Solid-propellant rocket

A simplified diagram of a solid-fuel rocket.

1. A solid fuel-oxidizer mixture (propellant) is packed into the rocket, with a cylindrical hole in the middle.
2. An igniter combusts the surface of the propellant.
3. The cylindrical hole in the propellant acts as a combustion chamber.
4. The hot exhaust is choked at the throat, which, among other things, dictates the amount of thrust produced.
5. Exhaust exits the rocket.

A simple solid rocket motor consists of a casing, nozzle, grain (propellant charge), and igniter.

The solid grain mass burns in a predictable fashion to produce exhaust gases, the flow of which is described by Taylor–Culick flow. The nozzle dimensions are calculated to maintain a design chamber pressure, while producing thrust from the exhaust gases.

Once ignited, a simple solid rocket motor cannot be shut off, because it contains all the ingredients necessary for combustion within the chamber in which they are burned. More advanced solid rocket motors can be throttled, and also be extinguished,[4] and then re-ignited by control of the nozzle geometry, or through the use of vent ports. Further, pulsed rocket motors that burn in segments, and that can be ignited upon command are available.

Modern designs may also include a steerable nozzle for guidance, avionics, recovery hardware (parachutes), self-destruct mechanisms, APUs, controllable tactical motors, controllable divert and attitude control motors, and thermal management materials.

A battery of Katyusha rocket launchers fires at German forces during the Battle of Stalingrad, 6 October 1942

Aerojet 260 motor test, 25 September 1965

The medieval Song dynasty Chinese invented a very primitive form of solid-propellant rocket.[5] Illustrations and descriptions in the 14th century Chinese military treatise Huolongjing by the Ming dynasty military writer and philosopher Jiao Yu confirm that the Chinese in 1232 used proto solid propellant rockets then known as "fire arrows" to drive back the Mongols during the Mongol siege of Kaifeng.[6][7] Each arrow took a primitive form of a simple, solid-propellant rocket tube that was filled with gunpowder. One open end allowed the gas to escape and was attached to a long stick that acted as a guidance system for flight direction control.[7][6]

The first rockets with tubes of cast iron were used by the Kingdom of Mysore under Hyder Ali and Tipu Sultan in the 1750s. These rockets had a reach of targets up to a mile and a half away. These were extremely effective in the Second Anglo-Mysore War that ended in a humiliating defeat for the British East India Company. Word of the success of the Mysore rockets against the British triggered research in England, France, Ireland and elsewhere. When the British finally conquered the fort of Srirangapatana in 1799, hundreds of rockets were shipped off to the Royal Arsenal near London to be reverse-engineered. This led to the first industrial manufacture of military rockets with the Congreve rocket in 1804.[8]

In 1921 the Soviet research and development laboratory Gas Dynamics Laboratory began developing solid-propellant rockets, which resulted in the first launch in 1928, that flew for approximately 1,300 metres.[9] These rockets were used in 1931 for the world's first successful use of rockets to assist take-off of aircraft.[10] The research continued from 1933 by the Reactive Scientific Research Institute (RNII) with the development of the RS-82 and RS-132 rockets, including designing several variations for ground-to-air, ground-to-ground, air-to-ground and air-to-air combat.[11] The earliest known use by the Soviet Air Force of aircraft-launched unguided anti-aircraft rockets in combat against heavier-than-air aircraft took place in August 1939, during the Battle of Khalkhin Gol.[11] In June 1938, the RNII began developing a multiple rocket launcher based on the RS-132 rocket.[12] In August 1939, the completed product was the BM-13 / Katyusha rocket launcher. Towards the end of 1938 the first significant large scale testing of the rocket launchers took place, 233 rockets of various types were used. A salvo of rockets could completely straddle a target at a range of 5,500 metres (3.4 mi). By the end of World War II total production of rocket launchers reached about 10,000.[13] with 12 million rockets of the RS type produced for the Soviet armed forces.[14]

In the United States modern castable composite solid rocket motors were invented by the American aerospace engineer Jack Parsons at Caltech in 1942 when he replaced double base propellant with roofing asphalt and potassium perchlorate. This made possible slow-burning rocket motors of adequate size and with sufficient shelf-life for jet-assisted take off applications. Charles Bartley, employed at JPL (Caltech), substituted curable synthetic rubber for the gooey asphalt, creating a flexible but geometrically stable load-bearing propellant grain that bonded securely to the motor casing. This made possible much larger solid rocket motors. Atlantic Research Corporation significantly boosted composite propellant I_sp in 1954 by increasing the amount of powdered aluminium in the propellant to as much as 20%.[15]

Solid-propellant rocket technology got its largest boost in technical innovation, size and capability with the various mid-20th century government initiatives to develop increasingly capable military missiles. After initial designs of ballistic missile military technology designed with liquid-propellant rockets in the 1940s and 1950s, both the Soviet Union and the United States embarked on major initiatives to develop solid-propellant local, regional, and intercontinental ballistic missiles, including solid-propellant missiles that could be launched from air or sea. Many other governments also developed these military technologies over the next 50 years.

By the later 1980s and continuing to 2020, these government-developed highly-capable solid rocket technologies have been applied to orbital spaceflight by many government-directed programs, most often as booster rockets to add extra thrust during the early ascent of their primarily liquid rocket launch vehicles. Some designs have had solid rocket upper stages as well. Examples flying in the 2010s include the European Ariane 5, US Atlas V and Space Shuttle, and Japan's H-II.

The largest solid rocket motors ever built were Aerojet's three 6.60-meter (260 in) monolithic solid motors cast in Florida.[16] Motors 260 SL-1 and SL-2 were 6.63 meters (261 in) in diameter, 24.59 meters (80 ft 8 in) long, weighed 842,900 kilograms (1,858,300 lb), and had a maximum thrust of 16 MN (3,500,000 lbf). Burn duration was two minutes. The nozzle throat was large enough to walk through standing up. The motor was capable of serving as a 1-to-1 replacement for the 8-engine Saturn I liquid-propellant first stage but was never used as such. Motor 260 SL-3 was of similar length and weight but had a maximum thrust of 24 MN (5,400,000 lbf) and a shorter duration.

Design begins with the total impulse required, which determines the fuel and oxidizer mass. Grain geometry and chemistry are then chosen to satisfy the required motor characteristics.

The following are chosen or solved simultaneously. The results are exact dimensions for grain, nozzle, and case geometries:

• The grain burns at a predictable rate, given its surface area and chamber pressure.[17]
• The chamber pressure is determined by the nozzle throat diameter and grain burn rate.
• Allowable chamber pressure is a function of casing design.
• The length of burn time is determined by the grain "web thickness".

The grain may or may not be bonded to the casing. Case-bonded motors are more difficult to design, since the deformation of the case and the grain under flight must be compatible.

Common modes of failure in solid rocket motors include fracture of the grain, failure of case bonding, and air pockets in the grain. All of these produce an instantaneous increase in burn surface area and a corresponding increase in exhaust gas production rate and pressure, which may rupture the casing.

Another failure mode is casing seal failure. Seals are required in casings that have to be opened to load the grain. Once a seal fails, hot gas will erode the escape path and result in failure. This was the cause of the Space Shuttle Challenger disaster.

Solid rocket fuel 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. Since the propellant volume is equal to the cross sectional area ${\displaystyle A_{s}}$ times the fuel length, the volumetric propellant consumption rate is the cross section area times the linear burn rate ${\displaystyle {\dot {b}}}$, and the instantaneous mass flow rate of combustion gases generated is equal to the volumetric rate times the fuel density ${\displaystyle \rho }$:

${\displaystyle {\dot {m}}=\rho \cdot A_{s}\cdot {\dot {b}}}$

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 other producing steady long burn, though has thermal difficulties, center of gravity (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.

The casing may be constructed from a range of materials. Cardboard is used for small black powder model motors, whereas aluminium is used for larger composite-fuel hobby motors. Steel was used for the space shuttle boosters. Filament-wound graphite epoxy casings are used for high-performance motors.

The casing must be designed to withstand the pressure and resulting stresses of the rocket motor, possibly at elevated temperature. For design, the casing is considered a pressure vessel.

To protect the casing from corrosive hot gases, a sacrificial thermal liner on the inside of the casing is often implemented, which ablates to prolong the life of the motor casing.

A convergent-divergent design accelerates the exhaust gas out of the nozzle to produce thrust. The nozzle must be constructed from a material that can withstand the heat of the combustion gas flow. Often, heat-resistant carbon-based materials are used, such as amorphous graphite or carbon-carbon.

Some designs include directional control of the exhaust. This can be accomplished by gimballing the nozzle, as in the Space Shuttle SRBs, by the use of jet vanes in the exhaust as in the V-2 rocket, or by liquid injection thrust vectoring (LITV).

LITV consists of injecting a liquid into the exhaust stream after the nozzle throat. The liquid then vaporizes, and in most cases chemically reacts, adding mass flow to one side of the exhaust stream and thus providing a control moment. For example, the Titan IIIC solid boosters injected nitrogen tetroxide for LITV; the tanks can be seen on the sides of the rocket between the main center stage and the boosters.[18]

An early Minuteman first stage used a single motor with four gimballed nozzles to provide pitch, yaw, and roll control.

An exhaust cloud engulfs Launch Pad 39A at NASA's Kennedy Space Center as the Space Shuttle Endeavour lifts off.

A typical, well-designed ammonium perchlorate composite propellant (APCP) first-stage motor may have a vacuum specific impulse (I_sp) as high as 285.6 seconds (2.801 km/s) (Titan IVB SRMU).[19] This compares to 339.3 s (3.327 km/s) for RP1/LOX (RD-180)[20] and 452.3 s (4.436 km/s) for LH₂/LOX (Block II RS-25)[21] bipropellant engines. Upper stage specific impulses are somewhat greater: as much as 303.8 s (2.979 km/s) for APCP (Orbus 6E),[22] 359 s (3.52 km/s) for RP1/LOX (RD-0124)[23] and 465.5 s (4.565 km/s) for LH₂/LOX (RL10B-2).[24]

Propellant fractions are usually somewhat higher for (non-segmented) solid propellant first stages than for upper stages. The 53,000-kilogram (117,000 lb) Castor 120 first stage has a propellant mass fraction of 92.23% while the 14,000-kilogram (31,000 lb) Castor 30 upper stage developed for Orbital Science's Taurus II COTS (Commercial Off The Shelf) (International Space Station resupply) launch vehicle has a 91.3% propellant fraction with 2.9% graphite epoxy motor casing, 2.4% nozzle, igniter and thrust vector actuator, and 3.4% non-motor hardware including such things as payload mount, interstage adapter, cable raceway, instrumentation, etc. Castor 120 and Castor 30 are 2.36 and 2.34 meters (93 and 92 in) in diameter, respectively, and serve as stages on the Athena IC and IIC commercial launch vehicles. A four-stage Athena II using Castor 120s as both first and second stages became the first commercially developed launch vehicle to launch a lunar probe (Lunar Prospector) in 1998.

Solid rockets can provide high thrust for relatively low cost. For this reason, solids have been used as initial stages in rockets (for example the Space Shuttle), while reserving high specific impulse engines, especially less massive hydrogen-fueled engines, for higher stages. In addition, solid rockets have a long history as the final boost stage for satellites due to their simplicity, reliability, compactness and reasonably high mass fraction.[25] A spin-stabilized solid rocket motor is sometimes added when extra velocity is required, such as for a mission to a comet or the outer solar system, because a spinner does not require a guidance system (on the newly added stage). Thiokol's extensive family of mostly titanium-cased Star space motors has been widely used, especially on Delta launch vehicles and as spin-stabilized upper stages to launch satellites from the cargo bay of the Space Shuttle. Star motors have propellant fractions as high as 94.6% but add-on structures and equipment reduce the operating mass fraction by 2% or more.

Higher performing solid rocket propellants are used in large strategic missiles (as opposed to commercial launch vehicles). HMX, C₄H₈N₄(NO₂)₄, a nitramine with greater energy than ammonium perchlorate, was used in the propellant of the Peacekeeper ICBM and is the main ingredient in NEPE-75 propellant used in the Trident II D-5 Fleet Ballistic Missile.[26] It is because of explosive hazard that the higher energy military solid propellants containing HMX are not used in commercial launch vehicles except when the LV is an adapted ballistic missile already containing HMX propellant (Minotaur IV and V based on the retired Peacekeeper ICBMs).[27] The Naval Air Weapons Station at China Lake, California, developed a new compound, C₆H₆N₆(NO₂)₆, called simply CL-20 (China Lake compound ^#20). Compared to HMX, CL-20 has 14% more energy per mass, 20% more energy per volume, and a higher oxygen-to-fuel ratio.[28] One of the motivations for development of these very high energy density military solid propellants is to achieve mid-course exo-atmospheric ABM capability from missiles small enough to fit in existing ship-based below-deck vertical launch tubes and air-mobile truck-mounted launch tubes. CL-20 propellant compliant with Congress' 2004 insensitive munitions (IM) law has been demonstrated and may, as its cost comes down, be suitable for use in commercial launch vehicles, with a very significant increase in performance compared with the currently favored APCP solid propellants. With a specific impulse of 309 s already demonstrated by Peacekeeper's second stage using HMX propellant, the higher energy of CL-20 propellant can be expected to increase specific impulse to around 320 s in similar ICBM or launch vehicle upper stage applications, without the explosive hazard of HMX.[29]

An attractive attribute for military use is the ability for solid rocket propellant to remain loaded in the rocket for long durations and then be reliably launched at a moment's notice.

Solid propellant rocket motors can be bought for use in model rocketry; they are normally small cylinders of black powder fuel with an integral nozzle and optionally a small charge that is set off when the propellant is exhausted after a time delay. This charge can be used to trigger a camera, or deploy a parachute. Without this charge and delay, the motor may ignite a second stage (black powder only).

In mid- and high-power rocketry, commercially made APCP motors are widely used. They can be designed as either single-use or reloadables. These motors are available in impulse ranges from "A" (1.26 Ns– 2.50 Ns) to "O" (20.48 kNs – 40.96 kNs), from several manufacturers. They are manufactured in standardized diameters and varying lengths depending on required impulse. Standard motor diameters are 13, 18, 24, 29, 38, 54, 75, 98, and 150 millimeters. Different propellant formulations are available to produce different thrust profiles, as well as special effects such as colored flames, smoke trails, or large quantities of sparks (produced by adding titanium sponge to the mix).

• Environmentally sensitive fuel formulations such as ALICE propellant
• Ramjets with solid fuel
• Variable thrust designs based on variable nozzle geometry
• Hybrid rockets that use solid fuel and throttleable liquid or gaseous oxidizer

• A. Davenas, ed. (1992). Solid Rocket Propulsion Technology. Pergamon. ISBN 978-0080409993.

• Robert A. Braeunig rocket propulsion page
• Astronautix Composite Solid Propellants
• Ariane 5 SRB
• Amateur High Power Rocketry Association
• Nakka-Rocketry (Design Calculations and Propellant Formulations)
• 5 cent sugar rocket
• Practical Rocketry Archived 2006-06-18 at the Wayback Machine
• NASA Practical Rocketry