Ballistic Missile - The Basics
Source: American Federation of Scientists 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 The Soviet and Russian military developed a system of
five range classes:
Strategic - over 1000 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
weap-ons 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 un-manned 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.
Missile Components
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.
* Thrust is 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.
Propulsion
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.
* Cryogenic Propellant - 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.
* Monopropellants - 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.
Guidance System
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 mea-suring acceleration. In their basic form such devices are
sim-ple. 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 dis-tances,
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 bas-ket. 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 sup-plies 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 informa-tion 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 genera-ted 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 alterna-tive 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.
Reentry Vehicle
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.
http://www.fas.org/nuke/intro/missile/basics.htm
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 kilometers
Operational-Strategic - 500 to 1000 kilometers
Operational - 300 to 500 kilometers
Operational-Tactical - 50 to 300 kilometers
Tactical - up to 50 kilometers