Ballistic Missile Basics
A ballistic missile (BM) is a a missile that has a ballistic trajectory over most of its flight path, regardless of whether or not it is a weapon-delivery vehicle. Ballistic missiles are categorized according to their range, the maximum distance measured along the surface of the earth's ellipsoid from the point of launch of a ballistic missile to the point of impact of the last element of its payload. Various schemes are used by different countries to categorize the ranges of ballistic missiles.
The United States divides missiles into four range classes.
Intercontinental Ballistic Missile ICBM over 5500 kilometers Intermediate-Range Ballistic Missile IRBM 3000 to 5500 kilometers Medium-Range Ballistic Missile MRBM 1000 to 3000 kilometers Short-Range Ballistic missile SRBM up to 1000 kilometersThe Soviet and Russian military developed a system of five range classes.
Strategic over 1000 kilometers Operational-Strategic 500 to 1000 kilometers Operational 300 to 500 kilometers Operational-Tactical 50 to 300 kilometers Tactical up to 50 kilometers
The 1987 Treaty on the Elimination of Intermediate-Range and Shorter-Range Missiles [INF Treaty] required elimination of all Soviet and American longer-range intermediate nuclear force (LRINF) missiles with ranges between 1,000 and 5,500 kilometers, as well as shorter-range intermediate nuclear force (SRINF) missiles with ranges between 500 and 1,000 kilometers. The Missile Technology Control Regime initially focused on missiles with ranges greater than 300 kilometers, the range of the Soviet SCUD missile.
Delivery systems vary in their flight profile, speed of delivery, mission flexibility, autonomy, and detectability. Each of these considerations is important when planning a chemical or biological attack.
Ballistic missiles have a prescribed course that cannot be altered after the missile has burned its fuel, unless a warhead maneuvers independently of the missile or some form of terminal guidance is provided. A pure ballistic trajectory limits the effectiveness of a chemical or biological attack because, generally, the reentry speed is so high that it is difficult to distribute the agent in a diffuse cloud or with sufficient precision to ensure a release under the shear layer of the atmosphere. In addition, thermal heating upon reentry, or during release, may degrade the quality of the chemical or biological agent. U.S. experience has shown that often less than 5 percent of a chemical or biological agent remains potent after flight and release from a ballistic missile without appropriate heat shielding.
A ballistic missile also closely follows a pre-established azimuth from launch point to target. The high speed of the ballistic missile makes it difficult to deviate too far from this azimuth, even when submunitions or other dispensed bomblets are ejected from the missile during reentry. Consequently, if the target footprint axis is not roughly aligned with the flight azimuth, only a small portion of the target is effectively covered.
A ballistic missile has a relatively short flight time, and defenses against a ballistic missile attack are still less than completely effective, as proved in the Allied experience during the Gulf War. However, with sufficient warning, civil defense measures can be implemented in time to protect civil populations against chemical or biological attack. People in Tel Aviv and Riyadh received enough warning of SCUD missile attacks to don gas masks and seek shelter indoors before the missiles arrived. Even with these limitations on ballistic missile delivery of airborne agents, Iraq had built chemical warheads for its SCUDs, according to United Nations' inspection reports.
Nuclear weapons differ markedly from chemical, biological, or conventional warheads. The principal difference is the size, shape, and inertial properties of the warhead. Generally, nuclear weapons have a lower limit on their weight and diameter, which determines characteristics of the delivery system, such as its fuselage girth. Though these limits may be small, geometric considerations often influence the selection of a delivery system. Chemical and biological weapons, which are usually fluids or dry powders, can be packed into almost any available volume. Nuclear weapons cannot be retrofitted to fit the available space; however, they can be designed to fit into a variety of munitions (e.g., artillery shells).
Nuclear weapons also have a different distribution of weight within the volume they occupy. Fissile material, the core of a nuclear weapon, weighs more per unit of volume than most other materials. This high specific gravity tends to concentrate weight at certain points in the flight vehicle. Since virtually all WMD delivery systems must fly through the atmosphere during a portion of their trip to a target, a designer has to consider the aerodynamic balance of the vehicle and the required size of control system to maintain a stable flight profile while carrying these concentrations of weight. Chemical, biological, and conventional weapons all have specific gravities near 1.0 gram/cc, so these materials may be placed further from the center of gravity of the vehicle without providing large compensating control forces and moments. In some special applications, such as ballistic missile reentry vehicles and artillery shells, the designer needs to include ballasting material-essentially useless weight-to balance the inertial forces and moments of the nuclear payload.
Because nuclear weapons have a large kill radius against soft and unhardened targets, accuracy is a minor consideration in the delivery system selection as long as the targeting strategy calls for countervalue attacks. Nuclear weapons destroy people and the infrastructure they occupy. They only require that the delivery system places the warhead with an accuracy of approximately 3 kilometers of a target if the weapon has a yield of 20 kilotons and to an even larger radius as the yield grows. Most unmanned delivery systems with a range of less than 500 kilometers easily meet these criteria. Often, as is the case with ballistic missiles, the quality of the control system beyond a certain performance does not materially change the accuracy of a nuclear warhead, because a large fraction of the error arises after the powered phase of the flight as the vehicle reenters the atmosphere. While this is true of chemical and biological warheads as well, with a nuclear warhead, there is less need to compensate for this error with such technologies as terminal guidance or homing reentry vehicles. To be effective, a delivery vehicle employed to spread chemical or biological agents must distribute the material in a fine cloud below a certain altitude and above the surface. It should be capable of all-weather operations and should not betray its presence to air defense assets.
Sir Isaac Newton stated in his Third Law of Motion that "every action is accompanied by an equal and opposite reaction." A rocket operates on this principle. The continuous ejection of a stream of hot gases in one direction causes a steady motion of the rocket in the opposite direction. A jet aircraft operates on the same principle, using oxygen in the atmosphere to support combustion for its fuel. The rocket engine has to operate outside the atmosphere, and so must carry its own oxidizer.
A rocket is a machine that develops thrust by the rapid expulsion of matter. The major components of a chemical rocket assembly are a rocket motor or engine, propellant consisting of fuel and an oxidizer, a frame to hold the components, control systems and a payload such as a warhead. A rocket differs from other engines in that it carries its fuel and oxidizer internally, therefore it will burn in the vacuum of space as well as within the Earth's atmosphere. A rocket is called a launch vehicle when it is used to launch a satellite or other payload into orbit or deep space. A rocket becomes a missile when the payload is a warhead and it is used as a weapon.
There are a number of terms used to describe the power generated by a rocket.
- Thrustis the force generated, measured in pounds or kilograms. Thrust generated by the first stage must be greater than the weight of the complete missile while standing on the launch pad in order to get it moving. Once moving upward, thrust must continue to be generated to accelerate the missile against the force of the Earth's gravity.
- The impulse, sometimes called total impulse, is the product of thrust and the effective firing duration. A shoulder fired rocket such as the LAW has an average thrust of 600 lbs and a firing duration of 0.2 seconds for an impulse of 120 lbsec. The Saturn V rocket, used during the Apollo program, not only generated much more thrust but also for a much longer time. It had an impulse of 1.15 billion lbsec.
- The efficiency of a rocket engine is measured by its specific impulse (Isp). Specific impulse is defined as the thrust divided by the mass of propellant consumed per second. The result is expressed in seconds. The specific impulse can be thought of as the number of seconds that one pound of propellant will produce one pound of thrust. If thrust is expressed in pounds, a specific impulse of 300 seconds is considered good. Higher values are better. Although specific impulse is a characteristic of the propellant system, its exact value will vary to some extent with the operating conditions and design of the rocket engine. It is for this reason that different numbers are often quoted for a given propellant or combination of propellants.
- A rocket's mass ratio is defined as the total mass at liftoff divided by the mass remaining after all the propellant has been consumed. A high mass ratio means that more propellant is pushing less missile and payload mass, resulting in higher velocity. A high mass ratio is necessary to achieve the high velocities needed for long-range missiles.
Most current long-range missiles consist of two or more rockets or stages that are stacked on top of each other. The second stage is on top of the first, and so on. The first stage is the one that lifts the missile off the launch pad and is sometimes known also as a "booster" or "main stage". When the first stage runs out of propellant or has reached the desired altitude and velocity, its rocket engine is turned off and it is separated so that the subsequent stages do not have to propel unnecessary mass. Dropping away the useless weight of stages whose propellant has been expended means less powerful engines can be used to continue the acceleration, which means less propellant has to be carried, which in turn means more payload can be placed onto the target.
Many different types of rocket engines have been designed or proposed. There are three categories of chemical propellants for rocket engines: liquid propellant, solid propellant, and hybrid propellant. The propellant for a chemical rocket engine usually consists of a fuel and an oxidizer. Sometimes a catalyst is added to enhance the chemical reaction between the fuel and the oxidizer. Each category has advantages and disadvantages that make them best for certain applications and unsuitable for others.
Liquid propellant rocket engines burn two separately stored liquid chemicals, a fuel and an oxidizer, to produce thrust.
A cryogenic propellant is one that uses very cold, liquefied gases as the fuel and the oxidizer. Liquid oxygen boils at 297 F and liquid hydrogen boils at 423 F. Cryogenic propellants require special insulated containers and vents to allow gas from the evaporating liquids to escape. The liquid fuel and oxidizer are pumped from the storage tanks to an expansion chamber and injected into the combustion chamber where they are mixed and ignited by a flame or spark. The fuel expands as it burns and the hot exhaust gases are directed out of the nozzle to provide thrust.
Hypergolic Propellant A hypergolic propellant is composed of a fuel and oxidizer that ignite when they come into contact with each other. No spark is needed. Hypergolic propellants are typically corrosive so storage requires special containers and safety facilities. However, these propellants are typically liquid at room temperature, and do not require the complicated storage facilities that are mandatory with cryogenic propellants.
Mono-propellants Monopropellants combine the properties of fuel and oxidizer in one chemical. By their nature, monopropellants are unstable and dangerous. Monopropellants are typically used in adjusting or vernier rockets to provide thrust for making changes to trajectories once the main stages of the rocket have burnd out.
Advantages of liquid propellant rockets include the highest energy per unit of fuel mass, variable thrust, and a restart capability. Raw materials, such as oxygen and hydrogen are in abundant supply and a relatively easy to manufacture. Disadvantages of liquid propellant rockets include requirements for complex storage containers, complex plumbing, precise fuel and oxidizer injection metering, high speed/high capacity pumps, and difficulty in storing fueled rockets.
The petroleum used as a rocket fuel is a type of kerosene similar to the sort burned in heaters and lamps. However, the rocket petroleum is highly refined, and is called RP-1 (Refined Petroleum). It is burned with liquid oxygen (the oxidizer) to provide thrust. RP-1 is a fuel in the first-stage boosters of the Delta and Atlas-Centaur rockets. It also powered the first stages of the Saturn 1B and Saturn V. RP-1 delivers a specific impulse considerably less than that of cryogenic fuels.
Cryogenic propellants are liquid oxygen (LOX), which serves as an oxidizer, and liquid hydrogen (LH2), which is a fuel. The word cryogenic is a derivative of the Greek kyros, meaning "ice cold." LOX remains in a liquid state at temperatures of minus 298 degrees Fahrenheit (minus 183 degrees Celsius). LH2 remains liquid at temperatures of minus 423 degrees Fahrenheit (minus 253 degrees Celsius). In gaseous form, oxygen and hydrogen have such low densities that extremely large tanks would be required to store them aboard a rocket. But cooling and compressing them into liquids vastly increases their density, making it possible to store them in large quantities in smaller tanks.
The distressing tendency of cryogenics to return to gaseous form unless kept supercool makes them difficult to store over long periods of time, and hence less satisfactory as propellants for military rockets, which must be kept launch-ready for months at a time. But the high efficiency of the liquid hydrogen/liquid oxygen combination makes the low-temperature problem worth coping with when reaction time and storability are not too critical. Hydrogen has about 40 percent more "bounce to the ounce" than other rocket fuels, and is very light, weighing about one-half pound (0.45 kilogram) per gallon (3.8 liters). Oxygen is much heavier, weighing about 10 pounds (4.5 kilograms) per gallon (3.8 liters).
The RL-10 engines on the Centaur, the United States' first liquid-hydrogen/liquid-oxygen rocket stage, have a specific impulse of 444 seconds. The J-2 engines used on the Saturn V second and third stages, and on the second stage of the Saturn 1B, also burned the LOX/LH2 combination. They had specific impulse ratings of 425 seconds. For comparison purposes, the liquid oxygen/kerosene combination used in the cluster of five F-1 engines in the Saturn V first stage had specific impulse ratings of 260 seconds. The same propellant combination used by the booster stages of the Atlas/Centaur rocket yielded 258 seconds in the booster engine and 220 seconds in the sustainer. The high efficiency engines aboard the Space Shuttle orbiter use liquid hydrogen and oxygen and have a specific impulse rating of 455 seconds. The fuel cells in an orbiter use these two liquids to produce electrical power through a process best described as electrolysis in reverse. Liquid hydrogen and oxygen burn clean, leaving a by-product of water vapor.
The rewards for mastering LH2 are substantial for spaceflight applications. The ability to use hydrogen means that a given mission can be accomplished with a smaller quantity of propellants (and a smaller vehicle), or alternately, that the mission can be accomplished with a larger payload than is possible with the same mass of conventional propellants. In short, hydrogen yields more power per gallon.
Hypergolic propellants are fuels and oxidizers which ignite on contact with each other and need no ignition source. This easy start and restart capability makes them attractive for both manned and unmanned spacecraft maneuvering systems. Another plus is their storability -- they do not have the extreme temperature requirements of cryogenics. The fuel is monomethyl hydrazine (MMH) and the oxidizer is nitrogen tetroxide (N2O4). Hydrazine is a clear, nitrogen/hydrogen compound with a "fishy" smell. It is similar to ammonia. Nitrogen tetroxide is a reddish fluid. It has a pungent, sweetish smell. Both fluids are highly toxic, and are handled under the most stringent safety conditions.
Hypergolic propellants are used in the core liquid propellant stages of the Titan family of launch vehicles, and on the second stage of the Delta. The Space Shuttle orbiter uses hypergols in its Orbital Maneuvering Subsystem (OMS) for orbital insertion, major orbital maneuvers and deorbit. The Reaction Control System (RCS) uses hypergols for attitude control. The efficiency of the MMH/N2O4 combination in the Space Shuttle orbiter ranges from 260 to 280 seconds in the RCS, to 313 seconds in the OMS. The higher efficiency of the OMS system is attributed to higher expansion ratios in the nozzles and higher pressures in the combustion chambers.
Solid propellant rockets are basically combustion chamber tubes packed with a propellant that contains both fuel and oxidizer blended together uniformly. The solid-propellant motor is the oldest and simplest of all forms of rocketry, dating back to the ancient Chinese. It's simply a casing, usually steel, filled with a mixture of solid-form chemicals (fuel and oxidizer) which burn at a rapid rate, expelling hot gases from a nozzle to achieve thrust.
The principal advantage is that a solid propellant is relatively stable therefore it can be manufactured and stored for future use. Solid propellants have a high density and can burn very fast. They are relatively insensitive to shock, vibration and acceleration. No propellant pumps are required thus the rocket engines are less complicated. Disadvantages are that, once ignited, solid propellants cannot be throttled, turned off and then restarted because they burn until all the propellant is used. The surface area of the burning propellant is critical in determining the amount of thrust being generated. Cracks in the solid propellant increase the exposed surface area, thus the propellant burns faster than planned. If too many cracks develop, pressure inside the engine rises significantly and the rocket engine may explode. Manufacture of a solid propellant is an expensive, precision operation. Solid propellant rockets range in size from the Light Antitank Weapon to the 100 foot long Solid Rocket Boosters (SRBs) used on the side of the main fuel tank of the Space Shuttle.
The Space Shuttle uses the largest solid rocket motors ever built and flown. Each reusable booster contains 1.1 million pounds (453,600 kilograms) of propellant, in the form of a hard, rubbery substance with a consistency like that of the eraser on a pencil. The four center segments are the ones containing propellant. The uppermost one has a star-shaped, hollow channel in the center, extending from the top to about two thirds of the way down, where it gradually rounds out until the channel assumes the form of a cylinder. This opening connects to a similar cylindrical hole through the center of the second through fourth segments. When ignited, the propellant burns on all exposed surfaces, from top to bottom of all four segments. Since the star-shaped channel provides more exposed surface than the simple cylinder in the lower three segments, the total thrust is greatest at liftoff, and gradually decreases as the points of the star burn away, until that channel also becomes cylindrical in shape. The propellant in the star-shaped segment is also thicker than that in the other three. A solid propellant always contains its own oxygen supply. The oxidizer in the Shuttle solids is ammonium perchlorate, which forms 69.93 percent of the mixture. The fuel is a form of powdered aluminum (16 percent), with an iron oxidizer powder (0.07) as a catalyst. The binder that holds the mixture together is polybutadiene acrylic acid acrylonitrile (12.04 percent). In addition, the mixture contains an epoxy-curing agent (1.96 percent). The binder and epoxy also burn as fuel, adding thrust. The specific impulse of the Space Shuttle solid rocket booster propellant is 242 seconds at sea level and 268.6 seconds in a vacuum.
Hybrid propellant rocket engines attempt to capture the advantages of both liquid and solid fueled rocket engines. The basic design of a hybrid consists of a combustion chamber tube, similar to ordinary solid fueled rockets, packed with a solid chemical, usually the fuel. Above the combustion chamber tube is a tank, containing a complementary reactive liquid chemical, usually the oxidizer. The two chemicals are hypergolic, and when the liquid chemical is injected into the combustion chamber containing the solid chemical, ignition occurs and thrust is produced. The ability to throttle the engine is achieved by varying the amount of liquid injected per unit of time. The rocket engine can be stopped by cutting off the flow of the liquid chemical. The engine can be restarted by resuming the flow of the liquid chemical. Other advantages of hybrid propellant rocket engines are that they provide higher energy than standard solid propellant rockets, they can be throttled and restarted like liquid propellant rockets, they can be stored for long periods like solid propellant rockets, and they contain less than half the complex machinery (pumps, plumbing) of standard liquid propellant engines. They are also less sensitive to damage to the solid fuel component than standard solid propellant system. Hybrid rockets control the combustion rate by metering the liquid component of the fuel. No matter how much solid component surface area is exposed, only so much can be burned in the presence of the liquid component. Disadvantages are that these engines do not generate as much energy per pound of propellant as liquid propellant engines and they are more complex than standard solid fueled engines. Hybrid propellant rocket engines are still in development and are not yet available for operational use.
The guidance system in a missile can be compared to the human pilot of an airplane. Every missile guidance system consists of an attitude control system and a flight path control system. The attitude control system functions to maintain the missile in the desired attitude on the ordered flight path by controlling the missile in pitch, roll, and yaw. The attitude control system operates as an auto-pilot, damping out fluctuations that tend to deflect the missile from its ordered flight path. The function of the flight path control system is to determine the flight path necessary for target interception and to generate the orders to the attitude control system to maintain that path.
The operation of a guidance and control system is based on the principle of feedback. The control units make corrective adjustments of the missile control surfaces when a guidance error is present. The control units will also adjust the control to stabilize the missile in roll, pitch, and yaw. Guidance and stabilization corrections are combined, and the result is applied as an error signal to the control system.
The heart of the inertial navigation system for missiles is an arrangement of accelerometers that will detect any change in vehicular motion. An accelerometer, as its name implies, is a device for measuring acceleration. In their basic form such devices are simple. For example, a pendulum, free to swing on a transverse axis, could be used to measure acceleration along the fore-and-aft axis of the missile. When the missile is given a forward acceleration, the pendulum will tend to lag aft; the actual displacement of the pendulum form its original position will be a function of the magnitude of the accelerating force. The movement of the mass (weight) is in accordance with Newton's second law of motion, which states that the acceleration of a body is directly proportional to the force applied and inversely proportional to the mass of the body.
Usually there are three double-integrating accelerometers continuously measuring the distance traveled by the missile in three directions--range, altitude, and azimuth. Double-integrating accelerometers are devices that are sensitive to acceleration, and by a double-step process measure distance. These measured distances are then compared with the desired distances, which are preset into the missile; if the missile is off course, correction signals are sent to the control system. If the missile speed were constant, the distance covered could be calculated simply by multiplying the speed by time of flight. But because the acceleration varies, the speed also varies. For that reason, the second integration is necessary.
When targets are located at great distances from the launching site, some form of navigational guidance must be used. Accuracy at long distances is achieved only after exacting and comprehensive calculations of the flight path have been made. Navigational systems that may be used for long-range missile guidance include inertial and celestial.
- Inertial guidance. The simplest principle for guidance is the law of inertia. In aiming a basketball at a basket, an attempt is made to give the ball a trajectory that will terminate in the basket. However, once the ball is released, the shooter has no further control over it. If he has aimed incorrectly, or if the ball is touched by another person, it will miss the basket. However, it is possible for the ball to be incorrectly aimed and then have another person touch it to change its course so it will hit the basket. In this case, the second player has provided a form of guidance. The inertial guidance system supplies the intermediate push to get the missile back on the proper trajectory. The inertial guidance method is used for the same purpose as the preset method and is actually a refinement of that method. The inertially guided missile also receives programmed information prior to launch. Although there is no electromagnetic contact between the launching site and the missile after launch, the missile is able to make corrections to its flight path with amazing precision, controlling the flight path with accelerometers that are mounted on a gyro-stabilized platform. All in-flight accelerations are continuously measured by this arrangement, and the missile attitude control generates corresponding correction signals to maintain the proper trajectory. The use of inertial guidance takes much of the guesswork out of long-range missile delivery. The unpredictable outside forces working on the missile are continuously sensed by the accelerometers. The generated solution enables the missile to continuously correct its flight path. The inertial method has proved far more reliable than any other long-range guidance method developed to date.
- Celestial Reference. A celestial navigation guidance system is a system designed for a predetermined path in which the missile course is adjusted continuously by reference to fixed stars. The system is based on the known apparent positions of stars or other celestial bodies with respect to a point on the surface of the earth at a given time. Navigation by fixed stars and the sun is highly desirable for long-range missiles since its accuracy is not dependent on range. The missile must be provided with a horizontal or a vertical reference to the earth, automatic star-tracking telescopes to determine star elevation angles with respect to the reference, a time base, and navigational star tables mechanically or electrically recorded. A computer in the missile continuously compares star observations with the time base and the navigational tables to determine the missile's present position. From this, the proper signals are computed to steer the missile correctly toward the target. The missile must carry all this complicated equipment and must fly above the clouds to assure star visibility. Celestial guidance (also called stellar guidance) was used for the Mariner (unmanned spacecraft) interplanetary mission to the vicinity of Mars and Venus. ICBM and SLBM systems at present use celestial guidance.
- Command Guidance Multi-source radio signals that allow a triangulation of position offer an alternative to acceleration measurements. Advanced missile powers dropped radio guidance in the 1960's and switched to autonomous inertial measuring units, which are carried onboard the missile. The United States considered radio guidance again in the late 1980's for mobile missiles but dropped the idea in favor of a Global Positioning System (GPS). A radio guidance system could transmit signals from the launch site, or an accurate transmitter array near the launch site to create the signals. Radio command and control schemes, because of the immediate presence of a radio signal when the system is turned on, alert defenses that a missile launch is about to occur. And performance for these systems degrades because of the rocket plume and radio noise. Also, these systems are very much subject to the effects of jamming or false signals.
Global Positioning System (GPS) and the Global Navigation Satellite System (GLONASS) are unlikely ever to be used in the control function of a ballistic missile. The best military grade GPS receivers produce positions with an uncertainty of tens of centimeters. If a missile has two of these receivers in its airframe spaced 10 meters apart, the best angular resolution is roughly in the centiradian range. Theater ballistic missiles [TBMs] require milliradian range angular accuracy to maintain control. However, GPS has significant application for an TBM outfitted with a post-boost vehicle (bus) or attitude control module that navigates a reentry vehicle to a more accurate trajectory.
Following the completion of the propulsive phase of the mission, the missile will typically align, inertially stabilize, and release a reentry vehicle [RV] on a trajectory toward a pre-selected target. During atmospheric reentry, the exterior of the RV is protected from aerothermodynamic heating by a thermal protection system (TPS).
The aerodynamic shape configuration (ballistic or lifting) of a reentry vehicle determines the severity, duration, and flight path of reentry experienced by the vehicle. This, in turn, affects the vehicle systems complexity and the heating loads on the payload. A lifting reentry vehicle has many operational advantages over a non-lifting vehicle. Primarily, the reentry loads can be minimized to almost any desired level, with flexibility in landing site selection. The vehicle has the ability to deviate its reentry trajectory to reach selected landing sites "cross range" from the orbital track, and to fine tune deorbit propulsion system errors. Spherical and ballistic vehicles can only deorbit to selected sites which are on the orbital ground track. A disadvantage of the lifting shape over the non-lifting shape lies in the complexity and high cost associated with guidance and control of the lifting vehicle. A failure of the guidance or control system could render the vehicle uncontrollable and cause it to diverge a great distance off course.
Methods which have been used to protect RVs in the past include:
- ablation (erosion of surface material, such as silicone elastomers); and
- radiative heat shield (e.g., ceramic-based surface insulation systems).
Either of these methods, or a combination of them, may be used to protect the RV against excessive heating. After the vehicle reenters the atmosphere, it will decelerate to below sonic speeds. In order to further reduce the velocity of the RV for delivery of chemical or biological agents, supplemental deceleration systems such as parachutes may be used.
RVs possess a tremendous amount of kinetic energy, which must be dissipated during reentry as the vehicles decelerate to their impact or landing velocity. The RV reenters the Earth's atmosphere at velocities of up to Mach (M) 25. As the RV passes through the atmosphere, atmospheric friction decelerates it to below M 1, and converts its kinetic energy primarily into thermal energy (heat). Within the stagnation zone, an area immediately in front of the RV, an area of compressed, extremely hot, ionized and stagnant air is formed. Heat from the hot gas is transferred to the surface of the RV.
The heat generated during reentry is not only dependent on atmospheric density, but is also inversely proportional to the square root of the radius of the RV's nose cone and proportional to the cube of its velocity. Hence, blunt nose RVs are heated less than slender ones; and lifting RV designs, which use the glider principle, produce less heat than ballistic hyperbolic descent designs because their velocity is typically lower. Thus, a full evaluation of thermal impacts during reentry is dependent on both vehicle- and mission-specific criteria.
Temperatures generated within the hottest area (the stagnation zone) during ballistic reentry may exceed 11,100°C (20,000°F). Heat generation is not as severe on vehicles which are capable of some degree of lift during reentry; the temperature of the Apollo capsule surface reached about 2,760°C (5,000°F). Thermal protection systems are required for RVs to ensure the vehicle does not burn up during reentry. The choice of systems to be used is dependent upon the vehicle design, the reentry temperatures the RV may be subject to, and mission-specific requirements of the warhead. Thermal protection systems for the exterior of RVs which may be feasible include ablation, radiative heat shield, heat sink, transpiration, and radiator. However, to date, heat sink, transpiration, and radiator systems have not been used to protect the exterior surface of RVs from the thermal stress of reentry.
Ablation cooling or simple ablation is a process in which heat energy is absorbed by a material (the heat shield) through melting, vaporization and thermal decomposition and then dissipated as the material vaporizes or erodes. In addition, high surface temperatures are reached and heat is dissipated by surface radiation, pyrolysis of the surface material causing formation of a "char," and the generation of chemical by-products which move through the char carrying heat outward towards the surface boundary. The rejected chemical by-products then tend to concentrate in the ablation boundary layer where they further block convective heating. These ablative materials may be chemically constructed or made from natural materials.
A common man-made ablative material in current use is a firm silicone rubber whose chemical name is phenolmethylsiloxane. It has a silicone elastomer base, with silica filler and carbon fibers for shear strength. Its primary use is in high shear, high heatflux environments; it is used on control surfaces and nose cones of hypervelocity vehicles, including some parts of the Space Shuttle. This material yields a carbonaceous char on pyrolysis, which is a glassy, ceramic-type material composed of silicon, oxygen, and carbon. An ablative material known as polydimethylsiloxane has been used on manned reentry capsules in the past, including the Mercury program. An elastomeric silicon ablative material was used in the Discoverer program. An example of a natural material is the oak wood heat shield used on the Chinese FSW reentry vehicles.
During reentry, the ablative processes begin in the upper atmosphere when the pyrolysis temperature of the material is reached resulting from an increase in atmospheric friction. At altitudes above 120 km (75 mi), atmospheric density is generally insufficient to cause the onset of ablation.
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