is usedto atomize the ethanol used as fuel and to create a spray that mixes well with the oxidizer withinthe combustion chamber. Inlet slot angle (90, 75, 60 and 45 degrees), swirl chamber length (15, 20and 25 mm) and outlet orifice diameter (3, 6 and 9 mm) were examined.Previous studies in swirl injectors show that CFD can be used to analyze the flow in such aninjector, furthermore theoretical models exist that can predict some of the general characteristicsof the flow. Previous studies have also simulated transient behavior and flow features effectingbreakup of fuel flowing through a swirl injector.A steady state simulation using Volume of Fluid (VOF) multiphase modeling and $k$-$\omega$ \emph{SST}turbulence modeling was used to simulate the swirl injector intended for the rocket engine. It wasfound that a wider outlet orifice would give a wider cone angle of spray. This is desirable in thecurrent rocket engine design as it will promote greater mixing of fuel and oxidizer higher up in thecombustion chamber. No large variances was observed when different inlet slot angles was simulated. Ashorter swirl chamber length reduced the amount of losses in energy due to viscous forces. The flowafter the outlet orifice was not simulated so the effect of turbulence kinetic energy and energylosses outside of the swirl injector have not been analyzed, previous studies have indicated thatturbulent kinetic energy does have an effect on the breakup and atomization of the fuel.It was concluded that using a wider outlet orifice of 9 mm gave the best results out of the differentgeometric parameters analyzed and the swirl chamber length should be a short as possible.

We study numerical methods for solving the nonlinear porous medium and Navier-Lame problems. When coupled together, these equations model the flow of exhaust through a porous medium, soil, and the effects that the pressure has on the soil in terms of spatial displacement. For the porous medium equation we use the Crank-Nicolson time stepping method with a spectral discretization in space. Since the Navier-Lame equation is a boundary value problem, it is solved using a finite element method where the spatial domain is represented by a triangulation of discrete points. The two problems are coupled by using approximations of solutions to the porous medium equation to define the forcing term in the Navier-Lame equation. The spatial displacement solutions can be used to approximate the strain and stress imposed on the soil. An analysis of these physical properties shows whether or not the material ceases to act as an elastic material and instead behaves like a plastic which will tell us if the soil has failed and a crater has formed. Analytical as well as experimental tests are used to find a good balance for solving the porous medium and Navier-Lame equations both accurately and efficiently.

No description available.

Energetic materials (i.e., explosives, propellants, and pyrotechnics) have been used for centuries in a wide variety of applications that include celebratory firework displays, the demolition of ‘immovable’ structures, mining resources from the earth’s crust, launching humans into outer space, and propelling munitions across the battlefield. Many different scientific and engineering domains have found unique value in their characteristic release of significant heat and pressure. While the rate at which energetic materials react is often dependent on the source of initiation, surrounding thermodynamic conditions, and formulation sensitivity, many applications aim for a controlled combustion process to produce large amounts of work output – solid and liquid rocket motors and gun-launched projectiles are a few key examples. Other energetic material systems are often inadvertently exposed to thermal insults, which can result in similar combustion behavior. To accurately model these systems, it is important to have a fundamental understanding of the chemical kinetics that control various aspects of the combustion process (e.g., changes in temperature (T), pressure (P), and species mole fractions (X)). Detailed chemical kinetic models are often used to understand and subsequently predict such behavior. Understanding the gas-phase reaction kinetics of energetic materials is essential when trying to predict critical performance parameters such as flame speeds, temperature and pressure profiles, and heat flux between material phases. These parameters can have significant impact on predictions of system-level performance (e.g., the specific impulse of solid rocket motors, propellant burn rates in projectile systems, and munition responses to thermal insult and extended temperature cycling). While the gas-phase reaction kinetics of energetic material combustion were heavily studied from the late 1970’s to the early 2000’s, research efforts beyond this time frame have primarily focused on condensed-phase chemistry as it is thought to be less understood. Over the past two decades, however, there have been significant advances in our understanding of small molecule reactions that have not yet been accounted for in many energetic material models. One such example are chemically termolecular reactions – a new class of phenomenological reactions that have not yet been considered for inclusion in any energetic material kinetic models. Recent studies have indicated that chemically termolecular reactions, mediated through ephemeral collision complexes, have significant impact on the global kinetics of certain combustion systems. This discovery has since prompted the question of which systems are significantly influenced by chemically termolecular reactions and should therefore account for their presence in gas-phase phenomenological models. Although a select number of systems have already been investigated, such as flame speed and ignition delay predictions in common hydrocarbon combustion scenarios, the influence of chemically termolecular reactions on the kinetics of energetic materials has not yet been explored. As an initial investigation into energetic materials, a case study for RDX was performed, for which abundant computational and experimental data are available. To aid in assessing the impact of chemically termolecular reactions, for which almost no data are available, this study leveraged an automated procedure to identify and estimate rate constants for potential chemically termolecular reactions based exclusively on data available for related reactions. Four detailed kinetics models for RDX were independently screened for potential chemically termolecular reactions. Model predictions including these chemically termolecular reactions revealed that they have significant potential impact on profiles of major species, radicals, and temperatures. T he analysis pinpointed ∼20-40 chemically termolecular reactions, out of the thousands of possibilities, estimated to have the largest impact. These reactions, including many mediated by ephemeral HNO** and NNH** complexes, are therefore worthwhile candidates for more accurate quantification via master equation calculations. More generally, just as the importance of including chemically termolecular reactions in hydrocarbon combustion models is becoming recognized, the present results show compelling evidence for the need for their inclusion in energetic material models as well. The investigation into chemically termolecular reactions yielded a secondary conclusion based on the observed influence of the small molecule C/H/N/O chemistry on overall predictions of energetic material combustion – updating the small molecule chemistry in RDX models produced significant changes to predictions of major species and temperature, suggesting that the development of a comprehensive gas-phase energetic material combustion model would be of great value and have broad utility as a foundational model for a great variety of C/H/N/O energetic materials. To begin developing such a model, all small molecule chemistry in current kinetic models was reviewed with the intent of identifying a sub-model in need of revisions and subsequently addressing its uncertainties using targeted experiments to improve overall predictions. The ammonia sub-model was selected as it is both highly uncertain and highly influential in many energetic material models. Ammonia (NH₃) has garnered substantial attention in recent years due to its importance across many scientific domains – including its potential use as a carbon-free fuel and long-term energy storage option, its use in reducing combustion-generated nitrogen oxide emissions, its role as a decomposition fragment of many energetic materials, and its presence as an important impurity during biofuel and biomass combustion that can affect overall system kinetics, among others. Yet, it is generally recognized that there are still significant gaps in the present understanding of ammonia kinetics -– in both experimental data sets and sub-models within the overall ammonia kinetic mechanism. For example, most experimental studies of ammonia oxidation have used molecular oxygen as the primary or sole oxidizer. While large mole fractions of molecular oxygen are encountered in many combustion scenarios, there are select systems where ammonia is more likely to be oxidized via nitrogen-containing species (e.g. N₂O and NO₂) and, more generally, there are relatively untested reaction sets that would be accentuated in such conditions. To address these gaps in available experimental data needed for the validation of ammonia kinetics models, jet-stirred reactor experiments were performed for mixtures of NH₃/N₂O/N₂ over an intermediate temperature range (850-1180 K). In these experiments, the mole fractions of NH₃, N₂O, and NO were measured using a combination of gas chromatography, chemiluminescence, electrochemical detection, and infrared absorption – where agreement among the different diagnostics (within 3% for N₂O and 7% for NO) ensured high confidence in the experimental measurements. Comparison of the experimental results and model predictions suggested deficiencies in commonly used models for nitrogen kinetics. Various modeling analyses pointed to the central role of the N₂O + NH₂ = N₂H₂ + NO reaction, on which recent kinetic models all rely on the same rate constant estimate that appears to have not been tested in previous validation data sets for NH₃ kinetics. A second set of jet-stirred reactor experiments were performed for mixtures of NH₃/NO₂/O₂/N₂ over a slightly different temperature range (700–1100 K). Agreement among different diagnostics (≤7% for NO₂ and ≤4% for NH₃) and excellent experimental repeatability confirmed high confidence in all species measurements. Measured mole fractions were compared to predictions from five recently developed kinetic models using flux analysis and uncertainty-weighted kinetic sensitivity analysis, both of which pointed to the importance of reactions involving H₂NO that are both influential in this system and highly uncertain. The measurements from the jet-stirred reactor experiments presented here were combined with comprehensive sets of experimental data and high-level theoretical kinetics calculations using the MultiScale Informatics (MSI) approach to unravel the large uncertainties present in current NH3 oxidation kinetic sub-models. Emphasis was placed on NH₃ oxidation via nitrogen-containing species as this chemistry has been shown to accentuate influential reactions (e.g., the NO₂+NH₂ and NH₂+NO reactions) that are known to be important during the combustion of many energetic materials (e.g., AN, ADN, and AP). The resulting MSI model accurately predicted nearly all of the experimental and theoretical target data within estimated or reported uncertainties. Additional predictions of two NH₃/NO₂ validation data sets, which were not included in the MSI framework, demonstrated its ability to accurately extrapolate predictions to untested T/P/X conditions, indicating that the converged MSI model demonstrates truly predictive behavior. The MSI NH₃ oxidation model presented here should be considered for inclusion in many energetic material models as the NH₃/NOₓ kinetic system is known to be important to the combustion of various propellant and explosive formulations. This sub-model will help to form a foundational gas-phase kinetic model relevant to many different energetic materials, including those that contain inorganic additives for increased energy density and blast effects.