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Basic Avionics part 1 (ATA Ch 22, 23, 31)


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1.1. General

Power flight control are employed in high- performance aircraft, and are generally of two main types, are known as: Power assisted and power operated. The choice of either system for a particular type of is governed by the forces required to overcome the aerodynamic load acting on the flight control surfaces. In basic form, both system are similar in that a hydraulically-operated a servo control unit, consisting of a control valve and an actuating jack, is connected between pilot’s control and relevant control surfaces. The major difference, apart from constructional features, is in the method of connecting actuating jack to control surfaces.

1.2. Power Assisted And Power Operated

In Power Assisted System, the pilot’s control is connected to the control surfaces via control lever. For example on pith control, when the pilot moves the control column to initiate a climb , the control lever pivots about point “x”, and accordingly commences moving the elevators up.

At the same time, the control valve pistons are displayed and this allows oil from the hydraulic system to flow to the left-hand side of the actuating jack piston, the rod of which is secured to the aircraft’s structure. The reaction of the pressure exerted on the piston causes the whole servo unit, and control level, to move to the left, and because of greater control effort produced the pilot is assisted in making further upward movement of the elevators.


Page 2 of 109 In a Power Operated System , the pilot’s control is connected to the control lever only, while the servo unit is directly connected to the flight control surface. The effort required by the pilot to move the control column is simply that needed to move the control lever and control valve piston. It does not vary with the effort required to move the control surface which as will be noted from the diagram, is supplied solely by servo unit hydraulic power. Since no forces are transmitted back to the pilot, he has no “feel’ of the aerodynamic load acting on the control surfaces. It is necessary therefore, to incorporated an “ artificial feel” device at the point between the pilot’s control and their connection to the servo unit control lever.

Figure 1.2. Power Operated

1.3. Fly by Wire System

Another system which may be considered under the heading of powering flight control, is referred to as a “ fly-by-wire” control system. Although not new in concept, complete redevelopment of the system was seen to be necessary in recent years, as a means controlling some highly sophisticated types of aircraft coming into service. The “fly by wire” system is a control system which wires carrying electrical signal from the pilot’s control , replace mechanical linkage entirely.

In operation, a movement of the control column and rudder pedals, and the power exerted by the pilot are measured by electrical transducers and the signal produce are then amplified and relayed to opera the hydraulic actuator units which are directly connected to the flight control surfaces.


Page 3 of 109 Figure 1.3. Fly By Wire simple Diagram

1.4. Artificial Feel System

When hydraulic actuator is used, an artificial feel system must be provided to prevent over-control by the pilot. In the case of the ailerons, a spring force is usually adequate. However, in dealing with elevators and rudders, it is common to have not only spring force but also to variable hydraulic force.

The hydraulic artificial feel is essentially varied as a function of airspeed. Artificial spring feel alone may be adequate at low speed, abut at high speeds greater resistance to cockpit control movement is needed to prevent overstressing the aircraft structure.

Artificial feel system serve another useful purpose. They position the cockpit control to a neutral position when the pilot releases the control wheel, column or rudder pedals. The neutral position in the case of the elevators is the position where the elevators are faired with the horizontal stabilizer.

The schematic of the feel computer shows how the hydraulic pressure on the hydraulic feel piston is varied as function of airspeed and horizontal stabilizer position.


Page 4 of 109

Figure 1.4. Artificial Feel System

2. AUTOPILOT 2.1. General

An autopilot is a mechanical, electrical, or hydraulic system used to guide a vehicle/ (aircraft) without assistance from a human being.

The first aircraft autopilot was developed by Sperry Corporation in 1912. The autopilot connected a gyroscopic heading indicator and attitude indicator to hydraulically operated elevators and rudder. It permitted the aircraft to fly straight and level on a compass course without a pilot's attention.

2.2. Simple Control System

A control system in which the measure value or a controlled condition is compared with a set value and correction dependent on their difference is applied to the correcting condition in order to adjust the controlled condition, without human intervention in the closed loop formed by the comparing and correcting chains of element and the process. This called Automatic Closed Loop process control system.

The physical elements are usually represented by Block Diagram, so this will be more easier in system performance analysis and design calculation. A remote-position control system of DC Motor servomechanism is given as example. The typical closed-loop features are:, feedback, comparison and error amplification.


Page 5 of 109 Figure 2.1. Block Diagram of A Remote Position Control System

2.2.1. Servo Mechanism

A servo mechanism is any control system used for the control of motion parameter such a displacement, velocity and acceleration. The objective of the control system is to displace the process in such manner that it follows a continually changing input or desired value (sometimes known as the demand signal ).

This system are inherently fast-acting, having very small time lags and response times in the order of milliseconds. Because of the fast response speed required, this type of system usually employs electrical or hydraulic actuation.


Page 6 of 109 2.2.2. Feedback Control System

The advantages of feedback control are: 1. The provision of stability

2. The adjustment of dynamic response, including reduction of lags and provision of desired or specified command/response relationship, especially as regard the improvement of linearity and the reduction of the effect of cross coupling forces. 3. The suppression of unwanted inputs and disturbances

4. The suppression of the effect of variations and uncertainties in the characteristic of the controlled element.

Feedback can improve the speed of response and may be used so as to enforce some desired correspondence between the input and output of system.

2.3. Automatic Flight Control System ( AFCS )

Automatic Flight Control System means a system which automatically controls an aircraft about its pitch, roll, and yaw axis or combination of these axis, and include related sub system such as, stability augmentation system, speed command, auto throttle and trim systems.

Basically AFCS need to know at all times: 1. The attitude of an aircraft in pitch and roll. 2. The position of flight control surfaces. 3. What maneuver is to be performed.

From these data, AFCS will maintain move flight control and trim to maintain altitude and perform commanded maneuver.

Some AFCS also can:

1. Fly and maintain altitude and selected heading.

2. Capture and follow a course a long radio guidance beam. 3. Fly at selected speed or mach number.

4. Fly to fixed point of latitude and longitude . 5. Land the aircraft.

For these purpose AFCS has to get input data from compass indicating, flight director, radio navigation , air data and inertial navigation system,


Page 7 of 109 Figure 2.3. Disposition of Flight Control Surfaces

Figure 2.4. Automatic Flight Control System

2.3.1. Classification of System

Based on the number of axes about which control is effected, it is usual to classify system in the following manner:

1. Single Axis in which attitude control is normally about the roll axis only. The control surfaces forming part of the one and only control loop, there is ailerons. 2. Two axis in which attitude control is about the roll and pitch axis. The control surfaces forming part of the two loops, there are ailerons and elevators.

3. Three axis in which attitude control about all three axes is carried out by specifically related control channels of an automatic flight control system.


Page 8 of 109 2.3.2. Inner Loop Stabilization

In a closed loop AFCS, there are four principal element which together are allocated the task of coping with inner loop stabilization. The principal element are:

1. Sensing of attitude changes of the aircraft about its principal axes by means or stable reference devices, e.g. Gyroscope and accelerometer.

2. Sensing of attitude changes in terms of error signal and the transmission of such signals.

3. Processing of error signal and their conversion into a form suitable operation of the servomotor forming the output stage.

4. Conversion of processed signal into movement of the aircraft flight control surfaces.

In basic mode operation, the function of an automatic control system is to hold an aircraft on desired flight path, by detecting and correcting any departure from the path, in other words, it functions as a stability augmentation system ( SAS )


Page 9 of 109 2.3.3. Outer Loop Control

In addition to performing the primary function of stabilization an automatic flight control system can also be developed to perform the task of modifying the stabilized attitude of an aircraft by computing the necessary maneuver from input such raw data as airspeed, altitude, magnetic heading, interception of radio beam from ground based aid, etc. Such data input constitute outer loop control. The number of inputs serving as an indication of the progressive development of automatic flight from the basic single axis wing-leveling type of autopilot to the highly sophisticated flight guidance system now used in many present-day transport aircraft.

Figure 2.6. Inner Loop stabilization and Outer Loop Control

2.3.4. Signal Processing 1. Limiting

Limit signal must be placed on commanded control signals to prevent excessive attitude changes and a harsh maneuvering. There are two limiting elements in the signal processing : a roll command rate limiter and a roll command limiter

2. Synchronizing

AFCS is usually matched to the position or condition of the aircraft before its engage to take control. Any mismatched in the system may result undesirable condition or cause unsafe movement of aircraft.

Two synchronizing element to eliminate any mismatched in the system, those are altitude datum and integrator network.


Page 10 of 109 3. Gaining

Altering the response of an automatic system to any given level of input signal, thereby obtaining the best signal ratio to the operation of system when working in combination. ( mechanical gearing system )

Figure 2.7. Signal Processing

2.4. Servo Motors

The power output of any automatic flight control system consists of servomotors, or servo-actuators, which is connected in the aircraft’s primary flight control system circuit. The number of servomotors employed is governed by the number of control loops required.

In general, servomotors operate on either electro-pneumatic, electro-mechanical, or electro-hydraulic principles. Servomotors may be connected either in series or in parallel with the normal flight control system of an aircraft. A series-connected servomotor is one which moves the flight control surfaces without moving the pilot’s control. A parallel-connected servomotor moves both the flight control surfaces and the pilot’s control.

2.4.1. Electro-Pneumatic Servomotor

A servomotor designed for used in one particular type of three-axis autopilot system consists of an electro-magnetic valve assembly, comprised of dual poppet valves which are connected via pressure ports and orifices to two cylinders containing pistons sealed against pressure loss by rolling diaphragms (‘roll-frams’), The valve


Page 11 of 109 are controlled by electrical command signal from the autopilot signal processing element, and the pressure for actuation of the pistons is supplied either from engine-driven pump or from a tapping at a turbine engine compressor stage. The operating pressure is determined by the control force characteristic of the aircraft in which the particular autopilot is installed. Typical pressure range is from 7.5 to 30 psi. The piston rods are designed to drive an output linkage assembly which is connected to the appropriate flight control system circuit through a cable drum.

With no command input, each valve is open for an equal period of time, and so there is equal pressure in both cylinders and no output torque is transmitted to the control system. When a control command signal is introduced, the open-time period of one valve is increased, while the open-time of the other valve is decreased. Thus a differential pressure is developed in the two cylinders causing one piston rod to be extended and the other to be retracted, thereby causing rotation of the output linkage and deflection of the control surface to which it is connected.

Figure 2.8. Electro-Pneumatic Servomotor

2.4.2. Electro-Mechanical Servomotor

Type of the servomotors are designed for used in automatic control system, are Direct current and Alternating current.

Direct current operated servomotor.


Page 12 of 109 A servomotor consists of a motor which is coupled to the flight control system via an electro-magnetic clutch, a gear train and sprocket and chain. The servomotor also carries a solid-stated servo amplifier which amplifies the error signal transmitted by the attitude sensing transducer. Feedback is provided by a potentiometer, the wiper of which is driven by motor.

Two Phase Induction Servomotor

The alternating current operated servomotors may be either of two-phase induction motor type or of the type using the principle of hysteresis as applied to the gyroscope of certain attitude sensing element.

Two-phase induction motor type of servomotor has its reference phase constantly supplied with 115 volts alternating current at a frequency of 400 Hz.

Figure 2.10. Two-Phase Induction Type Servomotor

The control phase is supplied by the output of the associated servo amplifier, the voltage carrying from zero to 240 volts. The motor drives an output pulley via gear train and an electromagnetic clutch, the pulley providing the coupling between the servomotor and cable of the aircraft’s flight control system. A CX synchro and a device known as a Tachogenerator, are also geared to the motor, their respective functions being to provide position and rate feedback signals.

Hysteresis Servomotor

A servomotor utilizing a hysteresis motor is shown below. It operates on the same fundamental principle as the gyroscope motor but whereas in the lattes stator is directly connected to a three-phase supply of 115 volts at 400 Hz. To produce a unidirectional rotating field, the three-phase stator in the example of servomotor illustrated, is fed from a single-phase supply, and field rotation in either direction is obtained by splitting the phases by means of capacitors. The single-phase supply is connected to, or disconnected from the stator by means of silicon controlled rectifier ( SCR). Activation or “firing” one or other SCR is achieved by connecting the firing circuit to those circuit supplying the command signal which determine the direction in which the stator field and hence the servomotor must rotate in order to apply corrective control. Coupling between the motor and the aircraft’s flight control


Page 13 of 109 system is by means of a gear train and an electromagnetic clutch, and feedback signal are supplied by a tachogenerator coupled to the motor gear train.

Figure 2.11. Servomotor Utilizing a Hysteresis Motor

2.4.3. Electro-Hydraulic Servo Control

An example of the elevator control system is schematically illustrated below. The principal components of the control unit directly associated with automatic control are: shut-off valve, transfer valve, engage cam and position transducer for supplying feedback signal.



3.1. General

A control wheel steering mode ( CWS ) is provided in some automatic flight control system, its purpose being to enable the pilot to maneuver his aircraft in pitch, roll through the automatic control system by exerting normal maneuvering forces on the control wheel. The pitch and roll forces applied by the pilot are sensed by force transducers which generate output voltage signal proportional to the forces. The signal are supplied to the pitch and roll channels of the automatic flight control system.

When the autopilot is operating in “ Control Wheel Steering “ mode, the transducers provide signal to the roll and pitch channel to operate that ailerons or elevators as desired by the pilot.

Figure 3.1. CWS Force Transducer

3.2. Pitch Computer CWS Mode

While the autopilot is in control wheel steering mode the flight director could be in any mode selected by the pilot. He can utilize his manual control of the autopilot to follow the commands of the command bars in the ADI.

Control Wheel Steering Mode ordinarily operates in conjunction with attitude hold mode. A control wheel steering maneuver will be initiated only if pilot applies sufficient force to the control column, on the order of ten pounds, to cause the control wheel steering level detector to operate its switch. In that case, it switches in the CWS signal from the force transducer. The signal goes to the command limiter, where it is limited so as not to cause excessive airplane attitude changes. From there it goes to the servo motor amplifier, is rate limited , and runs the servo motor.

As long as CWS signal is present, the servo motor continues to run, turning the control synchro rotor, which changes aircraft pitch attitude. As long as the control synchro rotor is turning, there is a difference between its signal and the pitch attitude signal.


Page 15 of 109 This difference signal causes the transfer valve amplifier to keep the autopilot actuator moved far enough from neutral so that it develops a follow-up signal to cancel the difference signal. The elevators are therefore displaced far enough to cause the pitch attitude to change fast enough to follow up on the changes in position of the control synchro rotor. As soon as the pilot stop applying force to the control column, the control wheel steering level detector opens its switch and the autopilot revert to attitude hold mode.

Figure 3.2. PITCH Computer

3.3. Roll Computer CWS Mode

The output of the control wheel force transducer is switched into the servo motor circuit ahead of the roll rate limiter, through the switch operated by the control wheel steering level detector.

As long as the pilot maintains s force on the control wheel, the servo motor continues to run, increasing the airplane bank angle. When he releases the control wheel, the autopilot reverts to attitude hold mode.

If the pilot wishes to return the airplane to wings level or bank in the opposite direction, he applies a force on the control wheel in the opposite direction. The reverses the phase of the control wheel steering signal, causing the servo motor to run its synchro rotors in the opposite direction.


Page 16 of 109 When the autopilot is in control wheel steering mode, the rest of the computer is disconnected from the autopilot system by the switch above bank angle limiter.

The signal into and out of the transfer valve, amplifier are shown in solid black lines because as long as the CWS level detector is operated, the ailerons are operated.

Figure 3.3. ROLL Computer

4. OPERATIONAL MODES 4.1. Roll Channel

4.1.1. Basic Attitude Stabilization Mode


Page 17 of 109 The diagram above shows the synchronizing action of computer card servo motor loop prior to engaging the autopilot, and the autopilot maintaining the existing bank angle at the time of engage.

When the autopilot is disengaged, the servo amp and servo motor re operative. Roll attitude information from the vertical gyro presents itself as a resultant field in the stator of the control synchro in the computer. The active servo motor loop maintains the control synchro rotor at a position perpendicular to the resultant field in the stator.

Therefore at any time prior to engage, the position of the control synchro rotor is function of the bank angle of the airplane. If the position shown represents wing level, then the airplane banks to the right, the control synchro rotor will turn a corresponding number of the degrees clockwise ( prior to engage ).

To visualize this on the schematic, suppose the airplane is banked 20⁰ to the right. The vertical gyro transmit synchro rotor would have to move 20⁰ clockwise with respect to this stator. The resultant field in the control synchro rotor is correspondingly moved 20⁰ clockwise. The servo motor loop causes the control synchro rotor to follow the field, and the rotor also is turned 20⁰ clockwise.

Any changes in airplane bank attitude after autopilot engaged causes the vertical gyro transmit synchro to move the field in the control synchro, developing a not null signal in the control synchro rotor. The phase of the signal developed will cause the ailerons to operate in the direction required to restore the roll attitude existing at the time of engage. It is position of the control synchro rotor which determines airplane roll attitude.

4.1.2. Turn Command Mode


Page 18 of 109 The diagram illustrates an arrangement whereby the autopilot, at the time of engage, causes the airplane to come to a wings level condition if it is not already holding that attitude. The condition prior to engage is the same as previous prior to engage, with the servo motor loop holding the control synchro rotor perpendicular to the field in the stators, and therefore at a position corresponding to the airplane bank angle.

And additional synchro utilizing a sine winding is driven by the servo so that, when the airplane wings are level, the output of the stator is a null. The two synchro rotors are on a common shaft and turn degree of degree.

At the time of engage, when the right wing was down 20⁰ , the control synchro rotor began to move toward the wings level position. As soon as it began to move, it developed an error signal because it was no longer perpendicular to its field. The error signal was of the phase which caused the ailerons to operate and roll the airplane toward wings level.

As the airplane rolled toward wings level, the vertical gyro transmit synchro rotor moved toward wings level, causing the field in the control synchro in the computer to follow the motion of the rotor.

The control synchro rotor output continues until both it and the airplane attitude are in wings level position. The filed is then perpendicular to the rotor. There is no input to the transfer valve amplifier, the ailerons are not displaced, and the airplane holds its wings level attitude. If the airplane deviates from wings level, the vertical gyro transmit synchro moves the control synchro field away from perpendicular to its rotor, developing an error signal which causes the transfer valve to return the airplane to wings level attitude.

4.1.3. Heading Select Mode


Page 19 of 109 The diagram above illustrates the operation of the autopilot in heading select mode. The first airplane position shows the autopilot maintaining the heading of 90⁰ , which is the selected heading shown by the heading select cursor ( triangle ) at 90⁰ on the compass card. The cursor is also at the upper lubber line of the Horizontal Situation Indicator, so the signal from the heading select synchro is null.

The second position of the airplane shown on the HSI, that the pilot has selected a new heading of 150⁰ by moving the heading select bug to 150⁰ position on the compass card. Since the heading select error signal into the autopilot is a direct function of the separation of the heading select bug from the upper lubber line, he has introduced a very large signal into the computer, calling for maximum bank angle to the right. The autopilot banks the airplane to its maximum bank angle, and the airplane turns to the right.

The compass card begins to rotate counterclockwise and the heading select bug with it. As the airplane approaches the new selected heading, the amplitude of the heading select signal diminishes, calling for less and less bank angle. When the heading select bug gets to the upper lubber line, the airplane is on its new selected heading and the autopilot is holding wings level.

If the airplane deviates from the selected heading, the bug moves away from the lubber line. The heading select error synchro then develops a signal causing the airplane to bank until the selected heading is restored.

4.1.4. VOR / LOC Mode


Page 20 of 109 The diagram above illustrates a typical capture of a VOR selected course. VOR or LOC capture is a mode switching function accomplished by a circuit is called “lateral beam sensor”.

The airplane is shown approaching the selected at an angle of 45⁰. The autopilot could be in heading mode, heading select mode or CWS mode. The autopilot mode select switch will have been moved to the VOR/LOC position. VOR capture in most autopilot occurs at about

one dot

(5 ⁰ ) of deviation from the selected radial.

In this illustration, the 270⁰ radial is captured. Capture of the 90⁰ radial while traveling the same direction on the other side of the VOR station would be the same operation.

At the time of capture, the intercept mode is automatically discontinued and VOR capture mode initiated. During VOR capture modes, the principal input signals are course select error and radio deviation. At the time capture (from the right side of the beam), radio deviation calls for the left turn, course select error calls for a right turn. Unless the intercept angle is unusually small, the course select error signal predominates, causing the airplane to turn toward the right, making a smooth approach to the radial. As the airplane gets closer to the selected path, the deviation signal diminishes and the course select error signal diminishes. The course select error signal will always predominate, however because if it did not, the deviation signal would diminish until the course select error did predominate. During the capture mode, the bank angle limit is typically 25⁰ or 30⁰ and roll rate limit is on the order of 4⁰ to 7⁰ per second.

4.2. Pitch Channel

4.2.1. Central Air Data Computer ( CADC )


Page 21 of 109 The computer consists of two pressure transducers, one for the measurement of Airspeed and the other for measuring Altitude. Each transducer is coupled to an inductive pick-off element the signals from which operate motors, gear trains and shafts, the rotation of the shafts being proportional to dynamic pressure and static pressure.

It’s necessary to correct the dynamic and static pressure inputs to CADC, for non linear characteristic, i.e. the airspeed square-law and the inverse pressure/altitude characteristic. Furthermore, corrections for errors arising from variations in airflow, which might possibly arise at a static vent location, must also be considered.

Corrections are accomplished by coupling accurately profiled cams to the appropriate output shafts driven by the motors of the pressure transducer pick-off elements. The cams are provided with the “followers” which actuate gear train, and further output shaft coupled to the rotors of CX synchros, the stators of which are connected to the CX synchros in an airspeed indicator, an altimeter and in the relevant hold mode circuits of the automatic flight control system. Thus, the rotation of the shafts and the synchro signals are converted into the required linear outputs.

4.2.2. Altitude Hold

Any changes of aircraft altitude about its pitch axis while in straight and level flight will be detected by the pitch attitude sensing element of the automatic control system, and the changes will be accordingly corrected.

Figure 4.6 Altitude Hold Sensor (1)

The sensor consists of a pressure transducer comprising an evacuated capsule assembly and E and I type of inductive pick-off element amplifier, and a two-phase induction type of chaser motor. The capsule assembly is subjected to changes of static pressure supplied to the case of sensor unit from the aircraft’s static pressure system, and its mechanically linked to the “I” bar of the pick-off element. A change of altitude produces a change of static pressure to cause the capsule assembly to expand or close up. Displaces the “I” bar and the signal is induced in the coil of the center limb of the “E” bar, the signal being a measure of the direction and rate of altitude change.


Page 22 of 109 Another example of an Altitude hold sensor which part of a Central Air Data Computer. In this case, the pressure transducer is connected to the cores in such a manner that they moved differentially within the winding of a differential transformer element to provide an altitude error signal from zero signal condition. The signal is amplified and drives the transducer capsule assembly in a direction opposite to that caused by an altitude change, so reducing the error signal to zero.

Figure 4.6 Altitude Hold Sensor (2)

The chaser motor is also connected to two solenoid-operated clutches, one engaging ganged potentiometers, and the other a CX synchro rotor. The potentiometers are in the signal line to the pitch servomotor, their function being to attenuate control signals as a function of sensed static pressure , and thereby adjust control loop gain for optimum operation. The function of the CX synchro is to transmit the altitude error signal to the pitch servomotor which will operate to return the aircraft to the altitude it is required to hold.

Another feature of this sensor is chaser motor damping to prevent oscillations as the motor and pick-off at the zero signal position. This is accomplished by feeding back an opposing signal from a rate generator driven by chaser motor.

The spring-loaded override assembly will open the linkage when a pressure change is applied.

4.2.3. Vertical Speed Selection and Hold

In climbing out after take-off, it is necessary for a particular rate of climb or vertical speed to be maintained and in order for this to be effected by an automatic control system, a vertical speed reference signal must first be established before engagement of the system.

In the diagram below, this rate signal is originated by a Tachogenerator driven by the altitude sensor of a central air data computer and is supplied to the pitch channel of the control system through a vertical speed mode select circuit which forms part of a


Page 23 of 109 pilot’s control unit. Circuits may vary between control systems, but the fundamentals of mode selection and operation.

The rate signal is applied to summing point 1, and after amplification it drives a vertical speed motor and generator, a gear train and an electrically operated clutch assembly. Since the signal must be synchronized before engagement of the pitch control channel, the clutch at this stage is energized and so through a further gear train and override mechanism, the clutch drives the vertical speed wheel of the controller in the “climb” direction and to the position corresponding to the prevailing vertical speed of the aircraft. The rate signal is also supplied to summing point 2 and the pitch

computer, via summing point 3 but as the pitch servomotor is not engaged no pitch control is applied. The vertical speed motor also drives a potentiometer wiper which feeds back a signal to summing point 1 and cancels the rate signal from the central air data computer when vertical speed wheel is positioned as the required speed.

Potentiometer signal is also supplied to summing point 2, and also cancelled the rate signal at this point, thus the vertical speed section of the pilot’s controller is in overall synchronism with the prevailing vertical speed of an aircraft.

Figure 4.7. Vertical Speed Selection And Hold

4.2.4. Airspeed Hold

Altitude sensor is required to measures only static pressure changes. Airspeed sensor is required to measure difference between static and dynamic pressure. The capsule assembly is opened to the source of dynamic pressure, and static pressure admitted to the sealed chamber in which the assembly is continued. The capsule expands or closes up under the influence of a pressure differential created by change of airspeed.

4.2.5. Mach Trim

In aircraft which are capable of flying at high subsonic speeds, and of transition to supersonic speed, larger than normal rearward movement of the wing center of pressure occur and in consequence larger nose-down pitching moments are produced. The compressibility effects arise which make the counteracting nose-up pitching moment


Page 24 of 109 produced by trimming the horizontal; stabilizer to a negative angle of attack position, less effective as aircraft speed increases.

Mach trim system is installed on aircraft which automatically senses increases of speed above the appropriate datum Mach number, by means of servo coupling, automatically re-adjust the position of the horizontal stabilizer thereby maintaining the pitch trim of the aircraft.

Figure 4.8 Mach Trim schematic diagram

4.3. Yaw Damping

All aircraft, particularly those having a swept-wing configuration, are subject to a yawing-rolling oscillation popularly is known as “Dutch Roll” but difference aircraft exhibit varying degrees of damping, i.e. the inherent tendency to reduce the magnitude of oscillation of eventual return to straight flight varies.

The natural damping of the Dutch-Roll tendency is dependent not only on the size of the vertical stabilizer and rudder, but also on the aircraft’s speed, the damping being more responsive at high speed than at low speed.

The system is designed that it can be operated independently of the automatic control system, so that in the event that the aircraft must be flown manually, Dutch Roll tendencies can still be counteracted.

The operating fundamentals of yaw damper system is generally may be understood from the picture below. The principal components of a system is the yaw damper coupler which contains a rate gyro powered directly from the aircraft power supply and the logic switching circuits relevant to filtering, integration, synchronizing,


Page 25 of 109 demodulation, and servo amplification. Servo amplifier output is supplied to the transfer valve of the rudder power control unit. This unit differs from those used for aileron and elevator control in that has an additional actuator ( yaw damper actuator ) and does not include the automatic control system engage mechanism.

An automatic flight control system may be used in all modes with the yaw damper system engaged, however the associated interlock circuit prevents the use of the control system when the yaw damper is engaged.

Figure 4.9 Yaw Damper schematic diagram

4.4. Flight Director

Flight Director System ( FDS ) or Integrated Flight System ( IFS ) means a system which integrates a number of signal inputs to provide an output to a display system. These inputs may include reference, heading, VHF Omni Range system ( VOR ), localizer, glide-slope, marker, radio altimeter, Inertial Navigation System ( INS ), Doppler, DME, RNAV, VLF Navigation system, Omega navigation system, speed control signal information.

A Flight director system developed in this manner comprises two principal display unit, they are variously called: Attitude Direction Indicator (ADI), Flight Director, or an approach horizon and Horizontal Situation Indicator ( HSI ) or a Course Deviation Indicator.


Page 26 of 109 Figure 4.10b Flight Director/ Attitude Director Indicator

Figure 4.11a Horizontal Situation Indicator


Page 27 of 109 4.5. Auto Throttle/ Thrust

An auto throttle (automatic throttle) allows a pilot to control the power setting of an aircraft's engines by specifying a desired flight characteristic, rather than manually controlling fuel flow. These systems can conserve fuel and extend engine life by metering the precise amount of fuel required to attain a specific target indicated air speed, or the assigned power for different phases of flight. A/T and AFDS (Auto Flight Director System) work together to fulfill the whole flight plan and greatly reduce pilots' work load.

4.5.1. Operation modes

In Speed mode the throttle is positioned to attain a set target speed. This mode controls aircraft speed within safe operating margins. For example, if the pilot selects a target speed which is slower than stall speed, or a speed faster than maximum speed, the auto throttle system will maintain a speed closest to the target speed that is within the range of safe speeds.

In Thrust mode the engine is maintained at a fixed power setting according to the different flight phases. For example, during Takeoff, A/T maintains a constant Takeoff power until Takeoff mode is finished. During Climb, A/T maintains a constant climb power; in Descent, A/T retards the throttle to IDLE position, and so on. When A/T is working in Thrust mode, speed is controlled by pitch (or the control column), and not protected by A/T. A Radar Altimeter feeds data to the auto throttle mostly in this mode.

4.5.2. Usage

On Boeing type aircraft, A/T can be used in all flight phases from Takeoff, Climb, Cruise, Descent, Approach, all the way to Land or Go-around, barring malfunction. Taxi is not considered as a part of flight, and A/T does not work for Taxi. In most cases, A/T mode selection is automatic without the need of any manual selection unless interrupted by pilots.

According to Boeing published flight procedures, A/T is engaged in BEFORE the takeoff procedure and is automatically disconnected 2 seconds after landing. During flight, manual override of A/T is always available. A release of manual override allows A/T to regain control, and the throttle will go back to the A/T commanded position except for 2 modes (Boeing type aircraft): IDLE and THR HLD. In these two modes, the throttle will remain at the manual commanded position.

4.6. Auto Landing 4.6.1. Description

In aviation, autoland describes a system that fully automates the landing phase of an aircraft's flight, with the human crew supervising the process.

Autoland systems were designed to make landing possible in visibility too poor to permit any form of visual landing, although they can be used at any level of visibility. They are usually used when visibility is less than 600 meters RVR and/or in adverse weather conditions, although limitations do apply for most aircraft—for example, for


Page 28 of 109 a Boeing 747-400 the limitations are a maximum headwind of 25 knots, a maximum tailwind of 10 knots, a maximum crosswind component of 25 knots, and a maximum crosswind with one engine inoperative of five knots. They may also include automatic braking to a full stop once the aircraft is on the ground, in conjunction with the autobrake system, and sometimes auto deployment of spoilers and thrust reversers.

Autoland may be used for any suitably approved Instrument Landing System (ILS) or Microwave Landing System (MLS) approach, and is sometimes used to maintain currency of the aircraft and crew, as well as for its main purpose of assisting an aircraft landing in low visibility and/or bad weather.

Autoland requires the use of a radar altimeter to determine the aircraft's height above the ground very precisely so as to initiate the landing flare at the correct height (usually about 50 feet (15 m)). The localizer signal of the ILS may be used for lateral control even after touchdown until the pilot disengages the autopilot. For safety reasons, once autoland is engaged and the ILS signals have been acquired by the autoland system, it will proceed to landing without further intervention, and can be disengaged only by completely disconnecting the autopilot (this prevents accidental disengagement of the autoland system at a critical moment). At least two and often three independent autopilot systems work in concert to carry out autoland, thus providing redundant protection against failures. Most autoland systems can operate with a single autopilot in an emergency, but they are only certified when multiple autopilots are available.

The autoland system's response rate to external stimuli work very well in conditions of reduced visibility and relatively calm or steady winds, but the purposefully limited response rate means they are not generally smooth in their responses to varying wind shear or gusting wind conditions – i.e. not able to compensate in all dimensions rapidly enough – to safely permit their use.


Page 29 of 109 4.6.2. Auto Landing Category

Instrument-aided landings are defined in categories by the International Civil Aviation Organization. These are dependent upon the required visibility level and the degree to which the landing can be conducted automatically without input by the pilot.

CAT I - This category permits pilots to land with a decision height of 200 ft (61 m) and a forward visibility or Runway Visual Range (RVR) of 550 m. Simplex autopilots are sufficient.

CAT II - This category permits pilots to land with a decision height between 200 ft and 100 ft (≈ 30 m) and a RVR of 300 m. Autopilots have a fail passive requirement.

CAT IIIa -This category permits pilots to land with a decision height as low as 50 ft (15 m) and a RVR of 200 m. It needs a fail-passive autopilot. There must be only a 10−6 probability of landing outside the prescribed area.

CAT IIIb - As IIIa but with the addition of automatic roll out after touchdown incorporated with the pilot taking control some distance along the runway. This category permits pilots to land with a decision height less than 50 feet or no decision height and a forward visibility of 250 ft (76 m, compare this to aircraft size, some of which are now over 70 m long) or 300 ft (91 m) in the United States. For a landing-without-decision aid, a fail-operational autopilot is needed. For this category some form of runway guidance system is needed: at least fail-passive but it needs to be fail-operational for landing without decision height or for RVR below 100 m.

CAT IIIc - As IIIb, but without decision height or visibility minimums, also known as "zero-zero".

Fail-passive autopilot: in case of failure, the aircraft stays in a controllable position and the pilot can take control of it to go around or finish landing. It is usually a dual-channel system.

Fail-operational autopilot: in case of a failure below alert height, the approach, flare and landing can still be completed automatically. It is usually a triple-channel system or dual system.


Page 30 of 109




1.1. General

Radio is the wireless transmission of signals through free space by electromagnetic radiation of a frequency significantly below that of visible light, in the radio frequency range, from about 30 KHz to 300 GHz. These waves are called radio waves. Electromagnetic radiation travels by means of oscillating electromagnetic fields that pass through the air and the vacuum of space.

Information, such as sound, is carried by systematically changing (modulating) some property of the radiated waves, such as their amplitude, frequency, phase, or pulse width. When radio waves strike an electrical conductor, the oscillating fields induce an alternating current in the conductor. The information in the waves can be extracted and transformed back into its original form. Radio systems used for communications will have the following elements. With more than 100 years of development, each process is implemented by a wide range of methods, specialized for different communications purposes.

1.2 Transmitter and Modulation

Each system contains a transmitter. This consists of a source of electrical energy, producing alternating current of a desired frequency of oscillation. The transmitter contains a system to modulate (change) some property of the energy produced to impress a signal on it. This modulation might be as simple as turning the energy on and off, or altering more subtle properties such as amplitude, frequency, phase, or combinations of these properties. The transmitter sends the modulated electrical energy to a tuned resonant antenna; this structure converts the rapidly changing alternating current into an electromagnetic wave that can move through free space (sometimes with a particular polarization).

Amplitude modulation of a carrier wave works by varying the strength of the transmitted signal in proportion to the information being sent. For example, changes in the signal strength can be used to reflect the sounds to be reproduced by a speaker, or to specify the light intensity of television pixels. It was the method used for the first audio radio transmissions, and remains in use today. "AM" is often used to refer to the medium wave broadcast band.


Page 31 of 109 Frequency modulation varies the frequency of the carrier. The instantaneous frequency of the carrier is directly proportional to the instantaneous value of the input signal. Digital data can be sent by shifting the carrier's frequency among a set of discrete values, a technique known as frequency-shift keying.

Figure 1.2. Frequency Modulation

FM is commonly used at VHF radio frequencies for high-fidelity broadcasts of music and speech. Normal (analog) TV sound is also broadcast using FM.

Angle modulation alters the instantaneous phase of the carrier wave to transmit a signal. It is another term for Phase modulation.

Figure 1.3. Phase Modulation Antenna

An antenna (or aerial) is an electrical device which converts electric currents into radio waves, and vice versa. It is usually used with a radio transmitter or radio receiver. In transmission, a radio transmitter applies an oscillating radio frequency electric current to the antenna's terminals, and the antenna radiates the energy from the current as electromagnetic waves (radio waves). In reception, an antenna intercepts some of the power of an electromagnetic wave in order to produce a tiny voltage at its terminals, that is applied to a receiver to be amplified. An antenna can be used for both transmitting and receiving.


Page 32 of 109

1.3. Propagation

Once generated, electromagnetic waves travel through space either directly, or have their path altered by reflection, refraction or diffraction, scattering, and fading. The intensity of the waves diminishes due to geometric dispersion (the inverse-square law); some energy may also be absorbed by the intervening medium in some cases. Noise will generally alter the desired signal; this electromagnetic interference comes from natural sources, as well as from artificial sources such as other transmitters and accidental radiators. Noise is also produced at every step due to the inherent properties of the devices used. If the magnitude of the noise is large enough, the desired signal will no longer be discernible; this is the fundamental limit to the range of radio communications.

1.3.1. Reflection

The reflection of electromagnetic waves by conducting medium. Some of radio energy will absorbed in the medium that the wave hit, some of it may pass through the material. Reflection coefficient is the ratio of the dielectric intensity of the reflected wave to that of the incident wave .

Figure 1.4 Reflection of RF

Figure 1.5 . Angle of reflection 1.3.2. Refraction

Refraction is bending of a wave as it passes from one medium into another. When a radio wave is transmitted into an ionized layer, refraction, or bending of the wave, occurs. Refraction is caused by an abrupt change in the velocity of the upper part of a radio wave as it strikes or enters a new medium. The amount of refraction that occurs depends on three main factors:

(1) the density of ionization of the layer (2) the frequency of the radio wave


Page 33 of 109

Figure.1.6. Refraction of RF 1.3.3. Scattering

All electromagnetic wave propagation is subject to scattering influences that alter idealized pattern.. Scattering of light and radio waves (especially in radar) is particularly important. Several different aspects of electromagnetic scattering are distinct enough to have conventional names. Major forms of elastic light scattering (involving negligible energy transfer) are Rayleigh scattering and Mie scattering. Inelastic scattering includes Brillouin scattering, Raman scattering, inelastic X-ray scattering and Compton scattering. The degree of scattering varies as a function of the ratio of the particle diameter to the wavelength of the radiation, along with many other factors including polarization, angle, and coherence

Figure 1.7. Scatter of RF 1.3.4. Fading

In radio wave communications, fading is deviation of the attenuation affecting a signal over certain propagation media. In the other word, fading is the fluctuation in signal strength at receiver and may be rapid or slow. The fading may vary with time, geographical position or radio frequency, and is often modeled as a random process. A fading channel is a communication channel comprising fading. In wireless systems, fading may either be due to multipath propagation, referred to as multipath induced fading, or due to shadowing from obstacles affecting the wave propagation, sometimes referred to as shadow fading.


Page 34 of 109 1.4. Radio Wave Segment


name Frequency Example uses

Very Low

Frequency 3–30 kHz

Navigation, time signals, submarine communication, wireless heart rate monitors, geophysics


Frequency 30–300 kHz

Navigation, time signals, AM long-wave broadcasting , RFID, amateur radio


Frequency 300–3000 kHz

AM (Amplitude Modulation) broadcasts, amateur radio, avalanche beacons


Frequency 3–30 MHz

Shortwave broadcasts, citizens' band radio, amateur radio and over-the-horizon aviation communications, RFID,

Over-the-horizon radar, Automatic link establishment (ALE) / Near Vertical Incidence Sky-wave (NVIS) radio communications,

Marine and mobile radio telephony Very High

Frequency 30–300 MHz

FM, television broadcasts and line-of-sight ground-to-aircraft and aircraft-to-aircraft communications. Land Mobile and Maritime Mobile communications, amateur radio, weather


Ultra High Frequency

300–3000 MHz

Television broadcasts, microwave ovens, microwave devices/communications, radio astronomy, mobile phones, wireless LAN, Bluetooth, Zig-Bee, GPS and two-way radios such as Land Mobile, FRS and GMRS radios, amateur radio Super

High Frequency

3–30 GHz

Radio astronomy, microwave devices/communications, wireless LAN, most modern radars, communications satellites,

satellite television broadcasting, DBS, amateur radio Extremely

High Frequency

30–300 GHz

Radio astronomy, high-frequency microwave radio relay, microwave remote sensing, amateur radio, directed-energy

weapon, millimeter wave scanner


This system is operated on frequency range 2 – 29.999 MHz, with channel spacing 1 KHz. Provides Transmission and reception in High Frequency Band and may be operated in LSB ( Lower Side Band) , USB( Upper Side Band ) or AM ( Amplitude Modulation) mode. This system is available for long distance communication, with a simple and an easy operation. The disadvantages of this system are : noise affected by other electronics devices and natural noise such as thunderstorm.


Page 35 of 109 2.2. Main Components

Basically, this system contain the components are:

a. HF Transceiver, is located on radio rack compartment.

Figure 2.1. HF transceiver

The following control and indicators are mounted in front of HF TxRx : # R/T FAULT RED LIGHT illuminates when any faults in TxRx

# KEY INTERLOCK RED LIGHT illuminates when any fault in associated with antenna Tuner

# SQL/LAMP TEST, when pressed, R/T fault and Key Interlock light will illuminate and disable receiver squelch

b. HF Control Panel, is located in the cockpit, installed either on the overhead instrument panel or pedestal.

Figure 2.2. HF Control Panel The Control Panel provides an indicator and the knobs are :

# Frequency Indicator to display to frequency selected

# Operational Mode Selector knob to select the operation mode on HF system will be operated.


Page 36 of 109 # Four Frequency Selector knobs provide selection of frequency required

# RF sensitivity/ squelch provides manual control for squelch/ sensitivity circuit. When sensitivity increases, the background noises will be increase.

c. HF Antenna Tuner

Figure 2.3. HF Antenna Tuner/coupler

The main function of antenna tuner is to match the output impedance of transceiver to the antenna impedance over the full transceiver frequency range. There Amber light is mounted on antenna tuner will illuminates during tuning process. The green light will illuminates along the HF operation after complete tune process.

d. HF Antenna

The main function of antenna is to convert electrical signal to electromagnetic wave and vise versa. Since horizontally polarized radio waves work better for sky wave propagation due of the greater ground absorption of vertically polarized waves, monopole antennas which have vertical polarization are not much used, and antennas based on horizontal dipoles are mostly used. The most common antennas in this band are wire antennas such as the rhombic antenna, in the upper frequencies, multi-element dipole antennas such as the Yagi, quad, and reflective array antennas. Powerful shortwave broadcasting stations often use large wire curtain arrays

e. Spark Gab

A simple spark gap consists of two conducting electrodes separated by a gap immersed within a gas (typically air). When a sufficiently high voltage is applied, a spark will bridge the gap, ionizing the gas and drastically reducing its electrical resistance. An electric current then flows until the path of ionized gas is broken or the current is reduced below a minimum value called the 'holding current'. In common speaking spark gab will protect the system from a thunderstorm.


Page 37 of 109 2.2. Block Diagram


This system is operated on air band frequency range 118 – 135.975 MHz, with channel spacing 50 KHz. Provides Transmission and reception in very high frequency Band which is frequency modulated. This system is available for short distance communication ( line of sight ) , not affected by other electronics devices or natural noise such as thunderstorm.

3.2. Main Components

Basically, this system contains the components are:

a. VHF Transceiver, is located on radio rack compartment.


Page 38 of 109 The following control and indicators are mounted in front of VHF TxRx :

# Transmit power light illuminates when operated . # Squelch Disable switch

b. VHF Control Panel, is located in the cockpit, installed either on the overhead instrument panel or pedestal.

Figure 3.2. HF Control Panel The Control Panel provides an indicator and the knobs are :

# Frequency Indicator to display to frequency selected.

# Frequency Selector knobs provide selection of frequency as required # Volume control knobs

# RF sensitivity/ squelch provides manual control for squelch/ sensitivity circuit. When sensitivity increases, the background noises will be increase.

c. VHF Antenna

The main function of antenna is to convert electrical signal to electromagnetic wave and vise versa



Page 39 of 109 3.3. Block Diagram


This system is Enables the Pilot and stewardess to address the passengers. It also Produce Chime when triggered from the passenger warning and stewardess call systems. This system contains PA amplifier and a number of speakers in the cabin and toilet.


Page 40 of 109 The power is applied to the passenger address amplifier when the electronic switch panel HI/OFF/LO switch is set to HI or LO. The amplifier has three audio input channels, only two which are used on this installation. Priority circuit within amplifier to ensure the input 1 has priority over the input 2.

Input 1 : Pilot or Copilot jack box, activated by engaging push button switch on audio selector panel and pressing the appropriate the PTT switch.

Input 2 : Stewardess handset. Activated when stewardess lift the handset from its hanged and the PA/IC switch is set to PA position.

There are two outputs from the amplifier. The main output fed the speaker and side tone output provides a side-tones signal to whichever the telephone is in used. The gain of amplifier is preset, but can be set to a level suitable prevailing noise condition by selection of HI or LO on HI/LO switch.


SELCAL is a selective-calling radio system that can alert an aircraft's crew that a ground radio station wishes to communicate with the aircraft. SELCAL uses a ground-based encoder and radio transmitter to broadcast an audio signal that is picked up by a decoder and radio receiver on an aircraft

The use of SELCAL allows an aircraft crew to be notified of incoming communications even when the aircraft's radio has been muted. Thus, crewmembers need not devote their attention to continuous radio listening. SELCAL is operates on the HF and VHF radio frequency band.

An individual aircraft has its own assigned SELCAL code. To initiate a SELCAL transmission, a ground station radio operator enters an aircraft's SELCAL code into a SELCAL encoder. The encoder converts the four-letter code into four designated audio tones. The radio operator's transmitter then broadcasts the audio tones on the aircraft's company radio frequency


Page 41 of 109 6. COCKPIT VOICE RECORDER ( CVR )

6.1. General

The system automatically records the audio output from the flight crew audio selector panel and from the microphone in the microphone monitor unit. Recording is performed on an endless loop tape which is automatically after one half hour of recording. There are 4 channel :

Channel 1 : not used in this installation Channel 2: copilot audio selector panel Channel 3: pilot audio selector panel

Channel 4: microphone monitor unit in the cockpit

Figure 6.1 CVR Unit figure 6.2 Microphone Monitor Unit

6.2. Operation

The audio from each channel is applied to separate amplifier which is combine with a recording bias signal before being applied to the recording head. Before reaching the recording head, the tape to be recorded passes to the erase head which erase a previous recording. The tape transit time between leaving the recording head and reach the erase head is about one and half hour.

At the end of flight, the whole tape may be erased by pressing the erase button on the microphone monitor unit. The tape will be erased with 8 seconds. During test, the 600 Hz audio tone will be available..


Page 42 of 109 6.3. Blok Diagram


Provides discharge path for static electrical charges accumulated by the Aircraft during flight. Installation of static discharge will provide minimized the interference of radio communication. Usually Static Discharges are mounted on each Aileron, each Elevator, and Rudder Trim.



8.1. General

A communications satellite or COMSAT is an artificial satellite sent to space for the purpose of telecommunications. Modern communications satellites use a variety of orbits including geostationary orbits, Molniya orbits, elliptical orbits and low (polar and non-polar Earth orbits).

For fixed (point-to-point) services, communications satellites provide a microwave radio relay technology complementary to that of communication cables. They are also used for mobile applications such as communications to ships, vehicles, planes and hand-held terminals, and for TV and radio broadcasting.

The Merriam-Webster dictionary Merriam-Webster Dictionary defines a satellite as a celestial body orbiting another of larger size or a manufactured object or vehicle intended to orbit the earth, the moon, or another celestial body.

Electronic communications devices like cell phones and computers on the internet utilize satellite communications (SATCOM). Today's satellite communications can trace origins all the way back to the moon. A project named Communication Moon Relay, was a telecommunication project carried out by the United States Navy. Its objective was to develop a secure and reliable method of wireless communication by using the Moon as a natural communications satellite.

8.2. Geostationary orbits

Figure 8.1 Geostationary Orbit

To an observer on the earth, a satellite in a geostationary orbit appears motionless, in a fixed position in the sky. This is because it revolves around the earth at the earth's own angular velocity (360 degrees every 24 hours, in an equatorial orbit).

A geostationary orbit is useful for communications because ground antennas can be aimed at the satellite without their having to track the satellite's motion. This is relatively inexpensive. In applications that require a large number of ground antennas, such as Direct TV distribution, the savings in ground equipment can more than outweigh the cost and complexity of placing a satellite into orbit.

The main drawback of a geostationary orbit is that all ground stations must have a direct line of sight to the satellite. This limits the ground area to 50-60 degrees of either side of the


Page 44 of 109 satellite's position, measured in both latitude and longitude; consequently, a geostationary satellite cannot service extreme northern and southern areas of the world. Another drawback is the height of the orbit, usually which requires more powerful transmitters, larger-than-normal (usually dish) antennas, and higher-sensitivity receivers on the earth. The large distance also introduces a significant delay, of ~0.25 seconds, into communications.

8.3. Low-Earth-orbiting satellites

Figure 8.2 Low Earth Orbit

A low Earth orbit (LEO) typically is a circular orbit about 200 kilometers (120 mi) above the earth's surface and, correspondingly, a period (time to revolve around the earth) of about 90 minutes. Because of their low altitude, these satellites are only visible from within a radius of roughly 1000 kilometers from the sub-satellite point. In addition, satellites in low earth orbit change their position relative to the ground position quickly. So even for local applications, a large number of satellites are needed if the mission requires uninterrupted connectivity.

Low-Earth-orbiting satellites are less expensive to launch into orbit than geostationary satellites and, due to proximity to the ground, do not require as high signal strength (Recall that signal strength falls off as the square of the distance from the source, so the effect is dramatic). Thus there is a tradeoff between the number of satellites and their cost. In addition, there are important differences in the onboard and ground equipment needed to support the two types of missions.

A group of satellites working in concert is known as a satellite constellation. Two such constellations, intended to provide satellite phone services, primarily to remote areas, are the Iridium and Globalstar systems. The Iridium system has 66 satellites. Another LEO satellite constellation known as Teledesic, with backing from Microsoft entrepreneur Paul Allen, was to have over 840 satellites. This was later scaled back to 288 and ultimately ended up only launching one test satellite.

8.4. Molniya satellites

Geostationary satellites must operate above the equator and therefore appear lower on the horizon as the receiver gets the farther from the equator. This will cause problems for extreme northerly latitudes, affecting connectivity and causing multipath (interference caused by signals reflecting off the ground and into the ground antenna). For areas close to the North (and South) Pole, a geostationary satellite may appear below the horizon. Therefore Molniya orbit satellite has been launched, mainly in Russia, to alleviate this problem.


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