Solid propellants combine fuel and oxidizer into a single cast grain that burns from an exposed surface into the chamber volume. The entire delivery system, the pumps and the injectors and the tank pressurization that a liquid engine requires, collapses into the grain itself and its case. The resulting engine is mechanically simple, storable for decades, and delivers a high density specific impulse at the cost of lower absolute specific impulse than a liquid engine and the loss of throttling and restart capability. This article treats solid propellant chemistry at the level the previous article establishes.

The design space of solid rocket propellants divides into two historical families and one composite-of-composites. The composite family, dominated by ammonium perchlorate oxidizer with a hydroxyl-terminated polybutadiene binder and aluminum fuel, delivers the highest specific impulse routinely achievable with solids and dominates large boost applications. The double-base family, dominated by nitrocellulose plasticized with nitroglycerin, carries the older munitions heritage and delivers modestly lower specific impulse. Composite modified double-base combinations blend the two chemistries and reach specific impulse values in the composite range while retaining some of the mechanical and manufacturing characteristics of double-base grains. The current research frontier pursues alternatives to ammonium perchlorate that eliminate the hydrogen chloride from the exhaust, and higher-energy binders and fuels that push delivered specific impulse toward the theoretical maximum for a solid.

Solid Rocket Motor Anatomy

A solid rocket motor consists of a pressure-containing case, an insulation layer that protects the case from the hot combustion gases, a cast propellant grain that combines fuel and oxidizer, an igniter at the forward end that starts combustion, and a nozzle at the aft end that accelerates the exhaust. The grain is bonded to the insulation, which is bonded to the case, so that no significant relative motion occurs during firing. A central perforation through the grain, or a set of shaped perforations, exposes the initial burning surface. Combustion propagates radially outward from that surface into the grain until the grain is consumed or the case fails.

The case is typically wound from glass, aramid, or carbon fiber in an epoxy or polyester matrix for the highest performance-to-mass ratio, or fabricated from steel for the highest mechanical robustness and lowest unit cost. The insulation layer is typically a filled elastomer, often ethylene propylene diene monomer rubber loaded with silica or Kevlar fibers, that ablates during firing at a controlled rate. The igniter is a smaller solid rocket motor of its own, whose exhaust impinges on the main grain surface to start ignition. The nozzle is typically a graphite or carbon-carbon throat insert set in a steel or composite housing, with the throat sized to produce the desired chamber pressure at the design mass flow rate.

The mechanical simplicity of this arrangement is the principal engineering attraction of solid propellants. A single pressure vessel with an igniter and a nozzle is all the vehicle needs to produce thrust. There are no turbopumps, no valves, no cooling passages, no injectors, and no propellant tanks in the ordinary sense. The engineering complexity moves into the chemistry of the propellant itself and into the grain design that shapes the thrust profile over the burn duration.

Composite Propellants

The composite family combines a granular crystalline oxidizer, a polymer binder that also serves as fuel, and typically a metal fuel additive, into a heterogeneous cast mixture. Ammonium perchlorate is the standard oxidizer for high-performance applications. Hydroxyl-terminated polybutadiene is the standard binder in modern formulations, having displaced the older polybutadiene acrylonitrile and polybutadiene acrylic acid binders that dominated through the nineteen sixties and nineteen seventies. Aluminum powder is the standard metal fuel additive.

A representative modern formulation contains approximately $68$ percent ammonium perchlorate by mass, $18$ percent aluminum by mass, and $12$ percent hydroxyl-terminated polybutadiene by mass, with the remaining $2$ percent divided among bond agents, curing agents, plasticizers, and ballistic modifiers. The specific fractions vary across manufacturers and applications by a few percentage points. The Space Shuttle solid rocket booster propellant, produced by Thiokol and later ATK, used $69.6$ percent ammonium perchlorate, $16$ percent aluminum, $12.04$ percent polybutadiene acrylonitrile, $1.96$ percent epoxy curing agent, and $0.4$ percent iron oxide as a burn rate catalyst. The Ariane 5 solid rocket booster propellant, produced by Europropulsion, uses approximately $68$ percent ammonium perchlorate, $18$ percent aluminum, and $14$ percent hydroxyl-terminated polybutadiene.

Ammonium Perchlorate

Ammonium perchlorate, chemical formula $NH_4 ClO_4$ and molecular weight $117.49$ grams per mole, is a crystalline salt that provides both nitrogen and oxygen to the combustion process. It decomposes at approximately $600$ kelvin through a two-step mechanism whose products depend on temperature and pressure. At low temperatures the decomposition produces chlorine oxides, hydrogen chloride, nitrogen oxides, water, and oxygen. At the temperatures of rocket combustion, above $2000$ kelvin, the decomposition proceeds essentially to the following net reaction with released oxygen available to burn the binder and the aluminum.

\[4 NH_4 ClO_4 \rightarrow 4 HCl + 2 N_2 + 5 O_2 + 6 H_2 O\]

Ammonium perchlorate is manufactured in controlled particle-size grades. Formulations typically use a bimodal or trimodal blend of coarse and fine grades to reach high packing density, which allows high oxidizer loading without compromising the grain’s mechanical integrity. A representative coarse grade has a mean particle diameter of approximately $200$ micrometers. A representative fine grade has a mean particle diameter of approximately $20$ micrometers. Ultrafine grades below $10$ micrometers are used to accelerate burn rate.

The presence of chlorine in ammonium perchlorate is the principal environmental drawback of composite propellants. The combustion exhaust contains approximately $20$ percent hydrogen chloride by mass. Hydrogen chloride combines with atmospheric water vapor near the launch site to form a corrosive acidic aerosol that damages ground equipment and vegetation in the immediate launch area. This effect motivates the research on chlorine-free oxidizers discussed later.

Hydroxyl-Terminated Polybutadiene

Hydroxyl-terminated polybutadiene, abbreviated HTPB, is a liquid prepolymer that cures into a solid elastomer through reaction with a diisocyanate or polyisocyanate curing agent. The polymer backbone consists of repeating butadiene units, chemical formula $(C_4H_6)_n$, terminated by hydroxyl groups at both ends. The typical molecular weight of the uncured prepolymer is approximately $2500$ to $3000$ grams per mole. The cured elastomer provides the mechanical binder that holds the oxidizer and metal particles in place, and the polybutadiene backbone contributes carbon and hydrogen to the combustion, so the binder serves simultaneously as structural matrix and as fuel.

The curing reaction crosslinks the prepolymer into a rubbery solid over several days at elevated temperatures, typically $60$ to $80$ degrees Celsius. During curing the propellant slurry is either poured into the motor case directly, called case bonding, or into a separate mandrel and transferred to the case after cure. The isocyanate curing agent is typically isophorone diisocyanate, abbreviated IPDI, chosen for its slow reaction rate that allows adequate pot life during casting. Toluene diisocyanate and diphenylmethane diisocyanate are used in some applications but are less common in modern practice because of the higher toxicity of their vapors.

Hydroxyl-terminated polybutadiene displaced polybutadiene acrylic acid and polybutadiene acrylonitrile as the dominant binder in the nineteen seventies and nineteen eighties. Hydroxyl-terminated polybutadiene delivers higher specific impulse, better mechanical properties at low temperatures, and longer shelf life than the older binders. Polybutadiene acrylonitrile retains a niche in legacy applications where the case-insulation-grain bond chemistry was qualified for that binder and requalification would be prohibitively expensive.

Aluminum Fuel

Metallic aluminum in powder form serves as the principal fuel additive in composite propellants. Aluminum combusts to aluminum oxide, chemical formula $Al_2 O_3$ and molecular weight $101.96$ grams per mole, releasing approximately $31$ megajoules per kilogram of heat.

\[4 Al + 3 O_2 \rightarrow 2 Al_2 O_3\]

That heat of combustion raises the chamber temperature by hundreds of kelvin above the value the binder alone would produce, and the higher chamber temperature translates directly to higher specific impulse.

Aluminum powder is used in particle sizes from approximately $5$ to $30$ micrometers. The particles are typically spherical or spheroidal, produced by atomization of molten aluminum in an inert atmosphere. The as-received surfaces carry a thin native oxide layer of approximately $2$ to $5$ nanometers that protects the metal from further oxidation during storage and mixing.

During combustion, the aluminum particles do not burn instantaneously. They melt at $933$ kelvin, and the molten droplets are carried by the gas flow away from the burning grain surface. The droplets ignite once the surface oxide is disrupted and the underlying aluminum contacts the oxidizer-rich gas. They burn as diffusion-limited spheres until the metal is consumed. The resulting aluminum oxide is present in the exhaust as condensed liquid droplets at the chamber temperature and as condensed solid particles once the exhaust cools below the melting point of aluminum oxide at approximately $2345$ kelvin. The transition from liquid to solid aluminum oxide occurs inside the nozzle for most engines.

The presence of condensed aluminum oxide particles in the exhaust produces a specific impulse penalty called the two-phase-flow loss. Gas expansion in the nozzle accelerates the gas but does not accelerate the condensed particles as efficiently, because the particles have thermal and mechanical inertia that resist rapid acceleration. The particles arrive at the nozzle exit at lower velocity than the gas, and the mass-weighted average exit velocity is lower than the pure-gas expansion would produce. The magnitude of the loss is typically $3$ to $8$ percent of ideal specific impulse, with larger losses at smaller aluminum loadings and at larger particle sizes.

Composite Propellant Combustion

Combustion proceeds through a coupled physical and chemical process at the grain surface. Heat from the combustion zone above the surface radiates and conducts back to the surface, raising the surface temperature to approximately $700$ to $1000$ kelvin. At this temperature the ammonium perchlorate crystals decompose and the hydroxyl-terminated polybutadiene binder pyrolyzes. Gaseous products diffuse away from the surface into a premixed reaction zone where the fuel-rich binder products react with the oxidizer-rich ammonium perchlorate products. Aluminum particles are liberated from the surface as the binder pyrolyzes around them, and the particles agglomerate on the surface before detaching to burn as droplets in the gas flow.

The steady-state burn rate depends primarily on chamber pressure. The empirical relationship, called Vieille’s law or Saint-Robert’s law after its independent nineteenth-century discoverers, expresses linear burn rate $r$ as a power function of chamber pressure $P_c$.

\[r = a \, P_c^n\]

The coefficient $a$ has units that depend on the units of $r$ and $P_c$ and is typically tabulated for a specific reference temperature. The pressure exponent $n$ is dimensionless and typically ranges from $0.2$ to $0.4$ for modern composite propellants. Stable combustion requires $n$ less than $1$. If $n$ were $1$ or greater, small pressure increases would accelerate combustion enough to further increase pressure without bound. The condition $n < 1$ produces a stable chamber pressure that any small perturbation naturally returns to.

Typical burn rates at chamber pressures near $70$ bar range from $5$ millimeters per second for slow-burning grain formulations to $30$ millimeters per second for fast-burning formulations. Iron oxide, copper chromite, and other transition-metal oxides in small percentages accelerate the burn rate. Oxamide slows it. Modern composite propellants can be formulated with burn rates spanning approximately a factor of ten from the slowest to fastest catalyzed variants.

Combustion also depends on the initial grain temperature. Cold grains burn more slowly than warm grains because the surface takes longer to reach ignition temperature. The temperature sensitivity coefficient $\pi_K$ expresses the fractional change in burn rate per kelvin of initial temperature change at constant chamber pressure.

\[\pi_K = \left( \frac{\partial \ln r}{\partial T_i} \right)_{P_c}\]

Typical values range from $0.001$ to $0.005$ per kelvin. Motors designed for wide temperature ranges must accommodate the burn rate variation from cold storage to warm storage, which propagates directly into chamber pressure and thrust variation.

Double-Base Propellants

Double-base propellants combine nitrocellulose and nitroglycerin into a gelatinous colloidal mixture. Neither ingredient requires separate oxidizer because both molecules contain oxygen in sufficient quantities to sustain their own combustion. Both are chemically related to organic nitrates and share the general behavior of that compound family.

Nitrocellulose, chemical formula approximately $C_6H_7N_3O_{11}$ for fully nitrated cellulose, is prepared by nitrating cellulose fiber with a mixture of nitric and sulfuric acids. The degree of nitration varies from approximately $10.5$ percent nitrogen for propellant-grade nitrocellulose to approximately $13.5$ percent for the highest energy grades. The nitrated cellulose retains the fibrous character of the parent cellulose but is now an energetic material that decomposes exothermically when heated.

Nitroglycerin, chemical formula $C_3H_5N_3O_9$ and molecular weight $227.09$ grams per mole, is prepared by nitrating glycerin with nitric and sulfuric acid. Pure nitroglycerin is a viscous liquid at room temperature and a shock-sensitive high explosive. Its combination with nitrocellulose gelatinizes the fibrous nitrocellulose into a plastic mass that can be extruded, rolled, or cast into propellant grains, and the gelatinization substantially reduces the shock sensitivity from the value of pure nitroglycerin.

A representative double-base propellant contains approximately $50$ percent nitrocellulose, $35$ percent nitroglycerin, and the remaining $15$ percent divided among diethyl phthalate as plasticizer, diphenylamine or centralite as stabilizer against slow acid-catalyzed decomposition, and lead compounds or other catalysts to control burn rate. The plasticizer softens the mixture during processing. The stabilizer absorbs nitrogen oxides that would otherwise catalyze further decomposition and shorten shelf life.

Double-base propellants are manufactured by two principal processes. Extruded double-base is prepared as a viscous dough that is forced through a die to produce cylindrical grains of the required diameter, then cut to length and cured. This process produces smaller grains suitable for tactical missiles, sounding rockets, and gun propellants. Cast double-base is prepared as a slurry of nitrocellulose fibers in a mixture of nitroglycerin, plasticizer, and stabilizer that is poured into the motor case and cured at low temperature for several days. This process accommodates larger grains but requires more elaborate handling because the pourable slurry is more shock sensitive than the extruded dough.

The delivered specific impulse of a double-base propellant ranges from approximately $210$ to $240$ seconds at sea level, depending on the specific composition and the chamber conditions. The chamber temperature is approximately $2600$ to $3000$ kelvin, lower than a composite propellant because the fuel-oxygen balance is more constrained by the fixed nitrocellulose and nitroglycerin chemistry. The exhaust contains no chlorine or aluminum, so the two-phase-flow loss is negligible and the delivered specific impulse comes closer to the theoretical value than in a composite formulation.

The principal disadvantages of double-base propellants are lower absolute specific impulse than composites, greater sensitivity to impact and thermal insult than modern composite formulations, and the slow acid-catalyzed decomposition that limits shelf life. The principal advantages are the absence of a corrosive exhaust, the finer control of combustion products for smoke-free operation, and the mature manufacturing base built up over more than a century of use in gun propellants and small rockets.

Composite Modified Double-Base Propellants

Composite modified double-base propellants, abbreviated CMDB, combine the nitrocellulose and nitroglycerin base of a double-base formulation with solid oxidizer particles and metal fuel particles suspended in the gelatinized matrix. A representative CMDB formulation contains approximately $30$ percent nitrocellulose, $25$ percent nitroglycerin, $20$ percent cyclotetramethylenetetranitramine, abbreviated HMX, as a solid energetic oxidizer, $15$ percent aluminum, and the remaining $10$ percent divided among plasticizer, stabilizer, and burn rate modifier.

The presence of HMX increases both the energy density and the chamber temperature above pure double-base values. HMX is a high explosive in its pure form and contributes carbon and hydrogen as fuel elements along with oxygen and nitrogen as oxidizer elements, all of which release energy during combustion. Aluminum contributes fuel and additional combustion energy in the same manner as in composite formulations. The nitrocellulose and nitroglycerin matrix serves both as the binder that holds the solid particles in place and as an additional energetic fuel-oxidizer combination.

CMDB propellants deliver specific impulse in the range of $250$ to $270$ seconds at sea level, comparable to composite formulations. They preserve the lower two-phase-flow loss of double-base because the aluminum loading is smaller than in a pure composite, and they preserve the smoke-free operation for the same reason. Their principal industrial niche has been strategic ballistic missile upper stages, where high specific impulse and reduced smoke are both required. American Navy and Air Force propulsion programs that developed classical CMDB in the nineteen seventies and nineteen eighties, along with modernized nitrate ester plasticized polyether descendants that displaced classical CMDB in newer strategic missile third stages, invested in the manufacturing base that persists today.

CMDB carries the shock sensitivity of double-base combined with the manufacturing complexity of composite. Its handling requirements are stricter than either pure family, and its cost per unit specific impulse is correspondingly higher. It is used where the specific-impulse advantage over pure composite justifies the additional cost, primarily in weight-critical strategic missile applications.

The Research Frontier

Two research directions occupy the current solid propellant frontier, both motivated by the environmental disadvantage of hydrogen chloride in composite exhaust and by the desire to push specific impulse beyond the composite ceiling.

Ammonium Dinitramide

Ammonium dinitramide, chemical formula $NH_4 N(NO_2)_2$ and molecular weight $124.06$ grams per mole, abbreviated ADN, is a crystalline oxidizer that produces exhaust containing water, nitrogen, and oxygen but no chlorine. It was developed in the nineteen seventies at the Zelinsky Institute in the Soviet Union and independently rediscovered at the SRI International research institute in the United States in the nineteen eighties. A composite propellant with ADN in place of ammonium perchlorate produces a chlorine-free exhaust and delivers specific impulse approximately $10$ seconds higher than a comparable ammonium perchlorate formulation, in exchange for the higher cost of ADN and the more limited production base. Its representative net decomposition proceeds according to the following equation.

\[NH_4 N(NO_2)_2 \rightarrow 2 N_2 + O_2 + 2 H_2 O\]

ADN has been used in production monopropellant systems as a liquid solution known as LMP-103S, treated in the next article on storable liquid propellants, but has seen only limited adoption in production solid propellants because the cost premium over ammonium perchlorate is not justified for most industrial applications absent a specific environmental driver.

Hydroxyl-Terminated Polyether

Hydroxyl-terminated polyether, abbreviated HTPE, is an alternative polymer binder chemistry with better mechanical properties at low temperatures than HTPB. HTPE grains retain useful mechanical strength at temperatures where HTPB grains become brittle. This makes HTPE binders attractive for applications with wide temperature ranges, particularly missile systems that must be stored at low temperatures and fired at any temperature within the operational envelope.

HTPE binders are being adopted incrementally in new missile programs but have not displaced HTPB in the large majority of solid propellant applications because HTPB delivers slightly higher specific impulse and has a mature manufacturing base.

Aluminum Hydride

Aluminum hydride, chemical formula $AlH_3$ and molecular weight $30.01$ grams per mole, abbreviated alane, is a metal hydride that contains approximately $10$ percent hydrogen by mass. As a solid propellant fuel additive it delivers substantially higher specific impulse than metallic aluminum, because the hydrogen contributes to the exhaust as low-molecular-weight water vapor rather than remaining as bound aluminum oxide. Theoretical specific impulse gains of $10$ to $20$ seconds over comparable aluminum formulations have been calculated.

Aluminum hydride has been studied since the nineteen sixties as a solid propellant additive but has not seen production use because the compound is unstable at room temperature and decomposes slowly to aluminum and hydrogen gas over months to years of storage. Grain shelf life has been the principal barrier. Research on stabilized formulations continues, particularly for high-performance missile applications where the specific-impulse advantage is worth substantial cost.

Grain Geometry and Thrust Profile

The exposed burning surface of the grain determines the mass flow rate of combustion products and therefore the thrust. As the surface recedes during burning, its area changes, and the thrust changes with it. Grain geometry design produces the thrust profile the mission requires, from the initial spike a booster needs to accelerate against gravity to the sustained neutral thrust an upper stage needs to burn efficiently at altitude.

Four principal thrust-versus-time behaviors are named. A progressive grain has an increasing burning area as it burns, producing increasing thrust. A regressive grain has a decreasing burning area, producing decreasing thrust. A neutral grain has a constant burning area, producing constant thrust. A dual-thrust grain has a boost phase of high thrust followed by a sustain phase of lower thrust, produced by a burning surface that regresses through two distinct area regimes.

A cylindrical end-burning grain, in which combustion propagates from the aft face toward the forward face with the sidewalls insulated, produces a neutral thrust profile because the burning area equals the case cross-section throughout the burn. A central-perforated cylindrical grain, in which combustion propagates radially outward from a cylindrical hole along the axis, produces a progressive thrust profile because the burning area increases with the perforation diameter. A star grain, in which the perforation has a star-shaped cross-section, produces a neutral thrust profile because the tips of the star burn back while the valleys burn forward, and the two rates roughly balance.

Modern boost applications typically use finocyl grains, in which the aft end has star-shaped fins for high initial thrust and the forward end is a plain cylindrical perforation for sustained thrust. The Space Shuttle solid rocket boosters used an eleven-point star aft configuration transitioning to a double-cone cylindrical configuration forward, producing an initial thrust spike of approximately $14.7$ meganewtons that tapered to approximately $10.7$ meganewtons at motor midpoint and recovered to approximately $11.7$ meganewtons before tailoff. This shape produced thrust proportional to atmospheric density during ascent, minimizing max-Q loading on the vehicle.

Performance and Density Specific Impulse

Modern composite solid propellants deliver specific impulse in the range of $240$ to $270$ seconds at sea level and $260$ to $290$ seconds in vacuum. The Space Shuttle solid rocket booster propellant delivered $237$ seconds at sea level and $268$ seconds in vacuum. The Ariane 5 solid rocket booster propellant delivers similar values. Modern strategic missile CMDB propellants deliver $250$ to $270$ seconds in vacuum. The theoretical maximum for a solid propellant, computed for an idealized combination that eliminates two-phase-flow losses, is approximately $310$ seconds vacuum, and no known formulation reaches this value.

Density specific impulse for solid propellants is exceptionally high because the propellant grain density is high. The density specific impulse of a propellant is the product of its bulk density and its specific impulse.

\[I_d = \rho_p \, I_{sp}\]

A typical composite propellant grain has a density $\rho_p$ of approximately $1780$ kilograms per cubic meter and a sea-level specific impulse of approximately $267$ seconds, giving a density specific impulse of approximately $476000$ seconds times kilograms per cubic meter. This value exceeds every liquid propellant combination and is one of the principal reasons solids dominate first-stage boost applications and tactical missiles where volume is more constrained than mass.

The characteristic velocity of a composite solid propellant is approximately $1520$ meters per second at typical chamber conditions. The thrust coefficient is comparable to a liquid engine of the same expansion ratio because the nozzle physics is identical once the combustion products are formed. The specific impulse is lower than a liquid engine of comparable chemistry primarily because the chamber temperature is lower and because the two-phase-flow loss reduces the effective exit velocity.

Tradeoffs

Solid propellants win over liquid propellants on storability, mechanical simplicity, and density specific impulse. They lose on absolute specific impulse, throttleability, restart capability, and exhaust cleanliness.

Storability is the strongest solid-propellant advantage. A well-designed composite motor stored in a controlled environment retains its performance for two to three decades. The principal decomposition mode is slow migration of plasticizer and slow oxidation of the polymer binder at the grain surface. Double-base motors show slower shelf life because of the nitroglycerin decomposition pathway, with typical usable life of one to two decades. This storability advantage is why every American ballistic missile in service uses solid propellant for all propulsion stages.

Mechanical simplicity is the second strong solid-propellant advantage. A solid motor consists of a case, an insulation layer, a grain, an igniter, and a nozzle. There are no pumps, no valves, no cooling passages, no injectors, and no propellant tanks. This makes solid motors substantially cheaper to manufacture per unit thrust and substantially more reliable in service, at the cost of throttling and restart flexibility.

Throttling is the strongest solid-propellant disadvantage. Once ignited, a solid motor produces the thrust its grain geometry dictates and cannot be reduced or shut down. The thrust profile is designed into the grain at manufacture. Some solids can be extinguished by rapid depressurization below a critical pressure, and thrust termination ports on ballistic missile third stages exploit this behavior for range control, but true throttling is unavailable.

Restart is impossible. Once a solid motor has burned out, its grain is consumed. Restart requires a second motor. This limitation prohibits solid propulsion from applications requiring multiple burns separated by coast phases, such as most upper-stage and spacecraft propulsion applications.

Exhaust cleanliness is the second-strongest solid-propellant disadvantage. Composite propellants produce hydrogen chloride and aluminum oxide particles that damage ground equipment near the launch pad, that add weight and volume to the exhaust plume, and that reduce specific impulse through two-phase-flow losses. Double-base propellants avoid the hydrogen chloride and the aluminum oxide but pay for the cleaner exhaust with lower absolute specific impulse. The composite modified double-base combinations recover most of the composite specific impulse advantage while accepting the aluminum oxide penalty.

Safety and detonation hazard is a mixed consideration. Modern composite propellants meet the Insensitive Munitions specifications that require resistance to bullet impact, sympathetic detonation, fuel fire, and slow cook-off. Double-base propellants are less resistant to these insults because of the nitroglycerin content. CMDB propellants require substantial care in production and handling because of the combination of high explosive constituents and metal particles. Absolute safety records depend as much on manufacturing quality and handling procedure as on chemistry.

Cost per unit specific impulse is lowest for composite propellants at production scale. Double-base propellants cost approximately twice as much per unit specific impulse because of the nitrocellulose and nitroglycerin manufacturing base’s higher unit costs. CMDB propellants cost three to five times as much because of the combined handling complexity and the high-explosive constituents. Advanced formulations using ammonium dinitramide or aluminum hydride cost substantially more than either standard family per unit specific impulse.

Applications and Industrial Base

Solid propellants dominate three application categories and appear in several others.

Large space launch boosters are one dominant category. The Space Shuttle solid rocket boosters, the Space Launch System boosters, the Ariane 5 solid rocket boosters, and the Vega first stage all use composite HTPB or PBAN propellants with ammonium perchlorate oxidizer and aluminum fuel. The Delta IV Heavy first stage and the Falcon 9 first stage are notable exceptions using liquid propellants, and the trend in next-generation launch vehicles is toward liquid first stages with smaller solid strap-on boosters where any solids appear at all.

Ballistic missiles are the second dominant category. Every American, Russian, French, British, and Chinese intercontinental ballistic missile in service uses solid propellant for all three stages. The Trident II D5, the Minuteman III, the Sineva, the M51, and the DF-31 all use composite or CMDB propellants. The choice of solid propulsion is driven by the storability advantage, since ballistic missiles must sit ready to fire for years or decades without maintenance.

Tactical missiles are the third dominant category. Air-to-air missiles, air-to-ground missiles, ship-to-ship missiles, and surface-to-air missiles use solid propellants almost exclusively. The AIM-9 Sidewinder, the AIM-120 AMRAAM, the AGM-114 Hellfire, the Harpoon, the Tomahawk boost phase, and the Patriot all use solid propellants. The choice is driven by storability and by the compact thrust-to-volume ratio that solids deliver.

Solid propellants also appear in sounding rockets, some upper stages, gas generators, ejection seat catapults, and hobbyist model rockets. In each case the mechanical simplicity and storability of the solid propellant is worth more than the specific impulse penalty against a liquid alternative.

The industrial base is dominated by a small number of manufacturers. Northrop Grumman, whose solid propulsion heritage runs back through Orbital ATK, ATK, and Thiokol, manufactures the American composite propellant products for space launch and strategic missile applications. Aerojet Rocketdyne manufactures a significant secondary line. Europropulsion, a joint venture of ArianeGroup and Avio, manufactures the European composite propellant products. The Russian manufacturers, formerly consolidated in the Soviet industrial complex, continue in production for Russian domestic requirements. The Chinese and Indian manufacturers serve their domestic requirements.

Conclusion

Solid propellants combine fuel and oxidizer into a single cast grain, producing a mechanically simple engine that delivers high density specific impulse and excellent storability at the cost of lower absolute specific impulse than a liquid engine and the loss of throttling and restart capability. The composite family, dominated by ammonium perchlorate oxidizer with hydroxyl-terminated polybutadiene binder and aluminum fuel, is the current industrial standard. The double-base family retains a smoke-free-exhaust niche. Composite modified double-base combinations serve strategic missile applications where the specific impulse advantage over pure composite justifies the additional cost. The research frontier pursues chlorine-free oxidizers and higher-energy fuels that promise to raise the ceiling on solid-propellant specific impulse in the next decade.

The next article, A219, covers cryogenic liquid propellants.

References