This is the second of four chapters dealing with solid propellant rocket motors. Here we describe several common solid rocket propellants, their principal categories, ingredients, hazards, manufacturing processes, and quality control. We also discuss liners and insulators, propellants for igniters, propellant tailoring, and propellants for gas generators.

Thermochemical analyses are needed to characterize the performance of any given propellant and specific methods are described in Chapter 5. Such analyses provide values for the effective average molecular mass, combustion temperature, average specific heat ratio, and characteristic velocity—these are all functions of propellant composition and chamber pressure. Specific impulses can also be computed for given nozzle configurations and exhaust conditions.

The term solid propellant has several connotations, including: (1) the rubbery or plastic‐like mixture of oxidizer, fuel, and other ingredients that have been processed (including curing) and constitute the finished grain; (2) the processed but uncured product; (3) a single ingredient, such as the fuel or the oxidizer. In this field, acronyms and chemical symbols are used indiscriminately as abbreviations for propellant and ingredient names; only some of these will be shown in this chapter.

13.1 CLASSIFICATION

Historically, the early rocket motor propellants used to be grouped into two classes: double‐base (DB)¹ propellants were the first production propellants and subsequently the development of polymers as binders made the composite propellants feasible. Processed modern propellants are more finely classified as described below. Such classifications are helpful but they are neither rigorous nor complete. Sometimes the same propellant will fit into two or more classifications.

1. Propellants are often tailored to and classified by specific applications, such as space launch booster propellants or tactical missile propellants, each having specific chemical ingredients, different burning rates, different physical properties, and different performance. Table 12–1 shows four rocket motor applications (each with somewhat different propellants), plus several gas generator applications and an artillery shell application. Propellants for rocket motors produce hot (over 2400 K) gases and are used for thrust, but gas generator propellants operate with lower‐temperature combustion gases (800 to 1200 K in order to use uncooled hardware) and are used to produce power, not thrust.
2. Double‐base (DB)¹ propellants form a homogeneous propellant grain, usually a nitrocellulose (NC)¹—a solid ingredient that absorbs liquid nitroglycerine (NG), plus minor percentages of additives. The major ingredients are highly energetic materials and they contain both fuel and oxidizer. Both extruded double‐base (EDB) and cast double‐base (CDB) propellants have found extensive applications, mostly in small tactical missiles of older design. By adding crystalline nitramines (HMX or RDX)¹ both performance and density can be improved; these are sometimes called cast‐modified double‐base propellants. Adding an elastomeric binder (rubber‐like, such as crosslinked polybutadiene) further improves the physical properties and allows more nitramine and thus increasing performance slightly. The resulting propellant is called elastomeric‐modified cast double‐base (EMCDB). These four classes of double‐base propellants have nearly smokeless exhausts. Adding some solid ammonium perchlorate (AP) and aluminum (Al) increases the density and the specific impulse slightly, but exhaust gases becomes smoky—such propellant is called composite‐modified double‐base propellant or CMDB.

   Two operational systems produced by ATK that use double‐based propellants are the AGM‐114 Hellfire (which uses XLDB, a minimum smoke crosslinked propellant) and the Hydra 70 rocket (with a plateau burning propellant, see Fig. 12–6).

3. Composite propellants form a heterogeneous propellant grain between oxidizer crystals and powdered fuel (usually aluminum) held together in a matrix of synthetic rubber (or plastic) binder, such as polybutadiene (HTPB).¹ Composite propellants are cast from a mix of solid (AP crystals, Al powder) and liquid (HTPB, PPG)¹ ingredients. The propellant is hardened by crosslinking or curing the liquid binder polymer with a small amount of curing agent, and curing it in an oven, where it becomes solid. In the past four decades composites have been the most commonly used class of propellant. Composites can be further subdivided:
   1. Conventional composite propellants, which usually contain between 60 and 72% AP as crystalline oxidizer, up to 22% Al powder as a metal fuel, and 8 to 16% of elastomeric binder (organic polymer) including its plasticizer.
   2. Modified composite propellant where an energetic nitramine (HMX or RDX) is added for obtaining some added performance and also a somewhat higher density.
   3. Modified composite propellant where an energetic plasticizer such as nitroglycerine (used in double‐base propellants) is added to give increased performance. Sometimes HMX is also added.
   4. High‐energy composite solid propellant (with added aluminum), where the organic elastomeric binder and the plasticizer are largely replaced by highly energetic materials and where some of the AP is replaced by HMX and RDX. Hexanitrohexaazaiso‐wurtzitane or CL‐20 is a recent propellant ingredient being used; it is produced outside of the United States (see Section 13–4 and Ref. 13–1). Some propellants are called elastomer‐modified cast double‐base propellants (EMCDB). Most are experimental propellants. Their theoretical specific impulse can be between 250 and 275 sec at standard conditions as explained below.
   5. Lower‐energy composite propellant, where ammonium nitrate (AN) is the crystalline oxidizer (not AP). These are used for gas generator propellants. When large amounts of HMX are added, they become minimum smoke propellants with fair performance.
4. Propellants may also be classified by the smoke density in the exhaust plume as smoky, reduced smoke, or minimum smoke (essentially smokeless). Aluminum powder, a desirable fuel ingredient for performance, is oxidized to aluminum oxide during burning, which yields visible, small, solid smoky particles in the exhaust gas. Most composite propellants (e.g., AP) are also smoky. By replacing AP with HMX and RDX and by using energetic binders and plasticizers to compensate for eliminating aluminum, the amount of smoke may be considerably reduced in composite propellants. Carbon (soot) particles and metal oxides, such as zirconium oxide or iron oxide, are also visible in high enough concentrations. This is further discussed in Chapter 20.
5. Safety ratings for detonation can distinguish propellants as a potentially detonable material (class 1.1) or as a nondetonable material (class 1.3), as described in Section 13.3. Examples of class 1.1 propellant are double‐base propellants and composite propellants containing a significant portion of a solid explosive (e.g., HMX or RDX), together with certain other ingredients.
6. Propellants can be classified by some of the principal manufacturing processes used. A cast propellant is made by mechanically mixing solid and liquid ingredients, followed by casting and curing; it is the most common process for composite propellants. Curing of many cast propellants takes place through a chemical reaction between binder and curing agent at above ambient temperatures (45 to 150°C); however, there are some that can be cured at ambient temperatures (20 to 25°C) or hardened by nonchemical processes such as crystallization. Propellants can also be made by a solvation process (dissolving a plasticizer in a solid pelletized matrix, whose volume is expanded). Extruded propellants are made by mechanical mixing (rolling into sheets) followed by extrusion (pushing through a die at high pressure). Solvation and extrusion processes are applied primarily to double‐base propellants.
7. Propellants have also been classified by their principal ingredient, such as the principal oxidizer (ammonium perchlorate propellants, ammonium nitrate propellants, or azide‐type propellants) or their principal binder or fuel ingredient, such as polybutadiene propellants or aluminized propellants. This classification of propellants by ingredients is further described later in Section 13.4 and Table 13–8.
8. Propellants with toxic and nontoxic exhaust gases are discussed in more detail in Section 13.3.
9. Experimental and/or production propellants. Such propellants are selected after extensive testing (preflight test, qualification tests) and demonstrated safety, life, and other essential properties. The culmination of any successful research and development (R&D) program is to have a propellant selected for production in a flight vehicle application.

Figures 13–1 and 13–2 show general regions for the specific impulse, burning rate, and density for the more common classes of propellants. The ordinate in these figures is an actual or estimated specific impulse at standard conditions (1000 psi chamber pressure and expansion to sea‐level pressure). These results do not reflect pressure drops in the chamber, nozzle erosion, or combustion losses and scaling assumptions. Composite propellants are shown to have a wide range of burning rates and densities; most of them have specific gravities between 1.75 and 1.81 and burning rates between 7 and 20 mm/sec. Composite propellants give higher densities, specific impulse, and a wider range of burning rates than others. Table 13–1 lists performance characteristics for several propellants. DB propellants and AN propellants have lower performance and density. Most composite propellants display similar performance and density but with a wider range of burning rates. The highest performance indicated is for a CMDB propellant whose ingredients are identified as DB/AP‐HMX/Al, though it is only 4% higher than the next.

Several of the above listed classifications tend to be confusing. The term composite‐modified double‐base propellant (CMDB) has been used for a DB propellant, where some AP, Al, and binder are added; alternatively, the same propellant could be classified as a composite propellant to which some double‐base ingredients have been added.

A large variety of chemical ingredients and propellant formulations have been synthesized, analyzed, and tested in experimental rocket motors. A typical solid propellant has between 4 and 12 different ingredients and this chapter discusses about a dozen basic propellant types, but many other types are still being investigated. Table 13–2 presents some advantages and disadvantages of selected propellant classes. Representative formulations for three types of propellant are given in Table 13–3. In actual practice, each propellant manufacturer uses its own formulation and processing procedures. The exact percentages of ingredients, even for a given propellant such as PBAN, can not only vary among manufacturers but often vary from one rocket motor application to another. The practice of adjusting mass percentages together with adding or deleting one or more of the minor ingredients (additives) is known as propellant tailoring. Tailoring is the practice of taking an established propellant and changing it slightly to fit a new application, different processing equipment, altered motor ballistics, storage life, temperature limits, and/or even a change in ingredient source.

New propellant formulations are typically developed using laboratory‐size mixers, curing ovens, and related equipment with the propellant mixers (1 to 5 liters), and operated by remote control for safety reasons. Process studies usually accompany the development of new propellant formulations to evaluate the “processability” and to guide the design of any special production equipment needed in preparing ingredients, mixing, casting, or curing it.

Historically, black or gun powder (a pressed mixture of potassium nitrate, sulfur, and an organic fuel such as ground peach stones) was the very first propellant to be used. Many other types of propellants ingredients have been used in experimental motors, including fluorine compounds, propellants containing powdered beryllium, boron, hydrides of boron, lithium, or beryllium, or new synthetic organic plasticizer and binder materials with azide or nitrate groups. Most of these have not yet been considered satisfactory or practical for production in rocket motors.

13.2 PROPELLANT CHARACTERISTICS

Propellant selection is critical to rocket motor design. Desirable propellant characteristics are listed below and further discussed in other parts of this book. Many requirements for particular solid propellant rocket motors will influence priorities for choosing these characteristics:

1. High performance or high specific impulse; this implies a high gas temperature and/or low exhaust gas molecular mass.
2. Predictable, reproducible, and initially adjustable burning rate to fit grain‐design needs and thrust‐time requirements.
3. For minimum variations in thrust or chamber pressure during burning, both the pressure or burning rate exponent and the temperature coefficient should be small.
4. Adequate physical properties (including bond strengths) over the intended operating temperature range with allowance for some degradation due to cumulative damage.
5. High density (resulting in a small‐volume rocket motor).
6. Predictable, reproducible ignition qualities (such as acceptable ignition overpressures).
7. Desirable aging characteristics and long life. Aging and life predictions depend on the propellant's chemical and physical properties, cumulative damage criteria with load cycling (see Section 12.4) and thermal cycling, and from actual tests on propellant samples and test data from failed motors.
8. Low moisture absorption, because moisture often causes chemical deterioration.
9. Simple, reproducible, safe, low‐cost, controllable, and low‐hazard manufacturing.
10. Guaranteed availability of all raw materials and purchased components over the production and operating life of the propellant, and acceptable control over undesirable impurities.
11. Low technical risk, such as a favorable history of prior applications.
12. Relative insensitivity to certain external energy stimuli as described Section 13.3, the hazards section.
13. Nontoxic and noncorrosive exhaust gases, also called green exhausts.
14. Not prone to combustion instability (see Chapter 14).
15. Equivalent composition, performance and properties with every new propellant batch.
16. No slow or long‐term chemical reactions or migrations between propellant ingredients or between propellant and insulator/liner.

Some of these desirable characteristics will also apply to all materials and purchased components used in solid rocket motors, such as the igniter, insulator, case, or safe‐and‐arm device but several of these characteristics can sometimes be in conflict with each other. For example, increasing the physical strength (more binder and or more crosslinker) will reduce propellant performance and density. So a modification of the propellant for one of these characteristics may cause changes in a few others.

Several illustrations will now be given on how characteristics of a propellant change when the concentration of one of its major ingredients is changed. Figure 13–3 shows calculated variations in combustion or flame temperature, average product gas molecular mass, and specific impulse as a function of oxidizer concentration for composite propellants that use a polymer binder [hydroxyl‐terminated polybutadiene (HTPB)] and various crystalline oxidizers; these results are taken from Ref. 13–4, based on a thermochemical analysis as explained in Chapter 5. The maximum values of and occur at approximately the same concentration of oxidizer. For practical reasons, this optimum percentage for AP (about 90 to 93%) and AN (about 93%) cannot be implemented because concentrations greater than about 90% total solids (including the aluminum and solid catalysts) cannot be processed in a mixer—a castable slurry that will flow into a mold requires more than 10 to 15% liquid content.

A typical composition diagram for a composite propellant is depicted in Fig. 13–4. It shows how the specific impulse varies with changes in the composition of the three principal ingredients: the solid AP, solid Al, and viscoelastic polymer binder.

For DB propellants variations of and are shown in Fig. 13–5 as a function of nitroglycerine (NG) concentration. The theoretical maximum specific impulse occurs at about 80% NG. In practice, NG, which is a liquid, is seldom found in concentrations over 60% because its physical properties are poor at the higher concentrations. Other major solid or soluble ingredients are also needed to make a usable DB propellant.

For CMDB propellants the addition of either AP or a reactive nitramine such as RDX allows for higher than with ordinary DB (where AP or RDX percent is zero), as shown in Fig. 13–6. Both AP and RDX greatly increase the flame temperature and thus make the effects of heat transfer more critical. The maximum values of occur at about 50% AP and at 100% RDX (an impractical propellant concentration because it cannot be manufactured and will not have reasonable physical properties). At high concentrations of AP or RDX, the exhaust gases contain considerable H₂O and O₂ (as shown in Fig. 13–7); these enhance erosion rates in carbon‐containing insulators or nozzle materials. In Fig. 13–7, the toxic HCl gas is present in concentrations between 10 and 20%, but in practical propellants it can seldom exceed 14%.

Nitramines such as RDX or HMX contain relatively few oxidizing radicals and the binder surrounding the nitramine crystals cannot be fully oxidized. As the binder decomposes at the combustion temperature, it releases gases rich in hydrogen and carbon monoxide (which reduce the molecular mass), cooling the gas mixture to lower combustion temperatures. The exhaust gases of AP‐based and RDX‐based CMDB propellant are shown in Fig. 13–7 where it can be seen that solid carbon particles disappear when the RDX content is above 85%.

13.3 HAZARDS

With proper precautions, training and equipment, all common propellants can be manufactured, handled, and fired safely. It is necessary and imperative to fully understand all hazards and methods for preventing dangerous situations from arising. Each material has its own set of hazards; some of the more common ones are described briefly below and also in Refs. 13–6 and 13–7. Not all apply to every propellant.

13.4 PROPELLANT INGREDIENTS

A number of relatively common propellant ingredients are listed in Table 13–6 for double‐base propellants and for composite‐type solid propellants in Table 13–7. They are categorized by major function, such as oxidizer, fuel, binder, plasticizer, curing agent, and so on, and each category is briefly described later in this section. However, several ingredients can have more than one function. These lists are not complete because over 200 other ingredients have been tried in experimental rocket motors. Most ingredients in Table 13–6 have a shelf life of over 30 years.

Table 13–8 shows a classification for modern propellants (including some new types that are still in the experimental phase) according to their binders, plasticizers, and solid ingredients; these solids may be an oxidizer, a solid fuel, or a combination or a compound of both.

Ingredient properties and impurities both may have a profound effect on propellant characteristics. A seemingly minor change in one ingredient can cause measurable changes in ballistic properties, physical properties, migration, aging, and/or ease of manufacture. When the propellant's performance or ballistic characteristics have tight tolerances, ingredient purity and properties must conform to equally tight tolerances and careful handling (e.g., no exposure to moisture). In the remainder of this section, a number of the important ingredients, grouped by function, are briefly discussed.

Inorganic Oxidizers

Some thermochemical properties of several oxidizers and radical‐containing oxygen compounds are listed in Table 13–9. Listed values depend on the chemical nature of each ingredient.

Designation	Binder	Plasticizer	Solid Oxidizer and/or Fuel	Propellant Application
Double base, DB	Plasticized NC	NG, TA, etc.	None	Minimum signature and smoke
CMDBa	Plasticized NC	NG, TMETN, TA, BTTN, etc.	Al, AP, KP	Booster, sustainer, and spacecraft
	Same	Same	HMX, RDX, AP	Reduced smoke
	Same	Same	HMX, RDX, azides	Minimum signature, gas generator
EMCDBa	Plasticized NC + elastomeric polymer	Same	Like CMDB above, but generally superior mechanical properties with elastomer added as binder
Polybutadiene	HTPB	DOA, IDP, DOP, DOA, etc.	Al, AP, KP	Booster, sustainer, or spacecraft; used extensively in many applications
	HTPB	DOA, IDP, DOP, DOA, etc.	AN, HMX, RDX, some AP	Reduced smoke, gas generator
	CTPB, PBAN, PBAA	All like HTPB above, but somewhat lower performance due to higher processing viscosity and consequent lower solids content. Still used in applications with older designs.
Polybutadyene with HMX or RDX	HTPB	DOA, IDP, DOP, DOA, etc.	AP, Al, HMX, or RDX	High energy vehicles
Polyether and polyesters	PEG, PPG, PCP, PGA, HTPE,b and mixtures	DOA, IDP, TMETN, DEGDN, etc.	Al, AP, KP, HMX, BiO3	Booster, sustainer, or spacecraft
Energetic binder (other than NC)	GAP, PGN, BAMO/ NMMO, BAMO/AMMO	TMETN, BTTN, etc. GAP‐azide, GAP‐nitrate, NG	Like polyether/polyester propellants above, but with slightly higher performance. Experimental propellant.

^a CMDB, composite‐modified double‐base; EMCDB, elastomer‐modified cast double‐base. For definition of acronyms and abbreviations of propellant ingredients see Tables 13–2 and 13–3.

^b HTPE, hydroxyl‐terminated polyether, binder for propellant developed for Insensitive Munitions by Orbital ATK.

Oxidizer	Chemical Symbol	Molecular Mass (kg/kg‐mol)	Density (kg/m3)	Total Oxygen Content (mass%)	Available Oxygen Content (mass%)	Remarks
Ammonium perchlorate	NH4ClO4	117.49	1949	54.5	34.0	Low n, low cost, readily available, high performance
Potassium perchlorate	KClO4	138.55	2519	46.2	40.4	Low burning rate, medium performance
Sodium perchlorate	NaClO4	122.44	2018	52.3		Hygroscopic, high performance, bright flame
Ammonium nitrate	NH4NO3	80.0	1730	60.0	20.0	Smokeless, medium performance, low cost

Ammonium perchlorate (NH₄ClO₄) is the most widely used crystalline oxidizer in solid propellants. It dominates the solid oxidizer field because of its desirable characteristics that include compatibility with other propellant materials, good performance, satisfactory quality and uniformity, low impact and friction sensitivities, and good availability. Other solid oxidizers, particularly ammonium nitrate and potassium perchlorate, have occasionally been used in production rockets but are now replaced by more modern propellants containing ammonium perchlorate. None of the many other oxidizer compounds that were investigated during the 1970s have reached production status.

The oxidizing potential of the perchlorates is generally high, which makes this material suited for high‐specific‐impulse propellants. Both ammonium and potassium perchlorate are only slightly soluble in water, a favorable propellant trait. All the perchlorate oxidizers generate gaseous hydrogen chloride (HCl) and other toxic and corrosive chlorine compounds in their reaction with fuels. In the firing of rockets, care is required, particularly for very large rockets, to safeguard operating personnel and/or communities in the path of exhaust gas clouds. Ammonium perchlorate (AP) is supplied in the form of small white crystals. Particle size and shape influences their manufacturing process and propellant burning rate. Therefore, close control of crystal sizes and size distributions present in a given quantity or batch is required. AP crystals are rounded (to nearly ball shape) to make for easier mixing than with sharp, fractured crystals. From the factory, they come in sizes ranging from about 600 µm (1 µm = 10⁻⁶ m) diameter to about 80 µm. Diameters below about 40 µm are considered more hazardous (e.g., can easily be ignited and sometimes detonate) and are not shipped; instead, propellant manufacturers take the larger crystals and grind them (at a motor factory) to smaller sizes (down to 2 µm) just before they are incorporated into a propellant.

Inorganic nitrates are relatively low‐performance oxidizers compared with perchlorates. However, ammonium nitrate (AN) is used in some applications because of its very low cost as well as its smokeless and relatively nontoxic exhaust. Its principal use is in low‐burning‐rate, low‐performance rocket and gas generator applications. Ammonium nitrate changes its crystal structure at several phase transformation temperatures. These changes cause slight changes in volume. One phase transformation at 32°C causes about a 3.4% change in volume. Repeated temperature cycling through this transition temperature creates tiny voids in the propellant, resulting in growth in the grain and a change in physical and/or ballistic properties. The addition of a small amount of stabilizer such as nickel oxide () or potassium nitrate () appears to change this transition temperature to above 60°C, a high enough value so that normal ambient temperature cycling will no longer cause recrystallization (Refs. 13–12 and 13–13). With such an additive, AN is known as phase‐stabilized ammonium nitrate (PSAN). Moreover, AN is hygroscopic and any moisture absorption will degrade and destabilize propellants containing AN.

Particle‐Size Parameters

The size, shape, and size distribution of solid particles like AP, Al, or HMX within the propellant can have a major influence on composite propellant characteristics. These particles are made spherical in shape to allow for easier mixing and to attain higher solid percentages in the propellant than shapes of sharp‐edged natural crystals. Normally, ground AP oxidizer crystals are rated according to particle size ranges as follows:

Coarse	400 to 600 µm (1 µm = 10−6 m)
Medium	50 to 200 µm
Fine	5 to 15 µm
Ultrafine	sub micrometer to 5 µm

Coarse and medium‐grade AP crystals are handled as class 1.3 materials, whereas the fine and ultrafine grades are considered as class 1.1 high explosives and are usually manufactured on‐site from medium or coarse grades, (see Section 13.3 for a definition of these explosive hazard classifications). Most propellants use a multimodal blend of oxidizer particle sizes to maximize the amount of oxidizer per unit volume of propellant, with the small particles filling part of the voids between the larger particles. A monomodal propellant has one size of solid oxidizer particles, a bimodal has two sizes (say, 20 and 200 µm), and a trimodal propellant has three sizes; multiple modes allow a larger mass of solids to be placed into a given volume. Problem 13–1 has a sketch depicting how voids between the largest particles are filled with smaller particles.

Figure 13–8 shows the influence of varying the ratio of coarse to fine oxidizer particle sizes on propellant burning rate together with the influence of a burning rate additive. Figure 13–9 shows that the effect of particle size of aluminum fuels on propellant burning rate is much less pronounced than that of oxidizer particle size (Fig. 13–8 also shows an effect of particle size). Particle size, range, and shape for both the oxidizer (usually ammonium perchlorate AP]) and solid fuel (usually aluminum) have significant effects on solid packing fractions and on rheological properties (associated with the flowing or pouring of viscous liquids) of uncured composite propellants. By definition, packing fraction is the volume fraction of all solids when packed to minimum volume (a theoretical condition). High packing fractions make mixing, casting, and handling during propellant fabrication more difficult. Figure 13–10 shows a resulting distribution of AP particle size using a blend of sizes; the shape of this curve can be altered drastically by controlling size ranges and ratios. Also, solid particle size, range and shape affect the solids loading ratio, which is the mass ratio of solid to total ingredients in the uncured propellants. Computer‐optimizing methods exist for adjusting particle‐size distributions to improve the loading of solids, which can be as high as 90% in some composite propellants. Even though high solids loadings are desirable for high performance, they often introduce complexity and higher costs into the processing of propellant. Trade‐offs among ballistic (performance) requirements, processability, mechanical strength, rejection rates, and facility costs are an ever‐present concern with many high‐specific‐impulse composite propellants. References 13–4 and 13–15 report on the influence of particle size on motor performance.

13.5 OTHER PROPELLANT CATEGORIES

Gas Generator Propellants

Gas generator propellants are used to produce hot gases, not thrust. They generally have a low combustion temperature (800 to 1600 K), and most do not require internal insulators when used with metal cases. Typical applications of gas generators are listed in Table 12–1. Of the large variety of propellants utilized to for gas generators only a few will be mentioned.

Stabilized AN‐based propellants have been used for many years with various binder ingredients. They give a clean, essentially smokeless exhaust and operate at a low combustion temperature. Because of their low burning rate they are useful for long‐duration gas generator applications, say 30 to 300 sec. Typical compositions are given in Ref. 13–13 and Table 13–10 describes a propellant representative of early gas generators. It is often of interest to add some fuel to AN that acts as a coolant.

Ballistic Properties
Calculated flame temperature (K)	1370
Burning rate at 6.89 MPa and 20°C (mm/sec)	2.1
Pressure exponent n (dimensionless)	0.37
Temperature sensitivity σp (%/K)	0.22
Theoretical characteristic velocity, (m/sec)	1205
Ratio of specific heats	1.28
Molecular mass of exhaust gas	19
Composition (Mass Fraction)
Ammonium nitrate (%)	78
Polymer binder plus curing agent (%)	17
Additives (processing aid, stabilizer, antioxidant) (%)	5
Oxidizer particle size, (µm)	150
Exhaust Gas Composition (Molar %)
Water (steam)	26
Carbon monoxide	19
Carbon dioxide	7
Nitrogen	21
Hydrogen	27
Methane	Trace
Physical Properties at 25°C or 298 K
Tensile strength (MPa)	1.24
Elongation (%)	5.4
Modulus of elasticity in tension (N/m2)	34.5
Specific gravity	1.48

One method of reducing flame temperatures is to burn conventional AP propellants hot and then to add water to bring combustion gas temperatures down to where uncooled metals can contain them. This is used on the MX missile launcher tube gas generator (Ref. 13–18). Another formulation uses HMX or RDX with an excess of polyether‐ or polyester‐type polyurethane.

For the inflation of automobile collision safety bags the combusted gas must be nontoxic, smoke free, have a low temperature (so as not to burn people), be quickly initiated; the propellant must be storable for rather long times without degrading and reliably available. One solution is to use alkali metal azides (e.g., or ) with an oxide and an oxidizer. The resulting nitrates or oxides are solid materials that are removed by filtering, and the gas is clean largely composed of moderately hot nitrogen. In one model, air can be aspirated into the air bag by the hot, high‐pressure gas (see Ref. 13–19). One particular composition uses 65 to 75% , 10 to 28% , and 5 to 16% as an oxidizer, a burn rate modifier, and a small amount of SiO₂ for moisture absorption. The resultant solid nitride slag is caught in a filter.

The ideal power delivered by a gas generator can be expressed as (see Chapters 3 and 5)

13–1

where is the mass flow rate, and are the enthalpies per unit mass respectively (at the gas generator chamber and exhaust pressure conditions), is the flame temperature in the gas generator chamber, is the gas constant, is the reciprocal of the pressure ratio through which these gases are expanded, and k the specific heat ratio. Because flame temperatures are relatively low, there is no appreciable dissociation and frozen equilibrium calculations are usually adequate.

13.6 LINERS, INSULATORS, AND INHIBITORS

Liners, insulators, and inhibitors represent three layer types that reside at grain interfaces and are defined in Section 12.3. Their materials do not contain any oxidizing ingredients but they may ablate, cook, char, vaporize, or disintegrate in the presence of hot gases. Many will actually burn if the hot combustion gases contain even small amounts of oxidizing species, but they will not usually burn by themselves. Liners, internal insulators, or inhibitors must be chemically compatible with the propellant and with each other to avoid migration (described below) and/or changes in material composition; they must have good adhesive strength so that they stay bonded to the propellant, or to each other. The temperatures at which they experience damage or large surface regressions should be adequately high. They should all have low densities in order to reduce inert mass. Typical materials are neoprene (specific gravity 1.23), butyl rubber (0.93), a synthetic rubber called ethylene propylene diene or EPDM (0.86), or a propellant binder such as polybutadiene (0.9 to 1.0); these values are low relative to propellant specific gravities of 1.6 to 1.8. For low‐smoke propellant these three rubber‐like materials should only give off some gases but few, if any, solid particles (see Ref. 13–22).

The liners principal purpose is to provide a proper bond between the grain and the case, or between the case and any internal thermal insulation. In addition to the desired characteristics listed in the previous paragraph, the liner should be a soft stretchable rubber‐type thin material (typically 0.02 to 0.04 in. thick with 200 to 450% elongation) to allow relative movement along the bond line between the grain and the case. This differential expansion is needed because the thermal coefficient of expansion of the grain is typically an order of magnitude higher than that of the case. A liner will also seal fiber‐wound cases (particularly thin cases), which are often porous, so that high‐pressure hot gases cannot escape. Typical liners for tactical guided missiles have been made from polypropylene glycol (about 57%) with a titanium oxide filler (about 20%), a di‐isocyanate crosslinker (about 20%), and minor ingredients such as an antioxidant. Rocket motor cases have to be preheated to about 82°C prior to liner application. Often used as polymer for liners is ethylene propylene diene monomer (EPDM) linked into ethylene propylene diene terpolymer to form a synthetic rubber; it adheres and elongates nicely.

In some present‐day motors, internal insulators not only provide for thermal protection of the case (from the hot combustion gases) but also often serve as a liner providing good bonding between propellant and insulator or insulator and case. Most motors today still have a separate liner and an insulating layer. Internal thermal insulation should also fulfill these additional requirements:

1. It must be erosion resistant, particularly at the motor aft‐end or blast tube. This may be achieved in part by using tough elastomeric materials, such as neoprenes or butyl rubbers that are chemically resistant to the hot gases and to the impact of particulates. Such surface integrity is also achieved when a porous black‐carbon layer formed on a heated surface (called a porous char layer) remains after some interstitial materials have been decomposed and vaporized.
2. It must provide good thermal resistance and have low thermal conductivity to limit heat transfer to the case and thus keep the case below its maximum allowable temperature, which is usually between 160 and 350°C for the plastic in composite material cases and about 550 and 950°C for most steel cases. These are accomplished by filling the insulator either with silicon oxide, graphite, Kevlar, or ceramic particles. Asbestos, an excellent filler material, is no longer used because of health hazards.
3. It should allow large‐deformations or strains to accommodate grain deflections upon pressurization or temperature cycling, and should transfer loads between the grain and the case.
4. Surface regression should be minimal so as to retain much of its original geometric surface contour and allow for thin insulation.

A simple relationship for the internal insulation thickness at any location in the rocket motor depends on the exposure time , the erosion rate (obtained from erosion tests at the likely gas velocity and temperature), and the safety factor f which can range from 1.2 to 2.0:

13–2

For small test rocket motors, some designers use the rule that the insulation depth should be twice the charred depth of the insulation.

Usually, the thickness of insulators cannot be uniform, varying by factors up to 20. It has to be thicker at locations near the nozzle where it is exposed to longer intervals and higher scrubbing velocities than at insulator locations protected by bonded propellant. Before making any material selection, it is necessary to evaluate the burned‐propellant flow field and thermal environment (combustion temperature, gas composition, pressure, exposure duration, and internal ballistics) in order to carry out proper thermal analyses (erosion predictions and estimated thickness of insulator). Evaluation of loads and deflections under loads at different motor locations are also needed to estimate shear and compression stresses. When high stresses or relief flaps are involved, structural analyses are also needed. Various software codes, such as those mentioned in Refs. 13–23 and 13–24, have been used.

Inhibitors are usually made of the same materials as internal insulators. They are applied (bonded, molded, glued, or sprayed) to grain surfaces that should not burn. In a segmented rocket motor, for example (see Fig. 15–2), where burning is allowed only on the internal port area, the faces of the cylindrical grain sections may be inhibited.

Migration describes the transfer of mobile (liquid) chemical species from the solid propellant to the liner, insulator, or inhibitor, and vice versa. Liquid plasticizers such as NG or DEGDN or unreacted monomers or liquid catalysts are known to migrate. This migratory transfer occurs very slowly but may cause dramatic changes in physical properties (e.g., the propellant next to the liner becomes brittle or weak), and there are several instances where nitroglycerine that migrated into an insulator made it flammable. Migration can be prevented or inhibited by using (1) propellants without plasticizers, (2) insulators or binders with plasticizers identical to those used in propellants, (3) a thin layer of an impervious material or a migration barrier (such as PU or a thin metal film), and (4) an insulator material that will not allow migration (e.g., PU) (see Ref. 13–25).

The graphite–epoxy case of the launch‐assist rocket motors used to boost the Delta launch vehicle utilizes a three‐layer liner: EPDM (ethylene propylene diene terpolymer) as a thin primer to enhance bond strength, a polyurethane barrier to prevent migration of the plasticizer into the EPDM liner, and a plasticized HTPB‐rich liner to prevent burning next to the case–bond interface. Composite AP–Al propellants also use the same HTPB binder.

Liners, insulators, and/or inhibitors may be applied to the grain in several ways: by painting, coating, dipping, spraying, or by gluing sheets or strips to the case or the grain. Often automated, robotic machines are used to achieve uniform thicknesses and high quality. Reference 13–22 describes the manufacture of particular insulators.

External insulation is often added to the outside of the motor case when it is part of the vehicle's structure, particularly in tactical missiles or high‐acceleration launch boosters. Such insulation reduces the heat flow from any hot boundary layers outside the vehicle surface (aerodynamic heating) to the case, and then to the propellant. It may thus prevent fiber‐reinforced plastic cases from becoming weak or propellants from becoming soft or, in extreme situations, from being ignited. Such insulator must withstand oxidation from hot air flows, have good adhesion, have structural integrity to loads imposed by the flight or launch, and must have reasonable cure temperatures. Materials ordinarily used as internal insulators are unsatisfactory because they burn in the atmosphere generating additional heat. The best is a nonpyrolyzing, low‐thermal‐conductivity refractory material (Ref. 13–26) such as certain high‐temperature paints. Any internal and external insulation also helps to reduce grain temperature fluctuations and thus thermal stresses imposed by thermal cycling, such as day–night variations or high‐ and low‐altitude temperature variations for airborne missiles.

13.7 PROPELLANT PROCESSING AND MANUFACTURE

The manufacture of solid propellant involves many complex physical and chemical processes. In the past, propellants have been produced by several different processes such as compaction or pressing of powder charges, extrusion of propellant through dies under pressure using heavy presses, and mixing with a solvent which is later evaporated. Even for the same type of propellant (e.g., double‐base, composite, or composite double‐base), fabrication processes are usually not identical for different manufacturers, motor types, sizes, or propellant formulation, and no single simple generalized process flowsheet or fabrication technique prevails. Most rocket motors in production today use composite‐type propellants and therefore more emphasis on this process is given here.

Figure 13–11 is a representative flowsheet for the manufacture of a complete solid rocket motor with a composite propellant made by batch processes. Processes marked with an asterisk are potentially hazardous, are usually operated or controlled remotely, and are usually performed in buildings designed to withstand potential fires or explosions. Mixing and casting processes are the most complex, being more critical than other processes in determining quality, performance, burn rate, and physical properties of the resulting propellant.

The rheological properties of uncured propellants (e.g., their flow properties in terms of shear rate, stress, and time) are all‐important to their processability and these properties usually change substantially throughout the path of the processing line. Batch‐type processing of propellants, including the casting (pouring) of propellants into motors that serve as their own molds, is the most common method. For very large motors several days are needed for casting up to perhaps 40 batches into a single case to form a single grain. Vacuum is almost always imposed on the propellant during the mixing and casting operations to remove air and other dispersed gases and to avoid air bubbles in the grain. Viscosity measurements of the mixed propellant (in the range of 10,000 to 20,000 poise) are made for quality control. Vacuum, temperature, vibration, energy input to the mixer, and time are some factors affecting the viscosity of the uncured propellant. Time is important in terms of pot life, that is, that period of time the uncured propellant remains reasonably fluid after mixing before it cures and hardens. Short pot lives (a few hours) require fast operations in emptying mixers, measuring for quality control, transporting, and casting into motors. Some binder systems, such as those using PVC, present a very long pot life and avoid the urgency or haste in the processing line. References 13–15, 13–11, and 13–27 give details on propellant processing techniques and equipment.

Double‐base propellants and modified double‐base propellants are manufactured by a different set of processes. The key goal here is the diffusion of liquid nitroglycerine into the fibrous solid matrix of nitrocellulose, thus forming, by means of solvation, a fairly homogeneous, well‐dispersed, relatively strong solid material. Several processes for making double‐base rocket propellant are in use today, including extrusion and slurry casting. In the slurry casting process the case (or the mold) is filled with solid casting powder (a series of small solid pellets of nitrocellulose with a small amount of nitroglycerine) and the case is then flooded with liquid nitroglycerine, which then solvates the pellets. Figure 13–12 shows a simplified diagram of a typical setup for a slurry cast process. Double‐base propellant manufacturing details are shown in Refs. 13–5 and 13–15.

Mandrels are used during casting and curing to assure the proper internal cavity or perforation pattern. They are made of metal in the shape of the internal bore (e.g., star or dogbone) and are often slightly tapered and coated with a nonbonding material, such as Teflon, to facilitate the withdrawal of the mandrel after curing without tearing the grain. For complicated internal passages, such as a conocyl, complex built‐up mandrels, which can be withdrawn through the nozzle flange opening in smaller pieces or which can be collapsed are necessary. Some manufacturers have had success in making permanent mandrels (that are not withdrawn but stay with the motor) out of a lightweight foamed propellant that burn very quickly once ignited.

An important objective in processing is to produce a propellant grain free of cracks, low‐density areas, voids, or other flaws. In general, voids and other flaws degrade the ballistic and mechanical properties of the propellant grain. Many small dispersed gas bubbles in a propellant grain may result in an abnormally high burning rate, one so high as to cause excessive pressure and catastrophic case failure.

The finished grain (or rocket motor) is usually inspected for defects (cracks, voids, and debonds) using X‐rays, ultrasonics, heat conductivity probes, and/or other nondestructive inspection techniques. Propellant samples are taken from each batch, tested for rheological properties, and cast into physical property specimens and/or small rocket motors, which are then cured and subsequently tested. Any determination of the sensitivity of rocket motor performance, including possible failure, to propellant voids and other flaws often requires the test firing of motors with known defects. Data from such tests are important in establishing inspection criteria for accepting and rejecting production rocket motors.

Special processing equipment is required in the manufacture of propellants. For composite propellants this includes mechanical mixers (usually with two or three blades rotating on vertical shafts agitating propellant ingredients in a mixer bowl under vacuum), casting equipment, curing ovens, and/or machines for automatically applying the liner or insulation to the case, see Ref. 13–11. Double‐base processing requires equipment for mechanically working the propellant (rollers, presses) or special tooling for allowing a slurry cast process. Computer‐aided filament winding machines are used for laying the fibers of fiber‐reinforced plastic cases and nozzles.

Casting of propellant into large boosters and strap‐ons very near the launch site and the vertical keeping of rocket motors has the potential to decrease problems associated with handling and transportation, as well as with propellant slump. This decreases the time from manufacturing to launch but is not recommended for long storage times in any vertical position.

PROBLEMS

1. Ideally, solid oxidizer particles in a propellant may be treated as spheres of uniform size. Three sizes of particles are available: Coarse at 500 µm, medium at 50 µm, and fine at 5 µm, each at a specific gravity of 1.95, and a viscoelastic fuel binder is available at a specific gravity of 1.01. Assume that these materials can be mixed and vibrated so that the solid particles will touch each other so that there are no voids in the binder, and that the particles occupy a minimum of space similar to the sketch of the cross section shown here. It is desired to put 94 wt‐% of oxidizer into the propellant mix, as this will give maximum performance. (a) Determine the maximum weight percentage of oxidizer if only coarse crystals are used or if only medium‐sized crystals are used. (b) Determine the maximum weight of oxidizer if both coarse and fine crystals are used, with the fine crystals filling the voids between the coarse particles. What is the optimum relative proportion of coarse and fine particles to give a maximum of oxidizer? (c) Same as part (b), but use coarse and medium crystals only. Is this better and, if so, why? (d) Using all three sizes, what is the ideal weight mixture ratio and what is the maximum oxidizer content possible and the theoretical maximum specific gravity of the propellant? (Hint: The centers of four adjacent coarse crystals form a tetrahedron whose side length is equal to the diameter.)

   

2. Suggest one or two specific applications (e.g., anti‐aircraft, intercontinental missile, space launch vehicle upper stage, etc.) for each of the propellant categories listed in Table 13–2 and explain why it was selected in comparison to other propellants.
3. Prepare a detailed outline of a procedure to be followed by a crew operating a propellant mixer. This 1‐m³ vertical solid propellant mixer has two rotating blades, a mixing bowl, a vacuum pump system to allow mix operations under vacuum, feed chutes or pipes with valves to supply the ingredients, and variable‐speed electric motor drive, a provision for removing some propellant for laboratory samples, and a double‐wall jacket around the mixing bowl to allow heating or cooling. It is known that composite propellant properties are affected by mix time, small deviations from the exact composition, temperature of the mix, the mechanical energy added by the blades, the blade speed, and the sequence in which the ingredients are added. It is also known that bad propellant would result if there are leaks that destroy the vacuum, if the bowl, mixing blades, feed chutes, and so on, are not clean but contain deposits of old propellant residue from a prior mix on their walls, if they are not mixed at 80°C, or if the viscosity of the mix becomes excessive. Usually, the sequence of loading ingredients needs to be: (1) prepolymer binder, (2) plasticizer, (3) minor liquid additives, (4) solid consisting of first powdered aluminum followed by mixed bimodal or trimodal AP crystals, and (5) finally the polymerizing agent or crosslinker. Refer to Fig. 13–11. Samples of the final liquid mix are taken to check for viscosity and density. Please list all the sequential steps that the crew should undertake before, during, and after the mixing operation. Each major step in this outline of the procedure for mixing a composite propellant has a purpose or objective that should be stated. Any specific data for each step (e.g., mass, duration, temperature, blade speed, etc.) is not required to be listed, but if known can be stated. Mention all instruments (e.g., thermometers, wattmeter, etc.) that the crew should have and identify those that they must monitor closely. Assume that all ingredients were found to be of the desired composition, purity, and quality.
4. Determine the longitudinal growth of a 24‐in.‐long freestanding grain with a linear thermal coefficient of expansion of 7.5 × 10⁻⁵/°F for temperature limits of −40 to 140°F.Answer: 0.32 in.
5. The following data are given for an internally burning solid propellant grain with inhibited end faces and a small initial port area:

   Length	40 in.
   Port area	27 in.2
   Propellant weight	240 lbf
   Initial pressure at front end of chamber	1608 psi
   Initial pressure at nozzle end of chamber	1412 psi
   Propellant density	0.060 lbm/in.3
   Vehicle acceleration	21.2g0

   Determine the initial forces on the propellant supports produced by the pressure differential and by the vehicle acceleration.Answer: 19,600 lbf, 5090 lbf.
6. A solid propellant rocket motor with an end‐burning grain has a thrust of 4700 N and a14 sec duration. Four different burning rate propellants are available, all with approximately the same performance and the same specific gravity, but different AP mix and sizes and different burning rate enhancing ingredients. They are 5.0, 7.0, 10, and 13 mm/sec. The preferred is 2.60, but values of 2.2 to 3.5 are acceptable. The impulse‐to‐initial‐weight ratio is 96 at an of 2.5. Assume optimum nozzle expansion at sea level. Chamber pressure is 6.894 MPa or 1000 psia and the operating temperature is 20°C or 68°F. Determine grain geometry, propellant mass, approximate hardware mass, and initial mass.
7. For the rocket in Problem 6 determine the approximate chamber pressure, thrust, and duration at 245 and 328 K. Assume the temperature sensitivity (at a constant value of ) of 0.01%/K does not change with temperature.
8. A fuel‐rich solid propellant for a gas generator drives a turbine of a liquid propellant turbopump. Determine its mass flow rate. The following data are given:

   Chamber pressure
   Combustion temperature
   Specific heat ratio
   Required pump input power	970 KW
   Turbine outlet pressure	0 psia
   Turbine efficiency	65%
   Molecular mass of gas	22 kg/kg‐mol
   Pressure drop between gas generator and turbine nozzle inlet	0.10 MPa

   Windage and bearing friction is 10 kW. Neglect start transients.Answer: .
9. A gas generator propellant the following characteristics:

   Burn rate at standard conditions	4.0 mm/sec
   Burn time	110 sec
   Chamber pressure	5.1 MPa
   Pressure exponent n	0.55
   Propellant specific gravity	1.47

   Determine the size of an end‐burning cylindrical grain.Answer: Single end‐burning grain 27.2 cm in diameter and 31.9 cm long, or two end‐burning opposed grains (each 19.6 cm diameter × 37.3 cm long) in a single chamber with ignition of both grains in the middle of the case.

REFERENCES

1. 13–1. O. Bolton et al., “High Power Explosive with Good Sensitivity: A 2:1 Cocrystal of CL‐20:HMX,” Crystal Growth, Design, Vol. 12, No. 9, August 2012, pp. 4311–4314. doi:10.1021/cg3010882.
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