Rocket Propulsion

The Hybrid Graphene Ion Missile (HGIM) represents a next-generation missile system that integrates hybrid propulsion, graphene-based stealth technology, and artificial intelligence (AI)driven control for enhanced operational effectiveness. By combining chemical and ion propulsion, leveraging graphene oxide (GO) micro-antennas, and utilizing plasma stealth modulation, this system offers superior range, adaptability, and survivability in modern warfare. This white paper details the technological advancements, design methodology, and strategic advantages of HGIM.

The aerodynamic performance and stability of a rocket during flight are significantly influenced by the design and orientation of its control surfaces. This study investigates the effects of varying grid fin cant angles on the aerodynamic forces and stability characteristics of a rocket. Using ANSYS Fluent and Open Rocket, simulations are conducted at a constant velocity of 100 m/s, with the cant angle ranging from 0° to 15° in increments of 2.5°. The primary objective is to analyze how different grid fin configurations impact drag force, lift characteristics, and overall flight stability. The results provide valuable insights into optimizing grid fin designs for improved aerodynamic efficiency and control, contributing to the enhancement of future highperformance aerospace vehicles.

Exploration Mission 2 (EM-2) will be NASA's first manned flight on the Space Launch System (SLS) and Orion Spacecraft. The mission has been changed from an SLS Block 1B configuration to Block 1. This change has necessitated a reexamination of the flight profile to determine what changes must be made in order to accommodate the reduced launch vehicle performance on the Block 1. Launch availability and orbital debris risk will be traded to find the best flight profile for both SLS and Orion.

Sonic nozzles are emerging as crucial reference devices in the calibration of flow meters designed for hydrogen service. In this research study we aim to investigate existing equations governing the critical flow factor (C*) for hydrogen and assess the achievable uncertainty in determining this vital property. An examination of the literature has been undertaken to target experimental measurements related to hydrogen mixtures, setting the stage for a comprehensive gap analysis. We introduced C* values and validated our calculations with two calibration gases: nitrogen, and methane and their standardised C* values. Then, the verified methodology has been utilised for generating C* values for hydrogen. This study concluded that the integration of precise experimental data and the utilisation of representative equations and optimised thermodynamic models is essential for enhancing the accuracy of C* calculations, particularly in the context of the expanding role of sonic nozzles in hydrogen flow meter calibration.

The fundamental physics of interstellar travel is a particular set of mathematical formalisms. Relativistic mathematics choreographs a new dance between velocity and energy to reveal the beauty of a perfectly-executed tango. Cosmic forces dance restlessly at the threshold of human ambition, captivating the viewer in quantum possibilities. Relativistic constraints may bind the voyager to cosmic law, but these constraints also allow direct examination of the energy requirements that illuminate the path through darkness. The right minds may transform this assessment of practical engineering limitations from a theory into a journey, but this revisitation of Manifest Destiny should be approached with humility and retrospective caution.

The convergent-divergent design of the nozzle of a rocket engine is the most defining part of it. The nature of the nozzle design enables it to convert the high pressure, temperature and relatively low velocity adiabatic expanding gaseous fluids into very high velocity and thrust gases. The rapidly exiting gas as a consequence of the rapid expansion process results in low pressure and temperature. It is critical in research to find out the influence of modification (grooves) would have in the application and design of a reaction device such as a rocket engine. The Computational Fluid Dynamics software, Analysis System (ANSYS) was employed together with necessary boundary conditions to design and simulate such a modified rocket engine nozzle. The interaction and relationship existing among the pressure, velocity, Mach number and temperature in the grooved divergent rocket nozzle was found and observed. From the plots and nozzle simulation contours, both velocity and Mach number increased rapidly along the nozzle area but decreased steadily. Pressure varied directly withtemperature and inversely with density. It is observed that divergent nozzle grooving influences the velocity, Mach number, pressure and temperature within the divergent nozzle area of a rocket nozzle.

The convergent-divergent design of the nozzle of a rocket engine is the most defining part of it. The nature of the nozzle design enables it to convert the high pressure, temperature and relatively low velocity adiabatic expanding gaseous fluids into very high velocity and thrust gases. The rapidly exiting gas as a consequence of the rapid expansion process results in low pressure and temperature. It is critical in research to find out the influence of modification (grooves) would have in the application and design of a reaction device such as a rocket engine. The Computational Fluid Dynamics software, Analysis System (ANSYS) was employed together with necessary boundary conditions to design and simulate such a modified rocket engine nozzle. The interaction and relationship existing among the pressure, velocity, Mach number and temperature in the grooved divergent rocket nozzle was found and observed. From the plots and nozzle simulation contours, both velocity and Mach number increased rapidly along the nozzle area but decreased steadily. Pressure varied directly withtemperature and inversely with density. It is observed that divergent nozzle grooving influences the velocity, Mach number, pressure and temperature within the divergent nozzle area of a rocket nozzle.

The nozzle is an important component of the rocket motor system, and a rocket’s overall performance is highly reliant on its aerodynamic design. The nozzle contour can be meticulously shaped to improve performance significantly. The design and shape of rocket nozzles have evolved over the last several decades as a result of extensive research. The nozzle design is composed of two components, an integrated throat, an entry and an exit cone, and a thermal protection system. The Bell Nozzle is designed to provide clearance space for placing the ITE and exit cone, with a cone inflection angle of 16 and a thermal protection system. This paper intends to review and summarize all such developments. Small-scale engine testing allows for the analysis of rocket nozzle materials, but the history of nozzle surface temperature and thermal stress may be adversely affected by side effects. The review focuses primarily on the nozzle shape which has the largest radiative flux past the neck, but the ...

The propellant used in space programs creates exclusively three ecological concerns: Ground-based Impacts which vary from groundwater contagion to explosions caused by inconvenient management of fuels. Atmospheric Impacts are generally due to the interface of propellant exhaust with the atmosphere. Biological Impacts include toxicity and corrosiveness of propellants. This leads to the emergence of Cryogenic and Semi Cryogenic engines where the main propellants are Liquid Hydrogen and Liquid Oxygen (very effective in controlling above mentioned three environmental concerns). The operation necessities designate a need for reliable and long-life LH 2 & LOX turbopumps. Existing turbopumps use rolling component bearings which restrict turbopump life. Foil Gas Bearings (FGB) can be effectively used in high-speed turbomachines to attain noteworthy consistency. The present work reports how FGB can successfully replace the conventional roller bearings and prove to be a very important contribution towards Green Propulsion Technology. FGB offers elimination of oil lubrication system (an environmental concern) resulting into condensed weight and improved temperature competence. Using FGB to cryo turbopumps would lengthen turbopump life well ahead of operation necessities. This work also focuses on how Foil Gas Bearing technology can be a great prospect for Indian Space Programmes.

This paper outlines the development and implementation of a Thrust Vector Control (TVC) system specifically designed for solid propellant rocket engines, aiming to enhance the portability of hybrid and liquid propellant rocket systems. The proposed system ensures trajectory stability and rapid response to flight emergencies by effectively counteracting external perturbations such as wind. Key components include gyroscopic (GYRO) sensors, a high-performance microcontroller, and servo motors, collectively referred to as thrust vectoring components, which dynamically adjust thrust direction opposite to the rocket's trajectory to maintain stability. Drawing from an extensive literature review of existing TVC system designs, with a focus on thrust characteristics and engine combustion timing, the design integrates geometric properties, kinematics, forces, energy requirements, safety, cost, and control methods, aligning with both mechanical and avionic design principles. The anticipated outcome is an advancement in rocket propulsion technology, characterized by improved angular speed control, trajectory linearity, and rapid response capabilities.

El objeto de este trabajo consiste en la determinación de maniobras evasivas óptimas de una aeronave contra misil, utilizando la teoría de control. Partiendo de las funciones de transferencia de cada uno de los bloques del sistema de guiado, aplicando la transformada de Laplace a las señales, y empleando la teoría de control, se obtiene un sistema de ecuaciones diferenciales que describe la dinámica del proceso, el cual una vez resuelto permitiría obtener los extremos relativos de la gráfica de distancia de paso. El principal problema en esta fase es conocer las constantes de tiempo que caracterizan la función de transferencia. Si los sensores encargados de la detección de amenazas del avión atacado, son capaces de asegurar, con elevada probabilidad, la identificación concreta del misil agresor y se conoce su función de transferencia (dentro de un catálogo conocido a través de técnicas de inteligencia), el sistema de ecuaciones diferenciales puede integrarse numéricamente y calcularse la distancia de paso misil-avión atacado. En caso de no conocer las constantes de tiempo, se considera proceder a usar un algoritmo genético, integrar el sistema de ecuaciones diferenciales y obtener una distancia de paso calculada. Si los sensores de detección de amenazas del avión atacado, obtienen la distancia de paso real en una fase inicial del vuelo, se puede comparar la distancia de paso real con la distancia de paso calculada y aplicando el algoritmo genético en un proceso de optimización iterativo, converger a unas constantes de tiempo donde la distancia de paso real y calculada estén muy próximas.

This study addresses the inadequacy of conventional failure analyses, which, in addition to regulatory and customer requirements,
often neglect organizational needs. It emphasizes the importance of a systemic approach to mitigating hazards in complex space
program management. This article proposes a new approach to addressing security issues that adds the management of security related
organizational needs to systemic engineering analysis. The case study of the catastrophic event involving SpaceX’s Starship
SN10 prototype used publicly available information to build the system-theoretic accident model and processes (STAMP) model
and identify organizational needs. The causal analysis based on systems theory (CAST) method was then applied to identify
possible causes. Finally, the system-theoretic process analysis (STPA) method was used to determine design-related organizational
needs and formulate recommendations for the design of the autogenous pressurization system. The presented method considered
organizational needs to identify the key elements involved in the accident, the primary causes, and the actions to mitigate the
associated hazards. This study proposed that managing organizational needs for system safety requires recognizing the current
situation and constructing prospective scenarios to prevent failures while emphasizing the importance of management’s proactive
measures, clear responsibilities, and active involvement of all members to ensure system reliability.

In the future, two important technological dreams will have become reality: fusion will be a viable power source, and human settlement on Mars will be feasible, desirable, and even necessary. Merging these two concepts is especially attractive for the aerospace ...

The development of aerospace technology in the future depends on supersonic and hypersonic vehicles. As the national defense and military strength become more and more crucial for national growth, communication, security and stability, the high-efficiency thrust vectoring nozzle, which can significantly enhance the maneuverability and stability of supersonic and hypersonic vehicles, has always been the focus of research for military powers. This thesis paper presents a numerical study on the flow characteristics and performance of baffled shock two dimensional vector nozzle. The baffled shock vector nozzle is a type of fluidic thrust vectoring nozzle that uses secondary injection to deflect the primary flow and generate a vector angle. Fluidic thrust vectoring technology is regarded as a key technology for the development of very low detectable vehicles because of its advantages such as fast response, light weight and good stealth performance. The main objectives of this study are to investigate the effects of various parameters such as slot interval distance, slot width, injection angle, nozzle pressure ratio, secondary flow pressure ratio and outflow Mach number on the deflection angle, thrust coefficient, thrust efficiency and secondary flow ratio of the nozzle. The numerical simulations are carried out using k−epsilon turbulence model, which is validated by comparing with experimental data. The results show that the nozzle performance can be improved by optimizing the slot interval distance and width, increasing the injection angle, adjusting the nozzle pressure ratio and secondary flow pressure ratio, and controlling the outflow Mach number. The results also reveal the complex flow phenomena inside the nozzle, such as shock wave interactions, flow separation and reattachment, and boundary layer effects. The study provides a comprehensive understanding of the flow characteristics and performance of baffled shock two dimensional vector nozzle and offers some guidance for its design and optimization.

We propose a conceptual study of prototype strained hydrocarbon molecules as high energy density additives for hybrid rocket fuels to boost the performance of these rockets without compromising safety and reliability. Use of these additives could extend the range of applications for which hybrid rockets become an attractive alternative to conventional solid or liquid fuel rockets. The objectives of the study were to confirm and quantify the high enthalpy of these strained molecules and to assess improvement in rocket performance that would be expected if these additives were blended with conventional fuels. We confirmed the chemical properties (including enthalpy) of these additives. However, the predicted improvement in rocket performance was too small to make this a useful strategy for boosting hybrid rocket performance. Detailed Description Hybrid rockets use a solid fuel with a liquid or gaseous oxidizer. They have advantages over solid fuel rockets for certain space applications because the solid fuel component is more stable than a combined solid fuel and oxidizer, and have advantages over liquid fuel rockets because the fuel does not require cryogenic storage or pumping and the burn rate can be controlled by regulating the flow of oxidizer. Ideal fuels should exhibit a fast regression rate, high density and good mechanical stability (so it will not break apart during burning). A current project at ARC is to evaluate hybrid rockets comprised of solid paraffin fuel and cryogenic nitrous oxide and LOX oxidizer and has shown that higher regression rates are needed. Strained cyclic hydrocarbons have positive enthalpy and can be blended with paraffins to make fuels for hybrid rockets, but to be effective the resulting solid must have physical and mechanical characteristics at least comparable to the pure paraffin fuels. The stable hydrocarbon molecules with the highest strain energy are comprised of triangular rings of carbon atoms (cyclopropane), but these are gases at normal temperatures and pressures. Recently, chemists have synthesized molecules (called ivyanes, shown in Figure 1) containing multiple cyclopropane rings bonded together. We propose investigating blends of ivyanes and paraffins to determine their Chem. 2010, 75, 4387) demonstrated thermochemical properties can be determined to an accuracy of 2 kJ/mol for hydrocarbon molecules containing 10-20 carbon atoms. In the second phase, the density, melting temperature and heat of sublimation of solid ivyanes and the blends will be determined from molecular dynamics simulations. Capabilities provided: Use of hydrocarbon additives could solve problems of fuel pellet break up encountered with aluminum or lithium additives without sacrificing higher regression rates (burn rates), and lead to enhanced safety and lower cost compared to solid and liquid rocket propulsion systems.

The Wind-US flow solver has been modified to include real gas thermodynamic properties and real gas fluxes of conserved variables. The real gas evaluations relevant to a specific molecule have been isolated to a single subroutine so that any substances can be implemented relatively easily. Molecular nitrogen and hydrogen have been implemented in the program to date. The real gas model has been implemented in the Wind-US Roe approximate Riemann solver at this time. The real gas nitrogen model has been used to predict converging-diverging nozzle flow at total pressures ranging from nominally 250 to 20,000 psi at a nominal stagnation temperature of 3000oR and exit Mach numbers of 8 to 14. A comparison of the 20,000-psi results with data is presented in this paper. The real gas hydrogen model is used to model high-pressure shock tube calculations relevant to analysis of light gas hypervelocity gun range facilities. I.

The design and analysis of propellant grain configurations is a crucial step in the design of solid propellant rocket motors. This is because the performance prediction relies on accurate calculations of grain geometrical properties. Knowing the grain's burn back phases makes solid rocket motor performance prediction a breeze. This research looked into grain burn back analysis for solid rocket motors using 3-dimensional star grain geometries. Using CATIA's parametric modeling capabilities and the dynamic factors that define the intricate arrangement, the design process begins. The grain configuration is defined by the initial geometry, which is a surface. At each web increment, new surfaces are made and geometrical parameters are calculated to ensure grain burn back. To determine the interior ballistics, the equilibrium pressure approach is employed. Using this method for preliminary grain configuration design is a breeze, and it works with any complicated geometry. This paper calculates the variation of thrust with respect to time as the propellant in the igniter burns, reducing the area of the remaining propellant and causing a change in pressure in the Solid Rocket Motor with respect to time. This change in pressure causes a variation in mass flow rate.The remaining grain area in the Solid Rocket Motor is determined by utilizing MATLAB with a 0.05 mm half set. We use MATLAB to double-check the CATIA numerical findings and make sure we're looking at the right area of the remaining propellant.

The Aerospace Corporation conducts extensive research regarding space traffic management (STM) and orbital debris issues, both at the behest of external customers, as well as internal, self-directed efforts. We are currently in the process of coordinating our disparate research efforts under an enterprise-wide rubric, with a goal of advancing a comprehensive research program, as resources permit. This multi-year effort will allow us to gain a synoptic view of the research terrain, with an eye towards discovering gaps in our efforts, and filling them. Aerospace will be examining STM, debris and space safety in a holistic fashion, since these elements have a number of dependencies on each other both reinforcing and opposing. We have conducted research in this area previously , but are now working to integrate many of the various research threads that we find across the corporation into a more integrated program of research, which is rather a new phase for us. Initially, we will deal with technical and policy analysis, with the output consisting of a series of studies and technical demonstrations that will guide and inform our subsequent research portfolio. The idea is not to develop a completely comprehensive, all-encompassing research program, but rather to leverage some of the unique strengths and capabilities that Aerospace has. Our starting points will be to expand assessments of large-scale approaches to STM with the goal of identifying integrated policy considerations and understanding the interactions between sensing and tracking capabilities and STM and the space environment. Topic areas we have identified for initial focus include: • Assessment of the interplay between short-term satellite safety and long-term environment evolution (see Sorge et al., • The interplay of mitigation and active debris removal (ADR) • The trade-offs between distributed satellite systems versus more capable, complex satellites on STM and environmental effects • Studies to provide quantitative data to define STM metrics and thresholds to enable definition of the performance levels needed for performing STM missions Subsequent phases of the program will investigate issues involving space tracking data integration, effective collision avoidance (COLA) system development and integration, mitigation technologies and techniques, and advanced collision prevention methods. The program will not be conducted in an entirely serial fashion, with one phase starting at the completion of the previous one, but rather in a phased approach, with sub-elements from a particular phase being initiated as we understand the logic and inter-connectedness of the various sub-elements. Although this is mainly an internal research effort, we feel the outcomes will be of interest to other researchers in the field, regulators, policy makers, and those working on relevant space safety and sustainability standards and best practices. We invite discussion and collaboration.

Past experience with ramjet engines suggests that combustion instabilities and associated flow oscillations may also occur in scramjet engines, although the problem has not been recognized and explored yet. The present work investigated thermoacoustic instabilities inside an ethylene-fueled scramjet combustor with a recessed-cavity flameholder. Detailed characterization was made of flame oscillation and acoustic motion using high-speed photography and pressure measurements. The effects of fuel/air equivalence ratio, fueling scheme, and simulated flight conditions on the stability characteristics in the engine flowpath were examined systematically. In addition, an analytical analysis based on the acoustic and convective feedback loops among the precombustion shock, fuel injection, and distributed flame was established to help understand and correlate the observed oscillation phenomena. The oscillation frequencies fall in a range of 100-400 Hz, with the upper bound dictated by the shockflame coupling mechanism and the lower bound by the injector-flame interaction.

Today the field of space transportation is of high interest, as a result of the growing competition world-wide. Of particular note are the new Space Companies in the US, which have increased the pressure that the global space transportation industry is under. With this is mind, DLR proposes a new ceramic rocket engine concept, with the goal of increased efficiency and reliability. Additionally, weight, cost and operational risk will be reduced, aiming to be competitive on the global scale in the future. Concerning these requirements, a specific rocket thrust chamber design, based primarily on the application of transpiration cooled porous and thermo-chemically resistant CMCs as inner combustion chamber liner material, is favored, aiming on the improvement of today's high performance standards, e.g. typical high performance main stage or upper stage propulsion systems. After more than twenty years of intensive technology development, DLR offers within its Black Engine technology program a large portfolio of fiber reinforced structure systems for functional components in rocket thrust chambers or entire rocket engines respectively, using additionally CFRP for load carrying structure components or transpiration lubricated CMC journal bearings for lifetime increase of rocket TPs. One major focus lies on the high ratio of thermal and mechanical load de-coupling capability, promising rocket engines with lifetime and maintenance standards comparable to aviation engines, which will be a significant progress for highly reusable main stages in the near future. The exclusively transpiration cooled ceramic rocket thrust chamber system, leads to several additional system improvements. First, evaluations are ongoing considering a new injection cooling method, where the porous CMC wall takes inherently both the function of cooling and injection. This new method is applied in a novel design concept, reducing significantly typical pressure loss in the combustion chamber process and promising an overall engine efficiency increase of more than 5 %. Not only is the increase in efficiency important, but also the increase in reliability by applying higher blow rates for higher inner wall safety. Technology approaches described above are a world first. Established international producers of launcher propulsion systems usually use metal designs for thrust chambers, with recent improvements coming from AM methods, increasing manufacturing efficiency. DLR's new system improvements using porous CMCs as key components in rocket thrust chambers as well as CMC journal bearings in rocket TPs predicts upcoming international competitiveness in the field of space propulsion.

PPRIME Institute and CNES have initiated research activities to investigate the propulsion behaviour of liquid storable propellants, and the related phenomena: atomization of propellants, ignition ability, combustion stability, propulsion efficiency. For this purpose, a specific facility "PERGOLA" has been designed with high modularity, allowing physical diagnostics in a pressurized combustor featuring relevant flow conditions (a few hundred g/s, 5 MPa). The first results demonstrate the capability of PERGOLA facility and characterize the behaviour of a reference storable bipropellant system: ethanol and hydrogen peroxide. The overall combustion performance is also characterized in terms of stability, characteristic velocity and efficiency.

We apply a particle code (Direct Simulation Monte Carlo, DSMC) to a nozzle flow expansion setup. We obtain good results with respect to Pitot pressure measurements based Mach number estimations at different axial distances from the nozzle exit plane. The relative deviation from the measured references is mostly about 10 %.

Grain size distribution during grain growth in austenitic stainless steels was investigated using three different methods, namely serial sectioning analysis, 3D Monte Carlo simulation, and Abbruzzese-Lucke statistical approach. The results showed that the volumetric grain distribution can be well estimated by the log-normal and gamma distribution models for the low and high number of grains, respectively. The abnormal grain growth was taken place in 304H steel during annealing, probably due to the presence of chromium carbide on the grain boundaries within the initial microstructure. The segregation resulted from dissolution of these carbides decreased the grain boundary mobility. A time-dependent equation was obtained for the mobility of grain boundary in this steel after the estimation of chromium content in the grain boundary areas. Moreover, a method was introduced for quantitative evaluation of the error during the measurement of grain size using the metallographic cross sections.

The rapid increase of space activities raises concerns about the safety, security and sustainability of future space operations. The EU has started working on an independent space surveillance and tracking (SST) capability, demonstrating leadership in space traffic management (STM). In this context, supported by the EU's H2020 research and innovation programme (GA 101004319), EUSTM ("Space Traffic Management for XXI Century Space Operations") is a comprehensive project that analyses the current STM support competences in Europe and defines the related needs for an autonomous European STM capability. It covers technology, governance, legal, regulatory, standardization, strategy and policy aspects and it will recommend STM guidelines and best practices. In addition, it

-Ce travail porte sur la simulation numérique de l'écoulement supersonique adapté dans les tuyères propulsives où le gaz d'essai supposé parfait est le gaz de combustion de l'Hydrogène. Il vise à déterminer les paramètres de l'écoulement Eulérien supersonique dans la tuyère convergente divergente. La méthode numérique utilisée pour la résolution de l'écoulement est basée sur une approche des volumes finis en coordonnées généralisées. L'intégration du système pour les équations de conservation d'Euler s'effectue sur un volume élémentaire quadrilatère. Dans cette étude, le traitement des flux convectifs est effectué en utilisant la méthode de Roe. Pour la discrétisation temporelle des équations, un schéma explicite de type Runge-Kutta du second ordre est utilisé. Abstract-This work focuses on the numerical simulation of supersonic flow in propelling nozzle. Test gas is the combustion gas of Hydrogen, it is assumed to be an ideal gas. The study aims to determine the parameters of the flow in the supersonic Euler converging-diverging nozzle. Discretisation and numerical solution of Euler equations are based on the theory of hyperbolic systems. The numerical method used for solving the flow is based on a finite volume approach in generalized coordinates with the use of structured meshes. Integration system for the conservation equations Euler is carried out on an elementary volume quadrilateral. In this study the treatment of convective flow is performed using the method of Roe. For the time discretisation of the equations, an explicit scheme of Runge-Kutta second order is used.

Hybrid rocket propulsion created a tremendous impact on rocketry after its inclusion in the rocket propulsion program. Simplicity, ease of handling, control of combustion, and ease of start and stop are some of the virtues of its rocket propulsion application. However, deficiencies in achieving the specific thrust and required energy for attaining that power have created concerns in the research domain about realizing such specific fuel consumption. In this paper, research has been demonstrated on various fuel compositions as one nature's product shellac and another one was with hydroxyl-terminated polybutadiene (HTPB) with one is one ratio. The laboratory-prepared fuel grain was studied in a static rocket motor using commercial oxygen gas as an oxidizer. The parameters like local regression rate, average regression rate, and average mass consumption rate with a selected mass flow rate of oxygen were evaluated and presented. Comparative results are demonstrated for virgin HTPB and HTPB with shellac in equal ratios with two different dimensions. The properties like heat of combustion were estimated using the bomb calorimetric method. All experimental data and simulated using theories are presented in the form of tables and figures for ready reference.

The results from calculating the apogee altitude for different 1u configuration payloads show that the maximum altitude expected is 21,000 ft. for a 1U payload and 11,000 ft. for a 6U payload. The motor chosen for the launch is the Pro 150 O5800 White Thunder with a total impulse of 30,601 Ns, average thrust of 5,802 N, and burn time of 5.27 seconds. OpenRocket software was used to simulate the flight profile and the loads expected during the launch. Previous Work

Commonly used motorcar fuels such as Petrol and Diesel contribute to a large amount of CO2 emission, directly linking it to global concern of climate change. Finding an alternative fuel or technology which can help reduce the Carbon-dioxide emission has therefore become a necessity to ensure sustainability of natural resources and fossil fuels for further future use. Before Hydrogen fuel came into existence, the EV's were a better alternative to reduce greenhouse gas emissions. Nonetheless EV's would seem straightforward from the front but they do come with a lot of, usually ignored, drawbacks involving Lithium-ion Battery usage, which itself produces a lot of carbon while its mining and extraction process, and can produce an amount of carbon as much as or more than the conventional gasoline and diesel vehicles do in combined. Electric Vehicles have a lot of benefits in terms of not using fossil fuel, but the fact that they do have many imperfections, all of which cannot be completely disregarded, makes them a not-so-good-choice.

NWNL has created a flexible deuterium-deuterium (DD) Fusion development paradigm to include Cooling, High Voltage, digital & analog telemetry, thermodynamic transport, shielding, mass flow control, and vacuum in a single system allowing for rapid iteration of fusor engine design and development. Each subsystem is assembled from off-the-shelf components, making maintenance and modification streamlined. This paper defines the advantages and tradeoffs in moving to off the shelf components during early research and development to save both time and money. This system meets the DOH radiation safety standards currently in place in Washington State.

Aviation has come a long way since its early days, and as technology has advanced, air vehicles have become
even more advanced and effective, especially on the battlefield. In addition, in-depth research and discoveries
in the field of aviation have a significant impact on the aviation industry, especially in the field of munitions.
In this context, aerial ammunition is an important component of aviation and is widely used in combat. Because
a malfunction that may occur during separation or premature detonation during separation from the aircraft
may cause fatal effects on the crew. That’s why the bomb must be safely separated from the air vehicles.
Modern bombs are therefore equipped with advanced safety features that guarantee a precise and safe
separation from the aircraft before explosion. In addition, the determination of the ballistic coefficient
according to the size of the desired effect reveals the importance of the choice of munition types. Otherwise, it
will not be possible to achieve the desired level of target destruction. This study aims to provide a
comprehensive understanding of the basic characteristics of aerial munitions, including the composition and
operation of generic aircraft bombs.

The paper summarizes the Space Technology Department at the Institute of Aviation goals for small spacecraft propulsion technology and link them to areas of propulsion research in Poland in general at the nearest future. A brief review of currently being investigated issue in the newly opened laboratory is presented as current and near term research, focused specifically on: concentrated solutions of H2O2 (stabilized and HTP - High Test Peroxide class) preparation, its long term storage possibilities, catalytic decomposition, “drop tests” for fuel ignition investigation, its use as “green” monopropellant or in hybrid rocket motors as the oxidant. By employing catalytic decomposition of HTP, auto-ignition of the solid fuel in a hybrid rocket is possible. When HTP is properly decomposed, the decomposition products alone release enough heat and the gases have suitable temperature (even over 700°C, depending on the concentration of the solution of peroxide) to provide efficient thrust. ...

There is evidence of past Near-Earth-Objects (NEOs) impacts on Earth and several studies indicating that even relatively small objects are capable of causing large local damage, either directly or in combination with other phenomena, e.g. tsunamis. This paper describes a space mission concept to demonstrate some of the key technologies to rendezvous with an asteroid and accurately measure its trajectory during and after a deflection maneuver. The mission, called SIROCO, makes use of the recently proposed ion beam shepherd (IBS) concept where a stream of accelerated plasma ions is directed against the surface of a small NEO resulting in a net transmitted deflection force. We show that by carefully selecting the target NEO a measurable deflection can be obtained in a few weeks of continuous thrust with a small spacecraft and state of the art electric propulsion hardware.

Comprehensive numerical analysis was conducted to elucidate the exhaust performance of rocket engine nozzles. The study focused on unravelling the intricate relationship between convergence and divergence angles and their impact on the exhaust performance parameters, including velocity coefficient (cv), angularity coefficient (Ca), and gross thrust coefficient (Cfg). In contrast to conventional studies that focus mainly on the divergent section, this research delved into both convergent and divergent aspects of nozzle geometry. For the convergent section, a range of angles from 20° to 45° was systematically examined. For the divergent section, a wide spectrum of angles was explored, ranging from small (10°-13°), medium (14°-19°) and large (20°-25°) divergent angles. Further, we venture beyond geometry, investigating the influence of nozzle pressure ratio (NPR) on these key metrics. Realisable 𝑘𝑘 -𝜀𝜀, enhanced wall traitement was used to simulate nozzle flow. The study identified the optimal convergent angle at 37.5°. The 15° diverging angle provides good overall performance, while the 23° angle strikes the ideal compromise: maximizing thrust and efficiency while minimizing weight and maintaining optimal performance.

NASA is providing key leadership in an international effort linking NASA and non-NASA resources to speed adoption of additive manufacturing (AM) to meet NASA's mission goals. Participants include industry, NASA's space partners, other government agencies, standards organizations and academia. Nondestructive Evaluation (NDE) is identified as a universal need for all aspects of additive manufacturing.

A skeletal chemical kinetic model for СН 4 /air combustion with 100 reactions and 24 chemical species was developed from the detailed mechanism, with 42 species and 298 reactions. The mechanism reduction was performed with the multi target reduction strategy realized in the in-house developed DLR RedMaster code. RedMaster is able to analyze the different chemical processes (ignition delay time and laminar flame speed) in the given time and height points. The obtained reduced model describes satisfactory experimental data for ignition delay and flame speed under conditions: p 5 = 1-50 bar, T 5 = 940K-210K, f = 0.5−2; p = 1-60 bar, T 0 = 300K, f = 0.6−1.4. Some problems related to reaction mechanism reduction are analyzed.

In the present work, SpaceX's Raptor engine combustion chamber has been investigated numerically to precisely understand the flow dynamics and variation of properties such as pressure, temperature and velocity distribution within the combustion chamber. The numerical results were compared to Raptor specifications, demonstrating a strong correspondence. This study shows that improving the correspondence between numerical results and Raptor technical specifications is possible by taking into account the influence of preburners on fuel and oxidizer inlet compositions. In this study, influence of equivalence ratio on engine performance has also been investigated. It has been shown that moving the equivalence ratio from 1.1 to 1.4 can improve engine performance by 5% but it also results in a 5% increase in pressure within the combustion chamber.

Resumen-Se presenta un modelo cero dimensional para estimar la masa ablacionada y el impulso en propulsores de plasma pulsantes ablativos de Teflón ® (PPPA), el cual es una modificación de modelos desarrollados con anterioridad. El modelo plantea el balance global de la energía, el cual considera la generación y transporte del pulso de corriente, la formación del arco, la disipación de energía por efecto Joule, la aceleración del plasma y la aceleración de los gases neutros. La ecuación resultante del balance de energía queda expresada en función de la masa ablacionada de Teflón ® y de los principales parámetros eléctricos y geométricos del propulsor. Una expresión explícita de la masa que se ablaciona se obtiene desde este balance. También, se lleva a cabo un estudio de sensibilidad del nuevo modelo respecto de variables tales como la resistividad y la emisividad del plasma en el arco. Se comparan los resultados numéricos con aquellos dados en la literatura especializada para distintos PPPA. Finalmente, se identifican las variables más relevantes respecto a la predicción de la masa ablacionada, el bit de impulso y el impulso específico.

Regolith volatiles analysis by SAM Orbital estimates of water in top 1-m of regolith via slow neutrons Water content of regolith vs. distance by DAN on Curiosity Rover 3 Kleinhenz, Julie E., and Aaron Paz. "An ISRU propellant production system for a fully fueled Mars Ascent Vehicle." 10th Symposium on Space Resource Utilization. 2017.

In recent years plenty of research is going on to produce effective manufacturing techniques which can produce complex unique designs with reduced cost and lead time for aerospace applications. In this aspect, Additive manufacturing is getting a wide range of popularity in the aerospace manufacturing domain due to its vast applications and advantages. In terms of geometric flexibility and processing time, additive manufacturing outperforms conventional manufacturing processes. Additive manufacturing has found its place in aerospace, defence, biomedical and automotive industries. However, new application areas like space technology are opening up. This method reduces the demand for industrial infrastructure and produces decentralized products. In this study, we have reviewed the manufacturing of rocket propulsion systems through additive manufacturing techniquesWire-arc additive manufacturing and powder bed fusion are the key areas of focus for us (WAAM). These techniques help produce fewer weight components with part consolidation, reducing tooling and assembly requirements. Combustion chambers and nozzles produced using this method meet the performance requirements, they also demonstrated significant cost and schedule savings for hardware delivery. The components in rocket engines operate in extreme and harsh environments. Thus, special materials and complex geometries are required to achieve high performance. Additive manufacturing techniques can enable this, which in turn enables designer freedom from geometric constraints commonly found using traditional manufacturing techniques. Many other components such as injector systems in rocket engines can be made using additive manufacturing within less time and into a single injector head without traditional welding and joining many components A review discusses manufacturing processes, materials used and properties. It also discusses various challenges faced by additive manufacturing.