Chapter 15 Guidance and Control

Guidance and Control



1. Understand the purpose of a guidance system.

2. Understand the three phases of guidance.

3. Be acquainted with the concept and the types of control-guidance systems.

4. Know and understand the methods of homing guidance.

5. Understand the application of monopulse methods to weapon homing systems.

6. Be acquainted with the function and the types of self-contained guidance systems.

7. Be acquainted with the basic operation and applications of accelerometers.

8. Be acquainted with present guided flight paths.

9. Know and understand the variable flight paths.


The term missile in the post- World War II era has generally been used synonymously with "guided missile," due to the wide impact of guided missile technology upon the weapons field. In the un-guided case, initial conditions (such as train, elevation, powder charge in naval guns) and exterior ballistic effects are paramet-ers that, along with normal distribution, affect the "fall of shot." As advances in technology permitted (paralleled by in-creasing threat complexity), the development of guided missiles made possible a significant increase in terminal accuracy of mil-itary weaponry. The application of automatic control is preva-lent in broad regions of missile technology including:

Underwater Homing Torpedoes

Surface-to-Surface Aerodynamic Guided Missiles

Intercontinental Ballistic Missiles

Air-to-Air Guided Missiles

Surface-to-Air Guided Missiles

Guided Projectiles

This chapter will deal primarily with the aerodynamic guided missile. Various aerodynamic missile flight paths will be intro-duced and a simplified analysis included to demonstrate the rela-tive advantages/disadvantages of each.


16.2.1 Purpose and Function

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.

It should be clear at this point that the concept of "Gui-dance and Control" involves not only the maintenance of a part-icular vehicle's path from point A to B in space, but also the proper behavior of the vehicle while following the path. A missile that follows a prescribed path half the way to a target and then becomes dynamically unstable is then incapable of re-maining upon the path (or else fails structurally due to aero-dynamic loading). Such a vehicle, in order to perform properly, must be "piloted" and capable of responding to control signals.

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 sur-faces to stabilize the missile in roll, pitch, and yaw. Guidance and stabilization corrections are combined, and the result is ap-plied as an error signal to the control system.

16.2.2 Sensors

The guidance system in a missile can be compared to the human pilot of an airplane. As a pilot guides his plane to the landing field, the guidance system "sees" its target. If the target is far away or otherwise obscured, radio or radar beams can be used to locate it and direct the missile to it. Heat, light, televi-sion, the earth's magnetic field, and Loran have all been found suitable for specific guidance purposes. When an electromagnetic source is used to guide the missile, an antenna and a receiver are installed in the missile to form what is known as a sensor. The sensor picks up, or senses, the guidance information. Mis-siles that are guided by other than electromagnetic means use other types of sensors, but each must have some means of receiv-ing "position reports."

The kind of sensor that is used will be determined by such factors as maximum operation range, operating conditions, the kind of information needed, the accuracy required, viewing angle, and weight and size of the sensor, and the type of target and its speed.

16.2.3 Accelerometers

The heart of the inertial navigation system for ships and mis-siles is an arrangement of accelerometers that will detect any change in vehicular motion. To understand the use of acceler-ometers in inertial guidance, it is helpful to examine the gen-eral principles involved.

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 dis-placement of the pendulum form its original position will be a function of the magnitude of the accelerating force. Another simple device might consist of a weight supported between two springs. When an accelerating force is applied, the weight will move from its original position in a direction opposite to that of the applied 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.

If the acceleration along the fore-and aft axis were con-stant, the speed of the missile at any instant could be deter-mined simply by multiplying the acceleration by the elapsed time. However, the acceleration may change considerably over a period of time. Under these conditions, integration is necessary to determine the speed.

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, a second integration is necessary.

The moving element of the accelerometer can connected to a potentiometer, or to a variable inductor core, or to some other device capable of producing a voltage proportional to the dis-placement of the element.

Usually there are three double-integrating accelerometers continuously measuring the distance traveled by the missile in three directions--range, altitude, and azimuth (figure 16-1). 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.

Accelerometers are sensitive to the acceleration of gravity as well as missile accelerations. For this reason, the acceler-ometers that measure range and azimuth distances must be mounted in a fixed position with respect to the pull of gravity. This can be done in a moving missile by mounting them on a platform that is stabilized by gyroscopes or by star-tracking telescopes. This platform, however, must be moved as the missile passes over the earth to keep the sensitive axis of each accelerometer in a fixed position with respect to the pull of gravity. These fac- tors cause the accuracy of the inertial system to decrease as the time of flight of the missile increases.

To eliminate unwanted oscillations, a damper is included in the accelerometer unit. The damping effort should be just great enough to prevent any oscillations from occurring, but still per-mit a significant displacement of the mass. When this condition exists, the movement of the mass will be exactly proportional to the acceleration of the vehicle.

Figure 16-2 shows a mass suspended by a spring in a liquid-damped system. If the case experiences an acceleration in the direction indicated by the arrow, the spring will offer a re-straining force proportional to the downward displacement of the mass, while the viscous fluid will serve to dampen any undesira-ble oscillations.

Figure 16-3 shows a system that is electrically damped. The mass (M) is free to slide back and forth in relation to the iron core (C). When the vehicle experiences an acceleration, the voltage (E), which is proportional to the displacement of the mass, is picked off and amplified. The current (I) (still pro-portional to the displacement) is sent back to the coil around the core. The resulting magnetic field around the coil creates a force on the mass, which damps the oscillations. In this system the acceleration could be measured by the displacement of the mass (X), by the voltage (E), or by the current (I).


Missile guidance is generally divided into three phases--boost, midcourse, and terminal. These names refer to different parts of the flight path. The boost phase may also be called the launch-ing or initial phase.

16.3.1 Boost Phase.

Navy surface-to-air missiles accelerate to flight speed by means of the booster component. This booster period lasts from the time the missile leaves the launcher until the booster burns its fuel. In missiles with separate boosters, the booster drops away from the missile at burnout. The objective of this phase is to place the missile at a position in space from where it can either "see" the target or where it can receive external guidance signals. During the boost phase of some missiles, the guidance system and the aerodynamic surfaces are locked in position. Other missiles are guided during the boost phase.

16.3.2 Midcourse Phase.

The second, or midcourse, phase of guidance is often the longest in both distance and time. During this part of the flight, changes may be required to bring the missile onto the desired course and to make certain that it stays on that course. During this guidance phase, information can be supplied to the missile by any of several means. In most cases, the midcourse guidance system is used to place the missile near the target, where the

system to be used in the final phase of guidance can take over. In other cases, the midcourse guidance system is used for both the second and third guidance phases.

16.3.3 Terminal Phase.

The last phase of missile guidance must have high accuracy as well as fast response to guidance signals. Missile performance becomes a critical factor during this phase. The missile must be capable of executing the final maneuvers required for intercept within the constantly decreasing available flight time. The maneuverability of the missile will be a function of velocity as well as airframe design. Therefore, a terminal guidance system must be compatible with missile performance capabilities. The greater the target acceleration, the more critical the method of terminal guidance becomes. Suitable methods of guidance will be discussed in later sections of this chapter. In some missiles, especially short-range missiles, a single guidance system may be used for all three phases of guidance, whereas other missiles may have a different guidance system for each phase.


Missile guidance systems may be classified into two broad categories: missiles guided by man-made electromagnetic devices, and those guided by other means. (See figure 16-6.) In the

first category are those missiles controlled by radar, radio de- vices, and those missiles that use the target as a source of e-lectromagnetic radiation. In the latter category are missiles that rely on electromechanical devices or electromagnetic contact with natural sources, such as the stars (self-contained guidance systems).

All of the missiles that maintain electromagnetic radiation contact with man-make sources may be further subdivided into two subcategories.

(1) Control guidance missiles

(2) Homing guidance missiles

16.4.1 Control Guidance

Control guidance missiles are those that are guided on the basis of direct electromagnetic radiation contact with friendly control points. Homing guidance missiles are those that guided on the basis of direct electromagnetic radiation contact with the tar-get. Control guidance generally depends on the use of radar (radar control) or radio (radio control) links between a control point and the missile. By use of guidance information transmit-ted from the control point via a radio or radar link, the mis-sile's flight path can be guided. This chapter will use radar control guidance as a model for discussion because it is by far the most common application of control guidance methods. The principles discussed may be readily applied to radio (including television) control guidance. Radar Control Guidance. Radar control guidance may be subdivided into two separate categories. The first category is simply referred to as the command guidance method. The second is the beam-rider method, which is actually a modification of the first, but with the radar being used in a different manner. Command guidance--The term command is used to describe a guidance method in which all guidance instructions, or com-mands, come from sources outside the missile. The guidance sys- tem of the missile contains a receiver that is capable of re- ceiving instructions from ship or ground stations or from air- craft. The missile flight-path control system then converts these commands to guidance information, which is fed to the attitude control system.

In the command guidance method, one or two radars are used to track the missile and target. Figure 16-7 is a block diagram of how this method works in actual practice. As soon as the radar is locked on the target, tracking information is fed into the computer. The missile is then launched and is tracked by the radar. Target and missile ranges, elevations, and bearings are continuously fed to the computer.

This information is analyzed and a missile intercept flight path is computed. The appropriate guidance signals are then transmit-ted to the missile receiver. These signals may be sent by vary-ing the characteristics of the missile-tracking radar beam, or by way of a separate radio transmitter. The radar command guidance method can be used in ship, air, or ground missile delivery sys-tems. A relatively new type of command guidance by wire is now operational in some short-range antitank-type weapons. These systems use an optical sight for tracking the target while the weapons emits a characteristic infra-red signature, which is used for tracking the weapon with an IR sensor. Deviation of the wea-pon from the line of sight (LOS) to the target is sensed, and guidance commands are generated that are fed to the weapon con- trol system in flight via a direct wire link. Each weapon con- tains wire spools that pay out as the warhead flies out the line of sight to the target. Current usage of these systems is in relatively lightweight, portable, short-range battlefield envir-onments against armored targets where their high accuracy and substantial warheads are most efficiently employed. Beam-rider Method--The main difference between the beam-rider method and the radar command guidance method is that the characteristics of the missile-tracking radar beam are not varied in the beam-rider system. The missile has been designed so that it is able to formulate its own correction signals on the basis of its position with respect to the radar scan axis. The technique is best understood after reviewing the principles of conical-scan tracking in Chapter 5. The missile's flight path control unit is sensitive to any deviation from the scan axis of the guidance radar and is capable of computing the proper flight path correction. An advantage of this type of system is that is requires only one radar. This radar must, of course, have a conical-scan feature in order to provide both target-tracking capability and a missile flight-path correction reference axis. A second advantage is that since the missile formulates its own directional commands, several missiles may be launched to "ride" the beam simultaneously, without the need for a cumbersome and complicated multiple-missile command system.

Figure 16-8 illustrates a simple beam-rider guidance system on a typical LOS course. The accuracy of this system decreases with range because the radar beam spreads out, and it is more difficult for the missile to remain in its center. If the target is moving very rapidly, the missile must follow a continuously changing path, which may cause it to undergo excessive transverse accelerations.

16.4.2 Homing Guidance

Homing guidance systems control the flight path by employing a device in the weapon that reacts to some distinguishing feature of the target. Homing devices can be made sensitive to a variety of energy forms, including RF, infrared, reflected laser, sound, and visible light. In order to home on the target, the missile or torpedo must determine at least the azimuth and elevation of the target by one of the means of angle tracking mentioned pre-viously. Active homing missiles will also have the means of de-termining range of the target if necessary. Tracking is perform-ed by a movable seeker antenna or an array with stationary elec-tronically scanned arrays in development for missiles and opera-tional in some torpedoes. Determination of angular error by amp-litude comparison monopulse methods is preferred over the older COSRO systems because of the higher data rate and faster response time; however, phase comparison monopulse or interferometer meth-ods have advantages in some applications. Homing guidance meth-ods may be divided into three types: active, semiactive, and passive homing (figure 16-9). These methods may be employed in seekers using any of the energy forms mentioned above, although some methods may be excluded by the nature of the energy form; for example, one would not build a passive laser seeker or an active or semi-active infrared seeker. Active Homing. In active homing, the weapon contains both the transmitter and receiver. Search and acquisition are conducted as with any tracking sensor. The target is tracked employing monostatic geometry in which the returning echo from the target travels the same path as the transmitted energy (figure 16-8). An onboard computer calculates a course to in-tercept the target and sends steering commands to the weapon's autopilot. The monostatic geometry allows the most efficient reflection of energy from the target, but the small size of the missile restricts the designer to high frequencies and low power output from the transmitter, resulting in short seeker acquis-ition range. Semiactive Homing. In semiactive homing, the target is illuminated by a tracking radar at the launching site or other control point. The missile is equipped with a radar receiver (no transmitter) and by means of the reflected radar energy from the target, formulates its own correction signals as in the active method. However, semiactive homing uses bistatic reflection from the target, meaning that because the illuminator platform and weapon receiver are not co-located, the returning echo follows a different path than the energy incident to the target. Due to its shape and composition, the target may not reflect energy efficiently in the direction of the weapon. In extreme cases the weapon may lose the target entirely, resulting in a missed in-tercept. This disadvantage is compensated for by the ability to use greater power and more diverse frequency ranges in an illum- ination device in a ship, aircraft, or ground station. Passive Homing. Passive homing depends only on the target as a source of tracking energy. This energy can be the noise radiated by a ship or submarine in the case of a passive homing torpedo, RF radiation from the target's own sensors in the case of an anti-radiation (ARM) weapon, heat sources such as ship, aircraft, or vehicle exhausts, contrast with the tempera-ture or visible light environment, or even the radiation all ob-jects emit in the microwave region. As in the other homing meth-ods, the missile generates its own correction signals on the basis of energy received from the target rather than from a con-trol point. The advantage of passive homing is that the counter detection problem is reduced, and a wide range of energy forms and frequencies are available. Its disadvantages are its sus-ceptibility to decoy or deception and its dependence on a certain amount of cooperation from the enemy. Retransmission Homing or Track Via Missile (TVM). Re-transmission homing is a blending of the characteristics of both command and semiactive homing guidance. In command guidance, missile steering commands are computed at the launch point using target position and missile position data derived from launch point sensors. In retransmission homing, the missile contains a semiactive seeker that determines the azimuth and elevation angle from the missile to the target, which is then coded and transmit-ted to the launch point via data link (down link). The fire-con-trol system at the launch point can use its own target tracking data, that of the missile (or both), and missile position data to compute steering commands, which are then transmitted to the mis-sile via an uplink. This technique is used in some new AAW mis-sile systems, including the U.S. Army Patriot system. Specific retransmission or TVM systems may vary somewhat from this ideal; however, they all will in some way use target angle data from the missile to compute steering commands at the launch point that are then transmitted to the missile.

16.4.3 Accuracy. Homing is the most accurate of all guidance systems because it uses the target as its source when used a-gainst moving targets. There are several ways in which the hom-ing device may control the path of a missile against a moving target. Of these, the more generally used are pursuit paths and lead flight paths, which are discussed in a subsequent part of this chapter. Because monopulse methods in weapons seekers are advantageous and are becoming the method of choice in current weapons, it is necessary to address the two basic types: Amplitude Comparison Monopulse. This method, described in Chapter 5, requires a gimballed seeker antenna covered by a radome at the nose of the weapon. Because of aerodynamic re-quirements, the radome shape is normally not optimal for radar performance. Very precise orders to the antenna are required to achieve target acquisition due to the single antenna's limited field of view. In these systems the size of the antenna directly determines the limits of the frequency range of the seeker. Its primary advantage is its consistent performance throughout the potential speed and maneuverability range of potential targets. Interferometer (Phase Comparison Monopulse). The in-terferometer eliminates the requirement for a movable antenna, having instead fixed antennas mounted at the edge of the airframe or on the wing tips, the result being reduced complexity and a wider field of view. As depicted in figure 16-10, two antennas separated by a known distance are installed for each mobility axis of the weapon. In the diagram the antennas A and B, separated by the distance d, receive energy emitted (passive homing) or reflected (semiactive homing) from the target.

Because the distance to the target is relatively large, it is assumed that the RF energy arrives as a series of planar waves with wavelength . In accordance with the discussion of electronic scanning in Chapter 7 and towed acoustic arrays in chapter 9 it is evident that for the geometry pictured, the phase sensed by antenna B will lag that sensed by antenna A by some phase angle which is proportional to d sin ; therefore:

= 2d sin

If is known and the phase angle can be determined, then the look angle, , can be calculated.

The interferometer provides the advantage of wide field of view, flexibility in airframe design, unobstructed use of weapon interior space, and the ability to cover broad frequency bands without constraints imposed by limited antenna size. The separation between the antennas governs the performance of the system, with missile body diameter or fin spread separation as the usual arrangement. The disadvantage of the interferometer is the angular ambiguity that may exist for wavelengths less than the separation between the antennas at a specific angle of incidence. If the distance between the antennas at an angle of incidence is d sin , and is less than d sin , then it is not possible to determine if the phase angle measured is just that or + n2 radians, where n is any integer. However, this is a minor problem in most homing systems because the absolute look angle is not as important as the rate of change of that angle.

The interferometer has an advantage in resolving multiple targets at twice the range of a typical amplitude comparison monopulse seeker in the same size weapon. This gives the missile twice the time to respond to the changeover from tracking the centroid of the group to tracking one specific target, thus in-creasing the hit probability.

16.4.4 Composite Systems. No one system is best suited for all phases of guidance. It is logical then to combine a system that has good midcourse guidance characteristics with a system that has excellent terminal guidance characteristics, in order to in-crease the number of hits. Combined systems are known as compos-ite guidance systems or combination systems.

Many missiles rely on combinations of various types of gui-dance. For example, one type of missile may use command guidance until it is within a certain range of a target. At this time the command guidance may become a back-up mode and a type of homing guidance commenced. The homing guidance would then be used until impact with the target or detonation of a proximity-fixed war-head.

16.4.5 Hybrid Guidance.

A combination of command guidance and semi-active homing guidance is a type of hybrid guidance. It achieves many advantages of both systems. It attains long-range capabilities by maintaining the tracking sensors on the delivery vehicle (ship, aircraft, or land base) and transmitting the data to the missile. By having the missile compute its own attitude adjustments, the entire mechanization of the fire control problem can be simplified.


The self-contained group falls in the second category of guidance system types. All the guidance and control equipment is entirely within the missile. Some of the systems of this type are: pre-set, terrestrial, inertial, and celestial navigation. These sys-tems are most commonly applicable to surface-to-surface missiles, and electronic countermeasures are relatively ineffective against them since they neither transmit nor receive signals that can be jammed.

16.5.1 Preset Guidance. The term preset completely describes one guidance method. When preset guidance is used, all of the control equipment is inside the missile. This means that before the missile is launched, all information relative to target loca-tion as well as the trajectory the missile must follow must be calculated. After this is done, the missile guidance system must be set to follow the course to the target, to hold the missile at the desired altitude, to measure its air speed and, at the cor-rect time, cause the missile to start the terminal phase of its flight and dive on the target.

A major advantage of preset guidance is that it is rela-tively simple compared to other types of guidance; it does not require tracking or visibility.

An early example of a preset guidance system was the German V-2, where range and bearing of the target were predetermined and set into the control mechanism. The earliest Polaris missile was also designed to use preset guidance during the first part of its flight, but this was soon modified to permit greater launch flex-ibility.

The preset method of guidance is useful only against sta-tionary targets of large size, such as land masses or cities. Since the guidance information is completely determined prior to launch, this method would, of course, not be suitable for use against ships, aircraft, enemy missiles, or moving land targets.

16.5.2 Navigational Guidance Systems. When targets are located at great distances from the launching site, some form of naviga-tional guidance must be used. Accuracy at long distances is achieved only after exacting and comprehensive calculations of the flight path have been made. The mathematical equation for a navigation problem of this type may contain factors designed to control the movement of the missile about the three axes--pitch, roll, and yaw. In addition, the equation may contain factors that take into account acceleration due to outside forces (tail winds, for example) and the inertia of the missile itself. Three navigational systems that may be used for long-range missile guidance are inertial, celestial, and terrestrial. 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 con-tact between the launching site and the missile after launch, the missile is able to make corrections to its flight path with amaz-ing 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 mis-sile 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. Figure 16-12 sketches the application of the system as it might be used for a guided mis- sile.

The missile must be provided with a horizontal or a vertical reference to the earth, automatic star-tracking telescopes to de-termine star elevation angles with respect to the reference, a time base, and navigational star tables mechanically or electric-ally 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 equip-ment 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.

16.5.3 Terrestrial Guidance Methods.

Prior to micro-miniaturization of computer circuits, the various methods of terrestrial guidance proposed had significant limita-tions. These proposed early systems included an inertial refer-ence system, a television camera to provide an image of the earth's surface, and a film strip of the intended flight path. The guidance system would compare the television picture with the projected film strip image and determine position by matching the various shadings in the two images. This method proved too slow in providing position data, even for a sub-sonic missile. Its other distinct disadvantage was that it required extensive low-level aerial photography of each potential missile flight path. The danger to flight crews and loss of the element of sur-prise involved in extensive pre-strike photo reconnaissance made such a system impractical.

With the availability of compact mass memory and vastly in-creased computational capability compatible with missile space and weight limitations, terrestrial guidance methods became prac-tical. The advent of small radar altimeters of high precision (Chapter 2) provided an alternative to photographic methods with the added advantage that weather and lighting conditions were relatively inconsequential. The radar altimeter provides a coarse means of detecting surface features by their height, which can then be compared with stored data concerning expected land contours along the missile flight path. The missile guidance system contains expected land-elevation values to the left and right of the missile's intended ground track. The guidance sys-tem will determine that the missile is located at a position where the stored data most closely matches the observed altitudes as pictured in figure 16-13. Once the direction of turn and the distance required to correct the error have been determined, the missile will turn to resume the intended track. This method is called Terrain Contour Matching or TERCOM. Even the most capable TERCOM system has insufficient memory to perform contour matching throughout a flight path of several hundred miles. Therefore, the missile will be provided with a series of small areas known as TERCOM maps along the route to the target. The number of TER-COM maps and their separation is determined by the quality of in-formation available on the area and the accuracy of the missile's inertial navigation system. Sufficient data is available from various sources to support TERCOM such that aerial reconnaissance of most target areas is not required prior to the engagement. TERCOM has sufficient accuracy to find, for example, a large military base within a region; however, it could not provide the accuracy to hit a specific section of that base, such as a group of hangars at an airfield. For this reason, a missile using some variation of TERCOM only would require a nuclear warhead.

Delivery of a conventional high-explosive warhead requires precision that can only be provided by some form of optical device in the terminal stage of flight. A cruise missile flies at altitudes and ranges that would prevent transmission of images back to the launch point. Advances in digitized imagery permit computer storage of grey-shaded scenes in the vicinity of the target. The digitized scene can be compared to data from a tele-vision camera in the missile and values of grey shading matched to determine actual position relative to desired position. The missile can correct its flight path to that desired and even fi-nally pick out its target. This method, called Digital Scene Matching Area Correlator or DSMAC, is sufficiently accurate to permit the use of a conventional high-explosive warhead. The DSMAC technique would be used only for the last few miles to the target, with the TERCOM method being used for the majority of the flight path. Both of the above methods are limited by the accur-acy of information used to create the digital TERCOM maps and DSMAC scenes that are loaded in the missile's memory. Building and formatting these data files for cruise missiles requires considerable support facilities and talented personnel.


A guided missile is usually under the combined influence of natural and man-made forces during its entire flight. Its path may assume almost any form. Man-made forces include thrust and directional control as shown in figure 16-14. The vector sum of all the forces, natural and man-made, acting on a missile at any instant, may be called the total force vector. It is this vector, considered as a function of time in magnitude and direction, that provides velocity vector control. Paths along which a guided missile may travel may be broadly classified as either preset or variable. The plan of a preset path cannot be changed in mid-flight; the plan of a variable path is altered according to conditions occurring during flight.

16.6.1 Preset Flight Paths.

Preset flight paths are of two types: constant and programmed. Constant. A preset guided missile path has a plan that has been fixed beforehand. This plan may include several different phases, but once the missile is launched, the plan cannot be changed. The phases must follow one another as originally planned. The simplest type of preset guided missile path is the constant preset. Here, the missile flight has only one phase.

The term constant preset may be broadened to include flights that are constant after a brief launching phase that is different in character from the rest of the flight. During the main phase of a constant preset guided-missile flight, the missile receives no control except that which has already been built into it. How-ever, it receives this control throughout the guided phase of flight. Often it is powered all the way. The nature of a con-stant preset guided-missile flight path depends on how it is pow-ered, and the medium through which it travels.

A torpedo fired from a submarine to intercept a surface tar-get, figure 16-15, may describe a straight line--a constant pre- set guided path. Programmed. A missile could be guided in a preset path against a fixed target; the joint effect of missile power and gravity would then cause the path to become a curve. A missile following a preset path may be guided in various ways--for in-stance, by an autopilot or by inertial navigation. The means of

propulsion may be motor, jet, or rocket. A more complex type of preset path is the programmed preset. Here, the weapon flight has several phases: for example: a torpedo, as illustrated in

figure 16-16, executing a search pattern. During the first phase, the torpedo, having been launched in some initial direc-tion other than the desired ultimate direction, gradually finds the desired direction by control mechanisms such as gyros and depth settings. The torpedo then maintains this direction for the remainder of this first phase, at the end of which it is presumed to be in the neighborhood of a target. During the sec-ond phase, the torpedo executes a search pattern, possibly a circular or helical path.

16.6.2 Variable Flight Paths.

The guided flight paths of greatest interest are those that can vary during flight. In general, the heading of the weapon is a function of target position and velocity. These parameters are measured by continuous tracking, and the resultant missile flight path is determined, assuming that the target motion will remain unchanged until new data is received. There are four basic types of variable flight paths in common use: pursuit, constant-bear-ing, proportional navigation, and line of sight. Pursuit. The simplest procedure for a guided missile to follow is to remain pointed at the target at all times. The mis-sile is constantly heading along the line of sight from the mis-sile to the target, and its track describes a pursuit path with the rate of turn of the missile always equal to the rate of turn of the line of sight. Pure pursuit paths are highly curved near the end of flight, and often the missile may lack sufficient

maneuverability to maintain a pure pursuit path in the terminal phase of guidance. When this is the case, the missile can be designed to continue turning at its maximum rate until a point is reached where a pursuit course can be resumed. The most common application of the pursuit course is against slow-moving targets, or for missiles launched from a point to the rear of the target.

Pursuit: Lead or deviated pursuit course is defined as a course in which the angle between the velocity vector and line of sight from the missile to the target is fixed. For purposes of illustration, lead angle is assumed to be zero, and only pure pursuit is described. (M = ). (See figure 16-17). Constant Bearing. At the opposite extreme to a pursuit path is a constant-bearing or collision path. The missile is aimed at a point ahead of the target, where both the missile and target will arrive at the same instant. The line of sight to this point does not rotate relative to the missile. The missile path is as linear as the effect of gravity and aerodynamic forces allow. If the target makes an evasive turn or if the target's velocity changes, a new collision course must be computed and the missile flight path altered accordingly. The outstanding feature

of this course is that for a maneuvering constant-speed target, the missile lateral acceleration never exceeds the target's lateral acceleration. The major drawback lies in the fact that the control system requires sufficient data-gathering and processing equipment to predict future target position.

Constant Bearing: A course in which the line of sight from the missile to the target maintains a constant direction in space. If both missile and target speeds are constant, a collision course results. (See figure 16-18.)

d = = 0

dt Proportional Navigation. The more advanced homing mis-siles will employ some form of proportional navigation. The mis-sile guidance receiver measures the rate of change of the line of sight (LOS) (bearing drift, if you will) and passes that informa-tion to the guidance computer, which in turn generates steering commands for the autopilot. The missile rate of turn is some fixed or variable multiple of the rate of change of the LOS. This multiple, called the navigation ratio, can be varied during mis-sile flight to optimize performance. A missile employing this method is said to use proportional navigation ratio may be less than 1:1 early in the flight to conserve velocity and increase range. As the flight proceeds, the navigation ratio will in-crease to 2:1, 4:1, or even more to ensure that the missile will be agile enough to counter target maneuvers in the terminal phase of flight.

Proportional: A course in which the rate of change of the missile heading is directly proportional to the rate of rotation of the line of sight from missile to target. (See figure 16-19.)

dM = Kd or M = K

dt dt Line of Sight. Here, the missile is guided so that it travels along the line of sight from the launching station to the target. This is, of course, the flight path flown by a beam-riding missile. An alternative form of a beam-riding path is the constant lead angle path. Here the beam that the missile follows is kept ahead of the line of sight by a constant offset. The major advantages of the line of sight path are its flexibility and the minimal complexity of the equipment that must be carried in the missile, since the major burden of guidance is assumed at the launching station.

Line of Sight: Defined as a course in which the missile is guided so as to remain on the line joining the target and point of control. This method is usually called "beam riding."(See figure 16-20.)

Table 16-1. Guidance Law Comparison

Law Advantage Disadvantage

Pursuit Simple Poor Against

Mechanization Maneuvering Targets

Launch Geometry


Fixed Lead Angle Simple Poor Against Maneuvering

Mechanization Targets

Line-of-Sight Less Geometry Lead Angle Correct for

Restricted Than Fixed Geometry Only


Constant Bearing Requires Minimum Requires Infinite Gain and Collision Maneuver Capability No-Time Lags--Most

Complex in Data

Gathering and


Proportional Good Against More Complex

Navigation Maneuvering Mechanization

All-Aspect Targets


Optimal Homing Solves Specific More Complex

Problems Such As Mechanization

Long Range

May Improve




Missile System Nears Production." International Defense

Review, Vol. 15 (1982): 132-36.

Table 16-3. Current Guidance System Examples

Weapon Guidance


Phoenix Semiactive/Active Command

Sparrow Semiactive

Sidewinder Passive (IR)


Harpoon Active Radar

Maverick T.V. (passive)

Walleye T.V. (passive)

Shrike Passive (RF)

Sram Inertial

Standard Arm Passive (RF)


Subroc Inertial

Asroc Ballistic


Sea Sparrow Semiactive

Hawk Semiactive

RAM Passive IR/RF (Interferometer)

Patriot Command, Semiactive (retransmission)

Standard MR (SM-1) Semiactive

Standard ER (SM-1) Semiactive

Standard SM-2 MR (Aegis) Command; Semiactive

Standard SM-2 ER Command; Semiactive


Poseidon Inertial

Tomahawk (Land Attack) Terrestrial (terrain following)

Trident Inertial

Standard Semiactive

Harpoon Active-Radar

Tomahawk (Antiship) Inertial/Active Radar

Battlefield Support

Lance Inertial

Tow Wire-Guided (command), Optically Tracked

Dragon Wire-Guided

Stinger Passive (IR); Proportional Navigation


The three phases of guidance are boost, midcourse, and terminal. The distinction between phases is primarily based upon a break-down of the flight path rather than in any change-over points in guidance methods. The terminal phase is, however, the most critical and demands peak performance form the guidance system.

Guidance systems are divided into two broad categories; those that use man-made electromagnetic devices and those that use some other means. The various subcategories of guidance systems are shown in table 16-2. Guided missile paths may be classified as either preset or variable. Preset guided paths have planned flight routines that cannot be changed in mid flight on the basis of updated data. A preset plan may be for a one-phase flight (constant preset) or a flight of several phases (programmed preset). Variable guided flight paths have plans that can be changed in mid flight; thus they make possible the successful intercept of a target that makes evasive maneuvers. Prediction of target position is con-tinuously reappraised and the missile course is recomputed in the light of new target data. Variable guided flight paths include pursuit, constant bearing, proportional navigation, and line of sight.

The interception by a missile of a moving target depends on the prediction of future target position, and requires certain assumptions. When using bullets, ballistic missiles, or preset guided missiles, it is assumed that the target motion measured while tracking will remain unchanged during missile flight. When using variable guided missiles, it is assumed that the target mo-tion, measured at any instant by almost continuous tracking, will remain unchanged over a short time interval.

Table 16-4. Guidance System Categories

Guidance Systems that Use

Man-Made Electromagnetic

Devices Self-Contained Guidance Systems

Radar Control Homing Preset Navigational

a. Command a. Active a. Inertial

b. Beam-Rider b. Semiactive b. Celestial

c. Modified c. Passive c. Terrestrial


Composite Systems

Hybrid (Command/Semiactive)


Bureau of Naval Personnel. Principle of Guided Missiles and Nuclear Weapons. NAVPERS 10784-B, 1st Rev. Washington, D.C.: GPO, 1972.

Commander, Naval Ordnance Systems Command. Weapons Systems Fundamentals. NAVORD OP 3000, vols. 2&3, 1st Rev. Washington, D.C.: GPO, 1971.

Gulick, J. F., and J. S. Miller, Missile Guidance: Interferometer Homing Using Body Fixed Antennas. Technical Memorandum TG1331 (TSC-W36-37), Laurel, Md.: Johns Hopkins University Applied Physics Laboratory.

Kaplan, Fred, "Cruise Missile: Wonder Weapon or Dud?" High

Technology, (Feb 1983).

Meller, Rudi. "Patriot, The U.S. Army's Mobile Air-defense Missile System Nears Production." International Defense Review, Vol.15 (1982): 132-36.