Rocket Propulsion Applications Research Articles

Test Activities on Hybrid Rocket Engines: Combustion Analyses and Green Storable Oxidizers—A Short Review

Hybrid rocket engines (HREs) offer a low-cost, reliable, and environmentally friendly solution for both launch and in-space applications. Hybrid propellants have been identified as green thanks to their use of non-toxic, non-carcinogenic oxidizers. Of particular relevance are storable oxidizers, namely high-concentration (≥90 wt.%) hydrogen peroxide (HP, H2O2) and nitrous oxide (N2O). This work provides a survey of experimental activities based on H2O2 and N2O for hybrid rocket propulsion applications. Open literature data are completed with original thermochemical calculations to support the discussion.

HYBRID ROCKET PROPULSION DEVELOPMENT AND APPLICATION.

Satellite technology is advancing. Satellite propulsion requires several desired rocket properties. This is necessary for 1 kg of payload. Launch vehicles no longer need rocket engines. Application includes satellite movement, orbit transfer, and probe and lander propulsion. Space travel is inventive and exciting. Research, production, and use costs must be decreased to make space transportation systems generally available. Security shouldn't be affected. Rocket propellants must be high-performing, non-toxic, and safe. Rocket engines have other parts. Restarting and throttling are crucial. These are inaccessible to solid rocket motors. Developing a liquid rocket engine is hard and expensive, but it can be restarted and throttled. It's being studied for use in hybrid rocket propulsion and other space applications. Space tourist vehicles, lunar and planetary landers, suborbital launch vehicles, satellite maneuvering systems, etc. are potential applications for microsatellites (including orbit transfer). More of these applications are pending approval. Hybrid propulsion is straightforward, safe, and permits regenerative braking, throttling, and restarting. Hybrid propellants are mostly non-toxic and storable. Separate oxidizer and fuel storage enhance safety. This study examines the history, development, current applications, and future of hybrid rocket propulsion. We'll discuss popular fuels and oxidizers. Explains hybrid rocket motor's low regression rate. The article includes the author research.

Measurements of Laminar Burning Velocities of Alternative Fuels: Application in Transport and Aerospace

The laminar burning velocity belongs to the fundamental combustion properties of fuels being a measure for their heat release, flame length, as well as reactivity and combustion stability, and thus, may impact the design of burners and combustion chambers. Also, these experimental data are needed for the validation and optimization within the construction and development of detailed chemical kinetic reaction mechanisms. Within this study, an overview of the different applications of fuel characterization regarding the specific area of interest (road transport, aviation, and aerospace) will be given. Depending on the application, effects of different molecular characteristics on the laminar burning velocity are evaluated: the presence of oxygen atoms and the grade of branching in a specific fuel molecule as well as the difference in the type of a chemical bond, here, single and double bonds. Examples of alternative fuels being discussed in the present study are: (I) oxymethylene ether (OMEn) in the field of road transport; (II) a paraffinic Alcohol-to-Jet fuel as sustainable aviation fuel; and (III) mixtures of ethane or ethene with nitrous oxide as green propellants for rocket propulsion applications.

Hydrogen peroxide – A promising oxidizer for rocket propulsion and its application in solid rocket propellants

This paper discusses the potential use of hydrogen peroxide as oxidizer for solid rocket propulsion. While hydrogen peroxide is a liquid in normal conditions, it may be used in solid rocket motor grains. Use of hydrogen peroxide of HTP class (High Test Peroxide) is proposed. It has been known for many decades but its utilization has been historically limited to liquid propellants. Until recently is application even in liquid state in rocket propulsion for space and defence has been limited due to storability and safety issues. With the availability of new grades of HTP with higher purity and concentrations, enhanced performance, safety and storability are possible. This results in HTP being considered for a wide range of rocket propulsion systems: as oxidizer in bipropellant and hybrid propulsion systems, as well a monopropellant. To ensure proper analysis of the potential of using HTP in next-generation solid rocket propellants, this paper reviews existing rocket propulsion applications of HTP. Modern use requires high performance and common current composite propellant compositions utilize ammonium perchlorate (AP) as oxidizer, what has several disadvantages, which are discussed. Alternative oxidizing compounds include ammonium dinitiramide (ADN), ammonium nitrate (AN), hydrazinium nitroformate (HNF), hexanitrohexaazaisowurtzitane (HNIW) and several other secondary explosives. Key properties of solid propellant oxidizers are listed and discussed. The need for further alternatives, despite numerous recent advances in solid rocket oxidizer technology is justified. The global push forward high performance green propulsion is one of the main motivations behind considering HTP for solid rocket propulsion. Theoretical performance of solid rocket motors using solid grains containing a high mass fraction of HTP is presented. This includes performance considering several fuels and additives and different oxidizer loadings. Up to date concepts of using hydrogen peroxide in solid propellants are reviewed. Solid cryogenic propellants using hydrogen peroxide are mentioned, but focus is given to solid propellants which could ensure flexible operations, thus use in state-of-the-art solid rocket motors. Challenges of HTP application as an oxidizer for solid propulsion are listed. This includes a discussion concerning its reactivity, thus limited compatibility with organic materials. Recommendations for further experimental work are proposed. Potential technology applications are listed with explanations why these particular niches may benefit from utilization of the new solid propellant technology.

Properties of Kerosene-Aluminium Nanofluid used to Estimate the Overall Heat Transfer Rates during Regenerative/Film Cooling of Thrust Chambers

Large heat transfer rates are always desired for rocket propulsion applications as high heat loads are associated at the nozzle exit. Different strategies have been employed in order to have high heat transfer coefficients including use of liquid nitrogen, spray cooling etc. ISRO has planned to use aluminium based nano-particles with kerosene in order to cool launching vehicles including GSLV Mk III as it is the heaviest rocket that can carry large payloads. Recently, ISRO has announced to install its own International Space Station (ISS) in future and in such applications larger payloads are to be carried by the rocket. In this work, an analytical study on the thermodynamic properties of the aluminium nano-particles based kerosene nanofluid has been done and an attempt has also been made to develop a temperature and pressure dependent correlation that can be used in computational analysis of thrust chambers while film/regenerative cooling.

The Aging Studies of the Low HCl Content Composite Rocket Propellant

AbstractThis paper presents technological research on the storage time of a composite rocket propellant with a low HCl content in combustion products, based on sodium nitrate. For research were prepared and casted in laboratory scale, a composite rocket propellant based on sodium nitrate and HTPB system, and as the reference sample, the classic composite rocket propellants based on the HTPB/AP system. Both propellants have been subjected to laboratory‐scale aging in accordance with the STANAG accelerated aging protocol, by 5 and 10 years. For each propellant, aged and not aged, parameters such as calorific value, hardness, sensitivity to the mechanical stimuli and decomposition temperature were determined. For the ballistic evaluation of the investigated propellants, the Laboratory Rocket Motor technique with the indirect burning rate determination method, was used. Based on the obtained data, was found that the composite rocket propellants based on sodium nitrate are technological and can be used in the rocket propulsion applications.

Tensile Behavior of Electron Beam-Welded and Post-Weld Vacuum-Annealed Nb-10% Hf-1% Ti Refractory Alloy Weldments

Nb-10% Hf-1% Ti refractory alloy is a high performance material extensively used for high temperature applications. Electron beam welding is one of the most widely used techniques to join refractory and reactive alloys. Bigger sizes of nozzles for rocket propulsion applications can be either made through deep drawing and flow turning route or by roll bending and welding route both using sheets/plates as input material for fabrication. The latter is a more economical option for mass production of the hardware using such exotic and expensive alloys. In view of this, both as-welded (AW) coupon and weld plus post-weld vacuum-annealed (AW + VA) coupon have been prepared to study their mechanical behavior. It has been observed that tensile strength and ductility have not been reduced in both these conditions vis-a-vis the base metal, confirming weld efficiency of the alloy to be ~ 100%. Microhardness is found to be 150-160 VHN in the base metal and 200-225 VHN in the weld fusion zone in AW condition, which became uniform (145-155 VHN) throughout the weldment in AW + VA condition. Microstructure of the weldment in AW condition is found to be consisting of grains solidified by epitaxial mode from base metal toward the weld centre. In AW + VA condition, improvement in tensile elongation is observed, which is found to be due to the presence of homogenized grains/more uniform microstructure near the heat-affected zone as compared to the steep gradient in grain size in different zones in the weld in AW condition.

Nitrous Oxide and Oxygen Mixtures (Nytrox) as Oxidizers for Rocket Propulsion Applications

Most of the oxidizers available to be used in chemical propulsion applications are highly hazardous materials that are either toxic or explosive in nature. Among the short list of oxidizing agents, liquid oxygen () and nitrous oxide () stand out as the most practical propellants due to their wide availability, broad base of use, cost effectiveness, and relatively benign nature. A new class of oxidizers (Nytrox) that are composed of equilibrium or nonequilibrium mixtures of nitrous oxide and oxygen are formulated in order to maximize the benefits of the pure components while retaining their practical advantages. Note that in the mixture serves as the pressurizing agent, whereas is the densifying component. The primary advantages of this new system over the pure oxidizers can be listed as 1) self-pressurization capability, 2) high density and density impulse, 3) noncryogenic operational temperatures, 4) higher performance compared with , 4) improved safety, 5) efficient gas-phase combustion, and 6) easier development of stable and efficient motors compared with liquid oxygen due to the exothermic decomposition of the molecule. Unlike the pure oxidizers, the mixture allows for two independent control variables (temperature and pressure), which can be fine tuned to optimize the system for a particular application.

Assessment of Catalysts for Hydrogen-Peroxide-Based Thrusters in a Flow Reactor

Hydrogen peroxide is a candidate propellant for rocket-propulsion applications with the potential to replace highly toxic propellants currently used. Decomposition of hydrogen peroxide yields a high-temperature oxygen-steam mixture, which can be used as monopropellant or as oxidizer in a bipropellant configuration. This work examines different types of cellular ceramic-based catalysts for hydrogen-peroxide decomposition at miniature scale of nominal mass flows of . An exhaustive investigation of different catalysts in a flow reactor configuration similar to a propulsion application is conducted. The test matrix includes honeycomb monoliths with different channel geometries, densities, lengths, different carrier materials, and wash-coating procedures, as well as different types of catalysts such as pellets and foams. Thirty nine catalyst configurations with a total of 121 catalysts have been experimentally investigated based on their transient and stationary performance at design mass-flow levels of .

IGNITION AND COMBUSTION OF NICKEL COATED AND UNCOATED ALUMINUM PARTICLES IN HOT POST-FLAME ENVIRONMENT

It is well known that aluminum particles require a high temperature to achieve ignition, often in excess of 1900 K. Lowering the ignition temperature of Al particles can increase the efficiency of solid propellant-based propulsion systems and thermobaric explosives. Current methods to reduce the ignition temperature include both reducing the particle size to the nano-scale and/or applying coatings that aid ignition. In this work, an experimental study of ignition and combustion of isolated 5 wt% Ni-coated and uncoated aluminum particles was conducted. Two particle sizes (nominally 32 μm and 9 μm in diameter) were examined. The ignition and combustion properties of these aluminum particles were observed in the postflame zone of a multi-diffusion flat-flame burner at atmospheric pressure. The results showed that the applied nickel coating decreased the ignition temperature of the Al particles by 750 K on average for 32 m particles and by 300 K for smaller 9 m particles. Measurements of particle burning times indicated that the application of the nickel coating does not affect the overall burning time of the particles. Thus, the application of nickel coatings on Al particle significantly decreases the ignition temperature while not affecting the overall combustion behavior. I. Introduction luminum is an energetic fuel additive for solid propellants and solid fuels in different propulsion systems, and also has effectiveness for particle damping in suppressing combustion instability in rocket motors. This has led aluminum particles to become widely employed in rocket propulsion applications. The importance of aluminum particle combustion in propulsion systems has led to a large number of combustion studies. Furthermore, aluminum particles have also been used in thermobaric explosive (TBX) applications where Al particles need to be ignited rapidly at low temperatures after the detonation of the high explosive charge. Beckstead 1 performed a comprehensive summary of aluminum particle combustion from a wide range of sources. He found a common ignition temperature of about 2,300 K, which is near the melting temperature of the aluminum oxide shell (2,327 K). Additionally, Beckstead's summary found that the ignition delay of Al particles is dependent on the thickness of the oxide shell. Burning time data of the Al particles was also compiled to determine if the D 2 law was applicable. It was found that the D n law was more applicable for larger particle sizes due to the oxide layer formation with the exponent being 1.5 or 1.8, depending upon the coefficient used in the correlation. Additionally, studies that compared the relative effectiveness of different oxidizer species including oxygen, water vapor, and carbon dioxide were compiled in Beckstead's work 1 . It was concluded that O2 was the most effective oxidizer, H2O was between 50 to 60 percent as effective, and CO2 was about 22 percent as effective. Beckstead's correlation is widely accepted for correlating aluminum combustion behavior for particle sizes above 20 m. Without consideration of the effective oxidizer mole fraction, pressure, and ambient gas temperature, Beckstead’s preliminary correlation indicated a D 1.99 law. When the diameter exponent is close to 2.0, the combustion can be regarded as a diffusion-limited process for relatively large particles beyond 20 m. Bazyn, et al. 2 conducted a study on the combustion of 10 μm-sized aluminum particles in the reflected region of a shock tube and found pressure exponents of −0.9, 0.2, and 0.3, for O2, H2O, and CO2 respectively, (i.e., tb  pi n where pi is the partial pressure of the i th

Large-Eddy Simulation of oxygen/methane flames under transcritical conditions

Large-Eddy Simulation of transcritical combustion is considered in this article. This analysis is motivated by rocket propulsion applications where cryogenic oxygen is injected at low temperature, below the critical value, in a high pressure environment where the pressure is above critical. Problems encountered in this range are related to the description of real-gas effects and in particular to the representation of the very large density changes which occur when a transcritical fluid evolves from low to high temperatures. The simulation also requires a suitable spatially filtered combustion model which can be used to determine the volumetric reaction rates in a situation where reaction essentially takes place in a non-premixed mode. A model is proposed and validated in the “H3” atmospheric flame configuration. This model is then used in a transcritical version of the AVBP LES flow solver. Results of calculations carried out for transcritical oxygen/supercritical methane injection are in good agreement with recent experimental data and provide additional information on the flame structure and flow configuration.

High Performance HTPB-Based Energetic Nanomaterial with CuO Nanoparticles

This work describes the first example to demonstrate the enhancement of performances of composite highly energetic materials by mean of employing standard CuO nano-powder as burning rate catalyst in comparison to micro-fillers. The solid composite propellants with CuO microparticles are less stable due to oversensitivity to pressure variations, but the nano-structured composite propellant yields high stable burning rates over a broad pressure range. In addition, the incorporation of CuO nanoparticles in the formulations of these energetic materials also improves their combustion and thermal properties, according to the characterization obtained by differential scanning calorimetry (DSC) and thermogravimetry analysis (TGA). These results indicate the excellent benefits found in using these nanoparticles as additive for solid rocket propulsion applications.

Modeling liquid hydrogen cavitating flow with the full cavitation model

Cavitation is the formation of vapor bubbles within a liquid where flow dynamics cause the local static pressure to drop below the vapor pressure. This paper strives towards developing an effective computational strategy to simulate liquid hydrogen cavitation relevant to liquid rocket propulsion applications. The aims are realized by performing a steady state computational fluid dynamic (CFD) study of liquid hydrogen flow over a 2D hydrofoil and an axisymmetric ogive in Hord's reports with a so-called full cavitation model. The thermodynamic effect was demonstrated with the assumption of thermal equilibrium between the gas phase and liquid phase. Temperature-dependent fluid thermodynamic properties were specified along the saturation line from the “Gaspak 3.2” databank. Justifiable agreement between the computed surface pressure, temperature and experimental data of Hord was obtained. Specifically, a global sensitivity analysis is performed to examine the sensitivity of the turbulent computations to the wall grid resolution, wall treatments and changes in model parameters. A proper near-wall model and grid resolution were suggested. The full cavitation model with default model parameters provided solutions with comparable accuracy to sheet cavitation in liquid hydrogen for the two geometries.

A Computer Model of Flamelet Distribution on the Burning Surface of a Composite Solid Propellant

The paper attempts to obtain an approximate distribution of oxidizer/fuel (O/F) flamelets on the burning surface of a composite solid propellant, mainly of the ammonium perchlorate (AP)-hydrocarbon (HC) binder variety commonly used in rocket propulsion applications today. A computer model is developed to simulate the random packing of oxidizer particles of different size distributions in the propellant. The focus of the present work is on the distinction between pockets of fine AP particles amidst the binder that may bum in a premixed O/F flame and the sufficiently large AP particles that bum with the surrounding vapor flow in an attached diffusion flame. A transition occurs between premixed burning and attached diffusion flamelet-burning over each oxidizer particle as the ambient pressure is varied. A simple criterion evolved earlier to delineate the two regimes of burning in the particle size-pressure domain has been employed to determine the flamelet type that prevails over each exposed particle of the model propellant at a given pressure. Delaunay triangulation is adopted to deal with regions of fine AP-binder corresponding to premixed burning. The results are compared with previously reported experimental observations of burning rate for specific propellant formulations that exhibited distinct jumps in their burning rate with variation in pressure. It was argued that the burning rate jumps corresponded to a concerted shift in the burning regime of a majority of fine particles. The present simulation supports this argument.

Effet du retrait sur la stabilisation des flammes cryotechniques

Mechanisms involved in cryogenic flame stabilization are important in liquid rocket propulsion applications. These mechanisms are studied experimentally in this article by considering a single jet-flame formed by a coaxial injector fed by liquid oxygen (LOX) and gaseous hydrogen (GH 2). The flame is analyzed by imaging the light emitted by hydroxyl (OH) radicals. To characterize the reaction zone structure one takes the average of the light emission signal. The result of this operation is then treated by numerical tomography (based on the Abel transform) to get the local volumetric light intensity distribution. This method is used to study the effect of the LOX tube recess on the structure of the flame in the vicinity of the injection plane. It is shown that the recess augments the flame expansion angle and the size of the volume where the reaction takes place.

Platonization in double-base rocket propellants

HE inclusion of small quantities of various lead compounds in the double-base propellant system results in increased burning rate at low pressure, followed by a plateau burning region, where the burning rate remains almost independent of variation in pressure, and the postplateau region, in which the burning rate/pressure relationship is similar to that of an unleaded propellant (Fig. 1). This phenomenon of independence of burning rate to pressure, generally known as platonizatio n, implies a low value of pressure index n in the plateau region. A low value of n permits the design of lighter rocket motors because of lower safety factors. The magnitude of n is one of the important factors in determining the suitability of a propellant for rocket propulsion applications. The importance of low-pressure ballistics was realized as early as in 1939 with the widespread use of rockets. Earlier studies centered around low-energy nitrocellulose propellants with a view toward minimizing nozzle erosion, burning rate, and pressure index. However, the real impetus to the platonization of double-base propellants came from the observation that the use of lead stearate as a lubricant in the extrusion of large rocket charges actually had modified their burning characteristics. This paper reviews the literature on platonization of double-base propellant with different inorganic and organic compounds of lead and other metals, and presents various theories to explain the mechanism of the plateau effect.