SOLID ROCKET PROPELANTS

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:

The following summarises the development of solid rocket propellants very briefly.

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:

Slightly higher empty mass for the rocket stage
Slightly lower performance than storable liquid propellants
Transportability issues: Solid propellants are cast into the motor in the factory, unlike liquid fuel rockets which can be fuelled at the launch pad. This means they have to either be: 1) limited in size to be transportable (as for the Delta and Ariane strap-on motors); 2) cast in segments, with the segments assembled at the launch base (as for Titan and the Space Shuttle); or 3) cast in a factory near the launch site (actually done for large test motors intended for Saturn V upgrades).
Once ignited, they cannot be easily shut down or throttled. Thereafter they have to be pre-cast or milled out for a specific mission.
Often catastrophic results in the event of a failure
Advantages of solid rocket motors, many of which make them ideal for military applications:
High density and low volume
Nearly indefinite storage life
Instant ignition without fuelling operations
High reliability

Progressive Development of Large Solid Rocket Motors. In the United States:

Early 1950’s: Hermes/ Sergeant (Army): 32 inch
March 1956: Polaris (Navy): 54 inch
Late 1950’s: Minuteman (Air Force): 65 inch (Thiokol)
1960-1963: USAF space development program – 86, 96, 100, 120 inch test motors
Early 1960’s: Titan 3 (DoD/NASA) – 120 inch (UTC)
1963-1965: Moon program (DoD/NASA) for Nova/Saturn vehicles
1. 44 inch (Aerojet), 65 (Thiokol), 120 inch (Aerojet) subscale motors
2. 156 inch (Lockheed and Thiokol), 260 inch (Thiokol and Aerojet) PBAN full scale motors.
Early 1970’s: Shuttle – 146 inch PBAN
1990’s: Titan 4B USRM – 126 inch HTPB

Fuel: Solid. Fuel Density: 1.35 g/cc.

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:

Slightly higher empty mass for the rocket stage
Slightly lower performance than storable liquid propellants
Transportability issues: Solid propellants are cast into the motor in the factory, unlike liquid fuel rockets which can be fueled at the launch pad. This means they have to either be: 1) limited in size to be transportable (as for the Delta and Ariane strap-on motors); 2) cast in segments, with the segments assembled at the launch base (as for Titan and the Space Shuttle); or 3) cast in a factory at the launch site (actually done for large test motors intended for Saturn V upgrades).
Once ignited, they cannot be easily shut down or throttled. Thereafter they have to be pre-cast or milled out for a specific mission.
Nearly always catastrophic results in the event of a failure

Advantages of solid rocket motors, many of which make them ideal for military applications:

High density and low volume

Nearly indefinite storage life

Instant ignition without fuelling operations
High reliability

Eng-engineslink	Thrust (vac)-kgf	Thrust(vac)-kN	Isp-sec	Isp (sealevel)-sec	Designed for	Status
15D305	1	0.01			First Stages	Out of Production
Star 5A	17		250		Upper Stages	In Production
NOTS-4	72	0.70	250	204	Upper Stages	Out of Production
Star 5C/CB	199		266		Upper Stages	In Production
Star 5CB	203	2.00			Upper Stages	In Production
Star 5C	208				Upper Stages	Out of Production
NOTS-3	231	2.26	250	204	Upper Stages	Out of Production
Star 6B	256		273		Upper Stages	In Production
Star 10	342		251		Upper Stages	In Production
Star 13	387	3.80	273		Upper Stages	Out of Production
PRD-22	400	3.92			First Stages	Out of Production
Martlet 4-3	550	5.39	300	210	Upper Stages	Study 1962
Star 12	567		252		Upper Stages	In Production
Star 13A	599		287		Upper Stages	In Production
Sergeant	680	6.66	235	214	Upper Stages	Out of Production
Star 12A	739		270		Upper Stages	In Production
Star 13B	775		286		Upper Stages	Out of Production
MIHT-4	1,000	9.80	295		Upper Stages	In Production
Star 17	1,116	19.60	286	220	Upper Stages	In Production
GCRC	1,179	11.60	230	210	Upper Stages	Out of Production
X-248	1,270	12.40	256	233	Upper Stages	Out of Production
X-248A	1,406	13.80	255	232	Upper Stages	Out of Production
Kartukov LL	1,500	14.70			First Stages	Developed 1946-48
Star 17A	1,633		287		Upper Stages	In Production
Star 24	1,891	20.00	283		Upper Stages	In Production
Mage 1	1,978	19.40	295	220	Upper Stages	Out of Production
Martlet 4-2	2,100	20.60	300	210	Upper Stages	Study 1962
Star 24C	2,189		282		Upper Stages	In Production
X-258	2,268	22.20	266	242	Upper Stages	Out of Production
Star 20B	2,495		289		Upper Stages	In Production
SPRD-99	2,500	24.50			First Stages	Out of Production
FW4-D	2,549	25.00	287	250	Upper Stages	Out of Production
Star 27	2,726	27.00	288		Upper Stages	In Production
SLV-4	2,736	26.80	283	60	Upper Stages	In Production
Star 30BP	2,753	27.00	292		Upper Stages	In Production
Star 20	2,767		287		Upper Stages	In Production
FW-4S,TEM640	2,800	27.40	280	255	Upper Stages	Out of Production
GF-02	2,957	29.00	230	200	Upper Stages	Out of Production
P6	3,000	29.40	211	211	Upper Stages	Out of Production
Iris	3,000	29.40	291	115	Upper Stages	Out of Production
Black Arrow-3	3,000	29.40	278	245	Upper Stages	Out of Production
Star 30C	3,329	1,647.00	287		Upper Stages	In Production
Star 26	3,402	39.10	271	220	Upper Stages	In Production
Pegasus-3	3,525	34.60	287		Upper Stages	In Production
Star 26B	3,531		272		Upper Stages	In Production
Star 26C	3,570		272		Upper Stages	In Production
Star 30E	3,608	1,780.00	291		Upper Stages	In Production
Star 37XFP	3,878	31.50	290		Upper Stages	In Production
Star 37	4,441	43.50	260	220	Upper Stages	Out of Production
Mage 2	4,638	45.50	293		Upper Stages	Out of Production
Star 37FM	4,819	47.90	290		Upper Stages	In Production
Star 37X	5,216	51.10	296	230	Upper Stages	Out of Production
M-V-4	5,300	52.00	298		Upper Stages	In Production
RSA-3-3	5,300	51.00	292		Upper Stages	Out of Production
NOTS-1	5,441	53.40	204	204	First Stages	Out of Production
X-254	6,169	60.50	256	233	Upper Stages	Out of Production
Star 48	6,848	67.20	287		Upper Stages	Out of Production
Martlet 4-1	6,900	67.70	300	210	First Stages	Study 1962
Star 48A s	7,863		283		Upper Stages	In Production
Star 48B s	7,863		286		Upper Stages	In Production
UM-129A	7,900	77.50	291	220	Upper Stages	In Production
H-1-3	7,900	77.00	291	220	Upper Stages	Out of Production
SRM-2	7,996	78.40	304	200	Upper Stages	In Production
Star 48B	8,044	66.00	286		Upper Stages	In Production
X-259A	8,239	80.80	295		Upper Stages	Out of Production
Star 31	8,391	80.00	294		Upper Stages	In Production
Star 63D	8,641		283		Upper Stages	In Production
SLV-3	9,249	90.70	277	190	Upper Stages	In Production
X-259	9,493	93.10	293	233	Upper Stages	Out of Production
Star 63F	10,669		297		Upper Stages	In Production
Star 63	10,931	107.20	282		Upper Stages	In Production
Pegasus-2	12,053	118.20	290		Upper Stages	In Production
M-3B-J	13,469	132.10	294		Upper Stages	Developed 1995-
M-3B-Mu	13,472	132.10	294		Upper Stages	In Production
Pegasus XL-2	15,653	153.50	290	240	Upper Stages	In Production
Kartukov Soyuz T – TM SAS 17k	17,500	171.00			First Stages	In Production
EPKM	18,000		292		Upper Stages	Hardware
P4	18,000	176.50	273	240	Upper Stages	Out of Production
Kartukov P-5	18,300	179.00			First Stages	Out of Production
SRM-1	18,508	181.50	296	115	Upper Stages	In Production
Star 75	20,511	242.80	288		Upper Stages	In Production
M56A-1	23,300	228.50	297	270	Upper Stages	Out of Production
MIHT-3	25,000	245.20	280	220	Upper Stages	In Production
TX-354-3	26,402	258.90	262	232	First Stages	In Production
SLV-2	27,227	267.00	267	216	Upper Stages	In Production
M33-20-4	29,164	286.00	247	232	First Stages	Out of Production
SPRD-30	30,000	294.00			First Stages	Out of Production
Kartukov P-35	30,000	294.00			First Stages	Out of Production
M34	30,000	294.20	301		Upper Stages	In Production
SB-735	33,430	327.80	263	238	First Stages	In Production
PSLV-3	33,519	328.70	291	160	Upper Stages	In Production
S-44	33,900	332.40	282		Upper Stages	In Production
PRD-15	40,000	392.00			First Stages	Out of Production
SPRD-15	41,000	402.00			First Stages	Out of Production
Castor 4	41,524	407.20	261	228	First Stages	Out of Production
Castor 4BXL	43,746	429.00	267		Upper Stages	In Production
Castor 4B	43,910	430.60	281	220	Upper Stages	In Production
Algol 3A	47,387	464.70	259	226	First Stages	In Production
Algol 1	48,022	470.90	236	214	First Stages	Out of Production
Algol 3	48,121	471.90	284	238	First Stages	In Production
Castor 4A	48,774	478.30	266	237	First Stages	In Production
Pegasus-1	49,447	484.90	285	180	First Stages	In Production
MIHT-2	50,000	490.30	280	220	Upper Stages	In Production
GEM 40	50,905	499.20	274	245	First Stages	In Production
RSA-3-1	51,000	500.00	273	230	First Stages	Out of Production
SLV-1	51,251	502.60	253	229	First Stages	In Production
RSA-3-2	53,000	519.00	284	220	Upper Stages	Developed -1995
M-23-Mu	53,433	524.00	285	220	Upper Stages	In Production
M-23-J	53,515	524.80	282		Upper Stages	Developed 1995-
Algol 2	57,537	564.20	255	232	First Stages	Out of Production
Pegasus XL-1	60,062	589.00	293	180	First Stages	In Production
Castor 4AXL	61,164	599.80	269		Upper Stages	In Production
GEM 46	62,000	608.10	274	242	First Stages	In Production
RSA-4-2	69,000	676.00	275	220	Upper Stages	Out of Production
SPB 7.35	70,360	690.00	263	240	First Stages	Out of Production
P9.5	70,360	690.00	263	240	First Stages	In Production
Kartukov Soyuz T – 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