Ammonium perchlorate composite propellant

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Ammonium perchlorate composite propellant (APCP) is a modern solid-fuel rocket used in rocket vehicles. It differs from many traditional solid rocket propellants such as black powder or zinc-sulfur, not only in chemical composition and overall performance, but also by the nature of how it is processed. APCP is cast into shape, as opposed to powder pressing as with black powder. This provides manufacturing regularity and repeatability, which are necessary requirements for use in the aerospace industry.


Ammonium perchlorate composite propellant is typically used in aerospace propulsion applications, where simplicity and reliability are desired and specific impulses (depending on the composition and operating pressure) of 180–260 seconds are adequate. Because of these performance attributes, APCP is regularly implemented in booster applications such as in the Space Shuttle Solid Rocket Boosters, aircraft ejection seats, and specialty space exploration applications such as NASA's Mars Exploration Rover descent stage retrorockets. In addition, the high-power rocketry community regularly uses APCP in the form of commercially available propellant "reloads", as well as single-use motors. Experienced experimental and amateur rocketeers also often work with APCP, processing the APCP themselves.



Ammonium perchlorate composite propellant is a composite propellant, meaning that it has both fuel and oxidizer mixed with a rubbery binder, all combined into a homogeneous mixture. The propellant is most often composed of ammonium perchlorate (AP), an elastomer binder such as hydroxyl-terminated polybutadiene (HTPB) or polybutadiene acrylic acid acrylonitrile prepolymer (PBAN), powdered metal (typically aluminum), and various burn rate catalysts. In addition, curing additives induce elastomer binder cross-linking to solidify the propellant before use. The perchlorate serves as the oxidizer, while the binder and aluminum serve as the fuel. Burn rate catalysts determine how quickly the mixture burns. The resulting cured propellant is fairly elastic (rubbery), which also helps limit fracturing during accumulated damage (such as shipping, installing, cutting) and high acceleration applications such as hobby or military rocketry.

The composition of APCP can vary significantly depending on the application, intended burn characteristics, and constraints such as nozzle thermal limitations or specific impulse (Isp). Rough mass proportions (in high performance configurations) tend to be about 70/15/15 AP/HTPB/Al, though fairly high performance "low-smoke" can have compositions of roughly 80/18/2 AP/HTPB/Al. While metal fuel is not required in APCP, most formulations include at least a few percent as a combustion stabilizer, propellant opacifier (to limit excessive infrared propellant preheating), and increase the temperature of the combustion gases (increasing Isp).

Common species


High energy fuels:

Low energy fuels acting as binders:

Special considerations

Though increasing the ratio of metal fuel to oxidizer up to the stoichiometric point increases the combustion temperature, the presence of an increasing molar fraction of metal oxides, particularly aluminum oxide (Al2O3) precipitating from the gaseous solution creates globules of solids or liquids that slow down the flow velocity as the mean molecular mass of the flow increases. In addition, the chemical composition of the gases change, varying the effective heat capacity of the gas. Because of these phenomena, there exists an optimal non-stoichiometric composition for maximizing Isp of roughly 16% by mass, assuming the combustion reaction goes to completion inside the combustion chamber.

The combustion time of the aluminum particles in the hot combustion gas varies depending on aluminum particle size and shape. In small APCP motors with high aluminum content, the residence time of the combustion gases does not allow for full combustion of the aluminum and thus a substantial fraction of the aluminum is burned outside the combustion chamber, leading to decreased performance. This effect is often mitigated by reducing aluminum particle size, inducing turbulence (and therefore a long characteristic path length and residence time), and/or by reducing the aluminum content to ensure a combustion environment with a higher net oxidizing potential, ensuring more complete aluminum combustion. Aluminum combustion inside the motor is the rate-limiting pathway since the liquid-aluminum droplets (even still liquid at temperatures 3000 K) limit the reaction to a heterogeneous globule interface, making the surface area to volume ratio an important factor in determining the combustion residence time and required combustion chamber size/length.

Particle size

The propellant particle size distribution has a profound impact on APCP rocket motor performance. Smaller AP and Al particles lead to higher combustion efficiency but also lead to increased linear burn rate. The burn rate is heavily dependent on mean AP particle size as the AP absorbs heat to decompose into a gas before it can oxidize the fuel components. This process may be a rate-limiting step in the overall combustion rate of APCP. The phenomenon can be explained by considering the heat-flux-to-mass ratio: As the particle radius increases the volume (and, therefore, mass and heat capacity) increase as the cube of the radius. However, the surface area increases as the square of the radius, which is roughly proportional to the heat flux into the particle. Therefore, a particle's rate of temperature rise is maximized when the particle size is minimized.

Common APCP formulations call for 30-400 µm AP particles (often spherical), as well as 2–50 µm Al particles (often spherical). Because of the size discrepancy between the AP and Al, Al will often take an interstitial position in a pseudo-lattice of AP particles.



APCP deflagrates from the surface of exposed propellant in the combustion chamber. In this fashion, the geometry of the propellant inside the rocket motor plays an important role in the overall motor performance. As the surface of the propellant burns the shape evolves (a subject of study in internal ballistics), most often changing the propellant surface area exposed to the combustion gases. The mass flux (kg/s) [and therefore pressure] of combustion gases generated is a function of the instantaneous surface area (m2), propellant density (kg/m3), and linear burn rate (m/s):

Several geometric configurations are often used depending on the application and desired thrust curve:

  • Circular bore: if in BATES configuration, produces progressive-regressive thrust curve.
  • End burner: propellant burns from one axial end to other producing steady long burn, though has thermal difficulties, CG shift.
  • C-slot: propellant with large wedge cut out of side (along axial direction), producing fairly long regressive thrust, though has thermal difficulties and asymmetric CG characteristics.
  • Moon burner: off-center circular bore produces progressive-regressive long burn though has slight asymmetric CG characteristics.
  • Finocyl: usually a 5 or 6 legged star-like shape that can produce very level thrust, with a bit quicker burn than circular bore due to increased surface area.

Burn rate

While the surface area can be easily tailored by careful geometric design of the propellant, the burn rate is dependent on several subtle factors:

  • Propellant chemical composition.
  • AP, Al, additive particle sizes.
  • Combustion pressure.
  • Heat transfer characteristics.
  • Erosive burning (high velocity flow moving past the propellant).
  • Initial temperature of propellant.

In summary, however, most formulations have a burn rate between 1–3 mm/s at STP and 6–12 mm/s at 68 atm. The burn characteristics (such as linear burn rate) are often determined prior to rocket motor firing using a strand burner test. This test allows the APCP manufacturer to characterize the burn rate as a function of pressure. Empirically, APCP adheres fairly well to the following power-function model:

It is worth noting that typically for APCP, 0.3<n<0.5 indicating that APCP is sub-critically pressure sensitive. That is, if surface area were maintained constant during a burn the combustion reaction would not runaway to (theoretically) infinite as the pressure would reach an internal equilibrium. This isn't to say that APCP cannot cause an explosion, but rather that the explosion would be caused by the pressure surpassing the burst pressure of the container (rocket motor).

Model/high-power rocketry applications

A high-power rocket launch using an APCP motor

Commercial APCP rocket engines usually come in the form of reloadable motor systems (RMS) and fully assembled single-use rocket motors. For RMS, the APCP "grains" (cylinders of propellant) are loaded into the reusable motor casing along with a sequence of insulator disks and o-rings and a (graphite or glass-filled phenolic resin) nozzle. The motor casing and closures are typically bought separately from the motor manufacturer and are often precision-machined from aluminum. The assembled RMS contains both reusable (typically metal) and disposable components.

The major APCP suppliers for hobby use are:

To achieve different visual effects and flight characteristics, hobby APCP suppliers offer a variety of different characteristic propellant types. These can range from fast-burning with little smoke and blue flame to classic white smoke and white flame. In addition, colored formulations are available to display reds, greens, blues, and even black smoke.

In medium- and high-power rocket applications, APCP has largely replaced black powder as a rocket propellant. Compacted black powder slugs become prone to fracture in larger applications, which can result in catastrophic failure in rocket vehicles. APCP's elastic material properties make it less vulnerable to fracture from accidental shock or high-acceleration flights. Due to these attributes, widespread adoption of APCP and related propellant types in the hobby has significantly enhanced the safety of rocketry.

Environmental and other concerns

The exhaust from APCP solid rocket motors contains mostly water, carbon dioxide, hydrogen chloride, and a metal oxide (typically aluminium oxide). The hydrogen chloride can easily dissolve in water and create corrosive hydrochloric acid. The environmental fate of the hydrogen chloride is not well documented. The hydrochloric acid component of APCP exhaust leads to the condensation of atmospheric moisture in the plume and this enhances the visible signature of the contrail. This visible signature, among other reasons, led to research in cleaner burning propellants with no visible signatures. Minimum signature propellants contain primarily nitrogen-rich organic molecules (e.g., ammonium dinitramide) and depending on their oxidizer source can be hotter burning than APCP composite propellants.

Regulation and legality

In the United States, APCP for hobby use is regulated indirectly by two non-government agencies: the National Association of Rocketry (NAR), and the Tripoli Rocketry Association (TRA). Both agencies set forth rules regarding the impulse classification of rocket motors and the level of certification required by rocketeers in order to purchase certain impulse (size) motors. The NAR and TRA require motor manufactures to certify their motors for distribution to vendors and ultimately hobbyists. The vendor is charged with the responsibility (by the NAR and TRA) to check hobbyists for high power rocket certification before a sale can be made. The amount of APCP that can be purchased (in the form of a rocket motor reload) correlates to the impulse classification, and therefore the quantity of APCP purchasable by hobbyist (in any single reload kit) is regulated by the NAR and TRA.

The overarching legality concerning the implementation of APCP in rocket motors is outlined in NFPA 1125. Use of APCP outside hobby use is regulated by state and municipal fire codes. On March 16, 2009, it was ruled that APCP is not an explosive and that manufacture and use of APCP no longer requires a license or permit from the ATF.[1]




  • Rocket Propulsion Elements. Sutton, George P.
  • Amateur Experimental Solid Propellants by Richard Nakka
  • Solid Propellant Burn Rate by Richard Nakka
  • Intro to Solid Propulsion by Graham Orr, Harvey Mudd College Experimental Engineering
  • BATFE Lawsuit Documents, 2002–Present, Tripoli Rocketry Association
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