Thermal Reactivity of Double-Base Rocket Propellants

Thermal decomposition studies on double-base propellants have been carried out in air and vacuum using differential scanning calorimetric, thermogravimetric and mass spectrometric techniques. Low temperature (90 °C–170 °C) decomposition seems to involve diffusion controlled process, the mechanism in air and vacuum being different. High temperature (170 °–205 °C) decomposition involves bond breakage.     

Nitramine Solid Rocket Propellants with Reduced Signature

Solid rocket propellants containing nitramine are considerably superior to doublebase propellants, both as regards their performance and mechanical properties. The pressure exponent of non-modified nitramine propellants is n ≥ 0.9. The possibility of changing the burning rate or, respectively, the pressure exponent has, however, only been realisable to a restricted extent up till now in propellants with an increased nitramine content. In the following propellant system containing nitramine, the effects of ammonium perchlorate on burning behaviour are studied:

• ammonium perchlorate (AP)
• hexogen (RDX)
• nitroglycerin (NG), or trimethylolethane trinitrate (TMETN)
• polyurethane binder (PU)

AP concentrations already as great as 10% produce considerable changes in the burning behaviour of the propellants described above. It is possible to reduce the pressure exponent by the addition of ammonium perchlorate from n ≥ 0.9 to n ≤ 0.65. The burning rates may also be influenced with AP concentrations ≤ 20% and by varying the AP particle size by the factor of 2. All the propellants prepared were easily castable and showed exceptionally good viscoelastic properties (strain at break ε_R > 200%) in the temperature range between −40 °C to + 50 °C. The thermal chemical stability is not influenced negatively in any way by the combination of nitric acid esters and ammonium perchlorate.     

Evaluation of Mechanical Properties of Solid Propellants in Rocket Motors by Indentation Technique

It is essential to evaluate the mechanical properties of propellants in a solid propellant rocket motor (SPRM) for structural integrity and its performance evaluation before the flight test. Conventionally, uni‐axial tensile testing on an universal testing machine (UTM) is used to evaluate the mechanical properties of solid propellant carton which is cast along with SPRM. Propellants in rocket motors experience various types of loading during storage, transportation, and environmental conditions over the period of time before actual flight whereas the propellant carton doesn’t experience the same as it is stored in magazine. At present, the mechanical properties of propellant cast in a carton are considered to be the mechanical properties of propellant in a rocket motor, which is not truly representative. Therefore, a non‐destructive indentation technique has been used to establish a method for evaluating the mechanical properties of solid propellants in rocket motors based on hydroxyl terminated polybutadiene. The test results obtained using the penetrometer indentation technique was analyzed comprehensively and compared with UTM results. The mathematical correlations were also developed using least square method and established by conducting the penetrometer indentation test on similar propellant composition. Further, the developed correlation was used to evaluate the mechanical properties of propellant in flight SPRM by penetrometer indentation technique.     

Adhesion in Solid Propellant Rocket Motors

Plasma treatment of EPDM-based rocket motor insulation materials may change the peel strength between these materials and polyurethane polymers by a factor of 0.1 to 10. The matching of surface energies seems to be important for this adhesion process. The surface tension of the components was measured to between 30 and 50 mNm^-1. The total surface energy of the insulation could be increased from 27.2 mNm^-1 to ca. 70 mNm^-1 by the plasma process. Maximum peel strength could be obtained by a treatment of less than 1 min, whereas in most cases longer times gave lower values. In some cases very long treatment also gave good strength, probably due to a rougher surface structure. The rate of cure of the polymer was important for the adhesion process as lower rates of cure correlated with higher peel strengths, which can be explained by diffusion of the polymer's components into the insulation.     

Sensitivity of Solid Rocket Propellants for Card Gap Test

Card gap test, which is standardized in Japan Explosives Society, was modified in order to apply it to solid rocket propellants and carried out to evaluate sensitivities against shock stimuli. Solid propellants tested here were mainly azide polymer composite propellants, which contained ammonium nitrate (AN) as a main oxidizer. Double base propellant, composed nitroglycerin and nitrocellulose (NC), and ammonium perchlorate (AP)-based composite propellants were also evaluated in order to compare with the azide polymer propellants. It is found that the sensitivity was dominated by the oxidizer characteristics. AP-and AN-based propellant had less sensitivity and HMX-based propellant showed higher sensitivity, and the adding of NC and TMETN were contributed to worse sensitive for the card gap test. Good relationship was obtained between the card gap sensitivity and the oxygen balance of propellants tested here.     

Combustion-Front Non-One-Dimensionality in Single- and Double-Base Propellants

Parameters of transverse waves propagating over the surface of specimens pressed from colloxylin and double-base propellants A and N are studied. By means of microvideofilming and thermocouple measurements, it is shown that the burning sites on the specimen surface are formed by a set of transverse waves. Under atmospheric pressure, the transverse-wave front has the form of a step 0.5–1.1 mm high and decreases with increasing pressure or initial temperature of the specimen. The front propagates with variable velocities in the horizontal and vertical directions. The mean velocity of the transverse wave is three to eight times higher than the normal burning rate of the specimen as a whole (with a wide spread of local values) and increases with increasing pressure. Behind the front, combustion can be interrupted till the arrival of the next transverse wave. As in the SHS process, the reason for the emergence of the transverse waves is combustion-wave spatial instability.     

Tactical Solid Rocket Motors Response to Bullet Impact

Vulnerability studies are very important for rocket motors development using high energy and minimum smoke solid propellant. Among the treated threats, the bullet impact is a well known test studied for many years. Nevertheless, during our experimentations against true rocket motor, the well-known scenarios failed. A new scenario has been identified and confirmed to take into account the new phenomenon. This scenario implied the solid propellant but also the geometry of the grain. The paper presents the experimental works and describes the physics of the phenomenon supported by the associated model studies.     

Organic Fillers for Smokeless Insulations of Double-Base Propellants

The number of organic fillers for smokeless polyurethan-insulations hitherto known is rather small. In this paper two general formulae, expanding the range of organic fillers, are described. The first formula describes molecules with linear structures, the second one those with cyclic structures. A common feature of the fillers mentioned above their high (N + O)/C ratio.     

Experimental study of Double-Base Propellants Combustion Mechanisms

Experimental results about r_b(P) laws of double-base propellants are presented. The influence of many parameters (calorific potential, fabrication process, additives) is studied. With surface structure examination, we show that the first super-rate (P < 100 bar) is in relation with the presence of a carbon layer which increases chemical reactions. The second super-rate (200 bar < P < 400 bar), more physical than chemical, is in relation with the presence of globules (probably PbO) which increase the thermal conductivity near the propellant surface.     

Burning Mechanism of Aluminized Solid Rocket Propellants Based on Energetic Binders

This paper reports results obtained from an experimental study of the combustion mechanism of aluminized propellants based on an energetic binder. The techniques used in this investigation include:     

固体火箭推进剂用燃烧催化剂研究新进展

从纳米燃烧催化剂、含能催化剂和双金属燃烧催化剂三方面综述燃烧催化剂的研究新进展。总结了各类催化剂的特点和发展过程中的技术难题,指出燃烧催化剂的发展方向,如纳米催化剂的团聚抑制和催化机理、含能催化剂能量与感度的统一以及双金属催化剂的结构表征和含能化等。附参考文献55篇。     

A Nonlinear Viscoelastic Theory for Solid Rocket Propellants Based on a Cumulative Damage Approach

A uniaxial nonlinear viscoelastic constitutive equation incorporating cumulative damage was developed and used to successfully model three highly filled composite solid propellants, two based on hydroxy-terminated polybutadiene with an ammonium perchlorate oxidizer and the third on a glycidyl azide polymer with a phase stabilized ammonium nitrate oxidizer. The nonlinear component of the model consists of a strain rate term, a damage term and a nonlinear exponent. The cumulative damage function itself is calibrated using data from two constant strain rate tests at opposing extreme values of strain rate. Four parameters in the nonlinear viscoelastic model are calibrated at a reference strain rate(low rate of strain) for various conditions of cumulative damage and two parameters are calibrated at the opposing extreme rate of strain for the condition of total damage(i.e., when cumulative damage equals unity). The theoretical predictions at intermediate strain rates are very encouraging, providing a good correlation with the experimental stress--strain data measured under uniaxial constant strain rate loading conditions. The incorporation of cumulative damage theory into the nonlinear viscoelastic constitutive equation enables the strength and failure time to be quantitatively defined in terms of the damage history. Predicted values of strength versus failure time and strength versus constant strain rate at the condition when the cumulative damage equals unity, agree reasonably well with the observed experimental results.     

固体火箭发动机装药端面燃烧燃速畸变现象分析

分析比较了引发固体火箭发动机装药燃速畸变的因素,如壁面附近的增塑剂、固化剂的迁移以及细氧化剂聚集等。认为固体装药燃速畸变现象产生的根本原因是壁面附近的氧化剂颗粒质量分数高于中心区域,氧化剂平均粒度低于中心区域。为验证这一理论分析,用基于粒状扩散火焰（GDF）模型得到的推进剂燃速计算公式,以氧化剂的含量和粒度为变量,计算了AP推进剂的燃速。结果表明,燃速计算值的变化趋势与理论分析结果相一致。     

Ballistic Modification of Composite Modified Double-Base Propellants containing ammonium perchlorate

The effect of certain lead and copper salts of organic acids and metallic oxides was studied on the burning rates and heat of explosion of CMDB propellants containing AP. Among lead and copper salts, basic lead salicylate gave increased burning rates, whereas lead methylene disalicylate and basic copper salicylate did not produce catalytic effect. In the case of metallic oxides, ferric oxide and cobalt oxide gave better catalytic activity than copper and chromium oxides. On crosslinking of nitrocellulose with an isocyanate, catalytic activity of ballistic modifiers was reduced. Burning rates obtained in a rocket motor (2-kg propellant) were higher by about 15%-20% than by strand burner (Crawford Bomb) method.     

A New Method for Determining the Vapor Pressure of Nitroglycerine Above Solid Rocket Propellants

Measuring the vapor pressure of compounds such as nitroglycerine (NG) which have a low volatility has not been a straightforward task in the past. There are a wide range of values in the literature for the vapor pressure of NG and furthermore, there is little information regarding the vapor pressure of this compound above solid rocket propellant formulations. We have devised a new method for determining the vapor pressure of NG both above the pure material and solids containing this nitrate ester. The values obtained for pure liquid NG are in good agreement with some previously published values. It was also found that the vapor pressure of this compound is slightly lowered when it is incorporated into solid rocket propellant formulations.     

Burning of Nano-Aluminized Composite Rocket Propellants

Several aluminum nanopowders were examined and compared with the final goal to evaluate their application in solid rocket propulsion. A detailed investigation of pre-burning properties by the Brunauer-Emmet-Teller method, electron microscopy, X-ray diffraction, and X-ray photoelectron spectroscopy was carried out. Ballistic properties and the combustion mechanism of several aluminized propellant formulations were investigated. In particular, aggregation and agglomeration of metal particles at and near the burning surface were analyzed by high-speed high-resolution color digital video recordings. All tested nano-powders are of Russian production; their physical characterization was carried out at the Istituto Donegani (Novara, Italy); ballistic studies were performed at the Solid Propulsion Laboratory (Milano, Italy) using laboratory and, for comparison, industrial composite propellants based on ammonium perchlorate as an oxidizer. Results obtained under a fair variety of operating conditions typical of rocket propulsion indicate, for increasing nano-Al mass fraction or decreasing nano-Al size, larger steady burning rates with essentially the same pressure sensitivity. While aggregation and agglomeration phenomena still occur, their significance may be reduced by using nano-Al instead of micro-Al. __________ Translated from Fizika Goreniya i Vzryva, Vol. 41, No. 6, pp. 80–94, November–December, 2005.     

Rocket Propellants with Reduced Smoke and High Burning Rates

Rocket propellants with reduced smoke and high burning rates recommend themselves for use in a rocket motor for high accelerating tactical missiles. They serve for an improved camouflage on the battle field and may enable guidance control due to the higher transmission of their rocket plume compared to traditional aluminized composite propellants. In this contribution the material based ranges of performance and properties of three non aluminized rocket propellants will be introduced and compared to each other. The selected formulations based on AP/HTPB; AP/PU/TMETN and AP/HMX/GAP/TMETN have roughly the same specific impulse of I_SP = 2430 Ns/kg at 70:1 expansion ratio. The burning rates in the pressure range from 10–18 MPa vary from to 26–33 mm/s for the AP/HTPB propellant, 52–68 mm/s for the formulation based on AP/PU/TMETN and 28–39 mm/s for the propellant based on AP/HMX/GAP. With 58% and 20% AP-contents the propellants with nitrate ester plasticizers create a much smaller secondary signature than the AP/HTPB representative containing 86% AP. Their disadvantage, however, is the connection of high performance to a high level of energetic plasticizer. For this reason, the very fast burning propellant based on AP/PU/TMETN is endowed with a low elastic modulus and is limited to a grain configuration which isn't exposed too much to the fast and turbulent airstream. The mechanical properties of the AP/HMX/GAP-propellant are as good or better as those of the AP/HTPB propellant. The first one exhibits the same performance and burn rates as the composite representative but produces only one fifth of HCl exhaust. For this reason it is recommended for missile applications, which must have high accelerating power together with a significantly reduced plume signature and smoke production.     

Determination of the Decomposition Behavior of Double-Base Propellants at Low Temperatures

Thermal and mechanical stability are very important characteristics of energetic materials used as ingredients of propellants compositions. For the determination of the thermal decomposition behavior of solid double base propellants, usually gravimetric, calorimetric, and chemical methods are being employed. At ICT, the decreasing rates of stabilizers determined by high-performance chromatography at different temperatures are successfully used as a mean to forecast the chemical life time of propellants containing nitric esters. On the other hand, the mechanical properties are evaluated by measuring the tensile strength with a more or less accuracy. In the last years, the increasing performance of the gel-permeation chromatography enabled the determination of the molecular weight and the molecular weight distribution of nitrocellulose in different propellants. In this connection a linear relationship between temperature and depolymerization rate was obtained. This relationship corresponds fairly well with the stabilizer depletion rate and the drop in the mechanical properties.     

Some Aspects of Time-Temperature Superposition Principle Applied for Predicting Mechanical Properties of Solid Rocket Propellants

Optimal test conditions for determining the mechanical properties of rocket propellants (temperatures and strain rates ranges) for delivering master curves were investigated. From master curves it is possible to predict the modulus, maximum stress and maximum strain in vide intervals of temperatures and strain rates, and especially the existing conditions during the ignition of rocket motor. Using the control experiments, at high strain rates, the good agreement between the results obtained from master curves was shown. The obtained results for composite rocket propellants (with carboxy-terminated polybutadiene, CTPB, as a binder), point out the drastic decreasing of maximum strain at high strain rates and low temperatures.     

Prediction and Comparison of Shelf Life of Solid Rocket Propellants Using Arrhenius and Berthelot Equations

The Arrhenius equation and the Berthelot equation for the prediction of shelf life of composite propellant formulations are compared. The elongation has a measurable variation with time and is taken as the fastest degrading parameter for HTPB/AP/Al based composite solid rocket propellants. An HTPB based aluminized composite propellant with 85 % solid loading and an initial elongation of 63.24 % is prepared. It is kept at an elevated temperature of 60 °C to achieve a higher rate of degradation for a prolonged time period (1 year). The elongation is monitored at regular intervals using JANNAF class C dog bone specimen in uni‐axial tensile mode. A reduction of the elongation to less than 50 % is taken as the end‐of‐shelf life of the propellant. The shelf life of the propellant is calculated to be 1.2 years at 60 °C. For the extrapolation of the shelf life at 60 °C to the shelf life at 27 °C, the results of both the Arrhenius equation and the Berthelot equation are compared. The activation energy (E) in the Arrhenius equation is obtained as 72.8 kJ mol⁻¹ and the 10 °C reaction rate rise factor (γ₁₀) is found to be 2.4. This comparison is independent of the propellant formulation and other researchers have reported a similar range of values for these parameters. The shelf life of this propellant formulation at 27 °C is conservatively predicted to be 20 years using both equations. In addition to estimation of shelf life by both equations using elongation as control parameter, this paper gives scaling curves, which are valid universally for predicting shelf life at 27 °C from data of shelf life at 60 °C. The use of scaling curves is independent of properties, propellant formulation and degradation mechanism considered for analysis.