Ammonium perchlorate composite propellant

Ammonium perchlorate composite propellant is a composite propellant, meaning that it has both fuel and oxidizer combined into an homogeneous mixture, in this case with a rubbery binder as part of the fuel. 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 aluminium), 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 aluminium 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. This includes the Space Shuttle missions, in which APCP was used for the two SRBs.

The composition of APCP can vary significantly depending on the application, intended burn characteristics, and constraints such as nozzle thermal limitations or specific impulse (I_sp). 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 I_sp).

Oxidizers:

• Ammonium perchlorate as the primary oxidizer
• Metal-oxide catalysts as thermite oxidizers

High energy fuels:

• Aluminium (high performance, most common)
• Magnesium (medium performance)
• Zinc (low performance)

Low energy fuels acting as binders:

• Hydroxyl-terminated polybutadiene (HTPB)
• Carboxyl-terminated polybutadiene (CTPB)
• Polybutadiene acrylonitrile (PBAN)

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 aluminium 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 changes, 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 aluminium particles in the hot combustion gas varies depending on aluminium particle size and shape. In small APCP motors with high aluminium content, the residence time of the combustion gases does not allow for full combustion of the aluminium and thus a substantial fraction of the aluminium is burned outside the combustion chamber, leading to decreased performance. This effect is often mitigated by reducing aluminium particle size, inducing turbulence (and therefore a long characteristic path length and residence time), and/or by reducing the aluminium content to ensure a combustion environment with a higher net oxidizing potential, ensuring more complete aluminium combustion. Aluminium combustion inside the motor is the rate-limiting pathway since the liquid-aluminium 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.

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) increases 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 ${\displaystyle A_{\text{s}}}$ (m²), propellant density ${\displaystyle \rho }$ (kg/m³), and linear burn rate ${\displaystyle b_{r}}$ (m/s):

${\displaystyle {\dot {m}}=\rho A_{\text{s}}b_{r}}$

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

• 
  Circular bore simulation
• 
  C-Slot simulation
• 
  Moon burner simulation
• 
  5-point finocyl simulation

• Circular bore: if in BATES configuration, produces progressive-regressive thrust curve.
• End burner: propellant burns from one axial end to another producing steady long burn, though has thermal difficulties, CG shift.
• C-slot: propellant with large wedge cut out of the 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.

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:

${\displaystyle b_{r}=ap^{n}}$

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