E. H. J. Pallett IEng, AMRAes. S. Coyle E. H.J. Pallett & S. Coyle Blackwell operation without some form of automatic flight control system comprising. Integrated Systems, First Edition, by Pallett, published by Pearson Education Limited, . I place E H J Pallett at system: for example, tne control and display unit (CDU) of a flight . flight of an aircraft and system's adjustments are automatic in. Author: E.H.J. Pallett Airplane Flight Dynamics And Automatic Flight Controls. Read more Automatic Control Systems, 9th Edition - Solutions Manual.
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download Automatic Flight Control, Fourth Edition on charmaudinamas.ml ✓ FREE SHIPPING on qualified orders. Aug 27, Author of Automatic flight control, Aircraft electrical systems, Aircraft Instruments, Instrumentos del Avion, Control Automatico de Vuelo. Aircraft Instruments & Integrated System by e.h.j Pallett - - Ebook download as PDF File .pdf), Text Download as PDF, TXT or read online from Scribd The system is also integrated with an aircraft's automatic flight control system (AFCS).
The pointer rotates against the tension of a hairspring which returns the pointer to its originally selected position when the Mach speed decreases to below the limiting speed. In addition to their basic indicating function. It has. The necessary computation is effected by calibrating the scales to logarithmic functions of pitot and static pressures.
The indicating element for this purpose is a servomotor-driven digital counter. In the Figure 2. It will be noted from Fig. In aircraft having an autothrottie system. In the example illustrated. A readout of the command speed is given on a digital counter which is also mechanically set by the command speed knob. A check on the operation of the failure monitoring and flag circuits. The dial presentation and mechanical features of a typical pneumatic type of altimeter are shown in Fig.
The resultant of both curves produces the linear scale as at curve 4. This conversion is represented by the graphical example shown in Fig.
Since the ISA also assumes certain temperature values at all altitudes. In standard conditions. The bi-metal compensator is simultaneously affected by the decrease in ambient temperature. In a similar manner. At higher altitudes the same effects on elasticity will take place. We may consider these errors by taking the case of a simple altimeter situated at various levels.
In practice. As far as altimeters are concerned. Assuming that at the sea-level airfield the pressure falls to The relationship between the various altitudes associated with flight operations is presented graphically in Fig. The altimeter will thus read a greater pressure Mop and will indicate an altitude greater than ft. The pressure of A 1B1 is. At point H. I 8 1- C'4: This may be seen from the three columns shown in Fig.. Variations in temperature cause differences of air density and therefore differences in weight and pressure of the air.
It will be apparent from the foregoing that. In order.. If the temperature of the air in part AB increases. Thus the altimeter. At point A the altimeter measures the pressure of the column AC.
At a the altimeter is assumed to be subjected to standard conditions. If now the altimeter is raised through ft as at The underlying principle of this may be understood by considering the setting device to be a millibar scale having a simple geared connection to the altitude pointer as shown in Fig. If the setting is then changed to. The deflected position of the capsules appropriate to whatever pressure is acting on them at the time will not be disturbed by rotation of the mechanism.
When the knob is rotated then. Likewise it will be noted that the setting knob is also geared to the sensing element mechanism body. Hg and the other in mb interconnected through gearing to a setting knob. In the altimeter shown in Fig. There are two code letter groups commonly used in connection with altimeter setting procedures. For this purpose. The requests and transmissions are adopted universally and form part of the ICAO 'Q' code of communication.
In order to make the settings flight crew are dependent on observed meteorological data which are requested and transmitted from air traffic control. The pressure set is a value reduced to mean sea-level in accordance with ISA. The zero reading is regardless of the airport's elevation above sea-level.
When used for landing and take-off. Any value is only valid in the immediate vicinity of the airport concerned. Since an altimeter with a QNH setting reads altitude above sea- level. QFE Setting the barometric pressure prevailing at an airport to make the altimeter read zero on landing at.
Height is the vertical distance of a level. SAS Transition altitude Height. Elevation is the vertical distance of a fixed point above or below mean sea-level. The following definitions. Where a runway is below the airport elevation.
QFE Altitude. It is used for flights above a prescribed transition altitude and has the advantage that with all aircraft using the same airspace and flying on the same altimeter setting. For altimeter settings the QFE datum used is the airport elevation which is the highest usable point on the landing area.
The transition altitude within UK airspace is usually ft to ft. Flight levels. Altitude is the vertical distance of a level. The other end of the metering unit is open to the interior of the case to apply static pressure to the exterior of the capsule.
Let us now see how the instrument operates under the three flight conditions shown in the diagram. An indicator mechanism is shown in schematic form in Fig. This is accomplished by incorporating a special air metering unit in the sensing system. The dial presentation is such that zero is at the 9 o'clock position. Since the rate at which the static pressure changes is involved in determining vertical speed. A typical example of this presentation is shown in Fig.
The reason for this is that a logarithmic scale is more open near the zero graduation. This tube serves the same purpose as the one employed in a pneumatic type of airspeed indicator. It is. Certain types of indicator employ a linear scale. A pneumatic type of indicator consists basically of three principal components: At the instant of commencing a descent. The pressure inside the case. Metering units are designed to compensate for the effects of the variables over the ranges normally encountered.
The construction of a typical indicator is shown in Fig. Apart from the changes of static pressure with changes of altitude.
In addition. It consists of a cast aluminium-alloy body which forms the support for I Rocking shaft assembly. During a climb. An adjustment device is provided at the front of the indicator for settiog the pointer to zero. The basic principle is illustrated in Fig.
The accelerometer comprises a small cylinder. When a change in vertical speed is initiated. The capsule displacement in turn produces instantaneous deflection of the indicator pointer over the descent portion of the scaie. Instantaneous vertical speed indicators IVS! These indicators consist of the same basic elements as conventional VSis. The flange of the metering unit connects with the static pressure connection of the indicator case. At initiation of an ascent. Displacements of the capsule in response to differential pressure changes are transmitted to the pointer via a balanced link and rocking-shaft assembly.
The purpose of the restrictor in the bypass line is to prevent any loss of pressure change effects created by displacements at the acceleration pump.
The cylinder is connected in a capillary tube leading to the capsule. The upper spring and its adjusting screws control the rate of descent calibration. The accelerometer response decays in each case after a few seconds.
The range of adjustment around zero depends on tne scale range of any one' type of indicator. At speeds below The temperature which would overall be the most ideal is that of air under pure static conditions at the flight levels compatible with the operating range of any particular type of aircraft concerned.
The measurement of static air temperature SAT by direct means is. As the helix expands or contracts. An example of this thermometer and its installation in one type of helicopter is shown in Fig. Details of the method by which this is normally accomplished will be given in Chapter 7. If the corresponding SAT value is to be determined and indicated. Various types of sensor may be adopted for the sensing of air temperature. The simplest type. The measurement of TAT requires a more sophisticated measuring technique.
This parameter is referred to as total air temperature TAT and is derived when the air is brought to rest or nearly so without further addition or removal of heat. The thermometer is secured through a fixing hole in the side window of a cockpit. For use in aircraft capable of high Mach speeds. The element is arranged in the form of a helix anchored at one end of a metal sheath or probe.
In this context.. The bled holes in the intake casing In flight. TAT sensors are of the probe type. The probe is in the form of a small" strut and air intake made of nickel-plated berylfa: It is secured at a pre-determined location in the front fuselage section of an aircraft typically at the side.
The heater dissipates a nominal W under in-flight icing conditions. A pure platinum wire resistance-type sensing element is used and is hermetically sealed within two concentric platinum tubes. The errors involved. A secord type of TAT probe is shown in Fig. An axial wire heating element. The principal differences between it and the one just described relate to the air intake configuration and the manner in which airflow is directed through it and the probe casing.
The probe has an almost negligible time- lag. The element is wound on the inner tube. The purpose of the engine bleed air injector fitting and tube is to create a negative differential pressure within the casing so that outside air is drawn through it at such a rate AIR fl. In addition to TAT. The internal arrangement of an LCD see page 15 type of indicator is schematically shown in Fig. This controls an 'OFF' flag which under normal conditions is held out of view by an energized solenoid.
TAT indicators can. The power supply to the computer is connected via supply. Air temperature indicators As in the case of other instruments. The purpose of this element is to transmit a signal to other systems requiring air temperature information. The motor then drives the counter drums. Detection of failure of the 26 V ac power to the indicator. The system is supplied with V ac which is then stepped down and rectified by a power supply module within the indicator.
In order to indicate whether temperatures are either positive or negative. The generation of the appropriate temperature signals is also accomplished by means of a de bridge circuit.
The probe element forms one part of a resistance bridge circuit. The circuit of a probe and a basic conventional pointer and scale type of indicator is shown in Fig. An example of this would be the airspeed measuring circuit of an ADC for the computing of true airspeed see Chapter 7.
The temperature data signals are transmitted from a digital type of ADC see Chapter 7 via a data bus and receiver to a microcomputer. In some cases. Some of the variations are illustrated in Fig. I Mechanical drive. TAT -,-,,--, ,-, ,-,. Function selector 0 push-button.
TAS, each of which can be selected in sequence by a push-button function select switch. When power is first applied, the indicator displays TAT, as in Fig.
Pushing the switch in for a third time returns the display to TAT. A test input facility is provided, and when activated it causes the display to alternate between all seven segments of each of the three digits , 'ON' for two seconds, and blank for one second.
Since this is normally done b: Details of the coloured display shown at d of Fig. Air data alerting and In connection with the in-flight operation of aircraft, it is necessary to warning systems impose limitations in respect of certain operating parameters compatible with the airworthiness standards to which each type of aircraft is certificated.
It is also necessary for systems to be provided which will, both visually and aurally, alert and warn a flight crew whenever the imposed operational limitations are being exceeded.
The number of parameters to be monitored in this way varies in relation to the type of aircraft and the number of systems required for its operation overall. As far as air data measuring systems are concerned, the principal parameters are airspeed and altitude, so let us now consider the operating principles of associated alerting and warning systems typical of those used in some of the larger types of public transport aircraft.
Mach warning system This system provides an aural warning when an aircraft's speed reaches the maximum operating value in terms of Mach number, i. Mmo a typical value is 0. The system consists of a switch unit which, as can be seen from Fig. It will also be noted that in lieu of a pointer actuating system, the sensing units actuate the contacts of a switch which is connected to a 28 V de power source..
At speeds below the limiting value, the switch contacts remain closed and the de passing through them energizes a control relay. The contacts of this relay interrupt the ground connection to an aural warning device generally referred to as a 'clacker' because of the sound it emits when in operation. When the limiting Mach speed at any given altitude is reached, the airspeed sensing unit causes the switch contacts to open, thereby de-energizing the control relay so that its contacts now complete a connection from the 'clacker' to ground.
Since the 'clacker' i:. A toggle switch that is spring-loaded to 'OFF' is provided for the When placed in the 'TEST' position, it allows de to flow to the ground side of the switch unit control relay, thereby providing a bias sufficient to de-energize the relay and so cause the 'clacker' to be activated.
In the exam: The other indicator, which is in the first officer's group of flight instruments, is also of the servo-operated The 'ciacker' units associated with the indicators are respectively designated as 'aural warning I' and 'aural warning 2'. The captain's indicator contains an overspeed circuit module that is supplied by the ADC with prevailing speed data and also the limiting V,,w and Mmo values appropriate to the type of aircraft.
The contacts of the switch unit in the first officer's indicator are connected to a relay, and since these contacts remain closed at speeds below maximum values, the relay is de-energized. When the maximum speed is reached, the relay coil circuit is interrupted and its contacts then change over to provide a ground connection for the de supply which activates 'aural warning 2' clacker unit. Test switches are provided for checking the operation of each clacker by simulation of overspeed conditions.
When switch 1 is operated de is applied to the overs peed circuit module in the captain's inr! The operation of switch 2 applies de to the relay coil such that it is shorted out against the standing supply from the closed airspeed switch; the relay 'is therefore de-energized to provide a ground connection for 'aural warning 2' clacker unit.
The indicators themselves provide visual indications of overspeed and these are discernible when the airspeed pointers become positioned coincident with pre-set maximum limit pointers see Figs 2. The selected altitude is set by means of a knob on the controller, and is indicated by a digital counter which is geared to the rotors of control and resolver synchros, so that they produce a corresponding signal.
The signal is compared with the pressure altitude signal. At predetermined values of rotor voltages of both synchros, two signals are produced and n. The sequence of alerting is shown at b of Fig. As an aircraft descends or climbs to the preselected altitude the difference signal is reduced, and the logic circuit so processes the input signals that, at a pre-set outer limit H 1 typically ft above or below preselected altitude, one signal activates the aural alerting device which remains on for two seconds; the annunciator light is also illuminated.
The light remains on until at a further pre-set inner limit H2 typically ft above or below preselected altitude, the second. As an aircraft approaches the preselected altitude, the synchro system approaches the 'null' position, and no further alerting takes place.
If an aircraft should subsequently depart from the preselected altitude, the controller logic circuit changes the alerting sequence such that the indications correspond to those given during the approach through outer limit Hi, i.
Angle of attack The angle of attack AoA , or alpha a angle, is the angle between sensing the chord line of the wing of an aircraft and the direction of the relative airflow, and is a major factor in determining the magnitude of lift generated by a wing.
Lift increases as a increases up to some critical value at which it begins to decrease due to separation of the slow-moving air the boundary layer from the upper surface of the wing, which, in turn, results in separation and turbulence of the main airflow.
The wing, therefore, assumes a stalled condition, and since it occurs at a particular angle rather than a particular speed, the critical AoA is also referred to as the stalling angle. The manner in which an aircraft responds as it approaches and reaches a stalled condition depends on many other factors, such as wing configuration, i. Other factors relate to the prevailing speed of an aircraft, which largely depends on engine power settings, flap angles, bank angles and rates of change of pitch.
The appropriate responses are pre-determined for each type of aircraft in order to derive specificaliy relevant procedures for recovering from what is, after all, an undesirable situation.
An aircraft will, in its own characteristic manner, provide warning of a stalled condition, e. It is, therefore, necessary to provide a means whereby a can be sensed directly, and at some value just below that at which a stalled condition can occur it can provide an early warning of its onset.
Stall warning systems The simplest form of system, and one which is adopted in several types of small aircraft, consists of a hinged-vane-type senso. The vane is protected against ice formation by an internal heater element. When a reaches that at which the warning unit has been preset.
Control switches for normal operation and for testing are also provided in this unit. In larger tyges of aircraft. Since the pitch attitude of an aircraft is also changed by the extension of its flaps. If the aircraft's attitude changes such that a increases. The circuit of a typical system is shown in basic form in Fig. When the aircraft is on the ground and electrical power is on.. Sensing relays and shock strut microswitches on the nose landing gear are included in the circuit of a system to permit operational change-over from ground to air.
Stick-shaking is accomplished by a motor which is secured to a control column and drives a weighted ring that is deliberately unbalanced to set up vibrations of the column. In normal level flight conditions. The complete unit is accurately aligned by means of index pins at the side of the front fuselage section of an aircraft.
Sensor signals. It consists of a precision counter-balanced aerod namic.. The only signal now supplied to the amplifier and demodulator is the modified a signal. The output is then supplied to a demodulator whose circuit is designed to 'bias off the ac voltage from the contacts of K 1.
The demodulator then produces a resultant voltage which triggers the switch SS 1 to connect a 28 V de supply direct to the stick-shaker motor. V to the circuit module amplifier. Bias of! In normal flight. During take-off. When such aircraft first get into a stalled condition then.
The positions are: The comparator is also supplied with signals from a central processor unit also within the module which processes a programme to determine maximum a angles based on the relationship between flap position and three positions of the leading edge slats. If the latter is higher than a computed maximum. The manner in In certain types of aircraft the sensor signals are transmitted to an air data computer.
A confidence check on system operation may be carried out by placing the circuit module control switch in the 'TEST' position. This energizes a relay which switches the sensor signal to the motor of an indicator.
Since the switch isolates the sensor circuit from the amplifier. In order to prevent the development of a deep stall situation. The aircraft then sinks rapidly in the deep Slalled attitude. In aircraft having computerized flight control systems. When selected for installation. Whenever stick-push is activated. Another type of indicator currently in use has a pointer which is referenced against horizontal yellow.
In some cases a conventional pointer and scale type of display is used. Indicators are connected to the alpha sensors of a stall warning system. Indicators There is no standard requirement for angle of attack indicators to be installed in aircraft. In the more sophisticated types of aircraft. The field differs from that of an ordinary magnet in several This partly explains the fact that the magnetic poles are relatively large areas.
That this is so is obvious from the fact that a magne. A plane passing through the magnet and the centre of the earth would trace out on the earth's surface an imaginary line called the magnetic meridian as shown in Fig. Terrestrial magnetism The surface of the earth is surrounded by a weak magnetic field which culminates in two internal magnetic poles situated near the North and South true or geographic poles. The origin of the earth's field is still not precisely known.
The operating principle of a direct-reading compass is based on established fundamentals of magnetism. As far as present-day aircraft are concerned. It would thus appear that the earth's magnetic field is similar to that which would be expected at the surface if a short but strongly magnetized bar magnet were located at the centre.
Its points of maximum intensity. If a map were prepared to show both true and magnetic meridians. The horizontal angle contained between the true and the magnetic meridian at any place is known as the magnetic variation or declination. Figure 3. Magnet'ic variation As meridians and parallels are constructed with reference to the true or geographic North and South pcles.. BB and CC are isoclinals.
I Terrestrial magnetism. While the variation differs all over the world. At some places on the earth. The angle the lines of force make with the earth's surface at any Figure 3.
Lines are drawn on the charts. Information regarding variation and its changes are given on special charts. It will not. These lines emerge vertically from the North magnetic pole. The angle of dip at all places undergoes changes similar to those described for variation and is also shown on charts of the world. The pivot point is above the centre of gravity of the magnet system which is balanced in such a way as to minimize the effects of angle of dip over as wide a range of latitudes North and South as possible.
This total force is resolved into its horizontal and vertical components. If stated as a relative value. As in the case of variation and dip. Earth's total force When a magnet freely suspended in the earth's field comes to rest. The earth's magnetic force may be stated either as a relative value or an absolute value. It is filled with a silicone fluid to make the compass aperiodic. The majority of compasses currently in use are of the card type.
Places on these charts having the same dip angle are joined by lines known. The relationship between these components and dip is shown in Fig. The system is pendulously suspended by an iridium-tipped pivot resting in a sapphire cup supported in a holder or stem. The bowl is of plastic Diakon and so moulded that it has a magnifying effect on the card and its graduations. Compass Direct-reading compasses have the following common principal construction features: Lines of equal H and Z forces are referred to as isodynamic lines.
The fluid also provides The bowl is in the form of a brass case which is sealed by a front bezel plate.
The compass shown at b of Fig. Its magnet system is similar to the one described earlier except that needle-type magnets are used. Values are quoted by manufacturers as part of the operating data appropriate to their equipment.
Changes in liquid volume are compensated by a capsule type of expansion device. Acceleration error This may be broadly defined as the error. Changes in volume of the fluid due to temperature changes. A small lamp is provided for illuminating the card of the magnet system. In this connection it is usual to apply the compass safe distance rule which.
Compass location The location of a compass in any one type of aircraft is of importance. Errors in indication The pendulous suspension of a magnet system.
Automatic Flight Control, 4th Edition
A permanent-magnet deviation compensator is located at the underside of the bowl. Compensation of the effects of deviation due to longitudinal and lateral components of aircraft magnetism see page 87 is provided by permanent magnet coefficient 'B' and 'C' corrector assemblies secured to the compass mounting plate.
There are two main errors that result from such components. The distance is measured from the centre of a compass magnet system to the nearest point on the surface of equipment. I hemisphere: RLY W!: In either the northern or southern hemispheres.
R c I s d s system when its centre of gravity is displaced from its normal position. The forces brought into play will be as shown in Fig. Consider now an acceleration on a northerly heading in the northern hemisphere. Since both the point P and centre of gravity are in the plane of the magnetic meridian.
The reaction to this force will be equal and opposite and must act through the centre of gravity. The reverse effects occur during a deceleration. The two forces constitute a couple which. When an acceleration occurs on an easterly heading in the northern hemisphere.
When accelerating or decelerating on any fixed heading. The extent and direction of the error is dependent upon the aircraft's heading. As soon as the system is tilted.
In order to form a clearer understanding of its effects. Turning from a nonherly heading towards east or west Figure 3. In the southern hemisphere the results will be reversed in each case. For a correctly banked tum. Turning errors During a turn. If the change in attitude is also accompanied by a change in speed. As The system's centre of gravity. Let us assume that a change in heading to the eastward is required.
If an aircraft flying level is put into a climb at the sa: One further point may be mentioned in connection with these errors. In the southern hemisphere diagram b the effects are somewhat different. We may again consider the case of an aircraft turning eastward from a northerly heading.
Since the centre of gravity is now north of point P.. The south magnetic pole is now the doll inant pole and so the offset dip angle of the magnet system changes to displace the centre of gravity to the north of point P The same effects will occur if the heading changes from N to W whilst flying in the northern hemisphere.
Turning from a southerly heading towards east or west If the turns are executed in the northern hemisphere Fig. Turning through east or west When turning from an easterly or westerly heading in either the northern or southern hemispheres diagrams e - h no errors will result because the centrifugal acceleration acts in a vertical plane through the magnet system's centre of gravity and point P.
In all the above cases. For this reason the term nonherly turning error is often used when describing the effects of centrifugal acceleration on compass magnet systems. Hard-iron magnetism is of a peramenent nature and is caused. Aircraft magnetism Magnetism is unavoidably present in aircraft in varying amounts. The two types of magnetism can be further divided in the same W'iJ.
The centre of gravity is merely deflected to the north or south of point P. In turning from a southerly heading in the southern hemisphere Fig. A point which may be noted in connection with turns from E or W is that when the N or S end of the magnet system is tilted up. Y that magnetic materials are classified according to their ability to be magnetized.
Components of hard-iron magnetism The total effect of this type of magnetism at a compass position may be considered as having originated from equivalent bar magnets lying longitudinally. The deviations caused by each of the components are set out in Table 3. Q and R. There is also a third type of magnetism.
Soft-iron magnetism is of a temporary nature and is caused by the metallic materials of an aircraft which are magnetically 'soft' becoming magnetized due to induction by the earth's field. The resulting deviations are termed easterly when positive. The effect of this type of magnetism. The components are respectively denoted as P. Such magnetism depends. The strength of these components does not vary with heading or change of latitude. The polarities and strengths of components X and Y vary with changes in aircraft heading relative to the fixed direction of the earth's component H.
Components of soft-iron magnetism The effect of this type of magnetism may be considered as originating from a piece of soft-iron in which magnetism has been induced by the earth's field. Components X. Component R effective only in the aircraft altitudes indicated.
A change in the polarity of component Z will only occur with a change in magnetic hemisphere. This field. Y and Z also change with geographical location because this results in changes in the earth's field strength and direction. Table 3. Total magnetic effect The total effect of the magnetic fields that produce deviating forces relative to each of the three axes of an aircraft is determined by algebraically summing the quantities appropriate to each of the related components.. Table Each of the three components produce three soft-iron components that are designatej aX The deviations caused by such components are set out in Table The polarities and direction of components cZ and jZ depend on whether an aircraft is in the northern or southern hemisphere..
In practice it is not necessary to distinguish between them. The coefficient is calculated by taking the average of the algebraic differences between deviations measured on a number of equidistant D and E.
The relationship between them and the components of aircraft magnetism is shown in Fig. There are five coefficients designated A. Coefficient A This represents a constant deviation and may be termed as either real A.
Coefficient C This represents the resultant deviation due to the presence. The coefficient is calculated from the formula: When these components are of like signs.
When of like and unlike signs these components cause deviations whose directions are the same as those caused by components P and cZ. Coefficient D This represents the deviation due to the presence. Deviation on W 2 Since components P and cZ cause deviation which varies as the sine of an aircraft's heading 8. In the case of A. These adjustments are effected by compensator or corrector magnet devices which.
A compensator forms an integral part of a compass see Fig. Adjustment for coefficient A is effected by repositioning the compass in its mounting by the requisite number of degrees. The total deviation on an uncorrected compass for any given direction of an aircraft's heading by compass may be expressed by the equation: One pair of magnets is positioned laterally to provide a variable longitudinal. It is calculated from the formula: Coefficient E This coefficient represents the deviation due to the presence of components bY and dX of like signs.
Deflection of the compass magnet system would be obtained in a similar manner with the aircraft heading west. Variation of the field strength by rotating the magnets will. It will be apparent from the foregoing operating sequences that maximum compensation of deviation on either side of cardinal headings is obtained when the magnets are in complete alignment.
The north-seeking pole of the compass magnet system will. The coefficient C compens. If the magnets are rotated so as to strengthen the field between poles N 1 and S 2.
This would also be the case if the aircraft was heading south. The manner in which compensation is carried out may be understood by considering the case of an adjustment having to be made for coefficient B.
Since tl1e intensity of a field varies in inverse proportion to the square of the distance from its source. When the appropriate compensator magnets are in the neutral position. The total field of the magnets. At b of Fig. The gears on which magnets are mounted are connected to operating heads which. Indication of the neutral position of magnets is given by aligning datum marks.
The gyroscope and As a mechanical device a gyroscope may be defined as a system Its properties containing a heavy metal wheel or rotor. There are three such instruments. The gimbal system is mounted in a frame as shown in Fig. Both these properties depend on the principle of conservation of angular momentum.
The three degrees of freedom are obtained by mounting the rotor in two concentrically pivoted rings. The system will not exhibit gyroscopic properties unless the rotor is spinning.
When the rotor is made to spin at high speed. The complete group constitutes what is tenned the 'basic six' arrangement. The whole assembly is known as the gimbal system of a free or space gyroscope.
The three additional instruments utilize a gyroscopic type of sensing element. It will also be found. I Elements of a gyroscope.
If a weight is now suspended from the inner gimbal ring with the rotor spinning it will be found that the ring will support the weight. The two gyroscopic properties may be more closely defined as follows: If we lift the front wheel off the ground. These rather intriguing properties can be exhibited by any system in which a rotating mass is involved.
Other familiar mechanical systems possessing gyroscopic properties are aircraft propellers. Figure 4. The property v. Determining the direction of precession The direction in which a gyroscope will precess under the influence of an applied force may be determined by means of vectors and by solving certain gyrodynamic problems.
Precession of a rotor will continue. The other segments will be affected in the same way. The angular change in direction of the plane of rotation under the influence of an applied force. The greater the force. The axis about which a force is applied is termetl the input axis.
Each segment has motion m in the direction of rotor rotation. The change in direction takes place. At a in Fig. Let us assume for a moment that the rotor is broken into segments and concern ourselves with two of them at opposite sides of the rim as shown at c.
In transmitting this force to the rim of the rotor. This motion is resisted by rigidity. It is done by representing all forces as acting directly on the rotor itself. The rate ot' precession also depends on three factors: At this point there will be no further resistance to the force and so precession will cease.
This time. I I X d el In the example illustrated in Fig. As in the previous case this results in the direction of motion changing to the resultant of motion m and force F.. F x--'1. These datums are established by using vertical and horizontal spin-axis gyroscopes respectively as shown in Fig. Both types utilize their fundamental properties in the following For this reason. It will also be noted from Fig. When a Limitations of a free In flight.
Each one has three degrees of freedom and. AXIS Dll!
Aircraft Electrical Systems Pallett
Thus, to an observer on the earth having no sense of the earth's rotation, the gyroscope would appear to veer or drift. This may be seen from Fig.
S a which illustrates a horizontal-axis gyroscope at a latitude A: At 'A', the input axis is aligned with the local N-S component of w,; therefore, to an observer at latitude A When the input axis is aligned with that of the earth 'B' , drift would also be apparent, but at a rate equal to w,, i.
In order to further illustrate drift, we may consider diagram b of Fig. If the same gyroscope were to be positioned so that its input axis ZZ 1 was aligned with the E-W component of w, at any point, its spin axis would then be vertical; in other words, it becomes a vertical-axis gyroscope.
Since the plane of rotation is coincident with that of the earth, there will be no apparent drift. Real drift Real drift results from imperfections in a gyroscope such as bearing friction and gimbal system unbalance. Such imperfections cause unwanted precession which can only be minimized by applying precision engineering techniques to the design and construction.
Transport wander Let us again consider a horizontal-axis gyroscope which is S!! In this position it will exhibit an apparent drift equal to w,. Assume now that it is carried to a lower latitude, ,and with its input axis aligned with the local vertical component of w,. During the period of transport it will have appeared to an observer on the earth that the spin axis has tilted in a vertical plane, until at the new latitude it appears to be in the position shown at c of Fig.
This apparent tilt, or transpon wander, would also be observed if, during transport, the input axis were aligned with either a local N -S component, or a local E-W component of w,. Transport wander will, of course, appear simultaneously with drift, and so for a complete rotation of the earth, the gyroscope as a whole would appear to make a conical movement.
The angular velocity or transpon rate of this movement will be decreased or increased depending on whether the E-W component of an aircraft's speed is.
The N -S component of the speed will increase the maximum divergence of the gyroscope axis from the vertical, the amount of divergence depending on whether the aircraft's speed has a N or S component and also on whether the gyroscope is situated in the northern or southern hemisphere.
The relationship between w. Transport wander u V i! If the input axis of a gyroscope were to be positioned such that its spin axis was vertical, then during transport it would only exhibit transport wander.
Control of drift and transport wander Before a free gyroscope can be of practical use, drift and transport wander must be controlled so that the plane of spin of the rotor is maintained relative to the earth; in other words, it requires conversion to what is termed an earth gyroscope.
The control of transport wander is normally achieved by using gravity-sensing devices which automatically detect tilting of the gyroscope's spin axis, and applying the appropriate corrective torques. The operation of some typical control methods will be described later under the headings of the appropriate flight instruments.
Aircraft Instruments & Integrated System by e.h.j Pallett -
Displacement Depending on the orientation of its gimbal system, a displacement gyroscope limitations gyroscope can be subject. Gimbal lock This occurs when the gimbal orientation is such that the spin axis becomes coincident with one or other of the axes of freedom which serve as attitude displacement references. Let us consider, for example, the case of the spin axis of a vertical-axis gyroscope shown in Fig.
If, in this 'locked' condition of the gimbal system, the gyroscope as a whole were to be mrned, then the forces acting on the gimbal system would cause the system to precess or topple. Gimbal error This is an error which is also related to gimbal system orientation, and it occurs whenever the gyroscope as a whole is displaced with its gimbal rings not mutually at right angles to each other.
The error is particularly relevant to horizontal-axis gyroscopes when used in direction indicating instruments see page Methods of operating There are two principal methods of driving the rotors of gyroscopic gyroscopic flight flight instruments: A typical vacuum system is shown schematically in Fig. A vacuum indicator, a relief valve, and a central air filter are also provided. In operation the pump creates a vacuum that is regulated by the valve at a value between 3. Some types of tum-and-bank indicator may operate at a lower value Vacuum-operated system.
Relief valve. Vacuum connection fl'; Each instrument case has two connections: When vacuum is applied, the pressure within the cases of the instruments is re,duced to allow surrounding air to enter' and emerge through the spinning jets. The jets are positioned adjacent to a series of recesses commonly called 'buckets' formed in the periphery of each gyroscope rotor, so that as the airstreams impinge on the 'buckets', the rotors are rotated at high speed.
An example of a relief valve is shown in Fig. During system operation the valve remains closed by compression of the spring, the tension of which is pre-adjusted to obtain the required vacuum so that air pressure acting on the outside of the valve is balanced against spring tension. If for some reason the adjusted value should be exceeded, the outside air pressure would overcome spring tension, thus opening the valve to allow outside air to flow into the system until the balanced condition was once again restored.
A pressure-operated system is, as far as principal components are concerned, not unlike a vacuum system but, as will be noted by comparing Figs 4.
Electric In electrically-operated instruments, the gyroscopes are special adaptations of ac or de motors that are designed to be driven from the appropriate power supply systems of an aircraft. In current applications, ac motors are adopted in gyro horizons, while de motors are more common to tum-and-bank indicators.
Gyroscopes used for the purpose of direction indicating can also be motor-driven, but they normally form part of a magnetic heading reference system, or of the more widely adopted flight director systems. These systems will be covered in later chapters.
Gyro horizon A gyro horizon indicates the pitch and roll attitude of an aircraft principle relative to its vertical axis, and so for this porpose it employs a Pressure regulator In-line fitter.
Pressure gauge. Pump Gyroscopic instruments. Supplementary indications of roll are presented by the position of a stabilized pointer and a fixed roll angle ,scale. Two methods of presentation are shown in Fig.
The gimbal system see Fig. In operation the gimbal system is stabilized so that in level flight the three axes are mutually at right angles.
When there is a change in an aircraft's attitude, it goes into a climb, say, the instrument case and outer ring wm move about the axis YY I of the stabilized inner ring. The horizon bar is pivoted at the side and to the rear of the outer ring, and engages an actuating pin fixed to the inner ring, thus forming a magnifying lever system. The pin passes through a curved slot in the outer ring. In a climb attitude the bar pivot carries the rear The front end of the bar is therefore moved downwards through a angle than that of the outer ring, and since the movement is relative to the symbolic aircraft element, the bar wili indicate a climb attitude.
Changes in the lateral attitude of an i. Process imaging for automatic control. Control Systems. Flight Control Systems: Practical Issues in Design and Implementation. Flight control systems: State Variable Methods in Automatic Control. Automatic Control of Aircraft and Missiles.
Adaptation and learning in automatic systems. Fuzzy Control Systems. Optimal Control Systems. Neural Systems for Control. Recommend Documents. Carstens, P. Your name. Close Send. Remember me Forgot password?Basic Electronics-Bemard Grob 4. The field differs from that of an ordinary magnet in several It is from the above mean sea-level values that all other corresponding values have been calculated and presented in what is Another method of indicating operating ranges is one that uses what are termed 'memory bugs'.
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