Solid Rocket Propellants Research Articles

Surface Erosion Analysis for Thermal Insulation Materials of Graphite and Carbon–Carbon Composite

A rocket uses fuel and oxidizers to generate propulsion by combustion and ejection, and is used for space exploration aircrafts, weapons, and satellite launches. In particular, the nozzle generating thrust of solid-propellant rockets is exposed to a high-temperature and high-pressure environment with erosion occurring from the combustion gas. When erosion occurs on the nozzle throat of such a rocket, it has a great impact on the flight performance such as reaching distance and flight speed. Many studies have been conducted to characterize erosion based on the thermochemical erosion model, since it has become important to choose nozzle materials suitable for such environments having robustness against combustion gasses of high temperature and high pressure. However, there is a limit to fully analyze the erosion characteristics only by the thermochemical erosion model. In this paper, we thus consider the mechanical erosion model with the thermochemical model for better understanding of erosion characteristics and investigate the thermochemical and mechanical erosion characteristics of nozzle throat heat-resistant materials made of graphite and carbon–carbon composites; the main factors affecting erosion are discussed by comparing the results of the experimental and theoretical models.

Dynamics of a cylindrical shell with a collapsing elastic base under the action of a pressure wave

In the dynamic and quasi-static statements, the issue of non-stationary deformation and stability of the solid propellant rocket engine (SPRE) was approximately solved. It is modeled by a thin, smooth cylindrical shell, inside of which, on a part of its length, there is an elastic base corresponding to a gradually burning powder charge. A pressure wave is moving along the outer surface of the body, simulated by the running load. The deformed state of the shell is considered axisymmetric and is determined on the basis of the moment theory of the shells. For diverse variants of mounting the ends of the shell in a closed form, expressions were obtained for the critical velocity of the load. Examples were considered.

Nano-Scale Copper Oxide: Preparation, Characterization and its Effect on the Thermal Behavior of Ammonium Perchlorate

A new generation of high energy materials depends on the use of Nano-particle oxides. Nano-scale copper oxide (nano-CuO) has large surface area and surface energy which is suitable for its application in the field of energetic materials. This manuscript reports a method for the synthesis of nano-CuO by a liquid-state reaction method. The prepared nano-CuO was characterized by transmission electron microscopy (TEM) and X-ray diffraction (XRD) to check the particles size, purity and morphology of the crystals. The effect of Nano-CuO on the thermal behavior of AP was tested by differential scanning calorimeter (DSC). The results proved that the average particle sizes of the nano-cuo particles are in the range of 10-20 nm. The thermal degradation rate of AP was increased by 23% in the presence of 1% nano-CuO and the heat release was increased by 51%. It was concluded that nano-CuO could have obvious effect on the burning behavior, performance and combustion characteristics of the solid rocket propellants.

Fabrication of hybrid aluminum nanoparticles with organosilicon surface by solvent-free coating approach

With the development of aerospace industry, hybrid nano-Al particles with considerable energy release and stable surface immune to the air and moisture have attracted much attention when they are applied as fuel to solid rocket propellants. In this study, nano-Al particles (80~120 nm) were covered with an organic single layer, which was obtained by solvent-free gas-phase-coating approach in the presence of organosilanes. The characterization results clearly indicated that the coated aluminum nanoparticles are resistant to the air and moisture. Note here that no significant loss of the active aluminum contents was observed after being coated with the organic single layer, meaning that negligible power loss occurs during the combustion of such coated aluminum particles as a fuel. Combining with the advantage that the wettability of the coated surface could be adjustable by using chlorotrimethylsilane (CTMS) and dimethoxydimethylsilane (DMODMS) in a given sequence, such approach opens a new door to produce a potential fuel with protected surface and considerable energy release for solid rocket propellant. Moreover, the possible mechanism for coating the surface of aluminum nanoparticles with CTMS and DMODMS was proposed.

Exploring New Ways to Solve Old Problems; the Hunt for Alternative Stability Tests

AbstractThe 80 °C self‐heat test (formally known as the silvered vessel test) is one of the methods the UK uses to assess if nitrate ester based solid rocket propellants are sufficiently stable for service use. However, it uses a large mass of propellant – of the order of 70 g – and so there have always been safety concerns with regards to this technique. Further to this, it is not an internationally recognised test. There has been a desire to find a replacement for the 80 °C self‐heat test since at least the 1980s. One of the key advantages of the method is that it uses a significant mass of material and so is more representative of stabiliser depletion reactions in a solid rocket motor compared with smaller scale ageing tests. It is also considered a reliable method as it will accurately show if a nitrate ester based propellant is sufficiently stable for service use or not. This research, using three nitrate ester based propellants, has shown that the technique which is colloquially named “extreme heat flow calorimetry” offers an alternative to the 80 °C self‐heat test. Further, this method appears to be a new way of assessing the stability of nitrate ester based propellants and it has also been shown to be able to detect changes upon ageing for a particular propellant.

Analysis of Thermal Decomposition of Solid Rocket Propellants

The paper presents results of thermal decomposition analysis of selected solid rocket propellants. Homogeneous propellant PAC and heterogeneous propellant H2 were subjected to simultaneous thermal analysis with the use of NETZSCH STA 2500 Regulus device with five heating rates of 2.5, 5, 7.5, 10 and 15 K/min. The method combines TG, DTG and DTA analytical techniques in a single measurement. The aim of the conducted experiments was to study thermal decomposition of these energetic materials as well as to determine activation energy of the decomposition process and the preconditioning factor from the TG curves. The tested materials properties and chemical composition along with a brief description of the experimental procedure are described. The inverse procedure of calculating the activation energy, based on the Ozawa-Flynn-Wall model is described. Finally, the results of thermal decomposition of two tested solid rocket propellants are presented along with maximum decomposition rates and percentage of mass loss.

Electric Solid Propellant Ablation in an Arc Discharge

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled, and extinguished) through the application and removal of an electric current. Electric solid propellants are also being considered for pulsed arc ablation electric thrusters, such as the pulsed plasma thruster. The focus of this work is the electrical and ablation characteristics of electric solid propellant within an arc discharge. Arc discharges of 5–20 J per pulse were created within a cylindrical cavity, and results for the electric solid propellant are compared with polytetrafluoroethylene (PTFE), which is a traditional propellant in ablative pulsed plasma thrusters. The data indicate that the electric solid propellant has higher specific ablation per pulse () relative to PTFE (), which quantitatively agrees with an ablation energy balance model. For both propellants, the equivalent circuit resistance and inductance of the plasma arc are 50 mΩ and 125 nH, respectively. Analyses are presented indicating that the physics of propellant ablation is similar for both propellants with the differences in the observed specific ablation owing to differences in the thermodynamic properties between propellants.

EPA proposes limit on perchlorate in water

The Trump administration is proposing several options for limiting perchlorate in drinking water, including leaving the chemical unregulated. Perchlorate interferes with thyroid metabolism, which can damage the nervous system of infants and young children. More than 16 million US residents have perchlorate in their drinking water. The substance is used in solid rocket propellant, fireworks, vehicle-airbag inflators, and flares. The US Environmental Protection Agencyproposed on May 23 to set a maximum contaminant level of 56 parts per billion of perchlorate in drinking water. In contrast, California set its maximum contaminant level for perchlorate at 6 ppb in 2007. In its proposal, the EPA is seeking feedback from the public on a limit of 56 ppb plus three other regulatory options: a limit of 18 ppb, a limit of 90 ppb, or simply not regulating perchlorate in drinking water. Eric Olsen of the Natural Resources Defense Counsel, an environmental group that

The excellent catalytic activity for thermal decomposition of ammonium perchlorate using porous CuCo2O4 synthesized by template-free solution combustion method

Spinel CuCo2O4 with 3D macro-/mesoporous structure has been successfully synthesized via template-free solution combustion method. X-ray diffraction (XRD), scanning electron microscope (SEM), transmission electron microscope (TEM), and N2 absorption/desorption have been done to obtain compositional and morphological information as well as BET surface area of the as-synthesized sample. Catalytic activities of the porous CuCo2O4 towards the thermal decomposition of ammonium perchlorate (AP) were investigated with differential scanning calorimetry (DSC) and thermogravimetry (TG) techniques. The results show that the addition of spinel CuCo2O4 with 3D macro-/nanoporousstructure to AP remarkably decreases the high temperature decomposition (HTD) from 470.92 °C for pure AP to 308.74 °C for AP containing CuCo2O4 additive with a mass ratio of 2%. The as-synthesized 3D macro-/nanoporous spinel CuCo2O4 is the promising type of catalytic materials for AP-based composite solid rocket propellants. The simple synthesis method is conducive to large-scale industrial production for application in AP thermal decomposition.

Анализ существующих воспламенительных устройств и разработка перспективной конструкции воспламенителя стартового ускорителя современных ракетно-прямоточных двигателей на твердом ракетном топливе

The article briefly reviews the existing designs and technical solutions for solid propellant igniters in rocket engines. The technical and design solutions in the development of solid-propellant rocket engine igniters are analyzed. The results of the development of a promising igniter design with gunpowder filler axially located in the internal channel of a propellant grain of a launching free-flowing booster of a rocket-ramjet engine are presented. Ensuring the required level of structure reliability and durability in the launch mode of the solid propellant rocket-ramjet engine, the igniter will improve the engine traction characteristics due to the combustion of the igniter aluminum shell as an additional energy source and using its combustion products in the stream of solid propellant combustion products.

Solid propellant-fuelled microrocket for space application

ABSTRACT The integration of propellant systems, namely nozzles, igniters and propellant chambers in microscale models, enables the realisation of a new class of devices called microrockets or microthrusters. With the accomplishment of microrocket for the control of the attitude or/and the station-keeping of very small satellite, course correction of satellites, clusters of the microrockets with sensors in deep space explorations and orbit adjustments can be explored. This paper details the design, fabrication, and testing of arrays of solid propellant rockets assembled in a polymeric substrate with electrical ignition system. The thrust generated from Type 1 microrocket is in the range of 1 mN to 2.4 mN and of Type 2 microrocket is 1.715 mN to 2.572 mN.

Numerical Research of Intra-Chamber Transient Processes Dynamics during Starting Operation of the Solid Propellant Rocket Engine with Special Arrangement

A conjugate problem statement of the computational experiment performing level is formulated. A complex of computer application programs was developed and tested for its implementation. The results of numerical study are given.

Combustion of HTPB Based Solid Fuels Containing Metals and Metal Hydrides with Nitrous Oxide

AbstractAn experimental investigation of the combustion characteristics of HTPB‐based solid fuels containing aluminum and aluminum hydride has been conducted. Aluminum is commonly used as a fuel supplement in solid rocket propellants because of its ability to increase the specific impulse of a propellant. Similarly, aluminum hydride, or alane, is another attractive fuel supplement because it can significantly (∼7–8 %) increase specific impulse even beyond that of aluminum. In this study, the regression rates of a baseline fuel (pure HTPB) were compared with that of fuels containing either aluminum or alane as additives using a counterflow combustion experiment with nitrous oxide (N2O) as the oxidizer. The additives were investigated in concentrations of 10, 20, and 40 % by weight of the fuels. The addition of aluminum at all loadings resullted in a decrease in the regression rate relative to the baseline, whereas the addition of alane resulted in at worst similar regression rates and at best (highest loading) approximately a 20 % increase in regression rate. The decomposition behavior of selected fuels was analyzed using traditional thermal analysis techniques; thermal gravimetric analysis and differential scanning calorimetry

Improving the Mechanical Properties of Glycidyl Azide Polymeric Matrix Used in Composite Solid Rocket Propellant by Adding Advanced Cross‐Linker

Glycidyl azide polymer (GAP) based binders have poor mechanical characteristics in comparison with hydroxyl terminated polybutadiene (HTPB) binders. In this study, advanced cross‐linker was used to improve the mechanical properties of GAP binder. GAP was prepared and characterized in comparison with HTPB prepolymer. Density, characteristics groups, nitrogen content, humidity, viscosity, and milligram equivalent of (OH) per binders were determined. A cross‐linker consists of trimethylol propane (TMP) and curing catalyst, dibutyltin dilaurate (DBTDL), was used as an additive to GAP polymeric matrix to enhance its functionality. Polymeric matrices based on GAP and HTPB were prepared with different curing ratio (NCO/OH) ranging from 0.7 to 1.5. Different weight percentages of cross‐linker were added to study its effect on the mechanical properties of GAP matrix. Five samples based on HTPB polymer and twenty samples based on GAP polymer were prepared. A LLOYD testing machine was used to determine the stress‐strain curves of all the studied samples. It was concluded that the cross‐linker used has significant influence on the characteristics of GAP polymeric matrix. Also the addition of 5 wt % of cross‐linker to GAP matrix at curing ratio = 1 produced optimum mechanical characteristics very close to that of HTPB matrix used in composite solid rocket propellants (CSRP). The optimum GAP polymeric matrix is candidate to replace the traditional HTPB binder in advanced CSRP.

THERMO-MECHANICAL ANALYSIS OF HETEROGENEOUS SOLID ROCKET PROPELLANT

The article presents results of thermo-mechanical analysis of heterogeneous solid rocket propellant H2, with special attention devoted to determining the glass transition temperature and softening temperature. Mechanical properties such as storage modulus (E’), loss modulus (E’’) and tan(δ) were measured using NETZSCH DMA 242C analyzer within temperature range from -120oC to +80oC at 2 K/min of heating rate and frequency of applied force f = 1 Hz. Relative thermal expansion as well as the Coefficient of Linear Thermal Expansion (CLTE) of H2 sample were determined using NETZSCH DIL 402C dilatometer within temperature range from 30oC to 80oC at 1 K/min of heating/cooling rate. The thermophysical properties including thermal conductivity, thermal diffusivity, and specific heat were determined within temperature range from -20oC to +80oC using KD2 Pro apparatus.

Initial Studies on Development of High-Performance Nano-structured Fe2O3 Catalysts for Solid Rocket Propellants

Space launch vehicles and strategic military vehicles commonly employ composite solid rocket propellants containing Fe2O3 as ballistic modifier. Nanoscale catalysts, including nanoscale Fe2O3, have been reported to exhibit superior activity in the thermal decomposition and combustion of composite rocket propellants. However, scalable methods to prepare such nano-structured catalysts with high performance as ballistic modifiers and systematic studies relating the synthesis parameters to the catalyst characteristics and consequently to the thermal and combustion properties of the composite propellant are scarce. In this paper, we report a novel and facile route to prepare nano-structured Fe2O3 with enhanced catalytic activity in the ballistic modification of ammonium perchlorate (AP)-based composite solid rocket propellant. A submerged spray precipitation method using air-assisted liquid-centered coaxial atomization has been developed to prepare these nano-structured Fe2O3 catalysts. The prepared Fe2O3 catalysts possess higher surface area and exhibit superior activity in the thermal sensitization of AP, leading to an 88% increase in the burning rate of AP-based composite solid rocket propellants, than the Fe2O3 catalyst prepared via traditional precipitation method. Merits of the developed preparatory route, influence of the atomization process on the nano-structure morphology and subsequent benefits of these nano-structured catalysts on the ballistic properties of AP-based propellants are demonstrated and discussed in this paper.

Study for Analysis of Effect of Machining Parameter & To Predict the Behaviour of Propellant Grain during Machining Operation

Solid Propellant Grain machining is one of the critical and hazardous operations in Solid Propellant Rocket Motor (SRM) processing. HTPB based Propellant grain surfaces are casted as per the feasible mandrel shape, but mostly the final grain surfaces are achieved by propellant machining operation, that generates various shapes and contours over casted Solid propellant grain. Some of these machined surfaces are directly exposed to the ignition and other acts as interface for further processing. These machined surfaces over propellant grain are generated by using CNC controlled Vertical Turn Mill Centre. Machining operation depends on three major machining parameters viz. speed, feed and depth of cut. Being a visco-elastic material machining of solid propellant surface is carried out with Special purpose Hollow Contouring Cutter with HSS Conical Insert. Solid Propellant in SRM is the prime mover to propel the payload to a particular range as per the mission requirement. Machining of extra propellant may lead to deviation from the target to be achieved. Due to visco-elastic behaviour of the propellant, during machining surface are compressed and after machining the machined surfaces are partially return back resultant the machined depths are varying with consecutive depth of cuts. This paper focuses on study of effect of depth of cuts on propellant grain surface keeping cutting speed and rotary table feed within a range. Study also analyzed to arrive at a conclusion to predict the machined depth over propellant grain, resulting in elimination of check cuts during propellant machining operation.

Оптимизация конструкции импульсных твердотопливных ракет- ных двигателей с металлическими корпусами путем объединения их в моноблочный отсек, изготовленный из высокопрочного пластика

Оптимизация конструкции импульсных твердотопливных ракет- ных двигателей с металлическими корпусами путем объединения их в моноблочный отсек, изготовленный из высокопрочного пластика

Effect of Coated Ammonium Dinitramide on the Properties of Nitrate-ester Plasticized Polyether Solid Rocket Propellants

Effect of Coated Ammonium Dinitramide on the Properties of Nitrate-ester Plasticized Polyether Solid Rocket Propellants

Cost estimating of commercial smallsat launch vehicles

Commercial launch service providers’ low-priced offerings have been a hotly debated topic. However, the strategies with which these firms reduce costs have seen little incorporation into the hardware Cost Estimating Methods (CEMs) and tools prevalent in the aerospace industry. This research changes this, by providing adaptations to agency-focused CEMs that befit a new commercial paradigm, with an emphasis on smallsat launch vehicles.A parametric model for estimating costs in an early phase of development was synthesized, with which it is possible to approximate the full life-cycle costs of small commercial liquid and solid propellant rockets, as well as their cost-based price per flight. Key elements included were reductions in cost achieved by commercial launch operators, by modeling reduced subcontractor management effort and profit retention experienced at lower subcontracting rates.Prices per flight of small commercial launch vehicles were approximated by combining a parametric cost estimating methodology used frequently in the context of space agencies such as ESA and NASA, called the T1 Equivalents method, with another parametric three-part estimate developed by Koelle for development, manufacture and operations phase costs. The first two phases were estimated through the T1 method, while the operations costs were modeled with TRANSCOST.Along with the newly developed methodologies, novel insights such as required launch rates have shone a light on small commercial launch systems’ cost feasibility in the age of public-private spaceflight partnerships.The model developed was able to approximate costs of development, manufacture and price per flight of three commercial rockets to within 20% of actual reported costs or prices. However, it is recommended the model is refined as more reference cost data, especially on a subsystem level, as well as pricing for these smaller rockets becomes available in the coming years.