Two Main Types Of Propellants Biology Essay

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A solid propellant consists of several chemical ingredients such as oxidizer, fuel, binder, plasticizer, curing agent, stabilizer, and cross-linking agent. The specific chemical composition depends on the desired combustion characteristics for a particular application. Two main types of propellants (homogeneous and heterogeneous) are distinguished by the condition in which their constituent ingredients are interconnected. In a homogeneous propellant, the ingredients are linked chemically and the resulting physical structure is homogeneous throughout. Typical examples of homogeneous propellants are single-base (NC nitrocellulose) or double-base (NC and NG nitroglycerine) propellants. In a heterogeneous or composite propellant, the ingredients are physically mixed, leading to a heterogeneous physical structure. It is composed of crystalline particles acting as oxidizer and organic plastic fuels acting as binder to adhere oxidizer particles together. The ingredients often used as oxidizers are ammonium perchlorate (AP), ammonium nitrate (AN), ammonium dinitramide (ADN), cyclotrimethylenetrinitramine (RDX), and cyclotetramethylenetetranitramine (HMX). The most commonly employed binders are either inert (typically HTPB, hydroxyl-terminated polybutadiene, with various plasticizers, ballistic modifiers, and crosslinking agents), or active (NG and NC, polyether polymer, and azide polymer such as GAP glycidyl azide polymer, BAMO bis-azidomethyl oxetane, and AMMO 3-azidomethyl-3-methyl oxetane). The quest for more energetic propellants with reduced pollutant emissions has resulted in the use of several non-AP ingredients in solid propellants. The ingredients belong to a wide spectrum of chemical families, but mostly fall into one of four categories:

Nitramines (RDX, HMX, HNIW hexanitrohexaazaisowurtzitane also known as CL-20, HNF hydrazinium nitroformate)

Azides (GAP, BAMO, AMMO)

Nitrate esters (NG, NC, BTTN 1,2,4-butane triol trinitrate, TMETN metriol trinitrate, DEGDN diethylene glycol dinitrate)

Nitrates (ADN, AN).


Fig. 1.1 shows the molecular structures of the above propellant ingredients. Their material densities, heats of formation and adiabatic flame temperatures are given in Table 1.1. The material densities are typically in the range of 1.2-2.0 g/cm3. The heat of formation varies widely and is instrumental in determining the flame temperature and the total energy released during combustion. Given the wide variety of energetic materials and possible compositions, large test matrices for their detailed characterization are difficult, time consuming, and expensive to develop. Hence, new opportunities to advance the field of energetic materials increasingly rest on the predictive capability of combustion models.

Fig. 1.1. Molecular structures of RDX, HMX, ADN, BTTN, NG, TMETN, GAP, BMMO

Accurate models can elucidate the interplay between the chemical and physical phenomena and the resulting mechanisms that produce the observed burning behavior and combustion characteristics as functions of pressure, initial temperature, external stimuli, and propellant chemical formulation. This will help us verify the proposed chemical pathways and identify the chemical kinetics involved. Sensitivity analysis using established models allows a natural link between experimental and modeling efforts, and can be used to design experiments and to identify key reactions and species that require further theoretical study. The development of such models is a significant step

towards more accurate and comprehensive predictions, and helps optimize chemical compositions to meet the required needs. Thus, the necessary level of insight can be attained and successfully embodied in the predictive models for effective propellant design, development, and testing activities.

Table 1.1. Monopropellant ingredient properties

Understanding the thermal decomposition of energetic materials and their ensuing combustion characteristics is one of the central objectives of the modeling efforts. The combustion characteristics of concern include pressure and temperature sensitivities of burning rate, propellant surface conditions, and spatial distributions of energy release, temperature, and species concentration.

In the next chapter we will discuss about the physiochemical processes involved in the combustion of solid propellant and, also the difficulties and challenges involved in studying solid propellant combustion and flame structures.




Combustion of a solid propellant involves an array of intricate physiochemical processes evolving from the various ingredients that constitute the propellant. Most of the individual ingredients in solid-propellant formulations burn as monopropellants. We first consider the example of self-deflagrating RDX monopropellant in a stagnant environment. The entire combustion- wave structure can generally be segmented into three regions: (1) solid phase, (2) subsurface two-phase and (3) gas phase. The underlying physical processes in each of these regimes are illustrated in Fig. 2.1. During burning, the propellant remains thermally stable in the solid phase until the temperature reaches the melting point at which thermodynamic phase transition occurs. Molecular degradation and evaporation then takes place in the liquid layer, generating bubbles and forming a subsurface two phase region, also referred to as a foam layer. The ensuing products subsequently undergo a sequence of rapid decomposition in the near field immediately above the foam layer. The propellant- burning surface is defined here as an interface between the foam layer and the gas-phase region, at which rapid gasification or decomposition takes place. In the gas-phase region, the species emanating from the burning surface react with each other and/or decompose to form other species. A wide variety of oxidation reactions continue to occur and release an enormous amount of energy in the gas phase with the final temperature reaching the adiabatic flame temperature. The heat feedback from the exothermic reactions occurring in the gas phase along with the condensed-phase heat release sustains the combustion process.

The specific processes in the condensed and gas phases depend on the particular ingredient under consideration. For example, in the case of HMX monopropellant, a polymorphic phase transition occurs from β-HMX to δ-HMX at about 460K. For AP, a similar transition from an orthorhombic to a cubic structure occurs at 513K. Owing to the low liquefaction temperature (478K for RDX and 558K for HMX), the solid-phase reactions are usually neglected in comparison to the much faster liquid- and gas-phase reactions.

This situation holds true for most ingredients, except for AP (Tmelt~725-825 K), which undergo significant exothermic reactions in the solid phase. Both RDX and HMX monopropellant combustion exhibits a visible, definite foam layer at low and moderate pressures. The thickness of the foam layer and the gas-phase reaction zone vary with pressure.

Fig. 2.1. Schematic illustration of combustion-wave structure of RDX monopropellant;

self-sustained combustion



In order to accurately model all the burning characteristics of concern for solid propellants, a detailed description of the gas-phase, condensed phase, and surface mechanisms is essential. In the gas phase, a variety of chemical species undergo a

array of chemical reactions coupled with the processes of molecular diffusion, convection, conduction, and radiation. In the condensed phase, decomposition reactions and/or evaporation occur in the foam layer, along with subsequent reactions

in the embedded bubbles (see Fig. 2.1).

Different opinions exist about the relative importance between the gas- and condensed-phase heat release in dictating the combustion characteristics of a solid propellant. A recent work claims that condensed-phase reactions dominate the burning behaviours of many propellant ingredients, according to thermocouple measurements in the pressure range of 1-70 atm. Such a statement, however, remains to be clarified. At a pressure above ~10 atm, the flame stands so close to the burning surface that the spatial resolution of thermocouple measurement becomes insufficient to accurately explore the detailed physiochemistry near the surface. For example, Zenin reported a flame standoff distance of ~0.3mm for HMX at 20 atm using thermocouples, whereas Parr and Hanson-Parr measured a value of~0.12mm at 12 atm using spectroscopic techniques. Because a higher flame standoff distance implies a lower heat transfer from the flame zone to the condensed phase, the thermocouple measurements often underpredict, especially at high pressures, the role of gas-phase reactions in determining the propellant-burning rate. Thus, one must consider both the condensed and gas-phase processes in an integrated manner to provide a high-fidelity description of the combustion of propellant ingredients. At low pressures, the flame stands off relatively farther from the surface, and the condensed-phase heat release may dominate the burning rate. At high pressures, however, the temperature gradient in the gas phase near the surface becomes much steeper, and the residence time in the condensed phase decreases considerably due to a thinner melt layer. As a consequence, the gas-phase heat release dominates the propellant combustion at high pressures.

Uncertainties as to the gas-phase processes currently center on the decomposition of large molecules and the rate coefficients of certain reactions in the desired ranges of pressure and temperature. There is a general lack of fundamental understanding of the condensed-phase processes. The knowledge of subsurface reactions is limited; not only in terms of pathways and rate coefficients, but also with respect to the very identities of both reactants and products involved in the condensed-phase reactions.

In the following chapter we will discuss about the detailed characteristics of the monopropellants and the propellants and then infer the importance of the various mechanisms involved. The emphasis is on the qualitative effects of flame structure on combustion mechanisms.



The two oxidizers used most often in modern solid propellants are ammonium perchlorate, AP, and HMX (or RDX). AN (ammonium nitrate) is a third oxidizer that has been used on a limited basis, but studying its combustion characteristics can give insight to combustion mechanisms in general. The two most common binder systems consist of either an inert binder system based on a polybutadiene polymer, HTPB, with various plasticizers, additives and cross linking agents or an active binder system based on energetic components such as nitroglycerin, NG, and nitrocellulose, NC. In general, comparing the combustion characteristics of monopropellants with the characteristics that occur when the monopropellants are combined with other ingredients to form propellants leads to an understanding of the physical mechanisms that are involved.



The relative burning rates of HMX, AP, and AN as monopropellants and in composite

propellant mixtures are shown in Fig. 3.1. As monopropellants the rate of HMX is slightly higher than that of AP and both are significantly higher than that of AN. All three have very comparable burn rate exponents (for typical solid propellant rocket pressures). The adiabatic flame temperatures are very different; for HMX ~3200 K; for AP ~1400 K; and for AN ~1247 K.

Although AP and HMX have very different flame temperatures, they have very similar combustion characteristics. This is apparently due to a higher reactivity of the chlorine oxides (the surface products from the condensed phase) of the AP compared to the nitrogen oxides (i.e. the typical surface products) of HMX. This observation is supported by the much lower rate of AN which has a flame temperature comparable to AP but does not give off the reactive chlorine oxides. The decomposition products from AN are more comparable to those of HMX, and one would expect similar reaction kinetics to be involved with AN and HMX, but not with AP. Therefore, the much lower rate of AN would correspond to the much lower flame temperature of AN compared to that of HMX.

AP and HMX appear to have similar monopropellant combustion characteristics, but when mixed with a fuel binder into a composite propellant, the resultant burn rates vary by as much as an order of magnitude. The burn rate of AP composite propellants exhibits a very strong particle size dependence.

Fig. 3.1. Burning rate characteristics of composite propellants

In contrast, the burn rates of HMX composite propellants are much lower than AP propellants, and show very little particle size dependence. Although the HMX monopropellant rate is approximately an order of magnitude greater than that of AN, both have comparable rates when mixed into composite propellants, and neither show as much particle size dependence as AP propellants.

Thermochemical calculations for AP composite propellants indicate a primary flame temperature on the order of ~2500-2800 K for typical concentrations. Thus, the AP diffusion flame is much more energetic than the AP monopropellant flame and leads to higher burning rates. HMX composite propellants have adiabatic flame temperatures on the order of 2000 K implying a less energetic diffusion flame than the monopropellant flame. Apparently the HMX diffusion flame robs energy from the monopropellant flame, suppressing the overall propellant burn rate. AN composite propellants have adiabatic flame temperatures on the order of ~1500- 2000 K. The AN diffusion flame is more energetic than the AN monopropellant flame, but apparently not enough to cause a significant increase in burn rate (most likely due to the lower reactivity of nitrogen oxides compared to the chlorine oxides from AP).

Because of the very large differences in burning rate of the different composite propellants it appears that the dominant mechanism in the propellant combustion must be related to the primary diffusion flame (i.e. as opposed to the monopropellant flame).


Double base propellants burn with characteristics significantly different than composite propellants. Figure 3.2 contains typical double base propellant burn rates which are compared to AP and HMX monopropellant burning rates. Double base propellant burn rates correlate very well with their heat of explosion (Hex); the more energetic propellants having the greater burn rates. Low energy propellants have rates lower than AP or HMX and high energy propellants have rates greater than with AP or HMX. The energy content of HMX and double base are comparable and both contain the same elements, CHON. Thus, it is not surprising that their rates are comparable. However, a double base propellant with an energy level equivalent to HMX will have a slightly higher rate than HMX. Apparently the composition of the DB surface products is more reactive than those from HMX. A key ingredient in HMX surface products that is not present in DB products is HCN. It could be that the reaction paths associated with HCN

may be related to the slightly lower reactivity of the HMX. In contrast, the lower energy content of AP gives rates comparable to either DB or HMX, apparently due to the more reactive chlorine oxides as discussed above.

Many modern double base propellants contain AP, HMX or a combination of the two oxidizers. Adding AP to a double base binder causes a significant effect on the burn rate due to the significant differences in the AP and double base chemistry. Thermochemical calculations varying the percent AP with double base binders of varying Hex show that the flame temperature is a very strong function of binder Hex and the percentage AP. For a low energy binder the flame temperature increases with the addition of AP, with a peak temperature occurring at about 65% AP. For a high energy binder the flame temperature increases only slightly, with a peak temperature at approximately 30% AP.

The addition of AP to double base binders causes a definite increase in burn rate with smaller particles causing a greater increase than larger particles. The very reactive products from the AP diffuse into the stream of double base decomposition products reacting in a primary diffusion flame analogous to the primary flame in composite propellants. As in composite propellants, adding AP can increase the propellant rate higher than the inherent rate of either the binder or the AP.

Adding HMX to a double base binder results in minimal changes in the burn rate. Data shows that adding different sized HMX particles at concentrations up to ~60% did not significantly alter the burn rate. This is not surprising considering that both HMX and double base binder each contain CHON elements and similar energy levels. Calculated flame temperatures varying the percent HMX for double base binders of

Fig. 3.2. Burning rate characteristics of double base propellants

Calculated flame temperatures varying the percent HMX for double base binders of varying Hex show for high energy binder there is little change in the flame temperature of the propellant. For low energy binder adding HMX increases the flame temperature. The calculated temperatures all increase towards the adiabatic flame temperature of HMX which is higher than any of the double base flame temperatures.

These observations make it apparent that the dominant combustion mechanism is the burning rate of the double base binder, and that the HMX does not contribute significantly to the propellant burn rate. Stoichiometrically, HMX is not a true oxidizer and decomposition products from HMX and double base are very similar; neither contains a powerful oxidizing species such as perchloric acid. It appears that the diffusion flame between the HMX and binder is very similar to the individual monopropellant flames, and causes small changes in the burn rate. This is very different from the AP/double base diffusion flame, wherein the AP primary diffusion flame is a dominant combustion mechanism competing with the double base monopropellant flame for control of the burn rate. The effect of the AP diffusion flame is related to both the increased temperature, but is probably more dependent on the increased reactivity of the AP surface products.


The initial decomposition reactions begin in the condensed phase. Most ingredients actually melt or go through a molten phase during combustion. Depending on the heat feed back rate from the flame, the condensed phase reactions continue until the reaction products are gaseous. Some of these reactions will be heterogeneous reactions, where gaseous, intermediate decomposition products react with the thin, molten, liquid phase. Gas phase decomposition products leave the condensed surface at the surface temperature and enter the gas phase as a boundary condition for the processes leading to the flame and the final combustion processes. The contribution of the condensed phase to the combustion process is to provide a concentration of reactive, gaseous intermediates at a given temperature (the surface temperature) to the gas phase which can then react in the gas flame. Depending on the initial pressure or temperature, the composition and temperature of the surface products will vary, and will provide a variation in the rate of reaction in the flame. Thus, although the condensed phase reactions are not the controlling mechanism in the combustion process, they do have a significant influence on burn rate, pressure exponent and temperature sensitivity.


Surface temperature is usually measured with fine thermocouples. This is difficult to do especially in crystalline monopropellants such as AP or HMX. However, a large number of investigators have obtained measurements of surface temperature for double base propellants, but only a few have obtained data in a systematic manner varying pertinent variables, such as pressure, initial temperature or propellant composition. The data from those systematic studies have been correlated and are plotted in Fig. 3.3. Other studies where the data are very limited have not been included in the correlations, but it is not likely that they would make a significant change in any of the conclusions. The surface temperature values for double base fall in the general range of 450 to 725 K, and increasing with increasing pressure.

Although data for pure AP are virtually nonexistent, various sources of surface temperature data for AP in composite propellants were found in the literature and used for a basis of comparison. The extensive data of Powlingll, using an IR detector to measure the surface temperature of AP composite propellants, appeared to be the most consistent, and those data have been included in Fig. 3.3. It is also very apparent that most of the data were obtained at burning rates and surface temperatures well below those of normal interest. AP has the highest surface temperature of the monopropellants considered with values measured up to 900 K.

Fig. 3.3. Comparison of calculated surface temperature and data

Mitani and Lengelle have reported surface temperatures for HMX (pressed with a small amount of fuel) which have been included in Fig. 3.3. The HMX data indicate values of 800 to 900 K, essentially parallel to the AP data. It is significant that the burn rate catalyst did not appear to have a significant influence on the surface decomposition characteristics. This seems to verify that the gas phase reaction is the controlling mechanism, and the surface decomposition simply accommodates the heat flux from the gas. The measured and calculated surface temperatures for AN vary between 500 and 600 K which is in the same range as double base propellants, but significantly lower than AP or HMX.


Some researchers that measured surface temperatures, also reduced their data to infer the condensed phase heat release, Qc particularly for double base propellant. It has been observed that Qc is proportional to the reaction temperature and decreases with increasing initial temperature. It would seem consistent that the amount of energy released would be related to surface temperature rather than pressure.

AN is a relatively simple compound due to the fact that it only contains three atoms, HON. This simplicity allows an estimation of the value of Qc. If it is assumed that free radicals are probably short lived in the condensed phase, then the number of intermediate products involved in the condensed phase reactions is probably limited. To estimate a value of Qc for AN, thermochemical calculations were made that correspond to five possible nitrogen oxidation states, as nitrogen is reduced from the nitrate ion, through the various nitrogen oxides, to N2. The results are summarized below where a heat of reaction and adiabatic reaction temperature are recorded for each oxidation state.

Table 3.1. AN Monopropellant Reactions

The initial decomposition of AN to HNO3 and NH3, and the subsequent reaction leading

to NO2 are both very endothermic and do not lead to an adiabatic reaction temperature. These are most likely the initial reactions occurring in the condensed phase. The reactions leading to NO and N20 are both exothermic with calculated reaction temperatures of approximately 600 K, which is the nominally measured surface temperature of burning AN. It is significant to note that the melting temperature of AN is 443 K. Therefore, it would appear that the nitrate is reduced at least to NO in the condensed phase and probable some fraction of the material reacts to N20, which is the predominant nitrogen product leaving the surface. Based on these calculations, Qc values of the order of 80 to 100 cal/gm, exothermic, should be expected for AN.

These basic processes will be similar for different propellant ingredients, whether oxidizer or binder. The actual composition and temperature of the products of the condensed phase will differ for different ingredients and conditions. The associated energy release will also vary for different ingredients.


Most ingredients used in solid propellants will actually burn as monopropellants (as

shown in Figs. 3.1 and 3.2). Figure 3.4 is a schematic of the combustion mechanisms and thermal profile involved in monopropellant combustion. As illustrated, the region of condensed phase reactivity can be relatively thick, especially at low pressures when the thermal wave penetrates deeply into the solid. The thickness of the condensed phase reaction zone is typically the same order of magnitude as the gas phase flame standoff distance. The decomposition products of the condensed phase reactions leave the surface and react in the gas phase to form a premixed, monopropellant flame. Because of the premixed nature of the flame, the thermal profile is very pressure dependent, as illustrated in the figure. The specific temperatures in the figure correspond to AP, but the general concepts apply to any monopropellant. The energy release has been measured for double base propellant. The energy release apparently occurs in a rather thick flame, but very close to the surface. Flame stand-off distances are on the order of 100 pm at low pressures decreasing to 10 pm or less at higher pressures. The actual energy release zone or 'flame' is even smaller. The very small dimensions make experimental measurements of flame zones very difficult.

Double base propellants and, to a lesser extent, HMX both exhibit a two stage flame. For double base propellants the inner flame is ~1200 to 1700 K depending on both pressure and binder energy. A dark zone separates the inner flame from a final flame of ~1500 to 3000 K, again depending on pressure and binder energy. The effect of the heat transfer from the flame on the burning rate of the material is dominated by the inner flame, which in turn, dominates the burning rate.

At high pressure (~ 100 atm) the two reaction zones merge, and only one flame is observed at higher pressures. The inner flame is apparently characterized by the reaction of NO2 to form NO. The NO is relatively stable, leading to the dark zone. The stability of the NO molecule has also been observed in studies of air pollution. The end of the dark zone is characterized by the reaction of NO to form N20 and ultimately N2. However, there is no observed dark zone and the inner zone disappears at pressures slightly above atmospheric.

Fig. 3.4. Monopropellant combustion mechanisms and thermal profile

These basic processes will be similar for different propellant ingredients, whether oxidizer or binder. The precise composition and temperature of the products will differ for different ingredients and for different conditions. The corresponding energy release will also vary. However, many of the actual reaction steps in the gas phase will be the same for differing ingredients. For example, the very important gas phase reactions involving NO2, NO, and N20 will be the same irrespective of the source. Only the composition and temperature of the reactants will vary for different ingredients. Thus, it should be possible to establish general reaction schemes that should be common for many ingredients.


Once the oxidizing and fuel species have diffused together, they can react in a diffusion

flame or the oxidizer can still react as a monopropellant., both flames can coexist. Trying to imagine what the flame structure is for composite propellants, varying pressure, oxidizer particle sizes, and everything else that can be involved in a propellant, can be a real problem.

Speculating on the flame structure of a composite propellant can be analogous. Different researchers have examined different aspects of combustion or of the flame structure, but none of us can actually 'see' what a flame looks like in an actual propellant environment. Figure 3.5 is a schematic of the solid propellant flame structure that has been proposed in the past. The concepts that are illustrated in Fig. 3.5 are specific to an AP composite propellant, but can be applied in general to most propellant types and ingredients. The figure illustrates the reaction zones that must be considered in order to understand the very complex combination of combustion mechanisms that occur in propellant combustion. However, all of these characteristics are speculative.

Fig. 3.5. Solid propellant flame structure and combustion mechanisms

The interactions between the different flames is critical in understanding the combustion

mechanisms, especially the pressure dependence. The diffusion flame is a dominant

mechanism (especially for strong oxidizers such as AP), and therefore, understanding diffusion flames is extremely important to understand the various mechanisms in propellant combustion. Even with weak oxidizers such as HMX or AN, the primary diffusion flame is a significant mechanism in the combustion process. In the case of HMX, the energy release in the primary flame can be a very reduced energy level from that of the HMX monopropellant flame. Thus, in the case of HMX composite propellants, the primary diffusion flame can actually 'rob' energy from the combustion process, reducing the effective energy release and the corresponding burning rate.


Previous diffusion flame analyses used in solid propellant modeling have typically been

based on Burke-Schumann type-models, and have assumed an infinite reaction rate in order to achieve an analytic solution to the problem. In actuality, the kinetic rates between the fuel and oxidizer are finite and need to be included as such. Although the geometry of a pure diffusion flame would be independent of pressure, the kinetic aspects of a practical diffusion flame are pressure dependent. Thus, pressure dependent combustion characteristics can be related to the characteristics of the diffusion flame. Also, in considering the interaction of acoustics or cross flow with the propellant flame structure, it is apparent that both will interact primarily with the final diffusion flame. Thus, although the final diffusion flame is usually of a secondary importance in determining the steady state burning rate, it can become a primary factor in determining erosive burning or the unstable combustion response.


When an oxidixer is incorporated into a propellant environment, the oxidizing species can react either in a monopropellant flame or in a diffusion flame with the fuel species from the binder. The trade-off between how these flames are established for varying conditions and how they interact, is critical to the understanding of propellant combustion. This not only requires an understanding of diffusion flames, including finite kinetics, but also of the interaction of the diffusion flame with the premixed monopropellant flame. The complex flame structure that evolves will also have a complex thermal environment associated with it.

Figure 3.6 contains a schematic of the temperature profiles that can exist in a typical propellant environment. At the edge of a crystal the surface sees a very hot diffusion flame with a rapidly rising temperature. This is essentially the flame profile that an embedded thermocouple might be exposed to. Slightly in from the edge of the crystal, the thermal profile could go through the monopropellant flame and then a diffusion flame. At the center of a crystal the thermal profile could again go through the monopropellant flame, probably having an extended distance of little activity and then a final diffusion flame. The profile would look much like the two stage flame of double base propellants. A thermocouple would not be able to measure a profile like this unless it could be embedded in the crystal. Data from thermocouples have been, and continue to be, an important source of information concerning the flame structure of burning propellants. However, care must be taken to recognize the limits of thermocouple measurements, especially the fact that they can only be placed in the binder, and

their measurements will always reflect the thermal profile at the edge of a crystal, but not the profile at the center of a crystal.

Fig. 3.6. Thermal profiles that can exist over a single burning crystal of AP

The primary diffusion flame can have a higher temperature than the monopropellant

flame, as illustrated for AP, or it can have a lower temperature, as would occur with HMX. In either case, the temperature profiles are very complex and will vary across the surface of an oxidizer crystal. The flame standoff distances of the flames relative to each other will also vary with particle size and with pressure, thus having a significant effect on the burning rate and the pressure exponent.


Because of the complexity of the flame structure, interpreting the effects of pressure on

the interaction of the monopropellant and diffusion flames is a very difficult, but important task if one is to understand how the flame structure changes with pressure. Figure 3.7 illustrates how the complexity of the flame structure may change for varying pressure.

Fig. 3.7. Pressure effects on propellant flame structure

At low pressure the monopropellant flame can have a standoff distance greater than the

diffusion flame, thus precluding the formation of a monopropellant flame. However, at

increasing pressure, the monopropellant flame can move closer to the propellant surface, minimizing the diffusion flame effects. This is due to the strong pressure dependence of the premixed flame versus the weak pressure dependence of the diffusion flame (due to the kinetic aspects of the flame). The changes with pressure obviously effect the characteristics of the pressure exponent, but also have a direct impact on the propellant's response to pressure perturbations (i.e. its response to combustion instability). A similar change in flame structure can result from changes in particle size, small particles having a structure similar to the low pressure structure, while larger particles will have a flame structure more similar to high pressure.


As a particle burns the amount of fuel binder that is available to burn in the diffusion

flame changes, Therefore, interpreting the effects of these geometrical changes on the

interaction of the monopropellant and diffusion flames.

Fig.3.8. Changes in propellant flame structure during the geometrical cycle of a burning particle

Figure 3.8 illustrates the geometrical change in flame structure that can occur as a particle burns. When the particle is first exposed to the combustion atmosphere, it is in a very fuel rich environment due to the fact that the particle is surrounded by fuel and only a small amount of oxidizer is initially exposed to the combustion front. Because of the fuel rich environment the initial flame that is established as the particle ignites, it most likely a diffusion flame. There is probably too much fuel available for the monopropellant flame to be established.

As burning progresses and more of the oxidizer is exposed, there is less fuel available

and the monopropellant can become established consuming part of the oxidizer decomposition products. This allows the classical flame structure to develop as shown in part 2 of the figure. Ultimately, the primary flame will become very oxidizer rich due to the very thin layer of fuel that surrounds the oxidizer. In this case the flame will bend over the fuel, causing the mushroom shaped structure shown in part 3 of the figure. These changes in flame structure will also depend strongly on particle size, small particles having a structure similar to the structure in part 1 of the figure, while the flame structure shown in part 3 will persist for most of the life time of larger particles. As burning continues the process will reverse as the availability of fuel changes going through a flame structure similar to part 2 and part 1 of the figure. As the particle burns

out it will most likely revert to the diffusion flame structure of part 1.




Upto now we have studied about the various compounds used in a solid propellant and discussed in detail the physiochemical processes involved in the combustion of solid propellants and the challenges and difficulties involved in studying solid propellant combustion and flame structures.

We also studied in detail the combustion mechanisms and burning characteristics of Monopropellants, Double-base propellants, and composite propellants and also compared the burning characteristics of a Monopropellant with Double-Base propellants and Composite propellants. Also the condensed phase characteristics of Monopropellants, Double-Base propellants, and Composite Propellants were studied in detail. Apart from the above the Monopropellant flame structure and the Composite propellant flame structure was studied in detail considering atmospheric and situational conditions that can affect the combustion of the solid propellants.

Further, for the next stage we will focus on the conditions and situations that affect the burning rate and how burning rate can be modified.