WO2014207405A1 - Composite solid propellant of which the aluminium reducing charge contains a low level of magnesium - Google Patents

Abstract

The present invention concerns a composite solid propellant of which the composition contains an oxidising charge and a reducing charge in a binder. The reducing charge characteristically consists of particles of at least one aluminium/magnesium alloy or consists of a mixture of particles of at least one aluminium/magnesium alloy and aluminium particles; said reducing charge containing a total quantity of magnesium that represents less than 3 %, and advantageously less than 1 %, of the mass of same. The presence of such a low level of magnesium helps reduce the thrust instabilities of a rocket engine operating with said propellant and helps increase the propulsive performances of said rocket engine.

Description

 COMPOSITE SOLID PROPERGOL WHOSE ALUMINUM REDUCTIVE LOAD HAS A LOW MAGNESIUM LEVEL

The main objects of the present invention are:

solid composite propellants;

 solid propellant charges constituted by said solid propellants; and

 the rocket motors closing said loadings.

 The invention is in the field of solid propellant propulsion and more particularly relates to aluminized composite solid propellants.

 The applications envisaged mainly concern rocket engines (accelerators or launchers stages), large ones, solid propellants for space launch vehicles or rocket engines, smaller in size, with solid propellants for tactical vehicles (missiles, rockets).

 The solid propellant rocket engines for space launchers existing to date, are of the type of the Ariane 5 rocket or the US space shuttle, large (typically, height ~ 20 m and diameter ~ 5 m). The solid propellant loads contained in this type of engine have a mass ranging from a few hundred kilograms to several hundred tons. Their operating time is of the order of a few tens of seconds to a few minutes.

 Rocket engines for tactical vehicles have smaller dimensions (typically, 0.5 m <height <10 m and 0.1 m <diameter <2 m). The solid propellant loads contained in this type of engine have a mass ranging from one kilogram to a few hundred kilograms. Their operating time is of the order of a few seconds per minute.

Solid propellants for these applications are composite propellants with an inert binder of the polyurethane type. They contain an oxidizing charge (generally consisting of ammonium perchlorate fillers (grains)) and a reducing charge (consisting of aluminum fillers (grains)). This propellant family is that concerned by the present invention. The mass ratios of these ingredients are typically about 68% for the oxidizing charge (usually ammonium perchlorate), about 20% for the reducing charge (aluminum) and about 12% for the binder.

 The burning rate (Vc) of a solid propellant depends on the pressure P prevailing in the combustion chamber of said solid propellant and conventionally follows a law (called the law of old) expressed in the form:

Vc = aP ⁿ .

 Said combustion rate Vc and the pressure exponent n of the propellant are fundamental parameters for the ballistic adjustment of a solid propellant engine (combustion time, thrust, combustion stability, etc.).

 The standard values of the ballistic parameters for the propulsion applications concerned by the present invention, using polyurethane-bonded aluminized composite propellants, are a combustion rate Vc of a few mm / s to 10 mm / s and a pressure exponent n = 0, 2 to 0.5, in an operating pressure range of 3 to 10 MPa.

 The person skilled in the art knows how to choose the particle sizes of the constituent raw materials of the solid propellant (in particular that of the oxidizing charge) to control the speed levels of the solid propellant.

 The aluminized composite propellants produce, during their combustion, gases and solid particles consisting mostly of alumina (about 30% of the mass ejected by the propellant).

 The combustion of aluminum in alumina during the combustion of composite propellants has been extensively studied. To improve this combustion, it has been recommended to associate magnesium with aluminum.

 The combustion of particles made of an aluminum / magnesium alloy has thus been studied and has been the subject of scientific publications. The teachings of the publications entitled:

"Combustion of individual aluminum-magnesium alloy particles in a flame of an oxidizer-fuel mixture", in Combustion explosion and shock waves, Vol. 17, No. 2, pages 186-190, April-June, 1971, and - "Combustion mechanism of aluminum-magnesium alloy particles", NTIS report FTD-HT-23-1238-74, 23 August 1974 show that the combustion of particles formed of an Al / Mg alloy is more complete and faster than that of aluminum particles ("pure"). Magnesium, entering at levels of 5% to 95% by weight in the Al / Mg alloy, has the role of exploding the droplets of molten metal into fine droplets, which promotes the good combustion of aluminum. thus better dispersed.

 US Pat. No. 3,044,911 (1962) discloses composite propellants incorporating a metallic filler whose particles, having an oblong shape or a needle shape, are made of aluminum ("pure") or an Al / Mg alloy. . The most satisfactory combustion properties (burning rate) are obtained for Al / Mg alloys incorporating from 15% to 35% by weight of magnesium.

 US Pat. No. 3,180,770 (1965) discloses solid propellants containing a solid filler consisting of an Al / Mg alloy, the Mg content in the alloy being between 5% and 70% by weight. These solid propellants are particularly interesting, compared to those that do not close Mg in the solid charge, because of their higher density, their better ignitability and combustion (better combustion due to the better combustion of metal particles) and their speed. higher combustion.

 US Pat. No. 3,350,245 describes composite propellants whose composition essentially contains an inorganic oxidant (ammonium or ammonium perchlorate) and an elastomeric polyurethane reducing binder. Within said composition, a part of said inorganic oxidant may be replaced by a powdered reducing metal, which may consist of a powder of an Al / Mg alloy containing from 5 to 15% by weight of Mg and of 95% by weight. at 85% by mass of AI.

This literature, dating from the years 1960-1970, therefore recommends the use of Al / Mg alloy particles, at mass levels of magnesium of at least 5%, in order to obtain an improved combustion of the particles (by compared to that of "pure" aluminum particles) and thus improved combustion properties of composite propellants incorporating said particles. The teaching of the VR publication Pai Verneker et al. in COMBUSTION AND FLAME, vol. 67, no. 2, 1987, pages 163-173 (ELSEVIER SCIENCE PUBLISHING CO.) Is in the same spirit. Later, the results of a study on the burning of aluminized propellants were published in Progress in Astronautics and Aeronautics, Vol. 90, pp. 479-513, 1984 ("Combustion of metallized propellants"). The phenomena involved in the burning surface of the propellant have been explained. It has been shown that the aluminum particles escaping from said burning surface of the propellant are likely, for a part of them, to agglomerate to form drops ("big drops") of a size much higher than that of the aluminum charges (introduced and present in the mass of the propellant). This study assumes that the formation of these "large drops" of aluminum is likely to affect the burning rate of the propellant, its combustion and generate instabilities of combustion.

 Since then, those skilled in the art have understood that in the combustion chamber of a solid propellant propellant, when dealing with an aluminized propellant, the combustion of aluminum is not an instantaneous phenomenon; the burning time of a drop of aluminum being all the more important that the size of said drop is important.

 The combustion of aluminum in the combustion chamber is thus a phenomenon which spreads out in space and lasts in time, insofar as the drops of liquid aluminum emitted on the surface of the burning propellant are entrained in the gas flow.

 This phenomenon is referred to as distributed combustion of aluminum in the combustion chamber.

 This distributed combustion induces vibration phenomena (pressure oscillations) in the combustion chamber. Theoretical work and laboratory tests have confirmed the impact of the distributed combustion of aluminum on the pressure oscillations in the combustion zone.

Distributed combustion in volume of aluminum thus induces a thermo-acoustic instability (which is reflected by pressure oscillations) in the combustion chamber, which can be described as follows: spatial and temporal fluctuations of pressure (fluctuations of the acoustic field) act on the (variable) phase shift between the speed of the gases and the speed of the liquid aluminum droplets, this phase shift induces a fluctuating energy production (production related to the combustion of aluminum), this fluctuation of energy production maintains (or even amplifies) fluctuations in the acoustic field, and therefore pressure in the combustion chamber.

 These rocket motor pressure oscillations cause thrust oscillations, sources of vibrations of the structure integrating said engine, which are likely to damage said structure and more particularly the payload of a launcher. It is therefore always sought to reduce these phenomena, especially in the case of rocket launchers engines carrying fragile payloads, such as satellites, to preserve said payloads.

 In the 2000s, thrust oscillations became a major concern for improving the supply of space launch vehicles, and were the subject of theoretical work (see, for example, "Aluminum combustion driven instabilities in solid rocket". motors ", in the Journal of Propulsion and Power, Vol 25, No. 2, March-April 2009 (pages 509-521)).

 Moreover, the person skilled in the art has always been concerned with improving the propulsive performance of solid propellants of rocket engines.

 The specific impulse (Isp) of a propellant reflects the intrinsic propulsive performance of said propellant and can be expressed according to:

 Isp = C * x Cf / g (s)

with:

 C *: called characteristic speed, which takes into account the thermochemical properties (gas temperature, gas density, molar mass ...) of the propellant combustion;

 Cf: referred to as the thrust coefficient, which takes into account the thermodynamic properties (P, nozzle neck area, gas density, etc.) of propellant combustion; and

g: coefficient of gravity.

The specific impulse indicates how long a kilogram of propellant produces a thrust to move a mass of one kilogram (or a force of about 9.81 N) into the Earth's gravitational field. The thrust of an engine incorporating a solid propellant is proportional to the Isp of said solid propellant and is written:

 F = Isp x q x g (N)

with q: mass flow rate of the propellant.

 A gain of a few tenths of a percent on the specific pulse of a solid propellant represents for the skilled person a significant improvement.

 The publication entitled "Powdered aluminum and oxygen rocket propellants: subscale combustion experiments", report NASA Technical Memorandum 106439 of 1993, is thus concerned (not with the combustion properties of a composite propellant incorporating, in its composition, a metallic charge Al / Mg but) propelling performance of metal powders Al / Mg in combustion with oxygen. This publication shows that the combustion with oxygen of a metal charge consisting of an Al / Mg alloy at a mass ratio of 9.8% Mg leads to a loss of propulsive performance compared to that a load of "pure" aluminum (see more particularly Figure 8 of this publication).

 The specific impulse loss (Isp, see above), induced by the incorporation of magnesium as a partial replacement of aluminum, is confirmed when the charges in question (in Al / Mg alloys) are used in composite propellants. . The applicant has more recently carried out thermodynamic calculations which show and quantify this loss (see the attached figure).

 Thus, the skilled person has not, because of the loss of propulsive performance induced, exploited composite propellants incorporating Al / Mg alloy particles, replacing pure aluminum particles. Aluminized composite solid propellants, which are currently used in rocket engine loadings, contain high purity aluminum.

The reducing charges of said composite solid propellants consist of an aluminum with a purity greater than 99.8%, or even 99.9% (by mass). In addition, there is alumina and the main impurities present are iron (at a content of about 0.16% by weight) and silicon (at a content of about 0.03% to about 0.01%). % by mass). Other metals likely to be present in the The composition of this high-purity aluminum is, in any event, less than or equal to 0.007% or even 0.001% by mass. As for the magnesium content, it is generally less than or equal to 0.001% by weight.

 Aluminum (high purity) reducing charges used in aluminized composite solid propellants are in the form of grains, more or less spherical, with a median diameter generally between 1 and 50 μηι.

 In such a context, the inventors have considered the phenomenon of the agglomeration of aluminum particles escaping from the combustion surface, that of the formation of "large drops", on an original plane. They highlighted the fact that the said "big drops" would be the cause of another problem, in addition to that of the instabilities of combustion (thrust oscillations): the problem of the incomplete combustion of the reducing charge in the chamber combustion of a rocket engine. Indeed, the transit time of "big drops" of aluminum in the combustion chamber of the rocket engine, before the expulsion of said "big drops" by the engine nozzle rocket, may be insufficient to ensure the total combustion of said "Big drops". Thus, a portion of the aluminum contained in the propellant charge of the engine can be expelled from the combustion chamber of the rocket engine by the nozzle without having completely burned.

 The inventors have carried out theoretical work and evaluated that: - in a rocket engine for tactical devices, of small dimensions, and in which therefore the transit time of the particles is short, the loss of thrust, with respect to the theoretical thrust (calculated with reference to the specific impulse of the propellant), induced by the incomplete combustion of the "big drops" of aluminum would be of the order of 2% to 3%;

- In a rocket engine for space launchers, of larger dimensions, the transit time in the engine of the particles is more important and the loss of propulsive performance would be only 0.5% to 1%.

The formation of "large drops" of aluminum on the combustion surface of the aluminized composite solid propellant is therefore likely to cause, on the one hand, thrust oscillations, the effects of which are more critical in the case of rocket engines for space launchers, and, on the other hand, a loss of performance of the rocket engine, loss of performance all the more marked as the dimensions of the rocket engine are less important.

 It is to the inventors' merit, in such a context, to have considered again the substitution, to a part of the aluminum of the reducing charge, of magnesium and to have brought to light that with a low rate of substitution. (see below), the propellant-specific impulse loss inherent in the replacement of aluminum (Al) with magnesium (Mg) (a specific impulse loss which was a real prejudice (see above)) is more damaging to the extent that it is compensated for, and even exceeded, by the increase in the thrust of the rocket motor, due to the more complete combustion of the reducing charge with magnesium (compared to the reducing charge without magnesium) , ie because of the bursting of the "big drops". Thus, although a solid propellant incorporating a low level of magnesium substituted for a portion of the aluminum has a specific impulse (Isp2) lower than that (Ispl) of the corresponding propellant containing pure aluminum (Isp2 < Ispl), a rocket motor incorporating it has a thrust (F2) greater than that (Fl) of said rocket engine incorporating the corresponding propellant incorporating pure aluminum (F2> F1). In addition, said bursting of "large drops" is advantageous with reference to the technical problem of thrust oscillations (see above). It is therefore the merit of the inventors to have shown the interest of such a substitution (Mg to Al), at a low rate, both on the reduction of thrust instabilities and on the propulsive performance of a rocket engine using the propellant involved.

The low levels of Al substitution by Mg recommended according to the invention (see below) are lower or even much lower levels than those described in the prior art (> 5% by weight), as far as where they are optimized with reference to the above compromise (loss of specific impulse (because magnesium is less energetic than aluminum) / thrust gain (due to better combustion of aluminum)). Incidentally note here that with rates below 5% in mass, but close to 5% in mass, the technical effect highlighted by the inventors is expressed, but with little intensity.

 The inventors thus propose aluminized solid propellants generating in the combustion chamber of a rocket engine, on their burning surface, fine drops of aluminum, burning in said combustion chamber in a short time, with reduction of the phenomena of thrust oscillations and increased rocket engine thrust. The propellants of the invention are, as indicated above, particularly interesting in that they therefore make it possible to increase the performance of the engines (in terms of thrust) and to reduce aerodynamic thrust oscillations.

 According to its first object, the present invention therefore relates to composite solid propellants. The composition of said composite solid propellants conventionally contains an oxidizing charge and a reducing charge in a binder. Typically, said reducing charge consists of particles of at least one aluminum / magnesium alloy or consists of a mixture of particles of at least one aluminum / magnesium alloy and aluminum particles; said reducing charge containing a total amount of magnesium which represents less than 3%, advantageously less than 1%, of its mass.

 The particles in question (Al / Mg or Al / Mg and Al) are particles (fillers) of the type of those of the prior art, ie grains, more or less spherical, whose median diameter is generally between 1 and 50 μ m

The reducing charge of composite solid propellants of the invention therefore contains magnesium (presence of Al / Mg alloy particles: all the particles are of this type (in the same alloy or in at least two alloys of this type) or at least only a part of them, the others being aluminum) and said magnesium, responsible for the bursting of "large drops", present in small quantities (less than 3%, advantageously less than 1%, by mass), has a beneficial effect on the thrust oscillations (by reduction of the distributed combustion of aluminum) and on the thrust itself (due to a more complete combustion (see above)) (beneficial effect on the thrust, despite an inevitable loss of specific impulse due to the substitution of a part of Al by Mg, insofar as the rate of said substitution remains low (less than 3% by weight, advantageously less than 1% by weight)).

 According to an alternative embodiment, the total amount of magnesium represents from 0.2 to 0.6% (limiting values included) by weight, advantageously from 0.35 to 0.45% (limit values included) by mass (very advantageously 0 , 4% by weight), of the mass of said reducing charge. Under these conditions, the compromise mentioned above is optimized, the combustion of aluminum is optimized with respect to the specific loss of impulse.

 According to another variant embodiment, which advantageously accumulates with the preceding one, the reducing charge consists of particles of at least one aluminum / magnesium alloy (it does not contain any aluminum particles), advantageously particles of aluminum / magnesium alloy (it does not contain aluminum particles and only one type of alloy is involved). It has been understood that said alloy (s) contain (s) less than 3% by weight of magnesium and, preferably, between 0.35 and 0.45% by weight of magnesium.

 With reference to the composition of the composite solid propellants of the invention, it is also possible to add the following.

The oxidizing charge, mixed with the reducing charge as specified above in the binder, generally comprises grains of ammonium perchlorate (NH ₄ CIO ₄ ). It is not excluded, however, that it includes or even consists of grains of an oxidant of another nature. According to an advantageous variant, said oxidizing charge consists of grains of ammonium perchlorate (see introduction of this text).

 The binder in question is usually a polyurethane binder (see also the introduction to this text). It advantageously consists of a polyurethane binder obtained by crosslinking a telechelic polybutadiene (PBHT) with an isocyanate-type crosslinking agent.

 The composition of composite solid propellants of the invention generally contains:

a reducing charge as defined above (ie particles of at least one aluminum / magnesium alloy or a mixture of particles of at least one aluminum / magnesium alloy and aluminum particles, said reducing charge containing a total amount of magnesium which represents less than 3%, (advantageously less than 1%) of its mass, for 12 to 20% of its mass;

 an oxidizing charge for 60 to 70% of its mass; and

 a binder for 12 to 20% of its mass.

 According to a variant, said composition contains:

 from 12 to 20% by weight of particles of an Al / Mg alloy, whose magnesium content is as specified above (is advantageously between 0.2 and 0.6% by weight);

 from 60 to 70% by weight of ammonium perchlorate grains; and

from 12 to 20% by weight of a polyurethane binder.

 For all practical purposes, it is specified that the ranges of values indicated above include their limit values.

 The composition of the solid propellants of the invention is moreover likely to contain (conventional) additives, such as:

 at least one plasticizer: said at least one plasticizer is preferably chosen from dioctyl azelate (DOZ), diisooctyl sebacate, isodecyl pelargonate, polyisobutylene and dioctyl phthalate (DOP);

 at least one additive of another type: said at least one additive (of another type) may especially consist of at least one adhesion agent between the binder and the oxidizing charge, such as, for example, bis oxide; (2-methylaziridinyl) methylamino phosphine (methyl BAPO) or triethylene pentamine acrylonitrile (TEPAN), at least one antioxidant selected from those of the rubber industry, such as ditertiobutylparacresol (DBC) or 2,2 ' -methylene-bis (4-methyl-6-tert-butyl phenol) (MBP5), at least one crosslinking catalyst, such as iron or copper acetylacetonate, tin dibutyldilaurate (DBTL), in at least one combustion catalyst, such as iron oxide, ...

 According to its second object, the invention relates to a solid composite propellant charge. Said loading consists of a composite solid propellant according to the first subject of the invention (material as described above).

According to its third object, the invention relates to a rocket engine containing a composite solid propellant charge, particularly suitable for space launchers or tactical gear. Typically, said engine contains a solid propellant charge according to the second subject of the invention, ie consisting of a composite solid propellant according to the first subject of the invention. It will be understood, in view of the above remarks, that said engine has reduced thrust oscillations and improved propulsive performance (with respect to said engine, the load of the same type consists of a composite solid propellant of the prior art (= whose charge is made of "pure" aluminum).

The invention is now described, in no way limiting, with reference to the attached figure (unique) and the example below.

 In said figure, it has been shown the ratio between the C * (characteristic speed, see above) of the same type of propellants whose reducing charge consists of an alloy containing Mg in substitution of aluminum (propellants type B of the example below) and the C * of a reference propellant (propellant A (of the same type) of the example below) whose reducing charge is composed of "pure" aluminum, as a function of the percentage of Mg replacing the aluminum in the Al / Mg alloy.

Example The reference propellant A (prior art) is a polyurethane-bonded composite solid propellant (12% by weight), comprising a charge (oxidizing) of ammonium perchlorate (68% by weight) and a (reducing) filler. "pure" aluminum (20% by weight).

 Propellant B, an example of the invention, is a composite solid propellant (of the same type) with a polyurethane binder (12% by weight), comprising a charge of ammonium perchlorate (68% by weight) and a metal charge (20%). % by weight) consisting of an Al / Mg alloy containing (a low content) 0.4% by weight of magnesium. The mass ratio of magnesium in the propellant, of 0.08%, is low.

The appended figure thus shows the loss of C *, and therefore of specific impulse (see the formula given above), of the propellant composite A reference, when the aluminum is partially replaced by magnesium.

 For propellant B (replacement of 0.4% by mass of Ai with Mg), the loss of propulsive performance translated by the decrease in the value of C * is 0.015%, and the loss on the impulse Specifically calculated for a combustion pressure of 70 bar and expansion in vacuum, is 0.035%.

 Table 1 below summarizes these results.

Table 1

According to the phenomenon described in the present description, a propellant of type B incorporating a low level of magnesium (as specified in the text above) allows, by bursting "large drops" of aluminum, a more complete combustion of the aluminum in a rocket engine, and therefore a thrust boost of said rocket engine from 0.5% to 3% (compared to the same rocket engine operating with a type A propellant).

 The type B propellant therefore has a certain advantage in terms of the propulsive performance of a rocket engine compared to type A propellant.

 It is recalled that with regard to the specific impulse of solid propellants A and B, we have:

 Isp (propellant B) <Isp (propellant A) (due to the substitution of one part of Al by Mg),

 but that, surprisingly, as regards the thrust of a rocket engine operating with said propellants A or B, we have:

F (propellant B)> F propellant (A) (due to more complete combustion). Moreover, the presence of Mg, favorable to the bursting of "large drops" of aluminum ensures faster combustion of aluminum, thus reducing the phenomenon of distributed combustion of aluminum in the propellant, and therefore undoubtedly an effect on the reduction of combustion instabilities.

 In conclusion, a rocket engine incorporating a propellant according to the invention (propellant of the type of propellant B above) has a superior propulsive performance (a greater thrust) to the same rocket engine incorporating the propellant of the corresponding prior art (propellant of the type of propellant A above), and in addition, it has lower combustion instabilities. It is therefore particularly well suited for space applications, for which the propulsive performance and the comfort of the payload are sought. It is also interesting for tactical applications because of the thrust gains achieved.

Claims

Composite solid propellant, the composition of which contains an oxidizing charge and a reducing charge in a binder, characterized in that said reducing charge consists of particles of at least one aluminum / magnesium alloy or consists of a mixture of particles of at least one aluminum / magnesium alloy and aluminum particles; said reducing charge containing a total amount of magnesium which represents less than 3%, advantageously less than 1%, of its mass. Composite solid propellant according to claim 1, characterized in that said total amount of magnesium represents from 0.2 to 0.6% by weight, advantageously from 0.35 to 0.45% by weight, of the mass of said reducing charge. Composite solid propellant according to claim 1 or 2, characterized in that said reducing charge consists of particles of an aluminum / magnesium alloy. Composite solid propellant according to any one of Claims 1 to 3, characterized in that the said oxidizing charge comprises grains of ammonium perchlorate, advantageously consisting of grains of ammonium perchlorate. 5. Composite solid propellant according to any one of claims 1 to 4, characterized in that said binder is a polyurethane binder. Composite solid propellant according to any one of Claims 1 to 5, characterized in that its composition contains:  said oxidizing charge, at a rate of 60 to 70% of its mass;  - said reducing charge, at a rate of 12 to 20% of its mass; and said binder, at a rate of 12 to 20% of its mass. 7. Composite solid propellant according to any one of claims 1 to 6, characterized in that its composition contains: from 60 to 70% by weight of ammonium perchlorate grains; - 12 to 20% by weight of particles of an aluminum / magnesium alloy, whose magnesium content is between 0.2 and 0.6% by weight; and  from 12 to 20% by weight of a polyurethane binder. 8. Composite solid propellant charge, characterized in that it consists of a composite solid propellant according to any one of claims 1 to 7. 9. Rocket engine containing at least one composite solid propellant charge, particularly suitable for space launchers or tactical devices, characterized in that said at least one composite solid propellant charge is a solid propellant charge according to claim 8. 10. Rocket motor according to claim 9, reduced thrust oscillations and improved propellant performance.