Solid rocket propellants differ from liquid propellants in that the oxidiser and fuel are embedded or bound together in a solid compound that is cast into the rocket motor casing. They began with black powder rockets in medieval times, progressed through double base propellants in the early 1900's, and finally achieved high performance as composite propellants from the 1940's. Composite motors were developed to a high degree of perfection in the United States in the 1950's and 1960's. In Russia, due to a lack of technical leadership and rail handling problems, serious use of composite propellants did not begin until the 1960's, and then primarily for military rockets. The detailed chemistry and development of solid propellants is provided by Andre Bedard in the following separate articles:
Solid propellant rockets, using black powder as the propellant, were introduced by the Chinese in the early 13th century. The next significant event occurred in the late 17th and 18th centuries when the development of nitro-cellulose, nitro-glycerine, cordite, and dynamite resulted in their consideration as a rocket propellant. Immediately before World War I, the French used nitro-cellulose as a propellant for artillery rockets.
In 1936, Dr. Theodore von Karman, and his associates at Caltech began a program that resulted in the first composite propellants using an organic matrix (asphalt) and an inorganic oxidiser (potassium perchlorate). Their work also covered the beginnings of understanding the associated interior ballistics, combustion, ignition, and related structural/materials issues. This was the start of modern solid propellant rocketry. Composite propellants virtually replaced double base propellants (based on mixtures of nitro-cellulose and nitro-glycerine) in most applications.
Following World War II many companies and agencies began propellant development programs involving a wide variety of oxidisers, fuels (binders) and processing methods. In this era, improvements in performance (quantified as specific impulse) were largely achieved by increasing oxidiser loading. Most of the binders were supplied by the rapidly expanding plastics industry.
The ever increasing number of potential missile programs resulted in growing pressure to provide other propellants that had improvements in terms of: performance, structural properties (strength, stability, impact resistance) thermal characteristics (temperature range, cycling), processing, cost, safety, quality, and reliability. In the early 1950s, Atlantic Research invented the use of up to 15 percent powdered aluminium to replace a like amount of oxidiser - giving a performance gain of about 15 percent. Propellant researchers began to understand the complete chemistry of solid propellants, and the need for molecular chain extensions and cross linking of the binders became apparent. The invention of bonding agents (as part of the fuel) greatly improved not only the mechanical properties, but also the resistance to ageing, humidity, and temperature cycling.
Two mainstream composite propellant/binder families emerged (Polyurethane and Polybutadiene), but these were accompanied by a large number of variations and evolutionary products. In addition, there were numerous associated/alternative formulations and concepts tailored to specific missile program requirements. Included among them were: Nitro-polymers, Fluorine based propellants, Beryllium additives, etc. At the same time double base propellants (based on mixtures of nitro-cellulose and nitro-glycerine) continued to evolve and compete. When double base propellants were used to replace conventional binders this resulted in the highest values of specific impulse ever attained.
Aerojet initially concentrated on Polyurethane (PU), and Thiokol favoured Polybutadiene (PB). Thiokol's work included PBAA, a copolymer of Butadiene and Acrylic Acid. This was replaced by PBAN, a terpolymer including Acrylic Acid and Acrylonitrile. Aerojet also conducted considerable development effort in this area, and PBAN was used in Aerojet's 260" space booster.
Several other companies also worked in these and other related areas. For example Phillips Petroleum with Rocketdyne developed Carboxy Terminated Polybutadiene (CTPB) using both a Lithium initiated polymerisation, and a free radical type. These propellants were widely used, but were later overtaken by Hydroxyl Terminated Butadiene (HTBD). By the 1990's Aerojet favoured HTBD and formulations thereof including double base binders.
In addition to the binder evolution, there was a variety of oxidisers to choose from: ammonium and potassium nitrates, perchlorates, and picrates. Perchlorates were generally favoured, but later environmental concerns were expressed at the amount of chlorine compounds (mainly hydrochloric acid) emitted into the atmosphere. One possible solution was the use of a hybrid (liquid and solid) system with a PBAN or similar grain and liquid oxygen as the oxidiser. This also provided a substantial cost saving, and allowed thrust variation and control features that were otherwise not easily achieved.
Paralleling the propellant formulation was development in the design of the propellant grain shape. In most asphalt rockets, the propellant was simply cast into the cylindrical motor chambers (or in some cases into a thin metal jacket which was then inserted into the chamber). Burning occurred only on the exposed aft end of the propellant, resulting in a constant level of thrust. The Aeroplex and other free-standing, rigid cylindrical grains (burning on the inner diameter and outer diameter.) also produced a constant thrust/time curve, because the increase in internal burning surface area just matched the decreasing external surface area.
Case-bonded propellants called for a different configuration of the burning surface. The outside of the propellant was bonded to the chamber and protected it from the hot gases. A simple cylindrical perforation down the centre of the grain would produce a steadily increasing pressure and thrust from very low at start to very high at completion of burning. The solution was to use a central star shaped perforation, which could produce an essentially flat thrust/time curve. The perforation was accomplished by casting the propellant around a core of the desired shape, which was removed after the propellant was completely cured. The tapered rays of the star provided an initial large burning surface, which decreased as the points burned away. Variations in the core geometry allowed a wide range of thrust/time characteristics, to match overall missile requirements.
Additional variations could be achieved by longitudinal variations in the core size and shape, as well as by casting layers of propellant having different characteristics. This latter concept was used for many tactical missiles requiring a boost/sustain thrust curve. For years, grain design was performed by manual geometric manipulation, but computer aided design greatly simplified the task.
The earliest production process for asphalt propellant was actually to hand-stir the ground oxidiser into the heated asphalt. Quality control and consistency were highly questionable, and the safety aspects were in hindsight, terrifying. The immediate solution was to use commercial bread dough mixers in steadily increasing size and robustness. For the more viscous propellant families, much more sturdy mixers were adapted from the tire industry. In addition, the commercially available oxidisers required grinding to achieve the desired fine grain sizes and grain size distribution.
Following fatal accidents in both propellant mixing (asphalt) and oxidiser grinding (potassium perchlorate), production processes were improved to include remote operation, modern instrumentation and control, and a host of other subsystems which significantly improved safety, versatility, and consistency.
The disadvantages of solid propellants in space applications include:
Advantages of solid rocket motors, many of which make them ideal for military applications:
In the United States:
Solid propellants have the fuel and oxidiser embedded in a rubbery matrix. They were developed to a high degree of perfection in the United States in the 1950's and 1960's. In Russia, development was slower, due to a lack of technical leadership in the area and rail handling problems. The disadvantages of solid propellants include:
Advantages of solid rocket motors, many of which make them ideal for military applications:
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TM SAS 73k||73,000||715.00||First Stages||In Production||LK-1||79,000||272||250||Lower Stages||Development||PRD-52||80,000||784.00||First Stages||Out of Production||Kartukov Soyuz SAS||80,100||785.00||First Stages||Out of Production||M55/TX-55/Tu-122||80,700||792.00||262||237||First Stages||Out of Production||GEM 60||86,830||851.50||275||245||First Stages||In Production||MIHT-1||100,000||980.60||263||238||First Stages||In Production||M24||126,984||1,245.30||288||203||Upper Stages||In Production||M-13||128,731||1,262.40||263||238||First Stages||In Production||RSA-4-1||155,000||1,520.00||263||238||First Stages||Out of Production||H-2-0||157,036||1,540.00||273||237||First Stages||In Production||H-2/J-1-1||158,730||1,556.60||273||248||First Stages||In Production||Castor 120||168,000||1,650.00||280||229||First Stages||In Production||S-40TM||212,500||2,083.90||272||204||Upper Stages||In Production||Peackeeper-1||224,796||2,204.40||282||250||First Stages||In Production||Peacekeeper 1||224,796||2,204.50||282||250||First Stages||In Production||SRB-A||230,000||280||First Stages||In Development||S-43||309,000||3,030.20||265||225||First Stages||In Production||S-43TM||327,000||3,206.70||276||170||Upper Stages||In Production||M14||385,488||3,780.30||276||246||First Stages||In Production||PSLV-1||495,590||4,860.00||264||237||First Stages||In Production||UA1205||596,474||5,849.30||263||238||First Stages||Out of Production||UA1206||634,977||6,226.90||265||240||First Stages||Out of Production||P230||660,000||6,472.30||286||259||First Stages||In Production||UA1207||725,732||7,116.90||272||245||First Stages||In Production||USRM||770,975||7,560.50||286||259||First Stages||In Production||UA-156||910,044||8,924.30||263||238||First Stages||Developed to 1966||AJ-260-1/3||1,030,455||10,105.00||275||First Stages||Design concept 1960's||200 inch solid, segment x 4||1,134,000||11,120.00||285||Upper Stages||Study, NASA, 1960||AJ-260X 1/3||1,136,300||11,143.00||263||238||First Stages||Design concept 1960's||SRB||1,174,713||11,519.80||269||237||First Stages||In Production||Redesigned SRM||1,174,736||11,520.00||269||First Stages||In Production||Thiokol 156||1,503,716||14,746.10||263||238||First Stages||Developed to 1966||Hercules||1,587,302||15,565.80||286||259||First Stages||Developed 1995-||AJ-260-2||1,804,460||17,695.30||263||238||First Stages||Developed to 1966||200 inch solid, segment x 6||2,857,000||28,017.00||263||238||First Stages||Out of Production||AJ-260X||3,608,918||35,390.70||263||238||First Stages||Developed to 1966||280 inch solid||4,712,000||46,208.00||265||238||First Stages||Study 1963||300 inch solid||6,485,000||63,595.00||263||234||First Stages||Study 1963||325 in solid||7,041,000||69,047.00||263||238||First Stages||Study General Dynamics 1963|