Synthesis and characterization of epoxy resins based on N,N'-aliphatic dicarboxyl bis(hydrazones)

A new class of epoxy resins having N-N bonds in their structure has been synthesized by reacting N,N'-aliphatic dicarboxyl bis(hydrazones) (the aldehyde/ketone derivatives of malonic, adipic, and sebacic dihydrazides) with epichlorohydrin. The reactivity of the[GRAPHICS] protons as a function of the substituent group and the number of methylene spacer groups present in the hydrazone has been examined. The resins obtained have been characterized by elemental and epoxy equivalent analyses and IR and NMR spectra. All these resins are found to have adequate viscosity and cure easily with amine curatives at elevated temperatures. Relevant properties for their use as binders in propellant formulations, such as thermal stability, heat of combustion, density, temperature dependence of viscosity, and mechanical strength of the composites, have been evaluated. (C) 1997 John Wiley & Sons, Inc.

Universal behaviour in erosive burning of solid propellants

This paper considers the extensive data and correlations on the erosive burning of solid propellants. A relatively simple nondimensional relationship between the ratio of the actual to nonerosive burn rate (eta) and a quantity g, which is the product of g(0)-the ratio of free stream mass flux to the mass flux from the surface for nonerosive condition-and Re-0(m), where Re-0 is the Reynolds number based on the nonerosive mass flux of the propellant and port diameter, is shown to correlate most data within the accuracies of the experiments with m = -0.125. This shows the above relationship to account for the effects of pressure, aluminum, even up to a proportion of 17%, burn rate catalysts, and motor size. It is concluded that the suggested correlation between eta and g may be adopted universally for most practical propellants. (C) 1997 by The Combustion Institute.

Modeling re-ignition and chuffing in solid rocket motors

A nonlinear model is developed to numerically simulate dynamic combustion inside a solid rocket motor chamber. Using this model, the phenomena of re-ignition and chuffing are investigated under low-L* conditions. The model consists of two separate submodels (coupled to each other), one for unsteady burning of propellant and the other for unsteady conservation of mass and energy within the chamber. The latter yields instantaneous pressure and temperature within the chamber. The instantaneous burning rate is calculated using a one-dimensional, nonlinear, transient gas-phase model previously developed by the authors. The results presented in this paper show that the model predicts not only the critical L*, but also the various regimes of L*-instabihty. Specifically, the results exhibit (1) amplifying pressure oscillations leading to extinction, and (2) re-ignition after a dormant period following extinction. The re-ignition could be observed only when a radiation heat flux (from the combustion chamber to the propellant surface) was included. Certain high-frequency oscillations, possibly due to intrinsic instability, are observed when the pressure overshoots during re-ignition. At very low values of initial L*, successive cycles of extinction/reignition displaying typical characteristics of chuffing are predicted. Variations of the chuffing frequency and the thickness of propellant burned off during a chuff with L* are found to be qualitatively the same as that reported from experimental observations.

New Energetic Epoxy Binders

A new class of epoxy resins having N-N bonds in the backbone has been synthesized with a view to explore their properties as energetic binders. The N-epoxidation of bis-dicarbonylhydrazones of adipic, azelaic and sebacic dihydrazides results in the formation of viscous resins having epoxide end groups. The resins have been characterized by the elemental and end group analyses, IR and NMR spectra. Relevant properties for their use as binders in solid propellants, such as thermal stability, heat of combustion, burn rate and performance parameters of AP-based propellant systems, have been evaluated. A significant increase in the burn rate of AP-based propellants noticed, is perhaps related to the exothermicity of the binder decomposition and the reactivity of N-N bonds with perchloric acid formed during the combustion of AP.

Solution Combustion Synthesis of Nanosized Copper Chromite and Its Use as a Burn Rate Modifier in Solid Propellants

Nano sized copper chromite, which is used as a burn rate accelerator for solid propellants, was synthesized by the solution combustion process using citric acid and glycine as fuel. Pure spinel phase copper chromite (CuCr2O4) was synthesized, and the effect of different ratios of Cu-Cr ions in the initial reactant and various calcination temperatures on the final properties of the material were examined. The reaction time for the synthesis with glycine was lower compared to that with citric acid. The synthesized samples from both fuel cycles were characterized by X-ray diffraction (XRD), X-ray photoelectron spectroscopy (XPS), BET surface area analysis, and scanning electron microscope (SEM). Commercial copper chromite that is currently used in solid propellant formulation was also characterized by the same techniques. XRD analysis shows that the pure spinel phase compound is formed by calcination at 700 degrees C for glycine fuel cycle and between 750 and 800 degrees C for citric acid cycle. XPS results indicate the variation of the oxidation state of copper in the final compound with a change in the Cu-Cr mole ratio. SEM images confirm the formation of nano size spherical shape particles. The variation of BET surface area with calcination temperature was studied for the solution combusted catalyst. Burn rate evaluation of synthesized catalyst was carried out and compared with the commercial catalyst. The comparison between BET surface area and the burn rate depicts that surface area difference caused the variation in burn rate between samples. The reason behind the reduction in surface area and the required modifications in the process are also described.

Experiments on thermal response of low aspect ratio packed beds at high Reynolds numbers with varying inflow temperature

An experimental study for transient temperature response of low aspect ratio packed beds at high Reynolds numbers for a free stream with varying inlet temperature is presented. The packed bed is used as a compact heat exchanger along with a solid propellant gas-generator, to generate room temperature gases for use in applications such as control actuation and air bottle pressurization. Packed beds of lengths similar to 200 mm and 300 mm were characterized for packing diameter based Reynolds numbers, Re-d ranging from 0.6 x 10(4) to 8.5 x 10(4). The solid packing used in the bed consisted of circle divide 9.5 mm and circle divide 5 mm steel spheres with suitable arrangements to eliminate flow entrance and exit effects. The ratios of packed bed diameter to packing diameter for 9.5 mm and 5 mm sphere packing were similar to 9.5 and 18 respectively, with the average packed bed porosities around 0.4. Gas temperatures were measured at the entry, exit and at three axial locations along centre-line in the packed beds. The solid packing temperature was measured at three axial locations in the packed bed. An average Nusselt number correlation of the form Nu(d) = 3.91Re(d)(05) for Re-d range of 10(4) is proposed. For engineering applications of packed beds such as pebble bed heaters, thermal storage systems, and compact heat exchangers a simple procedure is suggested for calculating unsteady gas temperature at packed bed exit for packing Biot number Bi-d < 0.1. (C) 2012 Elsevier Inc. All rights reserved.

Glycidyl azide polymer crosslinked through triazoles by click chemistry: curing, mechanical and thermal properties

Glycidyl azide polymer (GAP) was cured through click chemistry by reaction of the azide group with bispropargyl succinate (BPS) through a 1,3-dipolar cycloaddition reaction to form 1,2,3-triazole network. The properties of GAP-based triazole networks are compared with the urethane cured GAP-systems. The glass transition temperature (T-g), tensile strength, and modulus of the system increased with crosslink density, controlled by the azide to propargyl ratio. The triazole incorporation has a higher T-g in comparison to the GAP-urethane system (T-g-20 degrees C) and the networks exhibit biphasic transitions at 61 and 88 degrees C. The triazole curing was studied using Differential Scanning Calorimetry (DSC) and the related kinetic parameters were helpful for predicting the cure profile at a given temperature. Density functional theory (DFT)-based theoretical calculations implied marginal preference for 1,5-addition over 1,4-addition for the cycloaddition between azide and propargyl group. Thermogravimetic analysis (TG) showed better thermal stability for the GAP-triazole and the mechanism of decomposition was elucidated using pyrolysis GC-MS studies. The higher heat of exothermic decomposition of triazole adduct (418kJmol(-1)) against that of azide (317kJmol(-1)) and better mechanical properties of the GAP-triazole renders it a better propellant binder than the GAP-urethane system.

A characteristics study on the performance of a 2-stage light gas gun

In order to obtain an overall and systematic understanding of the performance of a two-stage light gas gun (TLGG), a numerical code to simulate the process occurring in a gun shot is advanced based on the quasi-one-dimensional unsteady equations of motion with the real gas effect,;friction and heat transfer taken into account in a characteristic formulation for both driver and propellant gas. Comparisons of projectile velocities and projectile pressures along the barrel with experimental results from JET (Joint European Tons) and with computational data got by the Lagrangian method indicate that this code can provide results with good accuracy over a wide range of gun geometry and loading conditions.

应用于固体火箭发动机喷管出口羽流温度测量的光学辐射高温计

利用辐射光谱测量温度的原理,设计并制造了一种新型的非接触式光学辐射高温计.详细介绍该温度计的基本原理和标定方法,并将该温度计应用于火箭发动机喷管出口羽流的温度测量中,初步的实验结果和数值模拟结果相一致.

Effects of anode temperature on working characteristics and performance of a low power arcjet thruster

An arc-heated thruster of 130–800 W input power is tested in a vacuum chamber at pressures lower than 20 Pa with argon or H2–N2 gas mixture as propellant. The time-dependent arc voltage-current curve, outside-surface temperature of the anode nozzle and the produced thrust of the firing arcjet thruster are measured in situ simultaneously, in order to analyze and evaluate the dependence of thruster working characteristics and output properties, such as specific impulse and thrust efficiency, on nozzle temperature.

Plume Parameters and Thruster Properties of a Low Power Arcjet

Abstract. A low power arcjet-thruster of 1 kW-class with gas mixture of H2-N2 or pure argon as the propellant is fired at a chamber pressure about 10 Pa. The nozzle temperature is detected with an infrared pyrometer; a plate set perpendicular to the plume axis and connected to a force sensor is used to measure the thrust; a probe with a tapered head is used for measuring the impact pressure in the plume flow; and a double-electrostatic probe system is applied to evaluate the electron temperature. Results indicate that the high nozzle temperature could adversely affect the conversion from enthalpy to kinetic energy. The plume flow deviates evidently from the LTE condition, and the rarefied-gas dynamic effect should be considered under the high temperature and low-pressure condition in analyzing the experimental phenomena.

Multispacecraft refueling optimization considering the J2 perturbation and window constraints

The optimization of a near-circular low-Earth-orbit multispacecraft refueling problem is studied. The refueling sequence, service time, and orbital transfer time are used as design variables, whereas the mean mission completion time and mean propellant consumed by orbital maneuvers are used as design objectives. The J2 term of the Earth's nonspherical gravity perturbation and the constraints of rendezvous time windows are taken into account. A hybridencoding genetic algorithm, which uses normal fitness assignment to find the minimum mean propellant-cost solution and fitness assignment based on the concept of Pareto-optimality to find multi-objective optimal solutions, is presented. The proposed approach is demonstrated for a typical multispacecraft refueling problem. The results show that the proposed approach is effective, and that the J2 perturbation and the time-window constraints have considerable influences on the optimization results. For the problems in which the J2 perturbation is not accounted for, the optimal refueling order can be simply determined as a sequential order or as the order only based on orbitalplane differences. In contrast, for the problems that do consider the J2 perturbation, the optimal solutions obtained have a variety of refueling orders and use the drift of nodes effectively to reduce the propellant cost for eliminating orbital-plane differences. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A new prototype of self-pressurizing fuel tank for micro and nano-satellites

To increase effective load, light-weight micro-propulsion system is necessary for micro-satellites. Traditional propulsion systems including large and heavy high-pressure vessels are difficult to be scaled down to fulfill the demand of micro-satellites. In this article, a novel self-pressurizing fuel tank without high-pressure gas vessel is proposed. When some liquid propellant is consumed, pressure is compensated with CO2 released by heating NH4HCO3 powder in the fuel tank. Comparing with other types of self-pressurizing liquid fuel tank, a gas generator with special and simple structure was designed to stop or continue the NH4HCO3 decomposition reaction easily, and consumed a small amount of energy to heat the powder effectively. Performance tests showed that this new prototype is very suitable for micro-thrusters.

微小卫星贮箱内两相流界面变形定位及控制管理

液化气推进技术是微小伴随卫星在轨飞行采用的一种新型推进技术. 微小伴星空间调姿、变轨过程需精确的推力控制, 因此必须了解卫星推进系统中推进剂的形位分布. 本文理论分析了空间微重力环境下液化气推进剂气/液界面的形位分布及变化, 并通过落塔实验验证了微重力环境下有效控制液体推进剂的管理方法.

Effect of Nozzle Temperature on the Performance of a 1 kW H2-N2 Arcjet Thruster

A 1 kW-class arcjet thruster was ¯red in a vacuum chamber at a pressure of 18 Pa. A gas mixture of H2 : N2 = 2.8 : 1.5 in volume at a total °ow rate of 4.3 slm was used as the propellant with an input power ¯xed at 860 W. The time-dependent thrust, nozzle temperature and inlet pressure of the propellant were measured simultaneously. Results showed that with the increase in nozzle temperature the thrust decreased and various losses increased. The physical mechanisms involved in these effects are discussed.