Bogdan FLORCZAK ? Institute of Industrial Organic Chemistry, Warsaw; Marek BIAŁEK, Mirosław SZCZEPANIK, Arkadiusz DZIK ? ZPS GAMRAT Sp. z o.o, Jasło

Abstract:
This paper presents the test results for the fundamental ballistic parameter of non-homogeneous solid rocket motors, that is, the linear burning rate and its dependence on  pressure. The burning rate was determined in the laboratory rocket motor system on the basis of the recorded characteristic curves p = f(t) obtained during the burning  procedure of the propellant cartridges of cylindrical shape with inhibited external lateral surface and the dimensions which provided the quasi constant level in the  combustion chamber. The change in the quasi constant pressure in the combustion chamber was obtained by burning the cartridges of the same shape but using nozzles  of various critical diameters.

Please cite as: CHEMIK 2013, 67, 13, 25-32

Introduction

Non-homogenous solid rocket propellants (also known as heterogeneous regarding their structure) are multi-component mixtures of liquid and solid chemical compounds, both organic and inorganic ones, whose properties depend on the composition and their ballistic properties also depend on the content of burning rate modifiers in propellants.

Non-homogeneous solid rocket propellants (hereinafter referred to as propellants) as high energy materials are used in the production of propellant cartridges of various shapes, types (insulated ones ? more rigid or bonded to the chamber wall in the rocket motor ? more flexible) and purposes (e.g. short and medium-range rockets). Taking into account their structures, propellants are the physical mixture of constant inorganic oxidiser, mainly ammonium chlorate (VII) (NA) and various types of solid and liquid chemical additives (organic and inorganic compounds and elements) which serve as technological, energy and ballistic additives. Ammonium chlorate (VII) and additives are introduced at the increased temperature into a liquid binding agent consisting of liquid synthetic rubber with functional groups, usually hydroxyl groups (?,??dihydroxyl polybutadiene known as HTPB) [1], plasticiser, hardening and cross-linking agents, which form the composite solid after the hardening process that occurs at the increased temperature of 50?80°C [2÷7].

Before hardening, propellants are a multi-component heterogeneous system usually containing two/three fixed components introduced into a three-component liquid phase (synthetic rubber, plasticiser, hardening agent) of a non-homogenous inner structure in which chemical reactions related to the hardening and cross-linking processes additionally take place, and which is consequently characterised by the defined pot life. It is a specific suspension having the high density of solid phase particles, whose viscosity depends, among other things, on: viscosity of the liquid phase, shape and sizes of particulates, particle size distribution, content of separate fractions of the solid phase (one, two or more), shear rate, temperature and time [8], and considering the technology for forming propellant elements from this suspension by the casting method under the reduced pressure, its viscosity should be ? 1.5 kPa [9]. As it has been mentioned above, propellants bonded to the wall of rocket motor chamber, which generally contain 84% of the solid phase, are among the propellant cartridges produced from this type of propellant.

The rocket engine consists of a motor casing, an insulator, a liner and propellant. After pouring propellant into the chamber and its hardening, it is bonded to the motor chamber wall through a flexible intermediate layer [10, 11] and it should be characterised by, among other things, adequate ballistic parameters. Linear burning rate and its dependence on pressure belong to the main ballistic parameters of propellants. This property can be modified by adding both organic an inorganic metallic chemical compounds as well as by using a suitable hardening agent. Generally, iron compounds [12], i.e. oxides [13÷15], nano-oxides [14, 15] or ferrocene [16] and its derivatives [17÷19] are metallic chemical compounds. The most commonly used hardening agents are: dimeryl diisocyanate (DDI), isophorone diisocyanate (IPDI) or toluene diisocyanate (TDI) [20]. A testing method in the laboratory rocket motor (LRM) system is one of the fundamental methods for indirect determination of this parameter [21, 22]. Such a method was applied by the authors in the course of conducted tests using the LRM of their own construction [23].

Experimental part

The mass of testing propellants of specified compositions (Tab. 1) was prepared using the planetary mixer with 6 dm3 volume by Drais Company.

The process of obtaining propellants for the tests included mixing the propellant components in the mixer at the increased temperature (65°C), under the reduced pressure of ca. 5 kPa (by adding the successive doses of particular solid components, i.e. Al powder and NA to the liquid mixture and additives, and DDI in the final phase of mixing) and for a specified time of more than one hour until the mass became liquefied and the mould was filled with the propellant mass under the reduced pressure specified above (Fig. 1) [24]. The filling process was based on putting the casting mould into a vacuum chamber, placing the inlet of the casting mould under the dosing unit, combining it through the bottom with the vibration generating unit ? the vibrator, placing it in the thermostat chamber, filling the feeder with semi-liquid propellant mass, creating vacuum and filling the chamber with semi-liquid propellant mass in a form of streams flowing off along the forming core. During the filling, the casting mould was heated at the constant temperature and subjected to vibrations using the vibrator to deaerate propellant more efficiently and to provide its uniform distribution in the mould. When the filling process was completed, the vacuum system was unsealed. As a result, the entering atmospheric air tightens propellant from the top, improving its compactness. After filling the mould with propellant, the mould was taken out of the vacuum and thermostat chambers and placed in the next thermostatic unit ? an incubator, to harden propellant. After hardening, the forming core was removed from the propellant cartridge, while the propellant cartridge in a form of the cylinder with a central duct was removed from the mould. Then, the propellant cartridge was mechanically treated and subjected to flaw detection testing. The cartridges prepared for testing as described above are presented in Photo 1.

Methodology

The method in the laboratory rocket motor (LMR) system was used for propellant tests to determine the linear relation between the burning rate and pressure [23]. This system consists of a replaceable chamber, a case, a nozzle chamber with a replaceable nozzle, a primer case, a chamber lock, a fuse plate and a perforated screw cap. The system consisting LSR, extensometer pressure transducers and digital amplifier MGC Plus allows the characteristic curve p = f(t) to be determined which is a basis for determining the burning rate of a tested cylindrical propellant with an inhibited external lateral surface and the dimensions providing quasi constant pressure in the combustion chamber. Owing to such a constructional solution, the linear burning rate of propellant depending on pressure controlled by replacing the nozzles can be determined indirectly. LMR is shown in the attached figure (Fig. 2), demonstrating the half-section of the laboratory rocket motor with the propellant cartridge in the isometric view.

The construction of the laboratory rocket motor according to the invention [23], besides retaining the characteristic features for the motors so far known [22] that have the protection system consisting of a diaphragm and a perforated securing nut, simplifies the motor construction as a result of placing the primer in the same duct where the diaphragm and the perforated securing nut are present, which can be opened in an emergency (ventilation duct) in case of overpressure of the combustion products in the motor chamber. The simplified construction of the frontal module of the rocket motor provides the possibility for performing only one flow duct in the front mounting (closure) of the motor. This duct placed outside the longitudinal axis of the motor is intended for the pressure transducer instead of so far applied two ducts situated outside the longitudinal axis of the motor, i.e. the one duct for the pressure transducer and the second one for the primer.

Test results

The tests on cartridges made from propellants (Tab.1) were carried out at the ballistic testing station in the LMR system as presented in Photo 2, at ZPS Gamrat Sp. z o.o. The measurements of the pressure were performed using the digital amplifier MGC Plus and extensometer pressure transducers. The measurement data were sent on-line and archived on a computer.

As a result of the conducted tests, the characteristic curves p = f(t) were recorded. They are shown in Figures 3 ÷ 6 (numbers in a legend indicate critical diameters of nozzles expressed in mm).

Methods of calculating the linear burning rate

The burning rate of a given rocket propellant dependent on the pressure was calculated on the basis of the recorded characteristic curves p = f(t) in the course of burning the cartridges in the laboratory motor by determining indirectly the burning rate of propellant for each recorded characteristic curve by calculating the average pressure (pressure integral divided by burning time tp = t2-t1) and burning time, and calculating the burning rate from the formula u = d/tp, where d is the thickness of the burning layer. In each tested case, the thickness of the burning layer was 16 mm. Pressure level in the motor chamber was regulated by replacing the nozzle (the lower the nozzle diameter was, the higher the pressure in the chamber was). The burning time was determined as a difference between the final (t2) and the initial (t1) burning time.

The time for which p = 0.5 pmax was assumed as the initial t1 and the final t2 time , using methods used by French firms SNPE and ONERA, German firm BYAERN-CHEMIE and Italian FIAT AVIO [25]. However in the first case it was assumed at the increasing initial curve of pressure, whereas in the second case ? at the declining curve of pressure (illustrated in Fig. 7).

Table 2 presents the values of average pressure, burning time and rate determined as specified above and the exponential dependence of burning rate on pressure u =A pn for the tested propellants calculated on their basis. The relationships between the burning rate and pressure u = f(p) are shown in Figure 8, where: u [mm/s], p [Mpa], A [mm/(sMPan)]

Summary

The designed laboratory system and the measurement method provide the indirect determination of the relationship between the burning rate and the pressure corresponding to more real conditions existing in the rocket motor. For higher pressures, the more convex changes in pressure during the cartridge burning in the motor chamber are observed. This can indicate that the cylindrical internal surface of the cartridge and the frontal surfaces are not ignited simultaneously and particularly, faster ignition from the side of the igniter than from the nozzle side (higher pressure and burning rate, shorter time for ignition the surface charge). This results in a progressive-degressive burning area instead of the quasi constant one.

Propellant with reduced content of Al powder (marked as propellant P3) demonstrates a higher burning rate. The further tests to obtain better measurement reliability are expected to involve several burning procedures at a known diameter of the nozzle. This research is financed as the development project from the funds allocated for science in years 2010÷2013.

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Bogdan FLORCZAK ? Ph.D. (Eng), graduated from the Faculty of Chemistry and Technical Physics at the Military University of Technology [WAT] in 1976. In 1990, he got his PhD at the Faculty of Chemistry and Technical Physics at the Military University of Technology. He is currently working in the Institute of Industrial Organic Chemistry. Research interests: chemistry and technology of high energy materials, particularly solid rocket propellants, materials technology. He is the author and co-author of 60 articles published in scientific and technical journals, an author and a co-author of 43 papers and posters presented at national and international conferences. He is co-author of 14 patents and 13 patent applications.
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Marek BIAŁEK ? M.Sc., graduated from the Faculty of Chemical Technology and Engineering at the Silesian University of Technology in 1980. He is currently working at ZPS Gamrat Sp. z o.o. Research interests: solid rocket propellants. He is co-author of several posters and patent applications.
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Mirosław SZCZEPANIK ? M.Sc., graduated from the Faculty of New Technologies and Chemistry at the Military University of Technology [WAT] in 2010. He is currently working at ZPS Gamrat Sp. z o.o. Research interests: high energy materials, rocket motors, thermodynamics.
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Arkadiusz DZIK ? Eng., graduated from the Faculty of IT and Administration at the University of Information Technology and Management in Rzeszów in 2008. He is currently working at ZPS Gamrat Sp. z o.o. Research interests: ballistic tests, rocket motors, numerical calculations and simulations, programming, 3D graphics.
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