ATA CHAPTER 22 AUTO FLIGHT SYSTEM 1. POWER OPERATED CONTROL SURFACES
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.
Figure 1.1. Power Assisted Page 1 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.
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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.
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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 4 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.
Figure 2.2 Electro-Hydraulic Powered Control
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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, Typical example is given in the next diagram.
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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 7 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 )
Figure 2.5. Inner Loop Stabilization
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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 9 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 10 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 enginedriven 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.
Figure 2.9. Direct Current Operated Servomotor Page 11 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 12 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.
Figure 2.12. Electro- Hydraulic Servo Control Page 13 of 109
3. CONTROL WHEEL STEERING 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 14 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 15 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
Figure 4.1. Hold Roll Attitude Page 16 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
Figure 4.2. Maintain Wings level
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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
Figure 4.3. Autopilot Heading Select Mode
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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
Figure 4.4. VOR ( LOC ) Capture
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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 )
Figure 4.5. Schematic Arrangement of Central Air Data Computer ( CADC ) Page 20 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 21 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 22 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 23 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 24 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.
Figure 4.10a Flight Director/ Attitude Director Indicator Page 25 of 109
Figure 4.10b Flight Director/ Attitude Director Indicator
Figure 4.11a Horizontal Situation Indicator
Figure 4.11b Horizontal Situation Indicator Page 26 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 27 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.
Figure 4.12. Phase of Auto Landing System
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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 failpassive 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.
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ATA CHAPTER 23 : COMMUNICATIONS SYSTEM 1. RADIO COMMUNICATION
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.
Figure 1.1. Amplitude Modulation Page 30 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 31 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 3) the angle at which the wave enters the layer. Page 32 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 33 of 109
1.4. Radio Wave Segment Band name
Frequency
Example uses
Very Low Frequency
3–30 kHz
Navigation, time signals, submarine communication, wireless heart rate monitors, geophysics
Low Frequency
30–300 kHz
Navigation, time signals, AM long-wave broadcasting , RFID, amateur radio
Medium Frequency
300–3000 kHz
AM (Amplitude Modulation) broadcasts, amateur radio, avalanche beacons
High Frequency
3–30 MHz
Very High Frequency
30–300 MHz
Shortwave broadcasts, citizens' band radio, amateur radio and over-the-horizon aviation communications, RFID, Over-thehorizon radar, Automatic link establishment (ALE) / Near Vertical Incidence Sky-wave (NVIS) radio communications, Marine and mobile radio telephony FM, television broadcasts and line-of-sight ground-to-aircraft and aircraft-to-aircraft communications. Land Mobile and Maritime Mobile communications, amateur radio, weather radio
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
30–300 GHz
Radio astronomy, high-frequency microwave radio relay, microwave remote sensing, amateur radio, directed-energy weapon, millimeter wave scanner
Extremely High Frequency
2. HF COMMUNICATION SYSTEM
2.1. General 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 34 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 35 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.
Figure 2.4. Spark gab Page 36 of 109
2.2. Block Diagram
3. VHF COMMUNICATION SYSTEM
3.1. General 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.
Figure 3.1. VHF transceiver Page 37 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 .
Figure 3.3. Antenna location
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3.3. Block Diagram
4. PASSENGER ADDRESS
4.1. General 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. 4.2. Operation
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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. 5. SELECTIVE CALLING ( SELCAL )
5.1. General 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
5.2. Block Diagram
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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..
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6.3. Blok Diagram
7. STATIC DISCHARGE
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.
Figure 7. Static discharge
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8. CUMMUNICATIONS SATELLITE
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 43 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-thannormal (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. Page 44 of 109
The first satellite of the Molniya series was launched on April 23, 1965 and was used for experimental transmission of TV signal from a Moscow uplink station to downlink stations located in Siberia and the Russian Far East, in Norilsk, Khabarovsk, Magadan and Vladivostok. In November 1967 Soviet engineers created a unique system of national TV network of satellite television, called Orbita, that was based on Molniya satellites. Molniya orbits can be an appealing alternative in such cases. The Molniya orbit is highly inclined, guaranteeing good elevation over selected positions during the northern portion of the orbit. (Elevation is the extent of the satellite's position above the horizon. Thus, a satellite at the horizon has zero elevation and a satellite directly overhead has elevation of 90 degrees). The Molniya orbit is designed so that the satellite spends the great majority of its time over the far northern latitudes, during which its ground footprint moves only slightly. Its period is one half day, so that the satellite is available for operation over the targeted region for six to nine hours every second revolution. In this way a constellation of three Molniya satellites (plus in-orbit spares) can provide uninterrupted coverage. 8.4. Structure Communications Satellites are usually composed of the following subsystems: 1. Communication Payload, normally composed of transponders, antenna, and switching systems 2. Engines used to bring the satellite to its desired orbit 3. Station Keeping Tracking and stabilization subsystem used to keep the satellite in the right orbit, with its antennas pointed in the right direction, and its power system pointed towards the sun 4. Power subsystem, used to power the Satellite systems, normally composed of solar cells, and batteries that maintain power during solar eclipse 5. Command and Control subsystem, which maintains communications with ground control stations. The ground control earth stations monitor the satellite performance and control its functionality during various phases of its life-cycle. The bandwidth available from a satellite depends upon the number of transponders provided by the satellite. Each service (TV, Voice, Internet, radio) requires a different amount of bandwidth for transmission. This is typically known as link budgeting and a network simulator can be used to arrive at the exact value.
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9. AIRCRAFT CUMMUNICATIONS ADDRESSING AND REPORTING SYSTEM
9.1. General Aircraft Communications Addressing and Reporting System (ACARS) is a digital datalink system for transmission of short, relatively simple messages between aircraft and ground stations via radio or satellite. The protocol, which was designed by Aeronautical Radio, Incorporated (ARINC) to replace their very high frequency (VHF) voice service and deployed in 1978, uses telex formats. SITA later augmented their worldwide ground data network by adding radio stations to provide ACARS service. Over the next 20 years, ACARS will be superseded by the Aeronautical Telecommunications Network (ATN) protocol for Air Traffic Control communications and by the Internet Protocol for airline communications. Prior to the introduction of datalink, all communication between the aircraft (i.e., the flight crew) and personnel on the ground was performed using voice communication. This communication used either VHF or HF voice radios, which was further augmented with SATCOM in the early 1990s. In many cases, the voice-relayed information involved dedicated radio operators and digital messages sent to an airline teletype system or its successor systems. .On the aircraft, the ACARS system was made up of an avionics computer called an ACARS Management Unit (MU) and a Control Display Unit (CDU). The MU was designed to send and receive digital messages from the ground using existing VHF radios. On the ground, the ACARS system was made up of a network of radio transceivers managed by a central site computer called AFEPS (Arinc Front End Processor System), which received (or transmitted) the datalink messages as well as routed them to various airlines on the network. 9.1.1. OOOI events One of the initial applications for ACARS was to automatically detect and report changes to the major flight phases (Out of the gate, Off the ground, On the ground, and Into the gate), referred to in the industry as OOOI. These OOOI events are determined by algorithms that use aircraft sensors (such as doors, parking brake and strut switch sensors) as inputs. At the start of each flight phase, a digital message is transmitted to the ground containing the flight phase, the time at which it occurred, and other related information such as amount of fuel on board or flight origin and destination. These messages are used to track the status of aircraft and crews. 9.1.2. Flight management system interface In addition to detecting events on the aircraft and sending messages automatically to the ground, initial systems were expanded to support new interfaces with other onboard avionics. During the late 1980s and early 1990s, a datalink interface between the ACARS MUs and Flight Management Systems (FMS) was introduced. This interface enabled flight plans and weather information to be sent from the ground to the ACARS MU, which would then be forwarded to the FMS. This feature gave the airline the capability to update FMSs while in flight, and allowed the flight crew to evaluate new weather conditions, or alternative flight plans.
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9.1.3. Maintenance data download The introduction of the interface in the early 1990s between the FDAMS / ACMS systems and the ACARS MU, resulted in datalink's gaining wider acceptance by airlines. The FDAMS / ACMS systems which analyze engine, aircraft , and operational performance conditions were now able to provide performance data to the airlines on the ground in real time using the ACARS network. This reduced the need for airline personnel to go to the aircraft to off-load the data from these systems. These systems were capable of identifying abnormal flight conditions and automatically sending real-time messages to an airline. Detailed engine reports could also be transmitted to the ground via ACARS. The airlines used these reports to automate engine trending activities. This capability enabled airlines to monitor their engine performance more accurately and identify and plan repair and maintenance their activities more rapidly. In addition to the FMS and FDAMS interfaces, the industry started to upgrade the onboard maintenance computers in the 1990s to support the transmission of maintenance-related information in real-time through ACARS. This enabled airline maintenance personnel to receive real-time data associated with maintenance faults on the aircraft. When coupled with the FDAMS data, airline maintenance personnel could now start planning repair and maintenance activities while the aircraft was still in flight. 9.1.4. Interactive crew interface All of the processing described above is performed automatically by the ACARS MU and other associated avionics systems, without flight crew intervention. As part of the growth of ACARS functionality, the ACARS MUs also interfaced directly with a control display unit (CDU), located in the cockpit. This CDU, often referred to as an MCDU or MIDU, provides the flight crew with the ability to send and receive messages similar to today’s email. To facilitate this communication, the airlines in partnership with their ACARS vendor defined MCDU screens that could be presented to the flight crew and enable them to perform specific functions. This feature provides the flight crew flexibility as to the types of information requested from the ground and the types of reports sent to the ground. As an example, the flight crew could pull up an MCDU screen that allowed them to send to the ground a request for various types of weather information. After the desired locations and type of weather information were entered, ACARS transmitted this information to the ground. In response to this request message, ground computers sent the requested weather information back to the ACARS MU, which was then displayed and/or printed. Airlines began adding new messages to support new applications (weather, winds, clearances, connecting flights, etc.) and ACARS systems were customized to support airline-unique applications, and unique ground computer requirements. This resulted in each airline having its own unique ACARS application operating on its aircraft. Some airlines have more than 75 MCDU screens for their flight crews, where other may have only a dozen different screens. In addition, since each airline’s ground computers were different, the contents and formats of the messages sent by an ACARS MU were different for each airline.
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A person or a system on board may create a message and send it via ACARS to a system or user on the ground, and vice versa. Messages may be sent either automatically or manually.
9.2. VHF sub network A network of VHF ground radio stations ensure that aircraft can communicate with ground end systems in real-time from practically anywhere in the world. VHF communication is line-of-sight, and provides communication with ground-based transceivers (often referred to as Remote Ground Stations or RGSs). The typical range is dependent on altitude, with a 200-mile transmission range common at high altitudes. Thus VHF communication is only applicable over landmasses which have a VHF ground network installed. 9.3. SATCOM and HF sub networks SATCOM can provide worldwide coverage. Depending on the satellite system in use, coverage may be limited or absent at high latitudes (such as needed for flights over the poles). HF datalink is a relatively new network whose installation began in 1995 and was completed in 2001. Aircraft with HF or global SATCOM datalink can fly in polar routes and maintain communication with ground based systems (ATC centers and airline operation centers). ARINC is the only service provider for HF datalink. 9.4. Datalink message types ACARS messages may be of three types: 1. Air Traffic Control (ATC) 2. Aeronautical Operational Control (AOC) 3. Airline Administrative Control (AAC) ATC messages are used to communicate between the aircraft and Air Traffic Control. These messages are defined in ARINC Standard 623. ATC messages are used by aircraft crew to request clearances, and by ground controllers to provide those clearances. AOC and AAC messages are used to communicate between the aircraft and its base. These messages are either standardized according to ARINC Standard 633 or defined by the users, but in the latter case they must meet at least the guidelines of ARINC Standard 618. Various types of messages are possible, for example, relating to fuel consumption, engine performance data, aircraft position, in addition to free text . 9.5. ACARS Component There are three major components to the ACARS datalink system: 1. Aircraft equipment 2. Service provider 3. Ground processing system
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9.5.1. Aircraft equipment The heart of the datalink system on board the aircraft is the ACARS Management Unit (MU). The older version of MU is defined in ARINC Characteristic 724B. Newer versions are referred to as the Communications Management Unit (CMU) and are defined in ARINC Characteristic 758. Aircraft equipment consists of airborne end systems and a router. End systems are the source of ACARS downlinks and the destination for uplinks. The MU/CMU is the router. Its function is to route a downlink by means of the most efficient air-ground sub network. In many cases, the MU/CMU also acts as an end system for AOC messages. Typical airborne end systems are the Flight Management System (FMS), datalink printer, maintenance computer, and cabin terminal. Typical datalink functions are: 1. FMS - sends flight plan change requests, position reports, etc. Receives clearances and controller instructions. 2. Printer - as an end system, can be addressed from the ground to automatically print an uplink message. 3. Maintenance computer - downlinks diagnostic messages. In advanced systems, inflight troubleshooting can be conducted by technicians on the ground by using datalink messages to command diagnostic routines in the maintenance computer and analyzing down linked results. 4. Cabin terminal - Often used by flight attendant to communicates special passenger needs, gate changes due to delays, catering, etc.
ACARS messages are transmitted over one of three air-ground sub networks. VHF is the most commonly used and least expensive. Transmission is line-of-sight so VHF is not available over the oceans or other vast expansions of uninhabited surface, such as the Amazon Basin. SATCOM is a fairly expensive service that provides (near) worldwide coverage. The Inmarsat satellite network does not cover the polar-regions. However Iridium became usable for ACARS transport in 2007 and provides excellent coverage in the polarregions. HF is a more recently established sub-network. Its purpose is to provide coverage in the polar-regions where Inmarsat coverage is unreliable. The router function built into the MU/CMU determines which sub network to use when routing a message from the aircraft to the ground. The airline operator provides a routing table that the CMU uses to select the best sub network.
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9.5.2. Datalink service provider The role of the datalink service provider (DSP) is to deliver a message from the aircraft to the ground end system, and vice versa. Because the ACARS network is modeled after the point-to-point telex network, all messages come to a central processing location. The DSP routes the message to the appropriate end system using its network of land lines and ground stations. Before the days of computers, messages came to the central processing location and were punched on paper tape. The tape was then physically carried to the machine connected to the intended destination. Today, the routing function is done by computer, but the model remains the same. There are currently two primary service providers of ground networks in the world (ARINC and SITA), although specific countries have implemented their own network, with the help of either ARINC or SITA. ARINC operates a worldwide network and has also assisted the Civil Aviation Administration of China (CAAC), as well as Thailand and South America with the installation of VHF networks. SITA has operated the network in Europe, Middle East, South America and Asia for many years. They have also recently started a network in the USA to compete with ARINC. Until recently, each area of the world was supported by a single service provider. This is changing, and both ARINC and SITA are competing and installing networks that cover the same regions. 9.5.3. Ground end system The ground end system is the destination for downlinks and the source of uplinks. Generally, ground end systems are either government agencies such as CAA/FAA, an airline operations headquarters, or, in the case of small airlines or general aviation consumers, a subscription based solution. CAA end systems provide air traffic services such as clearances. Airline and general aviation operations provide information necessary for operating the airline or flight department efficiently, such as gate assignments, maintenance, and passenger needs. In the early history of ACARS most airlines created their own host systems for managing their ACARS messages. Commercial off-the-shelf products are now widely available to manage the ground hosting.
9.5.4. Acronyms and glossary There has been rumor that the introduction of datalink into the airline industry originated as part of a contest to see how many acronyms could be developed around a specific technology. Whether this is true or not, the industry is at the point where acronyms are now nested within acronyms. For example, AOA is an acronym for ACARS Over AVLC, where AVLC itself is an acronym for Aviation VHF Link Control and VHF is also an acronym for Very High Frequency.
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ACARS
Aircraft Communications Addressing and Reporting System
ACMS
Aircraft Condition Monitoring System
AMS
ACARS Message Security, as specified in ARINC 823
AOA
ACARS Over AVLC. With the introduction of VDL Mode 2, the ACARS protocols were modified to take advantage of the higher data rate made possible by Mode 2. AOA is an interim step in replacing the ACARS protocols with ATN protocols.
ATN
Aeronautical Telecommunications Network. As air traffic increases, ACARS will no longer have the capacity or flexibility to handle the large number of datalink communications. ATN is planned to replace ACARS in the future and will provide services such as authentication, security, and a true internetworking architecture. Europe is leading the US in the implementation of ATN
AVLC
Aviation VHF Link Control. A particular protocol used for aeronautical datalink communications
CDU
Control Display Unit
CMF
Communications Management Function. The software that runs in a CMU, and sometimes as a software partition in an integrated avionics computer
CMU
Communications Management Unit. Successor to the MU, the CMU performs similar datalink routing functions, but has additional capacity to support more functions. CMU standards are defined in ARINC Characteristic 758
FDAMS
Flight Data Acquisition and Management System
FMS
Flight Management System. FMS standards are defined in ARINC Characteristic 702 and 702A.
HFDL
High Frequency Data Link is an ACARS communications media used to exchange data such as Airline Operational Control (AOC) messages, Controller Pilot Data Link Communication (CPDLC) messages and Automatic Dependent Surveillance (ADS) messages between aircraft endsystems and corresponding ground-based HFDL ground stations
HF
High Frequency. A portion of the RF spectrum
LRU
Line Replaceable Unit. An avionics "black box" that can be replaced on the flight line, without downing the aircraft for maintenance
MCDU
Multifunction Control Display Unit. A text-only device that displays messages to the aircrew and accepts crew input on an integrated keyboard. MCDU standards are defined in ARINC Characteristic 739. MCDUs have seven input ports and can be used with seven different systems, such as Page 51 of 109
CMU or FMS. Each system connected to an MCDU generates its own display pages and accepts keyboard input, when it is selected as the system controlling the MCDU. MIDU
Multi-Input Interactive Display Unit (often used as a third cockpit CDU).
MU
Management Unit. Often referred to as the ACARS MU, this is an avionics LRU that routes datalink messages to and from the ground
OOOI
Shorthand for the basic flight phases—Out of the gate, Off the ground, On the ground, In the gate
POA
Plain Old ACARS. Refers to the set of ACARS communications protocols in effect before the introduction of VDL Mode 2. The term is derived from POTS (Plain old telephone service) that refers to the wired analog telephone network
SATCOM Satellite Communications. Airborne SATCOM equipment includes a satellite data unit, medium to high power amplifier, and an antenna, possibly with a steerable beam. A typical SATCOM installation can support a datalink channel as well as one or more voice channels VDL
VHF Data Link
VHF
Very High Frequency. A portion of the RF spectrum, defined as 30 MHz to 300 MHz
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ATA CHAPTER 31 : INSTRUMENT SYSTEM 1. AIRCRAFT INDICATING SYSTEM
1.1. General An instrument is a device that measures or manipulates a physical quantity such as flow, temperature, level, distance, angle, or pressure. Instruments may be as simple as direct reading thermometers or may be complex multi-variable process analyzers. Instruments are often part of a control system in refineries, factories, and vehicles. From the operating point of view, we may regard an instrument as being made up of the following four principal elements : A. Detecting Element, which detects changes in value of the physical quantity or condition presented to it B. Measuring Element, which actually measures the value of the physical quantity or condition in terms of small translational or angular displacements C. Coupling Element, by which displacements are magnified and transmitted D. Indicating Element, which exhibits the value of the measures quantity transmitted by the coupling element, by relative positions of a pointer / index , and a scale.
Figure 1. Elements of Instrument
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1.1.1 Instrument Error Hysteresis Error
Hysteresis Error is the Instrument Indication which occurs after movement of the indicator because of a rise or fall in the force being measured by the instrument. The error occurs primarily because of backlash in gearing within the instrument, and / or viscosity of fluid in the capillary of a direct reading type of Indicator. To overcome the majority of error, Hairspring are used to help remove the backlash from the gearing and the use of a higher viscosity fluid assists in increasing the sensitivity of the instrument. A further development which has helped to reduce the Hysteresis effect of an instrument is to the instrument indicator. This overcomes to requirements for capillary lines to the director reading instrument, although there is still a Hysteresis Error between the transmitter and indicator. As the Hysteresis Error is reduced the instrument becomes more sensitive to changes in the measured force, consequently a compromise must be reached between the Hysteresis Error, and sensitivity, as an oversensitive instrument will fluctuate and overshoot the changes in the measured force causing difficulty in reading the instrument indication. Finally, the procedures during a calibration test will normally specify a direction ( either rising or falling as stated) for the indicator to more during the calibration. This will reduce the Hysteresis effect to a minimum, as the pointer will always be either moving upscale or downscale as required. Parallax Error The error that result from viewing the reference point and scale from any position which is not perpendicular in line on sight ( 900 direction of sight ).To minimized the error, it is important that our line of sight be perpendicular to the scale and reference point be as small as possible.
Figure 2. Instruments Error. Page 54 of 109
1.2. Variable Resistance System 1.2.1. Basic Theory of Analogue Circuit.
The potentiometer can be used as a voltage divider to obtain a manually adjustable output voltage at the slider (wiper) from a fixed input voltage applied across the two ends of the potentiometer. This is their most common use. The voltage across
can be calculated by:
1.2. 2. Wheatstone bridge
Figure 3. Wheatstone bridge circuit diagram.
A Wheatstone bridge is an electrical circuit used to measure an unknown electrical resistance by balancing two legs of a bridge circuit, one leg of which includes the unknown component. Its operation is similar to the original potentiometer. It was invented by Samuel Hunter Christie in 1833 and improved and popularized by Sir Charles Wheatstone in 1843. One of the Wheatstone bridge's initial uses was for the purpose of soils analysis and comparison.
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Operation In the figure, is the unknown resistance to be measured; , and are resistors of known resistance and the resistance of is adjustable. If the ratio of the two resistances in the known leg is equal to the ratio of the two in the unknown leg , then the voltage between the two midpoints (B and D) will be zero and no current will flow through the galvanometer . If the bridge is unbalanced, the direction of the current indicates whether is too high or too low. is varied until there is no current through the galvanometer, which then reads zero. Detecting zero current with a galvanometer can be done to extremely high accuracy. Therefore, if , and are known to high precision, then can be measured to high precision. Very small changes in disrupt the balance and are readily detected. At the point of balance, the ratio of
Alternatively, if , , and are known, but is not adjustable, the voltage difference across or current flow through the meter can be used to calculate the value of , using Kirchhoff's circuit laws (also known as Kirchhoff's rules). This setup is frequently used in strain gauge and resistance thermometer measurements, as it is usually faster to read a voltage level off a meter than to adjust a resistance to zero the voltage. Derivation First, Kirchhoff's first rule is used to find the currents in junctions B and D:
Then, Kirchhoff's second rule is used for finding the voltage in the loops ABD and BCD:
When the bridge is balanced, then IG = 0, so the second set of equations can be rewritten as:
Then, the equations are divided and rearranged, giving: Page 56 of 109
From the first rule, I3 = Ix and I1 = I2. The desired value of Rx is now known to be given as:
If all four resistor values and the supply voltage (VS) are known, and the resistance of the galvanometer is high enough that IG is negligible, the voltage across the bridge (VG) can be found by working out the voltage from each potential divider and subtracting one from the other. The equation for this is:
where VG is the voltage of node B relative to node D. Significance The Wheatstone bridge illustrates the concept of a difference measurement, which can be extremely accurate. Variations on the Wheatstone bridge can be used to measure capacitance, inductance, impedance and other quantities, such as the amount of combustible gases in a sample, with an explosimeter. The Kelvin bridge was specially adapted from the Wheatstone bridge for measuring very low resistances. In many cases, the significance of measuring the unknown resistance is related to measuring the impact of some physical phenomenon (such as force, temperature, pressure, etc.) which thereby allows the use of Wheatstone bridge in measuring those elements indirectly. The concept was extended to alternating current measurements by James Clerk Maxwell in 1865 and further improved by Alan Blumlein in about 1926. Here some of the popular variable resistance transducers that are being used for various applications: a. Sliding contact devices: In the sliding contact type of variable resistance transducers there is a long conductor whose effective length is variable.
Figure 4. Linear Potentiometer Page 57 of 109
b. Wire resistance strain gauge: This is an interesting devise used for the measurement of force, stress and strain. When the tension is applied to the electrical conductor, its length increases while the cross section area decreases, due to which the resistance of the conductor changes. This change in resistance can be measured easily and is calibrated against the input quantity.
Figure 5. Wire resistance stain gauge
c. Thermistors: Thermistors work on the principle that resistance of some material change with the change in their temperature. The commonly used thermistors are made up of the ceramic like semiconducting materials such as oxides of manganese, nickel and cobalt.
Figure 6. Thermistor
d. Resistance thermometers : also called Resistance Temperature Detectors (RTDs), are sensors used to measure temperature by correlating the resistance of the RTD element with temperature.
Figure 7. Resistance Thermometer Page 58 of 109
1.3. Synchros
A synchro is a type of rotary electrical transformer that is used for measuring the angle of a rotating machine such as an antenna platform. In its general physical construction, it is much like an electric motor. The primary winding of the transformer, fixed to the rotor, is excited by an alternating current, which by electromagnetic induction, causes currents to flow in three star-connected secondary windings fixed at 120 degrees to each other on the stator. The relative magnitudes of secondary currents are measured and used to determine the angle of the rotor relative to the stator, or the currents can be used to directly drive a receiver synchro that will rotate in unison with the synchro transmitter. In the latter case, the whole device may be called a selsyn (self synchronizing).
Figure 8. Synchro motor Synchro systems were first used in the control system of the Panama Canal in the early 1900s to transmit lock gate and valve steam positions, and water levels, to the control desks. Smaller synchros are still used to remotely drive indicator gauges and as rotary position sensors for aircraft control surfaces, where the reliability of these rugged devices is needed. Digital devices such as the rotary encoder have replaced synchros in most other applications.
Selsyn motors were widely used in motion picture equipment to synchronize movie cameras and sound recording equipment, before the advent of crystal oscillators and microelectronics. Large synchros were used on naval warships, such as destroyers, to operate the steering gear from the wheel on the bridge. 1.3.1. Types of Synchro Systems There are two types of synchro systems: Torque systems and control systems. In a torque system, a synchro will provide a low-power mechanical output sufficient to position an indicating device, actuate a sensitive switch or move light loads without power amplification. In simpler terms, a torque synchro system is a system in which the transmitted signal does the usable work. In such a system, accuracy on the order of one degree is attainable. In a control system, a synchro will provide a voltage for Page 59 of 109
conversion to torque through an amplifier and a servomotor. Control type synchros are used in applications that require large torques or high accuracy such as follow-up links and error detectors in servo, automatic control systems (such as an autopilot system). In simpler terms, a control synchro system is a system in which the transmitted signal controls a source of power that which does work. Quite often, one system will perform both torque and control functions. Individual units are designed for use in either torque or control systems. Some torque units can be used as control units, but control units cannot replace torque units 1.3.2. Operation
Figure 9. On a practical level, synchros resemble motors, in that there is a rotor, stator, and a shaft. Ordinarily, slip rings and brushes connect the rotor to external power. A synchro transmitter's shaft is rotated by the mechanism that sends information, while the synchro receiver's shaft rotates a dial, or operates a light mechanical load. Single and three-phase units are common in use, and will follow the other's rotation when connected properly. One transmitter can turn several receivers; if torque is a factor, the transmitter must be physically larger to source the additional current. In a motion picture interlock system, a large motor-driven distributor can drive as many as 20 machines, sound dubbed, footage counters, and projectors. Synchros designed for terrestrial use tend to be driven at 50 or 60 hertz (the mains frequency in most countries), while those for marine or aeronautical use tend to operate at 400 hertz (the frequency of the on-board electrical generator driven by the engines). Single phase units have five wires: two for an exciter winding (typically line voltage) and three for the output/input. These three are busses to the other synchros in the system, and provide the power and information to align the shafts of all the receivers. Synchro transmitters and receivers must be powered by the same branch circuit, so to speak; the mains excitation voltage sources must match in voltage and phase. The safest approach is to bus the five or six lines from transmitters and receivers at a common point. Different makes of selsyns, used in interlock systems, have different output voltages. In all cases, three-phase systems will handle more power and operate a bit more smoothly. The excitation is often 208/240 V 3-phase mains power.
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Synchro transmitters are as described, but 50 and 60-Hz synchro receivers require rotary dampers to keep their shafts from oscillating when not loaded (as with dials) or lightly loaded in high-accuracy applications. A different type of receiver, called a control transformer (CT), is part of a position servo that includes a servo amplifier and servo motor. The motor is geared to the CT rotor, and when the transmitter's rotor moves, the servo motor turns the CT's rotor and the mechanical load to match the new position. CTs have high-impedance stators and draw much less current than ordinary synchro receivers when not correctly positioned. 1.4. Resolver A resolver is a type of rotary electrical transformer used for measuring degrees of rotation. It is considered an analog device, and has a digital counterpart, the rotary (or pulse) encoder.
Figure 10. Resolver
The most common type of resolver is the brushless transmitter resolver (other types are described at the end). On the outside, this type of resolver may look like a small electrical motor having a stator and rotor. On the inside, the configuration of the wire windings makes it different. The stator portion of the resolver houses three windings: an exciter winding and two two-phase windings (usually labeled "x" and "y") (case of a brushless resolver). The exciter winding is located on the top; it is in fact a coil of a turning (rotary) transformer. This transformer induces current in the rotor without a direct electrical connection, thus there are no wires to the rotor limiting its rotation and no need for brushes. The two other windings are on the bottom, wound on a lamination. They are configured at 90 degrees from each other. The rotor houses a coil, which is the secondary winding of the turning transformer, and a separate primary winding in a lamination, exciting the two twophase windings on the stator. Page 61 of 109
The primary winding of the transformer, fixed to the stator, is excited by a sinusoidal electric current, which by electromagnetic induction induces current in the rotor. As these windings are arranged on the axis of the resolver, the same current is induced no matter what its position. This current then flows through the other winding on the rotor, in turn inducing current in its secondary windings, the two-phase windings back on the stator. The two two-phase windings, fixed at right (90°) angles to each other on the stator, produce a sine and cosine feedback current. The relative magnitudes of the two-phase voltages are measured and used to determine the angle of the rotor relative to the stator. Upon one full revolution, the feedback signals repeat their waveforms. This device may also appear in non-brushless type, i.e., only consisting in two lamination stacks, rotor and stator. Resolvers can perform very accurate analog conversion from polar to rectangular coordinates. Shaft angle is the polar angle, and excitation voltage is the magnitude. The outputs are the [x] and [y] components. Resolvers with four-lead rotors can rotate [x] and [y] coordinates, with the shaft position giving the desired rotation angle. Resolvers with four output leads are general sine/cosine computational devices. When used with electronic driver amplifiers and feedback windings tightly coupled to the input windings, their accuracy is enhanced, and they can be cascaded ("resolver chains") to compute functions with several terms, perhaps of several angles, such as gun (position) orders corrected for ship's roll and pitch. For the position evaluation, Resolver-to-Digital Converters are commonly used. They convert sine and cosine signal to binary signal (10 to 16 bit wide) that can more easily be used by the controller. Types Basic resolvers are two-pole resolvers, meaning that the angular information is the mechanical angle of the stator. These devices can deliver the absolute angle position. Other types of resolver are multi-pole resolvers. They have 2*p poles, and thus can deliver p cycles in one rotation of the rotor: electrical angle = mechanical angle * p. where p is the no. of pole pairs. Some types of resolvers include both types, with the 2-pole windings used for absolute position and the multi-pole windings for accurate position. Two-pole resolvers can usually reach angular accuracy up to about +/-5′, whereas multi-pole resolver can provide better accuracy, up to 10′′ for 16-pole resolvers, to even 1′′, for instance for 128pole resolvers. Multi-pole resolvers may also be used for monitoring multi-pole electrical motors. This device can be used in any application in which the exact rotation of an object relative to another object is needed, such as in a rotary antenna platform or a robot. In practice, the resolver is usually directly mounted to an electric motor. The resolver feedback signals are usually monitored for multiple revolutions by another device. This allows for geared reduction of assemblies being rotated and improved accuracy from the resolver system. Because the power supplied to the resolvers produces no actual work, the voltages used are usually low (<24 VAC) for all resolvers. Resolvers designed for terrestrial use tend to be driven at 50-60 Hz (mains power frequency), while those for marine or aeronautical used tend to operate at 400 Hz (the frequency of the on-board generator driven by the engines). Control systems tend to use higher frequencies (5 kHz).
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Other types of resolver include: Receiver resolvers These resolvers are used in the opposite way to transmitter resolvers (the type described above). The two diphased winding are energized, the ratio between the sine and the cosine representing the electrical angle. The system turns the rotor to obtain a zero voltage in the rotor winding. At this position, the mechanical angle of the rotor equals the electrical angle applied to the stator. Differential resolvers These types combine two diphased primary windings in one of the stacks of sheets, as with the receiver, and two diphased secondary windings in the other. The relation of the electrical angle delivered by the two secondary windings and the other angles is secondary electrical angle, mechanical angle, and primary electrical angle. These types were used, for instance, as analog trigonometric-function calculators. A related type is also the transolver, combining a two-phase winding like the resolver and a triphased winding like the synchro
1.4. Servo System A servomechanism, sometimes shortened to servo, is an automatic device that uses errorsensing negative feedback to correct the performance of a mechanism. The term correctly applies only to systems where the feedback or error-correction signals help control mechanical position, speed or other parameters.
Figure 11. Basic Servo Mechanism
1.4.1. Position control A common type of servo provides position control. Servos are commonly electrical or partially electronic in nature, using an electric motor as the primary means of creating mechanical force. Other types of servos use hydraulics, pneumatics, or magnetic principles. Servos operate on the principle of negative feedback, where the control input is compared to Page 63 of 109
the actual position of the mechanical system as measured by some sort of transducer at the output. Any difference between the actual and wanted values (an "error signal") is amplified (and converted) and used to drive the system in the direction necessary to reduce or eliminate the error. This procedure is one widely used application of control theory. 1.4.2. Speed control Speed control via a governor is another type of servomechanism. The steam engine uses mechanical governors; another early application was to govern the speed of water wheels. Prior to World War II the constant speed propeller was developed to control engine speed for maneuvering aircraft. Fuel controls for gas turbine engines employ either Hydromechanical or electronic governing. Positioning servomechanisms were first used in military fire-control and marine navigation equipment. Today servomechanisms are used in automatic machine tools, satellite-tracking antennas, remote control airplanes, automatic navigation systems on boats and planes, and antiaircraft-gun control systems. Other examples are fly-by-wire systems in aircraft which use servos to actuate the aircraft's control surfaces, and radio-controlled models which use RC servos for the same purpose. Many autofocus cameras also use a servomechanism to accurately move the lens, and thus adjust the focus. A modern hard disk drive has a magnetic servo system with sub-micrometre positioning accuracy. In industrial machines, servos are used to perform complex motion, in many applications. 1.4.3. Servomotor
Figure 12. Servomotor A servomotor is a specific type of motor and rotary encoder combination, usually with a dedicated, that forms a servomechanism. This assembly may in turn form part of another servomechanism. The encoder provides position and usually speed feedback, which by the use of a PID controller allow more precise control of position and thus faster achievement of a stable position (for a given motor power). Stepper motors are not considered as servomotors, although they too are used to construct larger servomechanisms. Stepper motors have inherent angular positioning, owing to their construction, and this is generally used in an open-loop manner, without an encoder. Servomotors are used for both high-end and low-end applications, although the mid-range is generally handled by stepper motors. Most servomotors, at least under this name, are precision industrial components. However the very cheap radio control servo, because it combines a free-running motor and a simple position sensor with an embedded controller, also qualifies as a servomotor. Page 64 of 109
1.5. Direct Torque 1.5.1. General Definition Torque, is a force to rotate an object about an axis, fulcrum, or pivot. Just as a force is a push or a pull, a torque can be thought of as a twist to an object. Mathematically, torque is defined as the cross product of the lever-arm distance and force, which tends to produce rotation. Common speaking, torque is a measure of the turning force on an object such as a bolt or a flywheel. For example, pushing or pulling the handle of a wrench connected to a nut or bolt produces a torque (turning force) that loosens or tightens the nut or bolt.
The symbol for torque is typically τ, the Greek letter tau. The magnitude of torque depends on three quantities: the force applied, the length of the lever arm connecting the axis to the point of force application, and the angle between the force vector and the lever arm. In symbols:
Where : τ is the torque vector and τ is the magnitude of the torque, r is the displacement vector (a vector from the point from which torque is measured to the point where force is applied), and r is the length (or magnitude) of the lever arm vector, F is the force vector, and F is the magnitude of the force, θ is the angle between the force vector and the lever arm vector. The length of the lever arm is particularly important; choosing this length appropriately lies behind the operation of levers, pulleys, gears, and most other simple machines involving a mechanical advantage. The definition of torque states that one or both of the angular velocity or the moment of inertia of an object are changing. And moment is the general term used for the tendency of one or more applied forces to rotate an object about an axis, but not necessarily to change the angular momentum of the object (the concept which in physics is called torque). . Page 65 of 109
1.5.2. Units Official SI literature suggests using the unit newton metre (N·m) or the unit joule per radian. The unit newton metre is properly denoted N·m or N m.. In Imperial units, "pound-force-feet" (lb·ft), "foot-pounds-force", "inch-pounds-force", "ounce-force-inches" (oz·in) are used, and other non-SI units of torque includes "metrekilograms-force". For all these units, the word "force" is often left out, for example abbreviating "pound-force-foot" to simply "pound-foot" (in this case, it would be implicit that the "pound" is pound-force and not pound-mass). This is an example of the confusion caused by the use of traditional units that may be avoided with SI units because of the careful distinction in SI between force (in newtons) and mass (in kilograms). 1.5.3. Moment arm A very useful special case, often given as the definition of torque in fields other than physics, is as follows:
The construction of the "moment arm" is shown in the figure to the right, along with the vectors r and F mentioned above. The problem with this definition is that it does not give the direction of the torque but only the magnitude, and hence it is difficult to use in threedimensional cases. If the force is perpendicular to the displacement vector r, the moment arm will be equal to the distance to the centre, and torque will be a maximum for the given force. The equation for the magnitude of a torque, arising from a perpendicular force:
For example, if a person places a force of 10 N at the terminal end of a spanner (wrench) which is 0.5 m long (or a force of 10 N exactly 0.5 m from the twist point of a spanner of any length), the torque will be 5 N-m – assuming that the person moves the spanner by applying force in the plane of movement of and perpendicular to the spanner. 1.5.4. Torque Measurement A torque sensor or torque transducer or torque meter is a device for measuring and recording the torque on a rotating system, such as an engine, crankshaft, gearbox, transmission, rotor, a bicycle crank or Cap Torque Tester. Static torque is relatively easy to measure.
Figure 13. Torque wrench Page 66 of 109
Dynamic torque, on the other hand, is not easy to measure, since it generally requires transfer of some effect (electric or magnetic) from the shaft being measured to a static system. Commonly, torque sensors or torque transducers use strain gauges applied to a rotating shaft or axle. With this method, a means to power the strain gauge bridge is necessary, as well as a means to receive the signal from the rotating shaft
Figure 14. Strain gauge One way to achieve this is to condition the shaft or a member attached to the shaft with a series of permanent magnetic domains. The magnetic characteristics of these domains will vary according to the applied torque, and thus can be measured using non-contact sensors. Such magnetoelastic torque sensors are generally used for in-vehicle applications on racecars, automobiles, aircraft, and hovercraft.
Figure 15. Finally, another way to measure torque is by way of twist angle measurement or phase shift measurement, whereby the angle of twist resulting from applied torque is measured by using two angular position sensors and measuring the phase angle between them. 1.6. Linear Variable Differential Transducer 1.6.1. General Definition The linear variable differential transformer (LVDT) also called just a differential transformer is a type of electrical transformer used for measuring linear displacement. LVDTs have been widely used in applications such as power turbines, hydraulics, automation, aircraft, satellites, nuclear reactors, and many others. The LVDT converts a position or linear displacement from a mechanical reference (zero, or null position) into a proportional electrical signal containing phase (for direction) and amplitude (for distance) information. The LVDT operation does not Page 67 of 109
require an electrical contact between the moving part (probe or core assembly) and the coil assembly, but instead relies on electromagnetic coupling.
Figure 16. LVDT
1.6.2. Operation The linear variable differential transformer has three solenoidal coils placed end-to-end around a tube. The center coil is the primary, and the two outer coils are the top and bottom secondary . A cylindrical ferromagnetic core, attached to the object whose position is to be measured, slides along the axis of the tube. An alternating current drives the primary and causes a voltage to be induced in each secondary proportional to the length of the core linking to the secondary. The frequency is usually in the range 1 to 10 kHz. As the core moves, the primary's linkage to the two secondary coils changes and causes the induced voltages to change. The coils are connected so that the output voltage is the difference (hence "differential") between the top secondary voltage and the bottom secondary voltage. When the core is in its central position, equidistant between the two secondary, equal voltages are induced in the two secondary coils, but the two signals cancel, so the output voltage is theoretically zero. In practice minor variations in the way in which the primary is coupled to each secondary means that a small voltage is output when the core is central. When the core is displaced toward the top, the voltage in the top secondary coil increases as the voltage in the bottom decrease. The resulting output voltage increases from zero. This voltage is in phase with the primary voltage. When the core moves in the other direction, the output voltage also increases from zero, but its phase is opposite to that of the primary. The phase of the output voltage determines the direction of the displacement (up or down) and amplitude indicates the amount of displacement. A synchronous detector can determine a signed output voltage that relates to the displacement. The LVDT is designed with long slender coils to make the output voltage essentially linear over displacement up to several inches (several hundred milli-meter) long. The LVDT can be used as an absolute position sensor. Even if the power is switched off, on restarting it, the LVDT shows the same measurement, and no positional information is lost. Its biggest advantages are repeatability and reproducibility once it is Page 68 of 109
properly configured. Also, apart from the uni-axial linear motion of the core, any other movements such as the rotation of the core around the axis will not affect its measurements.
1.7. Rotary Variable Differential Transducer A rotary variable differential transformer (RVDT) is a type of electrical transformer used for measuring angular displacement.
.
Figure 17. RVDT
More precisely, a Rotary Variable Differential Transformer (RVDT) is an electromechanical transducer that provides a variable alternating current (AC) output voltage that is linearly proportional to the angular displacement of its input shaft. When energized with a fixed AC source, the output signal is linear within a specified range over the angular displacement. RVDT’s utilize brushless, non-contacting technology to ensure long-life and reliable, repeatable position sensing with infinite resolution. Such reliable and repeatable performance assures accurate position sensing under the most extreme operating conditions. Most RVDT are composed of a wound, laminated stator and a salient two-pole rotor. The stator, containing four slots, contains both the primary winding and the two secondary windings. Some secondary windings may also be connected together.
2. PRESSURE MEASURING INSTRUMENTS
2.1. Definition Pressure is the effect of a force applied to a surface. Pressure is the amount of force acting per unit area. The symbol of pressure is p. Formula Mathematically:
Page 69 of 109
where: is the pressure, is the normal force, is the area of the surface on contact.
For liquids, the formula may be written:
where: is the pressure, is the density of the liquid, (the value is equal to the gravitational acceleration), is the depth of the liquid in meters.
Pressure is a scalar quantity. It relates the vector surface element (a vector normal to the surface) with the normal force acting on it. The pressure is the scalar proportionality constant that relates the two normal vectors:
2.2. Pressure Measurement
Absolute pressure sensor is zero-referenced against a perfect vacuum. This instrument measures the pressure relative to zero. atmospheric pressures, deep vacuum pressures, and altimeter pressures must be absolute. Gauge pressure sensor is zero-referenced against ambient air pressure, so it is equal to absolute pressure minus atmospheric pressure. Tire pressure and blood pressure are gauge pressures by convention. For most working fluids where a fluid exists in a closed system, gauge pressure measurement prevails. Pressure instruments connected to the system will indicate pressures relative to the current atmospheric pressure. Differential pressure sensor is the difference in pressure between two points. Differential pressures are commonly used in industrial process systems. Differential pressure gauges have two inlet ports, each connected to one of the volumes whose pressure is to be monitored. In effect, such a gauge performs the mathematical operation of subtraction through mechanical means, obviating the need for an operator or control system to watch two separate gauges and determine the difference in readings. Strain Gauge A strain gauge is a device used to measure the strain of an object. Invented by Edward E. Simmons and Arthur C. Ruge in 1938, the most common type of strain gauge consists of an insulating flexible backing which supports a metallic foil pattern. The gauge is attached to the object by a suitable adhesive, such as cyanoacrylate. As the object is deformed, the foil is deformed, causing its electrical resistance to change. This resistance change, usually Page 70 of 109
measured using a Wheatstone bridge, is related to the strain by the quantity known as the gauge factor.
Figure 18. Strain gauge based technology is utilized commonly in the manufacture of pressure sensors. The gauges used in pressure sensors themselves are commonly made from silicon, polysilicon, metal film, thick film, and bonded foil. Piezo-resistive Sensor The piezo-resistive effect of semiconductors has been used for sensor devices employing all kinds of semiconductor materials such as germanium, polycrystalline silicon, amorphous silicon, and single crystal silicon. Since silicon is today the material of choice for integrated digital and analog circuits the use of piezo-resistive silicon devices has been of great interest. It enables the easy integration of stress sensors with Bipolar and CMOS circuits. This has enabled a wide range of products using the piezo-resistive effect. Many commercial devices such as pressure sensors and acceleration sensors employ the piezoresistive effect in silicon. But due to its magnitude the piezo-resistive effect in silicon has also attracted the attention of research and development for all other devices using single crystal silicon. Semiconductor Hall sensors, for example, were capable of achieving their current precision only after employing methods which eliminate signal contributions due the applied mechanical stress. Piezo-resistivity has a much greater effect on resistance than a simple change in geometry and so a semiconductor can be used to create a much more sensitive strain gauge, though they are generally also more sensitive to environmental conditions (esp. temperature). Piezoresistors can be fabricated using wide variety of piezo-resistive materials. The simplest form of piezo-resistive silicon sensors are diffused resistors. Piezo-resistors consist of a simple two contact diffused n- or p-wells within a p- or n-substrate. As the typical square resistances of these devices are in the range of several hundred ohms, additional p+ or n+ plus diffusions are necessary to facilitate ohmic contacts to the device.
Page 71 of 109
Figure 19. Schematic cross-section of the basic elements of a silicon n-well piezoresistor.
Piezoresistive devices In silicon the piezo-resistive effect is used in piezo-resistors, transducers, piezo-FETS, solid state accelerometers and bipolar transistors. Frequency response of pressure sensors When fluid flows are not in equilibrium, local pressures may be higher or lower than the average pressure in a medium. These disturbances propagate from their source as longitudinal pressure variations along the path of propagation. This is also called sound. Sound pressure is the instantaneous local pressure deviation from the average pressure caused by a sound wave. Sound pressure can be measured using a microphone in air and a hydrophone in water. The effective sound pressure is the root mean square of the instantaneous sound pressure over a given interval of time. Sound pressures are normally small and are often expressed in units of microbar. Resonant Use the changes in resonant frequency in a sensing mechanism to measure stress, or changes in gas density, caused by applied pressure. This technology may be used in conjunction with a force collector, such as those in the category above. Alternatively, resonant technology may be employed by exposing the resonating element itself to the media, whereby the resonant frequency is dependent upon the density of the media. Sensors have been made out of vibrating wire, vibrating cylinders, quartz, and silicon MEMS. Generally, this technology is considered to provide very stable readings over time.
3. TEMPERATURE MEASUREMENT 3.1. Non Electrical Temperature Measurement 3.1.1 Alcohol Thermometer
The alcohol thermometer was the earliest, efficient, modern-style instrument of temperature measurement. As is the case with many early, important inventions, there are several people credited with its invention. These include Ferdinando II de' Medici, Grand Duke of Tuscany, who in 1654 made sealed tubes part filled with alcohol or urine, with a bulb and stem, depending on the expansion of a liquid, and independent of air pressure. Other sources, including the Encyclopedia Britannica, credit German scientist Daniel Gabriel Fahrenheit with
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inventing the alcohol thermometer in 1709. Fahrenheit was a skilled glassblower and his alcohol thermometer was the world's first reliable thermometer. 3.1.2 Mercury Thermometer The mercury-in-glass or mercury thermometer was invented by physicist Daniel Gabriel Fahrenheit in Amsterdam (1714). It consists of a bulb containing mercury attached to a glass tube of narrow diameter; the volume of mercury in the tube is much less than the volume in the bulb. The volume of mercury changes slightly with temperature; the small change in volume drives the narrow mercury column a relatively long way up the tube. The space above the mercury may be filled with nitrogen or it may be at less than atmospheric pressure, a partial vacuum. 3.1.3 Bi-metal mechanical Thermometer
A bimetallic strip is used to convert a temperature change into mechanical displacement. The strip consists of two strips of different metals which expand at different rates as they are heated, usually steel and copper, or in some cases steel and brass. The strips are joined together throughout their length by riveting, brazing or welding. The different expansions force the flat strip to bend one way if heated, and in the opposite direction if cooled below its initial temperature. The metal with the higher coefficient of thermal expansion is on the outer side of the curve when the strip is heated and on the inner side when cooled. The sideways displacement of the strip is much larger than the small lengthways expansion in either of the two metals. This effect is used in a range of mechanical and electrical devices. In some applications the bimetal strip is used in the flat form. In others, it is wrapped into a coil for compactness. The greater length of the coiled version gives improved sensitivity. 3.1.4. Galileo Thermometer A Galileo thermometer (or Galilean thermometer) is a thermometer made of a sealed glass cylinder containing a clear liquid and several glass vessels of varying densities. As temperature changes, the individual floats rise or fall proportion to their respective density. It is named after Galileo Galilei because he discovered the principle on which this thermometer is based—that the density of a liquid changes in proportion to its temperature—and invented a thermoscope based on this principle. The Galilean thermometer works on the principle of buoyancy. Buoyancy determines whether objects float or sink in a liquid, and is responsible for the fact that even boats made of steel float in water (while a solid bar of steel sinks). The only factor that determines whether a large object rises or falls in a particular liquid is the object's density relative to the density of the liquid. If the object is denser than the liquid then it sinks, as it is heavier than the liquid it Page 73 of 109
displaces. If the object is less dense than the liquid then it begins to sink until the weight of the displaced liquid becomes equal to the object's weight; then it floats at that depth. 3.1.5. Phosphor Thermometry Phosphor thermometry is an optical method for surface temperature measurement. The method exploits luminescence emitted by phosphor material. Phosphors are fine white or pastel-colored inorganic powders which may be stimulated by any of a variety of means to luminesce, i.e. emit light. Certain characteristic of the emitted light changed with temperature, including brightness, color, and afterglow duration. The latter is most commonly used for temperature measurement. 3.1.6. Liquid crystal thermometer Liquid crystal thermometer or plastic strip thermometer is a type of thermometer that contains heat-sensitive (thermochromic) liquid crystals in a plastic strip that change color to indicate different temperatures. Liquid crystals possess the mechanical properties of a liquid, but have the optical properties of a single crystal. Temperature changes can affect the color of a liquid crystal, which makes them useful for temperature measurement. The resolution of liquid crystal sensors is in the 0.1°C range. Disposable liquid crystal thermometers have been developed for home and medical use. For example if the thermometer is black and it is put onto someone's forehead it will change colour depending on the temperature of the person. Liquid crystal thermometers portray temperatures as colors and can be used to follow temperature changes caused by heat flow. They can be used to observe that heat flows by conduction, convection, and radiation. 3.2. Temperature Dependant Resistor 3.2.1. Resistance thermometers Resistance temperature detectors ('RTD's), are sensors used to measure temperature by correlating the resistance of the RTD element with temperature. Most RTD elements consist of a length of fine coiled wire wrapped around a ceramic or glass core. The element is usually quite fragile, so it is often placed inside a sheathed probe to protect it. The RTD element is made from a pure material, typically platinum, nickel or copper. The material has a predictable change in resistance as the temperature changes; it is this predictable change that is used to determine temperature There are three main categories of RTD sensors; Thin Film, Wire-Wound, and Coiled Elements. Thin film elements have a sensing element that is formed by depositing a very thin layer of resistive material, normal platinum, on a ceramic substrate; This layer is usually just 10 to 100 angstroms (1 to 10 nanometers) thick. This film is then coated with an epoxy or glass that helps protect the deposited film and also acts as a strain relief for the external lead-wires. Disadvantages of this type are that they are not as stable as their wire wound or coiled counterparts. They also Page 74 of 109
can only be used over a limited temperature range due to the different expansion rates of the substrate and resistive deposited giving a "strain gauge" effect that can be seen in the resistive temperature coefficient. These elements work with temperatures to 300 °C.
Figure 20. thin film element
Wire-wound elements can have greater accuracy, especially for wide temperature ranges. The coil diameter provides a compromise between mechanical stability and allowing expansion of the wire to minimize strain and consequential drift. The sensing wire is wrapped around an insulating mandrel or core. The winding core can be round or flat, but must be an electrical insulator. The coefficient of thermal expansion of the winding core material is matched to the sensing wire to minimize any mechanical strain. This strain on the element wire will result in a thermal measurement error. The sensing wire is connected to a larger wire, usually referred to as the element lead or wire. This wire is selected to be compatible with the sensing wire so that the combination does not generate an emf that would distort the thermal measurement. These elements work with temperatures to 660 °C.
Figure 21. wire-wound element
Coiled elements have largely replaced wire-wound elements in industry. This design has a wire coil which can expand freely over temperature, held in place by some mechanical support which lets the coil keep its shape. This “strain free” design allows the sensing wire to expand and contract free of influence from other materials; in this respect it is similar to the SPRT, the primary standard upon which ITS-90 is based, while providing the durability necessary for industrial use. The basis of the sensing element is a small coil of platinum sensing wire. This coil resembles a filament in an incandescent light bulb. The housing or mandrel is a hard fired ceramic oxide tube with equally spaced bores that run transverse to the axes. The coil is inserted in the bores of the mandrel and then packed with a very finely ground ceramic powder. This permits the sensing wire to move while still remaining in good thermal contact with the process. These Elements works with temperatures to 850 °C.
Figure 22. coiled element Page 75 of 109
3.2.2. Thermistor A thermistor is a type of resistor whose resistance varies significantly with temperature, more so than in standard resistors. The word is a portmanteau of thermal and resistor. Thermistors are widely used as inrush current limiters, temperature sensors, self-resetting over current protectors, and self-regulating heating elements. Thermistors differ from resistance temperature detectors (RTD) in that the material used in a thermistor is generally a ceramic or polymer, while RTDs use pure metals. The temperature response is also different; RTDs are useful over larger temperature ranges, while thermistors typically achieve a higher precision within a limited temperature range, typically −90 °C to 130 °C
.
Figure 23.
3.2.3. Thermocouple A thermocouple consists of two dissimilar conductors in contact, which produce a voltage when heated. The size of the voltage is dependent on the difference of temperature of the junction to other parts of the circuit. Thermocouples are a widely used type of temperature sensor for measurement and control and can also be used to convert a temperature gradient into electricity. Commercial thermocouples are inexpensive, interchangeable, are supplied with standard connectors, and can measure a wide range of temperatures. In contrast to most other methods of temperature measurement, thermocouples are self powered and require no external form of excitation. The main limitation with thermocouples is accuracy; system errors of less than one degree Celsius (°C) can be difficult to achieve. Any junction of dissimilar metals will produce an electric potential related to temperature. Thermocouples for practical measurement of temperature are junctions of specific alloys which have a predictable and repeatable relationship between temperature and voltage. Different alloys are used for different temperature ranges. Properties such as resistance to corrosion may also be important when choosing a type of thermocouple. Where the measurement point is far from the measuring instrument, the intermediate connection can be made by extension wires which are less costly than the materials used to make the sensor. Thermocouples are usually standardized against a reference temperature of 0 degrees Celsius; practical instruments use electronic methods of coldjunction compensation to adjust for varying temperature at the instrument terminals. Electronic instruments can also compensate for the varying Page 76 of 109
characteristics of the thermocouple, and so improve the precision and accuracy of measurements. Thermocouples are widely used in science and industry; applications include temperature measurement for kilns, gas turbine exhaust, diesel engines, and other industrial processes.
Figure 24. 3.2.4. Quartz thermometer The quartz thermometer is a high-precision, high accuracy temperature sensor. It measures temperature by measuring the frequency of a quartz crystal oscillator. The oscillator contains a specially cut crystal that results in a linear temperature coefficient of frequency, so the measurement of the temperature is essentially reduced to measurement of the oscillator frequency. Resolutions of 0.0001 °C, and accuracy of 0.02 °C from 0-100 °C are achievable. The high linearity makes it possible to achieve high accuracy over an important temperature range that contains only one convenient temperature reference point for calibration, the triple point of water. Introduced by Hewlett-Packard in 1965, the successor company, Agilent, has discontinued the Model 2804A Quartz Thermometer. 3.2.5. Pyrometer A pyrometer is a non-contacting device that intercepts and measures thermal radiation, a process known as pyrometry . This device can be used to determine the temperature of an object's surface. The word pyrometer comes from the Greek word for fire, "πυρ" (pyro), and meter, meaning to measure. Pyrometer was originally coined to denote a device capable of measuring temperatures of objects above incandescence Page 77 of 109
A pyrometer has an optical system and a detector. The optical system focuses the thermal radiation onto the detector. The output signal of the detector (temperature T) is related to the thermal radiation or irradiance j* of the target object through the Stefan–Boltzmann law, the constant of proportionality σ, called the Stefan-Boltzmann constant and the emissivity ε of the object.
This output is used to infer the object's temperature. Thus, there is no need for direct contact between the pyrometer and the object, as there is with thermocouples and resistance temperature detectors (RTDs). Pyrometers are suited especially to the measurement of moving objects or any surfaces that cannot be reached or cannot be touched.
4. QUANTITY INDICATING SYSTEM 4.1. DC Electrical Indication
Describe the construction and principles of operation of a typical DC fuel contents indication system with particular regard to the following: a. Conversion of float movement to electrical current: In most cases, a direct connection between the float and the indicator is not possible. A DC electrical indicator solves this problem. It converts mechanical motion of the float into varying direct current. This current then drives a mechanical indicator or is converted to a digital readout. The components of a float-type together with the methods of transmitting electrical signal. The float is attached to an arm pivoted to permit angular movement which is transmitted to an electrical element consisting of either a wiper arm and potentiometer, or a Desyn.
Figure 25. float to DC electrical
b. Wiper arm operation: For many years the most widely used fuel quantity measuring system has been the electrical resistance-type system. These systems use a sender, or Page 78 of 109
transmitter, that consists of a variable resistor mounted on the outside of the fuel tank and operated by an arm connected to float that rides on the surface of the fuel tank. Movement of the arm is transmitted through a bellows-type seal to operate the wiper of the resistor. The float rides on the fuel and drives the wiper across the resistance element. This type of system uses current-measuring instruments which are calibrated in fuel quantity. When the tank is empty, the float is at the bottom of the tank and the resistance is at maximum. This drives the pointer on the gauge to the EMPTY mark. When the tank is full, the wiper resistance is at its minimum and the gauge reads FULL. c. Resistance material: Instead of using a resistor, some units use a segment of composition resistance material. A wiper arm is driven by the float moves across the segment of composition resistance material, changing the circuit resistance.
d Ratiometer-type gauges: There are two types of indicators. Both use current -measuring instruments calibrated in fuel quantity. In the diagram below, the “empty” coil would be the coil on the left and the “full” coil the one on the right. Some of these units signal a full tank with maximum resistance while others do so with minimum resistance. They are common in modern small airplanes and cars.
Figure 26. Ratiometer Page 79 of 109
4.2. Digital Fuel Quantity Indication
The electronic-type (capacitance) fuel quantity gauge differs from the other types in that it has no movable devices in the fuel tank. Instead of floats and their attendant mechanical units, the dielectric qualities of fuel and air furnish a measurement of fuel quantity. Essentially, the tank transmitter is a simple electric condenser. The dielectric (or non-conducting material) of the condenser is fuel and air (vapor) above the fuel. The capacitance of the tank unit at any one time will depend on the existing proportion of fuel and vapors in the tank. The capacitance of the transmitter is compared to a reference capacitor in a rebalance-type bridge circuit. The unbalanced signal is amplified by the voltage amplifiers that drive a phase discriminating power stage. The output stage supplies power to one-phase of a two-phase ac motor that mechanically drives a rebalancing potentiometer and indicator pointer. The electronic type system of measuring fuel quantity is more accurate in measuring fuel level, as it measures the fuel by weight instead of in gallons. Fuel volume will vary with temperature (a gallon of gasoline weighs more when it is cold than when it is hot); thus, if it is measured in pounds instead of gallons, the measurement will be more accurate. In addition to the cockpit fuel quantity indicating system, some aircraft are provided with a means to determine the fuel quantity in each tank when the aircraft is on the ground. This is accomplished in several different ways. Some manufacturers use float operated, direct reading fuel gauges mounted in the lower surface of the wing. Another means is to use under wing bayonet gauges. There are two types in use, the drip gauge and the sight gauge. When using the drip gauge it is necessary to proceed slowly, using the trial and error method to find the exact fuel level. In large area tanks a proportionately large amount of fuel is represented by a fraction of an inch variation in fuel level. The long, hollow drip tubes require some time to drain once they are filled with fuel, and a substantial error in reading will be made if the diminishing drainage drip is mistaken for the steady drip that signifies that the tube is properly positioned. 5. STALL WARNING AND ANGLE OF ATTACK SYSTEM
5.1. Angle of attack
Figure 27. AOA .In fluid dynamics, angle of attack (AOA, or
(Greek letter alpha)) is the angle between a reference line on a body (often the chord line of an airfoil) and the vector representing the relative motion between the body and the fluid through which it is moving. Angle of attack is the angle between the body's reference line and the oncoming flow. This article focuses on the most common application, the angle of attack of a wing or airfoil moving through air. Page 80 of 109
In aerodynamics, angle of attack specifies the angle between the chord line of the wing of a fixed-wing aircraft and the vector representing the relative motion between the aircraft and the atmosphere. Since a wing can have twist, a chord line of the whole wing may not be definable, so an alternate reference line is simply defined. Often, the chord line of the root of the wing is chosen as the reference line. Another alternative is to use a horizontal line on the fuselage as the reference line (and also as the longitudinal axis). Some authors do not use an arbitrary chord line but use the zero lift axis instead - zero angle of attack corresponds to zero coefficient of lift. Some British authors have used the term angle of incidence instead of angle of attack. However, this can lead to confusion with the term riggers' angle of incidence meaning the angle between the chord of an aerofoil and some fixed datum in the aero plane. Relation between angle of attack and lift The relation between angle of attack coefficient and lift coefficient can be seen as curve below.
Figure 28. A typical lift coefficient curve .
The lift coefficient of a fixed-wing aircraft varies uniquely with angle of attack. Increasing angle of attack is associated with increasing lift coefficient up to the maximum lift coefficient, after which lift coefficient decreases. As the angle of attack of a fixed-wing aircraft increases, separation of the airflow from the upper surface of the wing becomes more pronounced, leading to a reduction in the rate of increase of the lift coefficient. The figure shows a typical curve for a cambered straight wing. A symmetrical wing has zero lift at 0 degrees angle of attack. The lift curve is also influenced by wing platform. A swept wing has a lower flatter curve with a higher critical angle. Critical angle of attack The critical angle of attack is the angle of attack which produces maximum lift coefficient. This is also called the "stall angle of attack". Below the critical angle of attack, as the angle of attack increases, the coefficient of lift (Cl) increases. At the same time, above the critical angle of attack, as angle of attack increases, the air begins to flow less smoothly over the upper surface of the airfoil and begins to separate from the upper surface. On most airfoil shapes, as the angle of attack increases, the upper surface separation point of the flow moves from the trailing edge towards the leading edge. At the critical angle of attack, upper surface flow is more separated and the airfoil or wing is producing its maximum coefficient of lift. As angle of attack increases Page 81 of 109
further, the upper surface flow becomes more and more fully separated and the airfoil/wing produces less coefficient of lift Above this critical angle of attack, the aircraft is said to be in a stall. A fixed-wing aircraft by definition is stalled at or above the critical angle of attack rather than at or below a particular airspeed. The airspeed at which the aircraft stalls varies with the weight of the aircraft, the load factor, the center of gravity of the aircraft and other factors. However the aircraft always stalls at the same critical angle of attack. The critical or stalling angle of attack is typically around 15° for many airfoils. Some aircraft are equipped with a built-in flight computer that automatically prevents the aircraft from increasing the angle of attack any further when a maximum angle of attack is reached, irrespective of pilot input. This is called the 'angle of attack limiter' or 'alpha limiter'. Modern airliners that have fly-by-wire technology avoid the critical angle of attack by means of software in the computer systems that govern the flight control surfaces. In takeoff and landing operations from short runways, such as Naval Aircraft Carrier operations and STOL back country flying, aircraft may be equipped with angle of attack or Lift Reserve Indicators. These indicators measure the angle of attack (AOA) or the Potential of Wing Lift (POWL, or Lift Reserve) directly and help the pilot fly close to the stalling point with greater precision. STOL operations require the aircraft to be able to operate close to the critical angle of attack during landings and at the best angle of climb during takeoffs. Angle of attack indicators are used by pilots for maximum performance during these maneuvers since airspeed information is only indirectly related to stall behavior. 5.2. Stall Warning And safety Devices Fixed –Wing Aircraft can be equipped with devices to prevent or postpone a stall or to make it less severe or to make recovery easier. An Aerodynamic twist can be introduced to the wing with the leading edge near the wing tip twisted downward. This is called washout and causes the wing root to stall before the wing tip. This makes the stall gentle and progressive. Since the stall is delayed at the wing tip, where the aileron are roll control is maintained when the stall begin. Stall Strip is a small sharp- edge device that , when attached to the leading edge of a wing, encourages the stall to start there in preference to any other location on wing. If attached closed to the wing root, it makes the stall gentle and progressive. If attached near the wing, it encourages to drop a wing when stalling. Stall fence is a small flat plate in the direction of the chord to stop separated flow progressing out along the wing. Vortex Generators, the tiny strips of metal or plastic placed on the top of the wing near the leading edge that protrude past the boundary layer into the free stream. As the name implies, they energize the boundary layer by mixing free stream airlfow with boundary layer flow thereby creating vortices, this increases the inertia of the boundary layer. By increasing the inertia of the boundary layer, airflow separation and resulting stall may be delayed. Page 82 of 109
An Anti- Stall Strake is a leading edge extension the generates a vortex on the wing upper surface to postpone the stall. A Stick Pusher is mechanical device that prevents the pilot from stalling an aircraft. It pushes the elevator control forward as the stall is approached, causing a reduction in the angle of attack. In generic term, a stick pusher is known as a stall identification device or stall identification system. Stick Shaker is a mechanical device the shakes the pilot’s control to warn of the onset of stall. Stall Warning is an electronic or mechanical device that sounds an audible as the stall speed is approached. The majority of aircraft contain some of this device that warns the pilot of an impending stall. The simplest such device is a Stall Warning horn which consist of either a pressure sensor or moveable metal tan that actuates a switch and produces an audible warning in response. 5.3. Angle Of Attack Indicator For a light aircraft , the “ Alpha System AOA” and a nearly identical “ Lift Reserve Indicator”, are both pressure differential instrument that display margin above stall and angle of attack on an instantaneous, continuous readout.
Figure 29A .AOA Vane / Transducer
Figure 29B. AOA Indicator
Figure 29C. Page 83 of 109
6. PITOT STATIC SYSTEM 6.1. International Standard Atmosphere ( ISA )
The International Standard Atmosphere (ISA) is an atmospheric model of how the pressure, temperature, density, and viscosity of the Earth's atmosphere change over a wide range of altitudes. It has been established to provide a common reference for temperature and pressure and consists of tables of values at various altitudes, plus some formulas by which those values were derived. The International Organization for Standardization (ISO) publishes the ISA as an international standard, ISO 2533:1975. Other standards organizations, such as the International Civil Aviation Organization (ICAO) and the United States Government, publish extensions or subsets of the same atmospheric model under their own standards-making authority. The ISA model divides the atmosphere into layers with linear temperature distributions. The other values are computed from basic physical constants and relationships. Thus the standard consists of a table of values at various altitudes, plus some formulas by which those values were derived. For example, at sea level the standard gives a pressure of 1013.25 hPa (1 atm) and a temperature of 15 Celsius, and an initial lapse rate of −6.5 °C/km (roughly −2 °C/1,000 ft). The tabulation continues to 11 km where the pressure has fallen to 226.32 hPa and the temperature to −56.5 °C. Between 11 km and 20 km the temperature remains constant.
Standard Atmosphere 1976 Layers in the ISA
Level Name
Layer
Base Base Base Geopotential Geometric Lapse Base Atmospheric Height Height Rate Temperature Pressure h (in km) z (in km) (in °C/km) T (in °C) p (in Pa)
0
Troposphere
0.0
0.0
−6.5
+15.0
101325
1
Tropopause
11.000
11.019
+0.0
−56.5
22632
2
Stratosphere
20.000
20.063
+1.0
−56.5
5474.9
3
Stratosphere
32.000
32.162
+2.8
−44.5
868.02
4
Stratospause
47.000
47.350
+0.0
−2.5
110.91
5
Mesosphere
51.000
51.413
−2.8
−2.5
66.939
6
Mesosphere
71.000
71.802
−2.0
−58.5
3.9564
7
Mesopause
84.852
86.000
—
−86.28
0.3734
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In the above table, geo-potential height is calculated from a mathematical model in which the acceleration due to gravity is assumed constant. Geometric height results from the assumption that gravity obeys an inverse square law. The ISA model is based on average conditions at mid latitudes, as determined by ISO's TC 20/SC 6 technical committee. It has been revised from time to time since the middle of the 20th century. 6.2. Pitot-static Aircraft system A pitot-static system is a system of pressure-sensitive instruments that is most often used in aviation to determine an aircraft's airspeed, Mach number, altitude, and altitude trend. A pitot-static system generally consists of a Pitot tube, a Static port, and the Pitot-static instruments.
Figure 30. Pitot head This equipment is used to measure the forces acting on a vehicle as a function of the temperature, density, pressure and viscosity of the fluid in which it is operating. Other instruments that might be connected are air data computers, flight data recorders, altitude encoders, cabin pressurization controllers, and various airspeed switches. Errors in pitotstatic system readings can be extremely dangerous as the information obtained from the pitot static system, such as altitude, is often critical to a successful flight. Several commercial airline disasters have been traced to a failure of the pitot-static system.
Figure 31. Diagram of a pitot-static
The pitot-static system of instruments uses the principle of air pressure gradient. It works by measuring pressures or pressure differences and using these values to assess the speed and altitude. These pressures can be measured either from the static port (static pressure) or the pitot tube (pitot pressure). The static pressure is used in all measurements, while the pitot pressure is only used to determine airspeed.
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Pitot pressure The pitot pressure is obtained from the pitot tube. The pitot pressure is a measure of ram air pressure (the air pressure created by vehicle motion or the air ramming into the tube), which, under ideal conditions, is equal to stagnation pressure, also called total pressure. The pitot tube is most often located on the wing or front section of an aircraft, facing forward, where its opening is exposed to the relative wind. By situating the pitot tube in such a location, the ram air pressure is more accurately measured since it will be less distorted by the aircraft's structure. When airspeed increases, the ram air pressure is increased, which can be translated by the airspeed indicator. Static pressure The static pressure is obtained through a static port. The static port is most often a flushmounted hole on the fuselage of an aircraft, and is located where it can access the air flow in a relatively undisturbed area. Some aircraft may have a single static port, while others may have more than one. In situations where an aircraft has more than one static port, there is usually one located on each side of the fuselage. With this positioning, an average pressure can be taken, which allows for more accurate readings in specific flight situations. An alternative static port may be located inside the cabin of the aircraft as a backup for when the external static port(s) are blocked. A pitot-static tube effectively integrates the static ports into the pitot probe. It incorporates a second coaxial tube (or tubes) with pressure sampling holes on the sides of the probe, outside the direct airflow, to measure the static pressure. When aircraft climbs, static pressure will decrease. Multiple pressure Some pitot-static systems incorporate single probes that contain multiple pressuretransmitting ports that allow for the sensing of air pressure, angle of attack, and angle of sideslip data. Depending on the design, such air data probes may be referred to as 5-hole or 7-hole air data probes. Differential pressure sensing techniques can be used to produce angle of attack and angle of sideslip indications.
6.3. Air Data Computer system Many modern aircraft use an air data computer (ADC) to calculate airspeed, rate of climb, altitude and Mach number. In some aircraft, two ADCs receive total and static pressure from independent pitot tubes and static ports, and the aircraft's flight data computer compares the information from both computers and checks one against the other. There are also "standby instruments", which are back-up pneumatic instruments employed in the case of problems with the primary instruments Air data computer (ADC) is an essential avionics component found in modern glass cockpits. This computer, rather than individual instruments, can determine the calibrated airspeed, Mach number, altitude, and altitude trend from input data from sensors such as an aircraft's pitot-static system, gyroscopes, GPS and accelerometers. In some very high speed aircraft such as the Space Shuttle, equivalent airspeed is calculated instead of calibrated airspeed. Air data computers usually also have an input of total air temperature. This enables computation of static air temperature and true airspeed. Page 86 of 109
In Airbus aircraft the air data computer is combined with altitude, heading and navigation sources in a single unit known as the Air Data Inertial Reference Unit (ADIRU). This has now been replaced by Global Navigation Air Data Inertial Reference System (GNADIRS). Air data module (ADM ) is a component of the navigation system on fly-by-wire aircraft, such as Airbus A320 or later Airbus. Each unit converts pneumatic (air pressure) information from a Pitot tube or a static port into numerical information which is sent on a data bus. This pressure information is received and processed by the Air Data Reference (ADR) component of the Air Data Inertial Reference Unit (ADIRU). This processed information is then sent to a display management computer(s) that present information on the cockpit's primary flight display. Airspeed information is also sent to the flight computers and other electronics. Airspeed information is also sent to the auto-flight subsystem (.e.g. flight management and guidance system). The air data module is a gas pressure sensor which converts mechanical forces created by gas pressure into digital signals that can be carried to the air data reference unit. ADM generally have a maintenance bus and communication bus, and a connector on the housing for a pressurized gas line that is connected to the Pitot tube or static ports. The maintenance bus can be EIA-485 and the communication bus can be ARINC 42.
7. ALTIMETER
The altimeter measures the height of the airplane above a given pressure level. Since it is the only instrument that gives altitude information, the altimeter is one of the most vital instruments in the airplane. To use the altimeter effectively, its operation and how atmospheric pressure and temperature affect it must be thoroughly understood. A stack of sealed aneroid (Aneroid—A sealed flexible container, which expands or contracts in relation to the surrounding air pressure. It is used in an altimeter or a barometer to measure the pressure of the air.) Wafers comprise the main component of the altimeter. These wafers expand and contract with changes in atmospheric pressure from the static source. The mechanical linkage translates these changes into pointer movements on the indicator. 7.1. Principle Of Operation
Figure 32. Principle of Operation of Altimeter Page 87 of 109
The pressure altimeter is an aneroid barometer that measures the pressure of the atmosphere at the level where the altimeter is located, and presents an altitude indication in feet. The altimeter uses static pressure as its source of operation. Air is denser at sea level than aloft, so as altitude increases, atmospheric pressure decreases. This difference in pressure at various levels causes the altimeter to indicate changes in altitude. The presentation of altitude varies considerably between different types of altimeters. Some have one pointer while others have two or more. The dial of a typical altimeter is graduated with numerals arranged clockwise from 0 to 9. Movement of the aneroid element is transmitted through gears to the three hands that indicate altitude. The shortest hand indicates altitude in tens of thousands of feet; the intermediate hand in thousands of feet; and the longest hand in hundreds of feet. This indicated altitude is correct, however, only when the sea level barometric pressure is standard (29.92 inches of mercury), the sea level free air temperature is standard (+15°C or 59°F), and the pressure and temperature decrease at a standard rate with an increase in altitude. Adjustments for nonstandard conditions are accomplished by setting the corrected pressure into a barometric scale located on the face of the altimeter. Only after the altimeter is set does it indicate the correct altitude. 7.2. Type of Altimeter 7.2.1. Three Pointer Drum type Altimeter The three-pointer altimeter is the most common type of instrument used in general aviation. It is named as such because it utilizes three pointers in order to display the current altitude. One pointer is used to display 100 Ft. increments. A second is used to display 1000 Ft. increments, and the third displays 10,000 Ft. increments. The Technical Information Section of this document provides instructions on how to properly read a three-pointer altimeter.
Figure 33. Three Pointer Drum type
7.2.2. Counter-Drum Altimeter The counter-drum altimeter is named as such because it displays altitude utilizing a single pointer and a rotating drum that displays digits. The drum displays ten thousand and one thousand foot increments. The pointer displays from 0 to 999 feet.
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Figure 34.Counter Drum type Altimeter
7.2.3. Encoding Altimeter An encoding altimeter can be of either the three-pointer or counter drum type of altimeter with an encoding module built into it. The encoding module takes the altitude information and converts that data into a digital code. This code is then sent via a set of wires to the aircraft transponder. A transponder is a radio device that reports the aircraft altitude to ground control radar.
Figure 35.Encoding Altimeter
7.3. Altimeter Measurement Indicated Altitude—That altitude read directly from the altimeter (uncorrected) when it is set to the current altimeter setting. True Altitude— The vertical distance of the airplane above sea level—the actual altitude. It is often expressed as feet above mean sea level (MSL). Airport, terrain, and obstacle elevations on aeronautical charts are true altitudes. Absolute Altitude—the vertical distance of an airplane above the terrain, or above ground level (AGL). Pressure Altitude— The altitude indicated when the altimeter setting window (barometric scale) is adjusted to 29.92. This is the altitude above the standard datum plane, which is a theoretical plane where air pressure. (Corrected to 15°C) equals 29.92 in. Hg. Pressure altitude is used to compute density altitude, true altitude, true airspeed, and other performance data. Density Altitude—This altitude is pressure altitude corrected for variations from standard temperature. When conditions are standard, pressure altitude and density altitude are the Page 89 of 109
same. If the temperature is above standard, the density altitude is higher than pressure altitude. If the temperature is below standard, the density altitude is lower than pressure altitude. This is an important altitude because it is directly related to the airplane’s performance. As an example, consider an airport with a field elevation of 5,048 feet MSL where the standard temperature is 5°C. Under these conditions, pressure altitude and density altitude are the same—5,048 feet. If the temperature changes to 30°C, the density altitude increases to 7,855 feet. This means an airplane would perform on takeoff as though the field elevation were 7,855 feet at standard temperature. Conversely, a temperature of -25°C would result in a density altitude of 1,232 feet. An airplane would have much better performance under these conditions. 7.4. Altimeter Error 7.4.1. Hysteresis Error This error is induced by the aircraft maintaining a given altitude for an extended period of time, then suddenly making a large altitude change. The resulting lag or drift in the altimeter is caused by the elastic properties of materials in which comprise the instrument. The error will eliminate itself with slow climbs or descents after maintaining a new altitude for a reasonable period of time. 7.4.2. Installation Error This error is caused by the change of alignment of the static pressure port with the relative wind. Improper installation or damage to the pitot static tube will also result in improper indication. 7.4.3. Barometric Error This error is caused by changes of atmospheric pressure and temperature. Sometimes the atmosphere pressure at any particular altitude departed of aircraft is different with destination. This kind of difference will make incorrect indication of aircraft altitude relative to the ground station destination. Variations in temperature will cause differences of air density and therefore in weight and pressure of the air. Blockage on static port or line will also give an improper indication on Altimeter. Blockage it can caused by ice-formation or by a small bug. 7.5. “Q” code Altimeter setting QNE : QFE :
QNH :
Setting the standard sea level pressure of 1013.25 mb ( 29.92 inHg ) to make the altimeter read the airfield elevation. Setting the pressure prevailing at an airfield to make the altimeter read zero on landing and take-off. The Aircraft height is relative to the airport accordingly Setting the pressure scale to make the altimeter read airfield height above sea level on landing and take-off.
7.6. Altimeter Test According to FAR 43 appendix E, every altimeter and static system of airplanes used for IFR flying must be checked every 24 calendar months. The test include: Page 90 of 109
Scale Error : The barometric scale is set to 29.92 in Hg and the instrument subjected to pressure corresponding to a series of test altitude. The instrument must not have a scale error in excess of that limitation is shown on table I. Hysteresis : This test is made to determine that the instrument will be within tolerance between a reading taken when the altitude is increasing and one taken when the altitude is decreasing. After-effect: This error shows up by the altimeter not returning to its original reading after the hysteresis test has been performed. Friction : All non-servo altimeters have enough friction that some of vibration is needed for their accurate reading. Case Leak : The case is tested at 18.000 feet pressure to be sure it does not leak more than 100 feet in one minute. Barometric scale error : This test determines that the movement of the barometric scale has the proper effect on the pointers.
Figure 36.
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8. VERTICAL SPEED INDICATOR
The vertical speed indicator (VSI), which is sometimes called a vertical velocity indicator (VVI), indicates whether the airplane is climbing, descending, or in level flight. The rate of climb or descent is indicated in feet per minute. If properly calibrated, the VSI indicates zero in level flight. 8.1 Principle Of Operation Although the vertical speed indicator operates solely from static pressure, it is a differential pressure instrument. It contains a diaphragm with connecting linkage and gearing to the indicator pointer inside an airtight case. The inside of the diaphragm is connected directly to the static line of the pitot-static system. The area outside the diaphragm, which is inside the instrument case, is also connected to the static line, but through a restricted orifice (calibrated leak).
Figure 37. VSI .
Both the diaphragm and the case receive air from the static line at existing atmospheric pressure. When the airplane is on the ground or in level flight, the pressures inside the diaphragm and the instrument case remain the same and the pointer is at the zero indication. When the airplane climbs or descends, the pressure inside the diaphragm changes immediately, but due to the metering action of the restricted passage, the case pressure remains higher or lower for a short time, causing the diaphragm to contract or expand. This causes a pressure differential that is indicated on the instrument needle as a climb or descent. When the pressure differential stabilizes at a definite ratio, the needle indicates the rate of altitude change. The vertical speed indicator is capable of displaying two different types of information: • Trend information shows an immediate indication of an increase or decrease in the airplane’s rate of climb or descent. • Rate information shows a stabilized rate of change in altitude. For example, if maintaining a steady 500-foot per minute (f.p.m.) climb, and the nose is lowered slightly, the VSI immediately senses this change and indicates a decrease in the rate of climb. This first indication is called the trend. After a short time, the VSI needle stabilizes on the new rate of climb, which in this example, is something less than 500 f.p.m. The time from the initial change in the rate of climb, until the VSI displays an accurate indication of the new rate, is called the lag. Rough control technique and turbulence can extend the lag period and cause erratic and unstable rate indications.
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8.2. Instantaneous Vertical Speed Indicator (IVSI) Some airplanes are equipped with an instantaneous vertical speed indicator (IVSI), which incorporates accelerometers to compensate for the lag in the typical VSI.
Figure 38. IVSI
Instrument Check—To verify proper operation, make sure the VSI is indicating near zero prior to takeoff. After takeoff, it should indicate a positive rate of climb
9. AIRSPEED INDICATOR 9.1 Principal of Operation The airspeed indicator is a sensitive, differential pressure gauge which measures and shows promptly the difference between pitot or impact pressure, and static pressure, the undisturbed atmospheric pressure at level flight. These two pressures will be equal when the airplane is parked on the ground in calm air. When the airplane moves through the air, the pressure on the pitot line becomes greater than the pressure in the static lines. This difference in pressure is registered by the airspeed pointer on the face of the instrument, which is calibrated in miles per hour, knots, or both.
Figure 39. ASI Page 93 of 109
9.2. Airspeed Measurement Indicated Airspeed (IAS)—The direct instrument reading obtained from the airspeed indicator, uncorrected for variations in atmospheric density, installation error, or instrument error. Manufacturers use this airspeed as the basis for determining airplane performance. Calibrated Airspeed (CAS)—Indicated airspeed corrected for installation error and instrument error. Although manufacturers attempt to keep airspeed errors to a minimum, it is not possible to eliminate all errors throughout the airspeed operating range. At certain airspeeds and with certain flap settings, the installation and instrument errors may total several knots. This error is generally greatest at low airspeeds. In the cruising and higher airspeed ranges, indicated airspeed and calibrated airspeed are approximately the same. Refer to the airspeed calibration chart to correct for possible airspeed errors. True Airspeed (TAS)—Calibrated airspeed corrected for altitude and nonstandard temperature. Because air density decreases with an increase in altitude, an airplane has to be flown faster at higher altitudes to cause the same pressure difference between pitot impact pressure and static pressure. Therefore, for a given calibrated airspeed, true airspeed increases as altitude increases; or for a given true airspeed, calibrated airspeed decreases as altitude increases. Groundspeed (GS)— the actual speed of the airplane over the ground. It is true airspeed adjusted for wind. Groundspeed decreases with a headwind, and increases with a tailwind.
Instrument Check—Prior to takeoff, the airspeed indicator should read zero. However, if there is a strong wind blowing directly into the pitot tube, the airspeed indicator may read higher than zero. When beginning the takeoff, make sure the airspeed is increasing at an appropriate rate.
10. AIR TEMPERATURE 10.1. Outside Air Temperature Indication Outside Air Temperature (OAT) or Static Air Temperature (SAT) refers to the temperature of the air around an aircraft, but unaffected by the passage of the aircraft through it is. The outside air temperature is used in many calculations pertaining to flight planning, some of them being takeoff performance, density altitude, cruise performance and go-around performance. The unit of temperature can be used for Mach Number calculation.
10.2. Total Air Temperature Stagnation temperature is known as total air temperature and is measured by a temperature probe mounted on the surface of the aircraft. The probe is designed to bring the air to rest relative to the aircraft. As the air is brought to rest, kinetic energy is converted to internal energy. The air is compressed and experiences an adiabatic increase in temperature. Therefore total air temperature is higher than the static (or ambient) air temperature. Page 94 of 109
Total air temperature is an essential input to an air data computer in order to enable computation of static air temperature and hence true airspeed. The relationship between static and total air temperatures is given by:
where:
Ts T total Ma
= Mach number
ᵧ
= ratio of specific heats = approx 1.400 for dry air
= static air temperature, SAT (kelvin or degree Rankine) = total air temperature, TAT (kelvin or degree Rankine)
11. MACH SPEED A Machmeter is an aircraft pitot-static system flight instrument that shows the ratio of the true airspeed to the speed of sound, a dimensionless quantity called Mach number. This is shown on a Machmeter as a decimal fraction. An aircraft flying at the speed of sound is flying at a Mach number of one, expressed as Mach 1.
Figure 40. Mach Speed
Some older mechanical Machmeters use an altitude aneroid and an airspeed capsule which together convert pitot-static pressure into Mach number. Modern electronic Machmeters use information from an air data computer system. At Standard Sea Level conditions (corresponding to a temperature of 15 degrees Celsius), the speed of sound is 340.3 m/s (1225 km/h, or 761.2 mph, or 661.5 knots, or 1116 ft/s) in the Earth's atmosphere. The speed represented by Mach 1 is not a constant; for example, it is mostly dependent on temperature and atmospheric composition and largely independent of pressure. Since the speed of sound increases as the temperature increases, the actual speed of an object traveling at Mach 1 will depend on the fluid temperature around it. Page 95 of 109
12. AIR DATA COMPUTER As we have already learned, pressure on which the operation of the primary flight instrument is dependent are transmitted through a system of pipelines. The length and quantity of the pipelines will vary according to the size of aircraft, and also the number of station at which indication of the relevant air data is required. In order to minimized plumbing arrangement, the concept of supplying the pressure to a special unit at some centralized location and then transmitting the air data electrically to wherever required, was developed and resulted in the design of unit designated as Central Air Data computer. Basically, a computer is an analogue device that produces an electrical signal equivalent of pitot and static pressure by the combined operation of mechanical and synchronous transmission devices. The final computed signals are then supplied to the appropriate indicators. Pressure sensing is accomplished by two pressure transducer, one sensing static pressure within the altitude module, while the other senses both pitot and static pressure within the computed airspeed module.
The Mach speed module and True Airspeed module are pure signal- generating devices, which are supplied with altitude and airspeed signal data from respective module. Static air temperature data required for TAS computation is sensed by a probe located outside the aircraft at some predetermined position and is routed through the mach speed module.
Figure 41.Block Diagram
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Another type of ADC.
Figure 42.Block Diagram
Digital Air Data Computer
Figure 43.Block Diagram of DADC
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13. GYROSCOPIC INSTRUMENTS 13.1. General Gyroscope is a device for measuring or maintaining orientation, based on the principle of angular momentum. Mechanically , a gyroscope is a spinning wheel or disc in which the axle is free to assume any orientation.
Figure 44. Gyro
13. 2. Characteristic of Gyro Rigidity is the property a rotating mass has a reluctance to change its plane of rotation in space unless acted by normal force. This property is dependent on three factor are mentioned below. 1. The mass of rotor 2. The speed of rotation 3. The distance at which the mass acts from the centre.
Figure 45.
Precession is the angular change of the direction of the plane of rotation under the action of external force. This property is dependent on also three factors 1. The strength and direction of the force applied. 2. The moment of inertia of the rotor 3. The angular velocity of the rotor Page 98 of 109
Figure 46.
13. 3. Type of Gyro Free Gyro : Tied Gyro : Earth Gyro : Rate gyro
:
A gyro having complete freedom in three planes at right angle to each other. This is also called as “space gyro”. A gyro having freedom in three planes at the right angle to each other but controlled by some external source. A gyro controlled by gravity to maintain its position relative to the earth A gyro having one plane of freedom at right angle to the plane of rotation, so constructed as to measure rate of movement about
13. 3. Gyro Horizon The gyro horizon or artificial horizon, indicates the pitch and bank attitude of an aircraft relative to the vertical, and for this purpose employs a displacement gyroscope whose spin axis is maintained vertical by gravity – sensing device. The rotation of rotor relative to the vertical axis.
Figure 47. Horizontal Gyro
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Erection System These system are provided for purpose of erecting the gyroscope to its vertical position, and maintain it in that position during operation. 1. Pendulous Vane Unit This unit employed with the air-driven instrument.
Figure 48. Pendulous Vane Unit
2. Ball type Erection Unit This unit utilizes the precessional forces resulting form the effect of gravity on a number of steel ball displaced with a rotating holder suspended from the gyro.
Figure 49.Ball Type Erection Unit
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3. Torque motor leveling switch.
Figure 50. Torque motor leveling unit
4. Electromagnetic Method for fast erection
Figure 51.Electromagnetic method
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13. 4. Directional Gyro The directional gyro provides a fixed reference against which aircraft’s heading is indicates. The instrument employs a horizontal axis gyroscope and being non-magnetic it us used in conjunction with magnetic compasses. It defines the short-term heading changes during turns, while the magnetic compass provides a reliable long term heading reference as in sustained straight and level flight.
Figure 52.DG
The modern directional gyros are slave to a magnetic sensor, called a flux gate. The flux gate provides a continuously sense to the earth’s magnetic field and servo mechanism constantly correct the heading indicator. These “slaved gyro” will reduce pilot workload by eliminating the need for manual realignment every ten or fifteen minutes. 13. 5. Gyro Wander Real Wander is actual movement of the spin axis caused by engineering imperfections such as friction and unbalance. Movement about the vertical axis away from its set position is reference to as real drift. Movement about the horizontal axis away from its set position is reference to as real topple. Apparent wander is the apparent movement of the spin axis away from the local vertical. The cause of this apparent movement is the rotation of the earth combined with gyroscope rigidity. Wander may also occur when a gyroscope is transported from one point on the Earth to another, is called Transport Wander. The apparent rate of displacement is reduced if the distance to the north or south poles is reduced. In fact, when gyro is located at either the north or south pole, the rotation of the earth will cause no apparent wander.
13. 6. Rate Gyro For the detection of rates of turn, direct use is made of gyroscope precession and in order to do this , the gyro is arranged in the manner shown. Such an arrangement is known as a rate gyroscope.
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Figure 53. Rate Gyro
It will be noted that the gyro differs in two respect from those employed in directional gyros and gyro horizon. It has only one gimbal ring and spring connected between the gimbal ring and casing to restrain movement about the longitudinal axis Y-Y1. When the instrument is in its normal operating position, due to the spring restrain the rotor spin axis will always be horizontal and the turn pointer will be at the zero datum mark. With the rotor spinning its rigidity will further ensure that the zero condition is maintained. A turn to the left causes of force to be applied at the front pivot of the gimbal ring, and this is the same as trying to push the rotor round at the point F on its rim. In the following this through 90 in the direction of rotation, precession will take place at point P, thus causing the gimbal ring and rotor to tilt about longitudinal axis. If the pointer were fixed to the gimbal ring, it would tilt through the same angle and would indicate a turn and also its direction. 13.7. Turn and Bank Indicator Turn and balance indicator (T/B) and the Turn coordinator (T/C) variant are essentially two aircraft flight instruments in one device. They each act as a rate of turn indicator that displays the rate the aircraft heading is changing and a balance indicator or slip indicator that displays the slip or skid of the turn.
Figure 54. Turn and Bank Indicator
The turn and balance indicator is often referred to under various names interchangeably, such as the turn and slip indicator or the turn and bank indicator. Simply put, both the turn and balance indicator (T/B) and the turn coordinator (T/C) use a gyro-driven system. The T/B uses a needle and a ball, while the T/C uses a rolling aircraft depiction and a ball. Although the turn and balance indicator is sometimes called the turn and bank indicator, the instrument does not give the aircraft's true bank angle. In fact, Page 103 of 109
neither the T/B nor the T/C actually give true bank angle information. True bank angle is calculated using the aircraft's speed and rate of turn.
Figure 55.
Turn indicator The turn indicator is a gyroscopic instrument that works on the principle of precession. The gyro is mounted in a gimbal. The gyro's rotational axis is in-line with the lateral (pitch) axis of the aircraft, while the gimbal has limited freedom around the longitudinal (roll) axis of the aircraft. Balance indicator Balance information of the aircraft is often obtained by an inclinometer, which is recognized as the "ball in a tube." An inclinometer contains a ball sealed inside a curved glass tube, which also contains a liquid to act as a damping medium. Historically, the balance indicator in early aircraft was merely a pendulum with a dashpot for damping. The ball gives an indication of whether the aircraft is slipping, skidding or in balanced flight. The ball's movement is caused by the force of gravity and the aircraft's precession forces. When the ball is centered in the middle of the tube, the aircraft is said to be in balanced flight. If the ball is on the inside (wing down side) of a turn, the aircraft is slipping. And finally, when the ball is on the outside (wing up side) of the turn, the aircraft is skidding. 13.8. Turn Coordinator Indicator
Figure 56.
The turn coordinator (T/C) is a further development of the turn and balance (T/B) indicator with the major difference being the display and the axis upon which the gimbal is mounted. The display is that of a miniature airplane as seen from behind. This looks similar to that of an attitude indicator. "NO PITCH INFORMATION" is usually written on the instrument to avoid confusion regarding the aircraft's pitch, which can be obtained from the artificial horizon instrument. In contrast to the T/B, the T/C's gimbal is pitched up 30 degrees from the lateral axis. This causes the instrument to respond to roll as well as yaw. This allows the instrument to Page 104 of 109
display a balance change more quickly as it will react to the change in roll before the aircraft has even begun to yaw. Although this instrument reacts to changes in the aircraft's roll, it does not display the roll attitude. The turn coordinator should be used as a performance instrument when the attitude indicator has failed. This is called "partial panel" operations. It can be unnecessarily difficult or even impossible if the pilot does not understand that the instrument is showing roll rates as well as turn rates. The usefulness is also impaired if the internal dashpot is worn out. In the latter case, the instrument is said to be under-damped and in turbulence will indicate large full-scale deflections to the left and right, all of which are actually roll rate responses. In this condition it may not be possible for the pilot to maintain control of the aircraft in partial-panel operations in instrument meteorological conditions. For this and other reasons, many highly experienced pilots prefer the "older" turn and bank indicator design.
Figure 57.
14. MAGNETIC COMPASS 14.1. Earth Magnetic Field Compasses are used to determine the direction of true North. However, the compass reading must be corrected for two effects. The first is magnetic declination, the angular difference between magnetic North (the local direction of the Earth's magnetic field) and true North. The second is magnetic deviation, the angular difference between magnetic North and the compass needle due to nearby sources of iron.
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.
Figure 58. Earth Magnetic Field
The picture above , illustrates the approximate directions of the earth’s field at different places around the earth. It can be visualized as something like that developed by huge cylindrical bar magnet buried in the interior of the earth. The field lines are in a pattern you would expect from iron filings scattered around a cylindrical bar magnet. The direction of the earth’s field at the north and south magnetic poles is vertical. At the magnet equator, it is horizontal and toward the poles. 14.2. Aircraft Compass Magnetic compass will indicate a direct heading of Aircraft. When aircraft headed to the north, the letter “N” will appear at the lubber lines fixed to the case in the pilot’s view. The compass needle magnet see only the horizontal component of the earth’s field. If the airplane is not wings level, the compass very well be horizontal. This introduces an error called : “ Northerly turning error.
.
. Figure 59. Aircraft Compass
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14.3. Deviation and Compensation A compass installed in an aircraft is subjected to disturbing influences due to the presence in its vicinity of iron and steel parts as well as electric circuit. It does not, therefore give the same a of bearing as it would if it were removed from all such influences, and were influenced solely by the magnetic field due to the earth There are some Definition on the Compass system which are : 1 Calibration means the measurement of residual deviation of a compass installed in an aircraft. 2 Compensation means the correction of deviation resulting from magnetism in aircraft. 3 Deviation means the angle required to be added algebraically to a compass reading to obtain the aircraft magnetic heading. 4 Direct reading compass means a compass which has the magnetic sensing element and heading indication located in the one instrument. 5 Standby compass means a direct reading compass which is not used as primary heading reference.
Compensation Devices There are two general types of compensating devices for correcting the usual aircraft magnetic compass deviation. One is the loose magnet type and the other is the screw type. Where the loose type magnet type of compensation is employed, small cylindrical compensating magnets are placed in holes in a small drawer provided in the top of the compass. One series of hole extends in the airplane’s fore and aft direction and another series of hole extend in the airplane’s athwart ship direction. Compensating magnet of sufficient intensity are placed in either direction, or in both direction, to counteract any pull which the magnetized parts of the airplane may exert upon the north seeking magnet of compass. The screw type compensator has two screws- adjustable magneto mounted on small rotatable pivots. The two adjusting screws are accessible by removing a panel in front and near the top of the compass case. One screw is marked N-S for north- south and the other screw is marked E-W for east- west. Error can be avoided by turning one or both screws with a non magnetic screwdriver.
Figure 60. Page 107 of 109
Compass Swing The process employed for determining deviation and making correction in a compass is called “ swinging compass”. To swing a compass it is necessary to use compass rose, which is circle marked on pavement with diameter lines marked for magnetic direction every 30 ⁰. The compass rose must be flat and have no surrounding magneticinfluences. A Compass check swing must be made when: a. b. c. d. e. f. g. h. i. j.
Aircraft has suffered heavy landing Aircraft has been struck by lighting If magnet material is stowed with 10ft of compass If aircraft has been standing on one heading for several days. If a modification is carried out involving magnetic characteristic aircraft. Installation of electrical equipment – radio – instrument –etc. in vicinity of compass. If changes in deviation on any particular heading is reported When compass receives a severe knock which could displace its mounting. After certificate of Airworthiness or complete overhaul. A newly installed compass.
Deviation coefficients.
Coefficient A
=
Deviation on N + E + S + W 4
Coefficient B
=
Deviation on E - Deviation on W 2
Coefficient C
=
Deviation on N – deviation on S 2
14.4. Remote Indicating Compass In their basic from remote indicating compasses currently in use are system in which magnetic detecting element monitors a gyroscopic indicating element. The principal component of any system is a flux detector unit, sometimes called a flux valve or fluxgate. It is located in an area relatively free from any disturbing magnetic fields of the aircraft itself so that the horizontal component of the earth magnetic field can be more accurately detected by sensing element within the unit. The sensing element forms part of synchro type of transmission system which in most compasses , is couple to a horizontal- axis directional gyro contained within either a heading display indicator or master gyro unit from which heading data is transmitted to a separate indicator.
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Figure 61 . Remote Indicating Compass
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