Solid Fuel Rocket Fundamentals

A solid fuel rocket is distinguished from a liquid fuel rocket by the type of fuel that it uses. It is more accurate to refer to the two basic types of rockets as solid propellant and liquid propellant rockets. Both types of rockets generate thrust by exhausting combustion products through a supersonic nozzle. Rocket propellants contain both a fuel and an oxidizer to produce energy. Both carry their own oxygen, so they can operate above the atmosphere and can be used in space.

A solid fuel rocket, commonly called a solid rocket motor, is a completely self-contained device that converts chemical energy into kinetic energy in a controlled way. Although there is much to know about the science, engineering, and manufacture of solid rockets, they are simple devices. Solid rockets consist of four main components: (1) propellant grain, (2) a case that is the thrust chamber that contains the pressurized combustion gases, (3) a nozzle for directing and accelerating the gases away from the motor, and (4) an igniter.

The propellant grain consists of the propellant charge shaped to deliver the desired thrust profile. The case contains the pressure of propellant combustion and is frequently a major portion of the vehicle airframe. Insulation is necessary to protect the case from the high-temperature combustion products. The igniter provides the heat necessary to initiate combustion at the propellant surface. The nozzle directs and accelerates the propellant exhaust gases.

The propellant is typically a solid rubberlike material, similar to a pencil eraser, that contains fuel and oxidizer particles. The propellant grain can have a wide variety of shapes, but it is generally a hollow cylinder designed to burn from the inside out, thereby not exposing the case to the extreme temperatures until near the end of the burn. The propellant is bonded to the case.

The rate at which the propellant burns and thereby generates thrust is designed into each propellant formulation. For instance, propellants that have fine oxidizer particles burn at higher rates than formulations that contain coarse oxidizer particles. Burn rate also varies with combustion chamber pressure, initial temperature of the propellant grain, and other factors.

The internal shape of the grain, which can vary the amount of exposed surface area, hence the burn rate, is typically established when it is cast (poured) and cured inside the case. An igniter, generally located in the head end and fires down the bore or grain center perforation, initiates the motor.

Hot combustion gases are ducted from the motor through a supersonic nozzle, which accelerates the gas and converts the pressure and temperature of propellant combustion to kinetic energy. Directional control is attained through a number of different means, but commonly the nozzle is jointed, allowing it to be vectored by mechanical actuators. Figure 1 illustrates a typical solid rocket booster.

History. Liquid fuel rockets date back only to Robert Goddard in the 1930s; solid fuel rockets have been around since the thirteenth century when the Chinese invented them. They were used in everything from fireworks to warfare. Modern solid fuel rocket history dates to the mid-1950s when the U.S. military started experimenting with a sulfur-based sealant made by Thiokol. This sealant and various solid fuels and oxidizers were used to make solid propellants. An early result of this work was the Sergeant missile. A scaled-down version of the Sergeant was clustered as the upper two stages of the Jupiter C (eleven Sergeants as the second stage and three as the third stage) that put the first U.S. satellite, Explorer 1, into orbit on 31 January 1958.

A typical solid rocket booster.

Figure 1. A typical solid rocket booster.

As the Cold War escalated, it became clear that solid propulsion offered several advantages for ballistic missiles: low cost, high reliability, low maintenance, long-term storability, packaging efficiency, and quick response. This led to a major development of solid fuel rocket capability in the United States and resulted in a long line of silo- and submarine-based ballistic missiles such as Minuteman, Peacekeeper, and the Poseidon/Trident family that have long-range capability.

In the 1960s with an eye to space access, two very large solid rockets were tested. One was 156 inches in diameter, and the other was an enormous 260 inches in diameter. These tests demonstrated, very early, the capability to make large, high-thrust solid rockets. This development led to large segmented solid booster systems for both the Titan and the Space Shuttle launch vehicles. Many space launch vehicles benefit from additional energy and thrust gained by strapping solid rockets around a liquid rocket core vehicle.

Because of their many advantages and inherent design flexibility, solid rockets are used in a wide range of applications besides launching. They have been used to transfer spacecraft from low Earth orbit to higher orbits, including geosynchronous altitudes. They have also been used as retrograde motors for stage separation and planetary descent. Solid rockets were used to reduce the descent velocity of the Sojourner spacecraft as it approached the Martian surface.

Many military tactical propulsion devices use solid rockets. These have a wide range of applications, including surface-to-surface, air-to-air, air-to-surface, and surface-to-air. These propulsion systems start with a diameter of approximately two inches and range to much larger, longer range systems. For example, solids power the Patriot antimissile system used successfully in the Gulf War in 1991.

Principle Of Operation. Solid rockets produce thrust by the same principle as a child’s balloon. Gases that exit the balloon in one direction push it in the opposite direction, illustrating Newton’s famous theorem of ”equal and opposite reaction.” Modern rockets take the principle of the balloon to a very high level of refinement (1). In a solid rocket, the gases are generated by burning a high-energy propellant. Unlike a balloon, solid rocket pressure vessels are rigid and allow much higher internal pressures. Modern rockets operate at 2000 psi, and chamber temperatures are 5000°F. The main component of thrust can be calculated as the product of the propellant mass flow rate and the velocity of the gases as they exit the nozzle. Under steady-state operating conditions, the propellant burning surface area, density, and burn rate determine the mass flow rate. Nozzle exit velocity is determined by propellant formulation, chamber pressure, nozzle geometry, and ambient pressure (2).

Manufacturing Flow. The type of motor case dictates the initial portion of the manufacturing flow. Solid rocket cases are of two types: filament-wound composite and metal. For filament-wound composite cases, the insulation is placed over a removable mandrel, cured, and machined, and the filament-resin mixture is wound over it. After the case is cared, the mandrel is removed. In a metal case motor, the case is fabricated first, and the insulation is laid up inside the case and then cured. This curing process also bonds the insulator to the case.

An elastomeric material, called a liner, is then applied to the inside of the insulated case to provide a bonding agent between the insulation and the propellant. A removable core is placed inside the case to provide the initial surface geometry for the propellant.

Propellant ingredients are mixed in bowls (up to 1800 gallons in size) similar to those used for making bread dough. When mixed, the propellant has the consistency of runny peanut butter. The propellant is then poured into the motor around the core and allowed to cure (become solid). Once the propellant has cured, the core tooling is removed. The loaded case is then ready for final assembly. In final assembly, the nozzle, igniter, and other mission-unique hardware such as cable raceways, thrust vector actuation systems, and interstages for connecting to payloads or other rocket stages are installed.


A solid rocket propellant consists of an elastomeric polymer that may be filled with as much as 91% of solid particles. The solid propellant may have 4 to 12 ingredients and is formulated using a combination of a fuel and an oxidizer that are intimately mixed on a microscopic level, and in some cases, are parts of the same molecule. Initially, the ingredients that make up the solid propellant are mixed together to form a thick liquid containing suspended solid particles. Once the propellant has been cast into the motor case, the mixture hardens enough to maintain its mechanical integrity and retains sufficient elasticity to prevent cracking as it experiences induced stresses from operation (high pressure) and storage (thermal expansion and contraction) (3).

Parameters for Formulating Propellants. Many interacting factors must be considered when formulating a solid rocket propellant. A primary consideration is for the propellant to provide sufficient energy to meet mission requirements. In addition to specific impulse (Isp), a measure of the amount of thrust the propellant provides per unit mass, solid rocket motors are often volume-limited. Hence propellant density must be considered. The operating pressure and mass flow rate of a solid rocket motor depend on the ballistic characteristics of the propellant, so these characteristics must be tightly controlled. Propellants typically use very energetic materials to achieve the performance objectives of the rocket motor. Energetic materials are often susceptible to ignition stimuli. They may be shock-sensitive, toxic, or apt to react explosively when damaged and ignited. Each of these factors must be considered when evaluating a combination of propellant materials.

There are other issues to be considered when formulating propellants, including their cost, maturity, reproducibility, manufacturability, compatibility, and service life. Finally, because rocket motors are generally not used immediately after they are manufactured, the propellant must maintain its properties over time and under the environmental conditions required for the given application.

Propellant Categories. A variety of different types of solid rocket propellants have been developed and deployed during the past several decades. Among other things, they can be grouped by signature, hazards, or application. The propellant exhaust signature is of major importance in military applications. Exhaust signature in this context refers to the visible plume that is emitted from the rocket booster when it is operating; the plume allows the adversary to track a vehicle visually back to its launch site. As a result, formulations are often designated as metallized (smoky), reduced smoke, or minimum smoke propellants. Reduced smoke propellants have very little primary smoke (smoke produced within the combustion chamber) but may form considerable secondary smoke (similar to the contrail which sometimes forms behind high-altitude aircraft) under certain atmospheric conditions. Minimum smoke propellants produce a minimum amount of both primary and secondary smoke. They are formulated to produce only gaseous combustion products.

Propellants may also be categorized by their potential for unplanned ignition. Those that have an explosion hazard are given a Class 1.1 designation. Those that have a mass fire hazard but not an explosion hazard are termed Class 1.3 propellants. Class 1.3 propellants are usually formulated using ammonium perchlorate (AP) as the principal oxidizer; Class 1.1 propellants often incorporate large quantities of the detonable nitramines HMX and RDX, and/or energetic nitrate esters such as nitroglycerin (NG) or butane triol trinitrate (BTN). Propellant Composition. A number of different materials are required in solid propellants to provide the desired properties. Table 1 gives a list of the categories of ingredients in solid propellants and the most common examples of each, along with brief comments. Not all propellants require ingredients from each category. By far the most common oxidizer in rocket propellants today is ammonium perchlorate (AP). It provides high energy, high density, excess oxygen, low detonability, low cost, ballistic tailorability, good mechanical properties, and good aging characteristics. Nitrates generally suffer from insufficient energy, poor ballistic characteristics, and moisture sensitivity. Ammonium nitrate (AN) is the cheapest oxidizer available for use in solid propellants, but its other drawbacks have prevented its widespread use. Newly developed oxidizers have not yet been extensively used due to their higher cost, immaturity, or limited availability.

Binders that provide mechanical integrity to propellants include poly-butadienes, polyethers, and polyesters. In most cases, the raw materials are low molecular weight polymers that contain reactive functional groups. They wet and suspend the solid materials and then cross-link (using an appropriate curative). Energetic plasticizers are typically used in combination with oxygenated binders. They provide both increased energy and improved mechanical properties. High-energy propellants contain as much as three times more plasticizer than polymer.

Metal fuels are incorporated into solid rocket propellants to provide increased specific impulse (Isp) and density. Isp is a measure of the theoretical performance of a propellant and is defined as the calculated specific impulse at a chamber pressure of 1000 psi expanded to 1 atmosphere with an optimum nozzle expansion ratio. Beryllium is the metal that provides the highest Isp. However, it has toxicity problems that limit its application. Boron and magnesium have been used in some systems, but aluminum is by far the most common metal fuel. Its use provides a theoretical Isp density improvement of about 10% compared to a nonmetallized formulation.

Table 1. Common Propellant Materials and their Functions

Functional category Common examples Comments
Solid oxidizers Ammonium perchlorate (AP), other perchlorates, AP is used in all but minimum smoke propellants.
ammonium nitrate (AN), other nitrates, AN is low cost but has numerous drawbacks,
ammonium dinitramide (ADN), hydrazinium including moisture sensitivity, little ballistic
nitroformate (HNF) tailorability, phase transitions, and aging concerns. ADN, HNF are still immature in U.S.
Energetic Nitramines: Cyclotrimethylenetrinitramine (RDX), Nitramines are used in most Class 1.1 and some
monopropellants cyclotetramethylenetetramintramine (HMX), Class 1.3 propellants. Provide increased 7sp.
hexanitrohexaazaisowurtzitane (CL-20) These three nitramines provide similar Isp in aluminized propellants. CL-20 is the densest and most oxygen-rich, but least mature.
Binders Hydroxyl-terminated polybutadiene (HTPB), HTPB, CTPB, PBAN are the most common Class 1.3
carboxyl-terminated polybutadiene (CTPB), propellant binders. PEG, PPG, NC are generally
polybutadiene acrylonitrile (PBAN), polyethylene plasticized with nitrate esters. NC and GAP are
glycol (PEG), polypropylene glycol (PPG), energetic.
nitrocellulose (NC), glycidyl azide polymer (GAP)
Curatives Isocyanates, epoxides Isocyanates are used to cure hydroxyl-terminated polymers; epoxides cure PBAN, CTPB
Fuels Beryllium, aluminum, magnesium Beryllium gives the highest Isp, but is toxic. Aluminum has better Isp and density than magnesium but is not as easily ignited.
Plasticizers Dioctyl adipate (DOA), dioctyl phthalate (DOP), DOA, DOP, and similar esters are used with
triacetin, other inert esters, nitroglycerin (NG), nonpolar binders such as HTPB. Triacetin is used
butanetriol trinitrate (BTTN), trimethylolethane as a desensitizer for nitrate ester propellants.
trinitrate (TMETN), other nitrate esters Nitrate esters provide increased 7sp. NG is highest density and highest performance nitrate ester. Other nitrate esters are used to decrease detonability (compared with NG) or to give better low-temperature properties.
Stabilizers A02246, p-rc-methyl nitroanaline (MNA), A02246 prevents oxidative cross-linking of HTPB;
nitrodiphenylamine (NDPA) MNA, and NDPA stabilize nitrate esters
Ballistic modifiers Iron oxide, aluminum oxide, oxamide Iron oxide; aluminum oxide accelerates burn rate; oxamide and other coolants slow burn rate.

Polymers, plasticizers, fuels, and curatives constitute the bulk of most solid propellant formulations, but several other ingredients are used to tailor one or more of their properties. These ingredients include ballistic modifiers such as iron oxide (which increases the burn rate), combustion stabilizers, chemical stabilizers (for nitrate esters and some polymers), processing aids, cure catalysts, and bonding agents. Bonding agents enhance the bond between oxidizer particles and the binder and greatly improve mechanical properties. Ballistic Properties. The ballistic (or combustion) characteristics of a propel-lant include burn rate as a function of pressure and temperature, combustion stability, and the completeness of combustion. Propellants that exhibit large changes in burn rate as a function of pressure are not often used. The dependence of propellant ballistics on temperature is particularly important for tactical rockets, which must operate over a wide temperature range (— 50°F to 150°F).

Ballistic properties can be controlled by a number of formulation variables. Among the most significant variables that affect burn rate are AP particle size (smaller = faster burn rate), total solids loading (higher = faster burn rate), the presence of ballistic modifiers (can increase or decrease burn rate, depending on the modifier), the curative used, and the metal content and particle size. Propellant Case Bond. One other factor that must always be considered when developing rocket motors is the means for attaching the propellant to the case wall. A thin layer of adhesive, or liner, is usually applied to the motor case before manufacturing the propellant. The liner is usually cured to some degree before the propellant is cast and then is cured more completely along with the propellant. The liner is quite significant because it must bond to both the inner case wall, which may be insulation or case material, and to the propellant. Unless it is to function also as a barrier to the diffusion of propellant ingredients such as plasticizers or curatives, it is usually formulated as an elastomeric material that has a backbone chemically similar to that of the propellant.

Propellant Grain Design

System constraints and requirements such as allowable length, volume, maximum pressure, and thrust profile largely drive the grain design. Additional factors such as clearance for the nozzle, the thrust vector control (TVC) method, clearance for the igniter, and location of the ignition system are also important considerations.

Design Considerations. Figure 2 shows an example of some features used in grain design. Grain structural loads often require stress relief features such as flaps or slots. Features such as radial or longitudinal slots are incorporated into the propellant grain to obtain the desired motor performance. Grain structural requirements drive such features as allowable web fraction (grain thickness/available distance) and the presence or absence of flaps and stress relief slots. A flap is an elastomeric piece bonded to the grain that disconnects the grain from the case. Flaps typically do not have a large impact on the ballistic performance of the motor; however, stress relief slots can present significant design complications. Stress relief slots are slots located in the grain solely for structural reasons.

Features used in grain design.

Figure 2. Features used in grain design.

Structural analysis of the solid propellant grain is important because structural failure of the propellant can result in catastrophic failure of the motor. A crack in the propellant or a separation in the propellant-to-case bond will increase the surface area, causing increased pressure, which may burst the case. A crack or bond separation may also allow flame to reach the case sooner than planned, resulting in case burn-through. Structural loads that are likely to cause propellant cracking or bond separation include cure and thermal shrinkage, internal pressurization, and acceleration.

In designing a motor, the stresses and strains induced in the propellant by these loads must be kept below the propellant’s capability. Most methods for reducing grain stresses and strains reduce the amount of propellant in the motor, so careful consideration must be given to both propellant loading and structural integrity. When designnig a propellant grain, it is assumed that propellant burns normal to the grain surface. The burning surface area as a function of distance burned is then used to determine the propellant mass flow rate as a function of time. Grain Manufacture. Typically, the grain forming tooling is placed in the motor case, and the propellant is cast between the case and the casting tooling. The propellant is then cured at elevated temperature, and when cure is complete, the tooling is removed from the motor. The two primary methods for introducing the propellant into the motor case are (1) pressure casting and (2) vacuum casting. In pressure casting, the propellant is forced through a tube into the motor. In vacuum casting, a vacuum is introduced in the interior of the motor case, and the propellant is pulled into the motor. Casting tooling may be removable; the tooling is removed from the motor after propellant cure is complete or may be left in place. Sometimes both types of grain forming tooling are used in the same motor.

Structural requirements may require curing the propellant under pressure. The combination of vacuum casting and pressure cure results in grains that have few defects such as voids. Grains may be cast with simple tooling, and more complicated slots are machined into the grain after propellant cure. This results in an additional manufacturing step, but changing the grain design is as simple as altering the machining program, rather than altering expensive hard tooling.

Another type of grain fabrication is extrusion. The propellant is forced through a die and cut to length. This process is not typically used for large rocket motors due to the difficulty of retaining the grain in the rocket motor case.

Motor Case

The solid rocket motor case is the pressure vessel that contains the solid propellant and provides a structural interface to external components. The case is designed to contain the high pressures generated by the burning propellant during motor operation. It is also the airframe that transfers thrust from the motor to the launch vehicle or missile system. As such, the case must incorporate features for attaching the pressure vessel to other stages, payloads, launch support equipment, or other support structures. Basic loads during motor operation are shown in Fig. 3.

Metal or composite materials are generally used in case design. Metals are generally less costly, more damage tolerant, and better characterized than composites. Composites generally weigh less due to the high strength to weight ratio but are less stiff for a comparable thickness. A composite case can be as much as five times lighter than a metal case. This weight reduction requires less propellant for equivalent vehicle performance, which ultimately can lead to a less expensive motor.


Internal insulation is the heat barrier between the case and the propellant. It protects the case from reaching temperatures that would endanger its structural integrity. Insulation also serves to (1) buffer the transmission of case stresses to the propellant, (2) inhibit burning on designated propellant surfaces, (3) provide a pressure seal for the case, and (4) limit the diffusion of chemical components to or from the propellant. Insulation may also be located on the exterior of the rocket motor to protect the case from aeroheating and to provide damage tolerance and protection from the elements.

Basic loads during motor operation.

Figure 3. Basic loads during motor operation.

Insulator materials are organic compounds that consist of reinforcing fillers contained in a binder. Fillers contribute to char-layer strength and include silica, asbestos, carbon or Kevlar fibers, nylon, and glass cloth. Elastomers and plastics are two classes of binders.

Internal insulator design predominantly involves provision for material char and ablation. The general equation for computing design thickness is



ET=exposure time

MAR = material affected or char rate

SF = safety factor

TP = thermal protection thickness

MT = manufacturing tolerance


The functions of a nozzle are to provide a ballistic throat for the motor, control motor pressure, direct subsonic gases into the throat, additional thrust, and thrust vector control (TVC). Solid rocket motors use ablative nozzles. Nozzle design requires knowledge of gas dynamics, heat transfer, combined thermal-structural loads, and material science. The nozzle environment is harsh; temperatures can reach 6000°F. Pressures can exceed 3500 psi, and exposure time can reach 300 seconds. Temperatures, pressures, and gas velocities vary greatly from the inlet region of a nozzle through the throat and into the exit cone. Typical Nozzle Designs. Nozzle configurations are either external or submerged. The external nozzle extends aft beyond the nozzle-to-chamber interface. The submerged nozzle extends forward into the case. Submerged nozzles provide a separated flow region in the aft motor that results in smoother gas flow into the throat region and decreased erosion of the motor insulator. Length-constrained motors employ nozzles submerged 10-40%.

Function. The nozzle inlet directs subsonic gases smoothly into the throat where they transit to sonic flow. The throat controls motor pressure by the initial size and material erosion rate interactively with the propellant burn rate. The exit cone expands and controls the supersonic gases. Contoured exit cones provide efficient and optimum expansion or thrust in the shortest length (4,5). Nozzles provide motor pressure control, efficient gas flow, and thermal protection for primary structural components by erosion and thermal degradation of ablative insulators.

Propellant exhaust gases heat up the nozzle surface of the ablative material via radiation and convection. Energy from the surface is conducted into the ablative material at rates dependent on the ablative reinforcing fiber type, the binder material (resin system), and the phenolic resins, or elastomeric binders. Heat conducted into the ablative material causes resin or binder degradation resulting in pyroltic gases that percolate through the porous char layer and provide internal “cooling.” Surface erosion is the result of excessive material temperatures (i.e., melting), mechanical erosion, or chemical reaction between the ablative constituents and the exhaust gases (6,7).

Nozzle materials. Materials consist of structural or ablative/insulators. Structural materials include metals fibers-reinforced phenolic or epoxy composites. Ablative selection requires knowledge of gas pressures, temperatures, velocities, and propellant exhaust characteristics. The properties of commonly used ablative or insulators, as well as some throat materials, can be found in Reference 4.

Carbon or graphite ablatives have been mostly derived from cellulosic or polyacrylonitrile (PAN) fiber precursors, subsequently carbonized or graphitized, woven into fabrics, and preimpregnated with phenolic resins. Silica and glass phenolics consist of melted glass filaments subsequently combined into fiber strands woven into fabrics and impregnated with phenolic resin.

Molding compounds consist of glass, silica, or carbon fibers or woven fabrics chopped into short lengths and mixed into phenolic resin. Elastomeric-based ablators or insulators combine chopped fibers or other fillers into the basic Buna-N, silicone, or ethylene-propylene diene monomer (EPDM) rubber compound. Most ablatives used in the inlet region must conform to structural component deformations and match the motor aft insulator erosion characteristics. They are typically Buna-N, silicone, or EPDM-based elastomeric material reinforced with short chopped glass, carbon, or Kevlar fibers. External nozzles often incorporate glass, silica, or carbon-reinforced phenolics because they provide better erosion resistance than the elastomeric-based materials.

Throat material selection is based on motor performance requirements and/or type of propellant. A high-performance motor with aluminized propellant often requires a low eroding throat material. Typical materials range from carbon phenolic used on the large Space Shuttle nozzle throat (8) to low eroding carbon-carbon or noneroding tungsten, which is used on high-performance tactical motors.

Exit cone ablatives require more erosion resistant material than inlets and typically incorporate silica-, carbon-, or graphite-reinforced phenolics. Reinforced elastomers do not have sufficient char strength to be used in the supersonic flow region of the exit cone where gas velocities or particle impingement tend to wash off low-strength charred materials.

Nozzle structures must withstand high loads due to combinations of motor pressures, ablative thermal expansion loads, thrust vector control (TVC) actuation, and external aerodynamic loads. Materials most often used are high-strength, aerospace-grade steel, aluminum, and titanium alloys.

Where weight is critical and stiffness is needed, polyacrylilonitrile (PAN)-based reinforced epoxies are incorporated as structural components. These can be filament-wound, tape-wrapped, or hand laid-up.

Thrust Vector Control

Thrust vector control (TVC) uses external means (typically mechanical or fluidic) to alter the direction of the thrust, thus changing the direction of the missile’s flight path. Figure 4 shows how one type of TVC, a movable nozzle, creates a turning moment on the missile that is equal to the thrust times the moment arm. TVC systems can be configured to provide control in all three axes: pitch (up and down motion), yaw (side to side motion), and roll (rotation about the longitudinal axis of the missile). TVC can be thought of as the steering system or “rudder” of a missile (8-10).

One type of TVC, a movable nozzle.

Figure 4. One type of TVC, a movable nozzle.

Fixed-Nozzle TVC. Fixed-nozzle systems use thrusters, secondary fluid injection, or mechanical deflectors to change the thrust vector direction. Thrusters are multiple discrete nozzles placed in strategic locations on the missile to create the desired turning forces. The gas may be supplied by a pressurized bottle (cold gas), a gas generator (warm gas), or may be taken directly from the combustion chamber of the rocket motor (hot gas). Thrusters have very fast response time, and they also create a shock wave on the outside of the missile that can amplify the resulting force. Disadvantages include the expensive exotic materials required for hot gas valves and the weight and volume penalty of packaging. Thrusters are used on the fourth stages of Minuteman III and Peacekeeper.

In secondary injection TVC, a fluid (liquid or gas) is injected into the exit cone through the wall. This creates side forces from a combination of the thrust of the injectant jet and pressure imbalances from shock waves. The injectant can be an inert or reactive liquid, in which a chemical reaction results in additional side force. Gas injectants may come from gas generators (warm), or they may be bled directly from the motor combustion chamber (hot). Advantages include fast response and the thrust addition to the main flow. Disadvantages are the large packaging volume required and limited thrust deflection (about 6°). Liquid injection systems have been in production on the Titan III, Minuteman III, and Polaris. Gas injection systems suffer from material problems in the severe environment and have never reached production status.

TVC systems that use mechanical deflection include jet vanes, jet tabs, and jetevators. Jet vanes are aerodynamic fins inside the nozzle that rotate to provide pitch, yaw, and roll control. Roll control is one advantage of jet vanes. Another advantage is the low torque required to actuate a jet vane. Disadvantages include the exotic materials, required for the severe flow environment, the length penalty from packaging, and thrust losses due to the aerodynamic drag on the vanes. Jet vanes were used on the German V-2 missile and on the United States Sergeant, Talos, and Pershing missiles. They are currently used on the United States RIM-7 Seasparrow, the Vertical-Launch ASROC, and the AIM-9X.

Jet tabs are retractable surfaces mounted at the aft end of the exit cone that pivot in and out of the gas flow, creating shock waves and pressure imbalance. An advantage over jet vanes is that no thrust losses occur when the tabs are out of the flow. Jet tabs also have a low actuating force requirement. However, jet tabs do not provide any roll control, have high thrust losses, and also require exotic materials. Jet tabs were used on the MK-106 Tomahawk booster rocket motor, which was later replaced by the MK-111 motor with a movable nozzle.

The jetevator consists of a spherical ring mounted around the nozzle exit cone that can be rotated into the supersonic gas stream. This rotation creates a steering side force on the missile. Jetevators were operational on the Polaris, BOMARC, and SUBROC missiles. One advantage is a side force that is linear with the deflection angle. Disadvantages include large weight and volume, severe environment requiring exotic materials, and large thrust losses. Movable-Nozzle TVC. In movable-nozzle systems, the nozzle is mechanically pivoted, which turns the hot supersonic flow of gases, thus changing the thrust vector. Movable nozzles are further subcategorized according to the location of the joint. If the entire nozzle and exit cone pivot as a unit, it is called a subsonic splitline. In the supersonic splitline, only the aft part of the exit cone pivots. Each has advantages and disadvantages; however, the supersonic splitline has never been in production due to manufacturing challenges.

Movable nozzles have lower thrust losses than the other types of TVC. However, a single movable nozzle cannot provide any roll control. Roll requires at least two nozzles. The Minuteman uses four hinged movable nozzles for pitch, yaw, and roll control, and the Space Shuttle has two booster motors with movable nozzles. Further categorization of movable nozzles is based on the type of joint. The flexible joint (also known as a flexible bearing or “flexbearing”) typically consists of alternating layers of an elastomeric material for flexibility and a rigid material (steel or composite) for strength and stiffness. The predictable and re-peatable nature of the actuating force required to move a flexbearing nozzle is considered an advantage. Flexbearing nozzles are the most widely used TVC systems, as shown by application to the Space Shuttle boosters, Ariane, Trident, and Peacekeeper. The Space Shuttle RSRM flexbearing (Fig. 5) is the largest TVC system in production.

The ball-and-socket joint, also known as a trapped ball, has a spherical socket that rides inside a mating spherical ball surface. The ball-and-socket nozzle has certain advantages over the flexbearing, including higher vector angles, higher motor pressure capability, and less pivot point shift. Disadvantages include an unpredictable stick-slip friction force, susceptibility to contamination, and it requires an antirotational device to prevent the nozzle from “rolling” in the socket. Two Navy surface-launched motors that employ cold-trapped ball nozzles are the Tomahawk MK-111 booster motor and the Mk-72 Aegis ER booster motor.

The fluid bearing/rolling seal, also known by the patented name Techroll®, is composed of a pair of rolling elastomeric convolutes that contain a fluid. The greatest advantage is the low actuating force required to move the bearing. The main disadvantages are the low structural stiffness and resulting large misalignment of the nozzle. This bearing is used on the Air Force Inertial Upper Stage (IUS) space motors.

Space Shuttle RSRM flexbearing—the largest TVC system in production.

Figure 5. Space Shuttle RSRM flexbearing—the largest TVC system in production.

The hinged movable nozzle is supported on thrust pins that ride in journal bearings. The hinged nozzle has the advantages of high deflection capability and low actuating force. The main disadvantage is that it provides control in only one direction (pitch or yaw) because it rotates only about one axis. The most well-known application of hinged nozzles is the Minuteman, which has four hinged nozzles on the first stage and thus provides all axes of control (pitch, yaw, and roll).

The gimbaled nozzle is an extension of the hinged nozzle; the thrust pins about which the nozzle pivots are themselves mounted in a rotating assembly that pivots about another set of thrust pins which are located 90° around the nozzle from the first set. This gives the nozzle omniaxial motion capability. This approach also has high deflection capability but requires a large envelope for packaging the gimbal mechanism. The gimbaled nozzle has been tested on the ground and in flight but has not been in production.

The rotatable nozzle is a canted nozzle mounted on a rolling bearing so that it can pivot about the motor centerline. Its main advantage is the low actuating force required. However, it is limited to motors that have multiple nozzles because movement of the nozzle induces pitch, yaw, and roll moments that must be balanced by the other nozzle. Another drawback is that the rotational deflection required is much larger than the actual vector angle achieved. This was an operational system for the Polaris missile.


The purpose of the igniter is to provide the heat and pressure rapidly needed to start propellant combustion. Figure 6 shows a typical solid propellant rocket motor that has an axial flow igniter and some of the factors that affect motor ignition. These factors include igniter location, mass flow rate and action time of the igniter, impingement of the igniter output on the propellant surface, propellant grain geometry, combustion chamber free volume, and nozzle throat size.

A typical solid propellant rocket motor with variables that affect ignition and grain heating.

Figure 6. A typical solid propellant rocket motor with variables that affect ignition and grain heating.

There are four main components in an ignition system: the safety device, the electroexplosive device, the booster charge, and the igniter main propellant charge. The safety device provides the electrical and mechanical safety features needed to keep stray electrical current from causing inadvertent ignition. For planned ignition, electrical arming commands are sent to the safety device that mechanically and electrically arm or align it. In this armed condition, an electrical firing signal can then start a chain of events that culminates in motor ignition. The electrical firing signal is sent to an electroexplosive device called an initiator. The output from the initiator typically ignites a booster charge. The output from the booster charge then ignites the igniter main propellant charge that in turn ignites the solid propellant in the rocket motor.

Several methods are used for initial igniter sizing. One simple method uses an igniter coefficient. This is a ratio of the mass flow rate of the igniter gases to the throat area of the motor. Values of 0.25 lbm/sec per square inch of throat area are typical.

Future of Solid Propellant Rockets

Solid rocket motors have a long history of service because of their inherent characteristics: they are low cost, simple, reliable, compact, storable, and deliver very high thrust/weight. They have historically demonstrated the capability to support a wide range of propulsion applications. The entire U.S. ICBM and Submarine Launched Ballistic Missile fleet converted to solid propulsion in the 1950s and 1960s because of their high performance, high reliability, responsiveness, low cost, and packaging efficiency. Similarly, all land, air, and sea launched tactical missiles are solid propulsion based.

As science, engineering, and manufacturing have advanced, so have these rockets. Today, these rockets have very highly developed performance. However, in the future, new structural materials will offer higher temperature capability and lighter weight. New propellants will offer more energy. New analytical techniques, combined with ongoing laboratory work continues to enhance our understanding of these motors. New tooling approaches for case, nozzle, and propellant grain will reduce cost and allow rapid design changes. Improved process controls will improve reliability, repeatability, and reduce cost. Improved case, nozzle, and insulating materials will increase reliability, reduce cost, and increase performance. New propellants will be more energetic, less sensitive, and less costly. Because ofthe basic strengths ofsolid propellants and because science and technology continues to build on these strengths, solid propellants can be expected to serve well into the future.

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