11w-FligInst.pdf

POLITECNICO DI MILANO - DIPARTIMENTO DI INGEGNERIA AEROSPAZIALE AIRCRAFT SYSTEMS – LECTURE NOTES, VERSION 2004 Chapter 11 – Flight instruments and navigation systems

Chapter 11

Flight instruments and navigation systems

These lecture notes are available for the students of the Polytechnic of Milan for free download. No commercialisation allowed. Queste dispense possono essere gratuitamente scaricate da Internet dagli studenti del Politecnico di Milano. E’ vietata la commercializzazione.

11.1


11.1 Introduction An aircraft has a number of navigation instruments depending on its category, ranging from the few basic elements of a glider to the sophisticated and complex systems of modern airliners and combat aircraft. The number of instruments and avionic systems has also increased in the decades: the first experimental airplanes had no instruments at all; the first instrument to be introduced was the magnetic compass, when flight became something more than an experiment; then fuel level and few engine indications were introduced, followed by altimeter, airspeed indicators and, later on, systems aimed to assist pilot in long flights, low visibility operations, position evaluation etc. Almost all the instruments are collected in cabin panels, clearly visible by the crew and rationally grouped. A complete differentiation between flight instruments and avionic systems is becoming hard, since nowadays they are often linked together. Following the traditional definitions, the flight instruments described in this chapter are as follows: • magnetic compass; • altimeter; • airspeed indicator; • vertical speed indicator; • attitude gyro; • turn rate gyro; • directional gyro; • gyrocompass; and the avionic systems as follows: • radar; • ADF; • VOR/DME; • hyperbolic navigation systems; • GPS; • instrumental landing systems; • Doppler navigation systems; • inertial systems.

11.2 Magnetic Compass The magnetic compass is one of the oldest instruments that man uses for travelling. Based on the fact that a reasonably constant magnetic field is generated in the Earth

atmosphere, a magnetic needle can easily indicate the direction of the flux lines. When the needle is mounted on a graduated rose and this is sighted by an index line fixed on the vehicle, the orientation of the vehicle with respect of the magnetic field is indicated, as shown in fig. 11.1.

Fig. 11.1 – Magnetic compass

The first drawback is due to the direction of the magnetic flux lines at high latitudes, i.e. near the magnetic poles. In fact in these regions the lines are no more parallel to the Earth surface because they tend to bend and converge into the pole. A simple


11.2


compass made of a needle and rose suspended on its centre of gravity would follow completely the flux line and become unreadable at high latitudes. This error is called dip and can be easily compensated by suspending the needle equipment above its

centre of gravity, as schematically shown in fig. 11.2 and as was actually drawn in fig. 11.1. This compensation generates another error: every time the aircraft is in manoeuvre, the inertia forces bring the compass equipment to swing. This error cannot be erased, but limited if the equipment is immersed in a liquid, that usually is an oil-alcohol mixture. Declination error is another important characteristic of the magnetic compass; this is due to the fact that the

Earth magnetic axis is not coincident with the Earth geographic axis and that the flux lines are disturbed by the mineral components in the subsurface. This error can be only compensated by knowing the declination in the actual flight area, which is reported on the flight maps together with its annual variation.

Fig. 11.2 – Dip error correction

FLUX LINES CORRECTED

NOT CORRECTED

Finally a similar disturbance, called deviation, is generated by the metallic and electric components on board. This error, that changes with aircraft heading, is periodically evaluated at ground with use of ground sights and can be reduced by mounting compensating magnets in the compass housing. The residual deviation is reported in a compass compensation table as a function of heading. Modern systems are based on magnetic compass and gyroscope, and will be described after the gyroscope sections.

11.3 Altimeter The classic measure of altitude is derived from the measure of external static pressure. Pressure is a function of altitude as indicated by the plot in fig. 11.3; the

curve can easily be obtained by the following considerations. Let’s consider an air particle of density ρ and height dz, subject to the gravity g and pressure p and p+dp respectively on the lower and upper surface; if one writes the equation of equilibrium along the vertical directions, easily the following result is obtained:

1200

e

dzgdp ⋅⋅−= ρ .

The air is an ideal gas, then being described by the well-known law:

TRp ⋅⋅= ρ ,

wtTa

TNQE

0

5000

10000

15000

20000

25000

30000

35000

40000

020

040

060

080

010

00

PRESSURE [mb]

ALT

ITU

DE

[ft]

Fig. 11.3 – Pressure vs. altitud

here R is the thermodynamic constant of ideal gases of 287 m2/s2·K and T the emperature. he temperature of steady air is experimentally found to decrease with altitude ccording to a gradient a of 6.5 °C/km:

hese lecture notes are available for the students of the Polytechnic of Milan for free download. o commercialisation allowed. ueste dispense possono essere gratuitamente scaricate da Internet dagli studenti del Politecnico di Milano. ’ vietata la commercializzazione.

11.3


dzadT ⋅−= . By combining the three laws, one can quickly obtain a differential equation that links pressure and altitude and, integrated from ground level to a generic altitude z, is as follows:

aRg

Tza

pp ⋅

⋅−=

00

1 ,

where p0 and T0 are, of course, the ground values of pressure and temperature respectively. This law shows that altitude can be evaluated from a measure of pressure, provided that reference values of ground pressure and temperature are set. The altimeter is then an instrument that measure pressure, then converting it to an

altitude indication. In its basic construction, it is schematically shown in fig. 11.4. The static pressure, coming from a pitot-static probe, is brought inside a chamber where an aneroid (vacuum) capsule is fixed, made of two corrugated plates sealed together. Depending on the pressure, the capsule will expand or shrink. This movement is picked up at the centre of the non-constrained

capsule plate, where it is more intense; then it is amplified and converted into an altitude reading by means of mechanical links. Modern altimeters have an electric pick-up and conversion to altitude signal. Normally more capsules are linked in series, to get a larger deformation.

Fig. 11.4 - Altimeter

STATIC PRESS

The reference pressure p0 is set with a knob near the display. The temperature T0 is set by the manufacturer at the standard value of 288 K. This means that the altimeter will be practically always in error. The reason why this systematic error is not dangerous is in the way the reference pressure is set. Three settings are used, called QNE and QNH and QFE. The QNE requires a 1013 mb pressure set and is used at cruising stage or anyway at high altitude, with good clearance from any ground obstacle; if all the airplanes have the same pressure setting and are assigned to different flight levels, they simply cannot conflict. The QNH is used during take-off, landing and any low altitude operation; in this case the equivalent local sea level pressure is set, in such a way that the altimeter, when the aircraft is at ground, will indicate the local real altitude from sea level. In the QFE mode, the altimeter is set to the ground pressure and then will indicate null altitude when at ground. In other words the indicated altitude for QFE and QNH is correct at ground only; nevertheless for low flight operation the pilot can request the QNH altitudes of relevant surrounding obstacles (mountains, towers, antennas), being then able to perform a safe flight.

11.4 Airspeed indicator The total pressure in a stagnation point of a fluid flow of density ρ, pressure p and velocity v in non-compressible conditions is given by:


11.4


2

21 vppT ρ+= ,

meaning that one can measure the velocity of the flow by knowing total and static pressure. An airspeed indicator can be based on this principle, taking pressure signals from a pitot tube and being mechanically sensitive to the differential pressure, as schematically indicated in fig. 11.5.

In compressible conditions the law is more complex but still shows the possibility of getting the airspeed from the static and total pressures. Moreover at high speeds the Mach number becomes important for the critic aerodynamic loads of the airplane. Mach number is given by airspeed and speed of sound, the latter being function of pressure and density, and these ones being

functions of altitude. As a matter of fact a mach meter is an instruments that reads airspeed and altitude with two separated capsules, then combining the signals.

Fig. 11.5 – Airspeed indicator

TOTAL PRESS

STATIC PRESS

A more complete speed instrument is finally the mach/airspeed indicator, which combines all information on one single display. The actual airspeed of the vehicle is called true airspeed (TAS) and is not of immediate measurement because it depends strictly on air density, i.e. pressure and temperature. The airspeed directly given by the differential pressure is called indicated airspeed (IAS). This indication is subject to positioning errors of the pitot and static probes, airplane attitude and instrument systematic defects. The airspeed corrected by these errors is called calibrated airspeed (CAS). Depending on altitude, the critic airspeeds for manoeuvre, flap operation etc change because the aerodynamic forces are function of air density. An equivalent airspeed VE (EAS) is defined as follows:

0ρρVVE = ,

where: V = true airspeed; ρ = density; ρ0 = ground density. Then EAS indicator performs the important role of showing an airspeed whose critic values are independent from altitude.

11.5 Rate-of-climb indicator


11.5

This instrument shows the vertical speed and is based on a measure of the static pressure variation. This measure can be obtained by a system made of a case housing a deformable capsule, like that schematically shown in fig. 11.6. The static pressure is carried both to the case and the capsule, but a metering unit throttles the tube to the case. This means that, when the airplane changes altitude and then the external pressure and density changes, the pressures and densities in the case and capsule will follow the variation with different transients; in particular the capsule will restore immediately the static pressure and density when they change, but the case


will be in delay. During this delay a flow is generated in the orifice that tends to restore the static pressure and density. Then for instance when the airplane climbs, the external pressure decreases and the capsule pressure follows this variation; the differential pressure between the case and the external pressure generates a flow in the metering unit that will tend to restore

equilibrium, but in the meanwhile the capsule is subject to a differential pressure and then shrinks.

Fig. 11.6 – Rate-of-climb indicator

A second more sophisticated type of instrument is the instantaneous rate-of-climb indicator, which makes use of an accelerometer to give immediate indication of the vertical velocity.

11.6 Gyroscopic instruments An airplane has a number of gyroscopic instruments for the determination of orientation and manoeuvre conditions.

A gyroscope is a rigid body with one principal moment of inertia larger than the other two, in high spin around its axis of max inertia. Mechanically this is well represented by a disc rotor, with most of the mass distributed far from the axis, as shown in fig. 11.7. A gyroscope has two fundamental properties, which may be demonstrated with a rather long manipulation of the mechanical laws of rigid bodies: rigidity and precession. Rigidity is the resistance of the spin axis to tilting, i.e.

the need of a torque higher than that necessary to tilt a simple rigid body; rigidity has a mathematical definition, being given by the spin moment of inertia Iz times the spin velocity ωz:

Fig. 11.7 - Gyroscope

y

z x

zzIH ω⋅=

Precession is the fact that, when a torque is applied along a direction perpendicular to the spin axis, the gyroscope will not tilt in the same direction of the torque, but in a direction perpendicular to it and the spin axis; to be more precise, the gyroscope, when subject to the torque, will tilt in such a way to tend to align its spin axis with the torque. Both the properties are represented by the solutions of the final differential equations of the three degrees of freedom gyroscope, neglecting the part of the solution that concerns the nutation, which is an additional oscillation of the spin axis:

HMHM

xy

yx

=

−=

ω

ω, (eq. 11.1)

where: ωx,y = angular velocity around x- and y-axis (see fig. 11.7 for the reference system);


11.6Mx,y = torque around x and y-axis.


Gyro rotors can be operated by airflow from the pneumatic system or by electric motors.

11.7 Attitude indicator Also called gyro horizon, artificial horizon and bank-pitch indicator, this is made by a 3-degree-of-freedom gyroscope with spin axis normally aligned with gravity and plays the important role of showing the bank and pitch angle of the vehicle. In principle the system is simple: the gyro is suspended by axes passing through its

centre of gravity with very low friction bearings; since no torques are applied to it, it will tend to keep its spin axis direction constant, meaning that the manoeuvring airplane will rotate around the gyro. If the gyro is aligned with the

horizon and its orientation can be displayed in some way, it will give an indication of vehicle bank and pitch angles.

Fig. 11.8 – Gyro horizon

zy

xEXTERNAL GIMBAL

INTERNAL GIMBAL

As a matter of fact, the way of building up the constraint and display system is slightly complicated; the system is schematically shown in fig. 11.8 together with a photo of the display. The gyro rotor is constrained in a case, the internal gimbal, by spin axis bearings (z-direction); the case is linked to an external gimbal by another couple of hinge bearings along an axis (y-direction) orthogonal to the spin axis; the external gimbal is linked to the instrument housing, and then to the aircraft, by a final couple of hinge bearings that define an axis (x-direction) orthogonal to the previous ones, so that all the rotational degrees of freedom are obtained. All three axes pass through the centre of gravity of the rotor. The two gimbals are mechanically linked to a movable figure depicting a horizon line and a bank pointer, which are sighted through a fixed display glass with markers. The typical error that affects a 3 degree-of-freedom gyroscope is drift, that can be of three types: real, apparent and transport drift. Real drift is due to construction imperfections that generate torques. Current technology significantly limits this problem. Apparent drift is due to Earth rotation: since the gyroscope tends to keep its axis orientation with respect to a fixed reference system, a gyroscope located somewhere on the Earth surface will tilt according to a local observer. Transport wander is due to the fact that a levelled flight is parallel to the Earth surface, which is not flat, bringing to an effect similar to the apparent precession. All these errors are compensated by erection systems: pendulum-like devices fixed on the internal gimbal align with the local (apparent) gravity and activate torque motors, or air nozzles, that slowly re-align the gyro spin axis. Of course such a technique if affected by the important error of being sensitive not only to gravity but also to any inertia acceleration, meaning that it is deceived during any airplane manoeuvre. To avoid this problem, erection systems are designed to discriminate drift misalignment from manoeuvre effect, simply based on the fact that drift progression is much more moderate than manoeuvre effect.


11.7


11.8 Turn indicator The rate-of-turn indicator is a 2-degree-of-freedom gyroscope, a configuration often referred to as rate gyro; its spin axis is in the direction of the airplane pitch axis and its rotation degree of freedom around the yaw direction is suppressed. The system looks schematically as indicated in fig. 11.9, where the gimbal is linked to the

instrument housing by a couple of hinge bearings that define an axis parallel to the airplane roll axis (x). Following the reference system indicated, its behaviour may be understood from the second of eq. 11.1, i.e. considering that an angular velocity ωy around the y-axis (yaw) generates a torque Mx around the x-axis (roll). This torque, proportional to the angular velocity, is

compensated by a spring; the gimbal will have an angle of rotation proportional to the rate of turn and can be mechanically linked to the pointer of the display.

Fig. 11.9 – Rate-of-turn gyro

yx

zSPRING

This instrument also is affected by an error, because the angular velocity is measured around an axis that is perpendicular to the gimbal; the gimbal is not levelled with the horizon during a turn, because the airplane has a bank angle and also because the gimbal is tilted by the gyroscopic effect; as a matter of fact, the gyro spin direction is such that the gimbal tilting is opposite to airplane rolling: when the airplane turn to the left, the gimbal precession is to the right and vice versa, but normally the gimbal will not remain perfectly horizontal. In other words this instruments gives an approximate measure of the turn rate, being reliable only to indicate a no-turn condition and the standard 2 minutes turn (3°/s). The turn indicator contains also another instrument, the slip indicator; this is a simple inclinometer, which is a pendulum system, usually made of a ball in a glass tube, curved as a circular arc. A liquid in the tube damps the ball movement. Being a pendulum, it is subject to gravity and inertia accelerations, keeping a neutral position during straight flight or coordinated turns.

11.9 Directional gyro and gyrocompass The magnetic compass, for its construction, is sensitive to inertia forces. This means that a compass is a reliable heading instrument in the long term, but in manoeuvre conditions it may swing and be hardly readable. To provide a more precise heading instrument in these conditions, a directional gyro is used. Like the horizontal gyro, it

is a 3-degree-of-freedom system, but with horizontal axis; due to gyroscopic rigidity, it will keep its orientation during manoeuvres, but is affected by drift errors that again can be compensated by erection devices.

Fig. 11.10 – Directional gyro display

The display of the directional gyro is similar to a compass display, as shown in fig. 11.10. In general the indication of the directional gyro is reliable in the short term (10 to 20 minutes); for this reason the crew must reset it to the compass reading during levelled flight conditions, when the compass is reliable, and most of all before starting


11.8


a manoeuvre. More modern systems have a gyrocompass, or a directional gyro connected or slaved to a remote compass, that cancel the long-term drift automatically. In these systems the compass is typically an electronic compass, or flux valve, made of a series of coils that sense the direction of the magnetic field, located near the wing tips in order to minimise the disturbances from magnetic and electric fields.

11.10 Non mechanical gyroscopes As seen in the previous paragraphs, the gyro rotors are suspended by a system of hinges and gimbals that allow building up 2- or 3-degree-of-freedom gyroscopes, depending on the fact they must measure angles or angular velocities. A fundamental requirement of the hinges is the very low friction, in order to generate low perturbation to the gyro sensing. This is obtained by using precise ball bearings, but nevertheless the mechanical properties of gyroscopic instruments are quickly deteriorated and require intensive maintenance, in particular the air gyros because the small solid particles that are in suspension in the air tend to clog the mechanisms. For this reason the newest generation of gyroscopes is based on solid-state inertial systems. The most common solutions are the ring laser gyros and micro-electro-mechanical silicon sensors. In the first case a laser beam is split into two beams that are conveyed by optical fibres into two identical circular paths, but with opposite directions. At the end of the paths the beams are compared in phase; if there is an angular velocity of the system around the laser ring, the two beams will be in a phase displacement proportional to the velocity (Sagnac effect). In the second case an oscillating silicon structure, when subject to an angular rotation perpendicular to the oscillating velocity, generates a Coriolis acceleration, which is proportional to the angular velocity; this acceleration is measured and converted into the angular velocity indication.

11.11 Avionics: radio bands Avionic instruments make use of radio transmissions. Very roughly speaking, a radio signal is generated by an electric oscillator, then amplified and emitted by an aerial. This signal, called carrier, is characterised by its frequency (or wavelength). Radio frequencies are grouped in bands, as indicated in tab. 11.1.

Band Designation Frequency Wavelength

Very Low Frequency VLF 3 - 30 kHz 100 - 10 km Low Frequency LF 30 - 300 kHz 10 - 1 km

Medium Frequency MF 300 - 3000 kHz 1000 - 100 m High Frequency HF 3 -30 MHz 100 - 10 m

Very High Frequency VHF 30 - 300 MHz 10 - 1 m Ultra High Frequency UHF 300 - 3000 MHz 100 - 10 cm

Super High Frequency SHF 3 – 30 GHz 10 - 1 cm Extremely High Frequency EHF 30 - 300 GHz 10 - 1 mm

Tab. 11.1 – Radio bands The choice of the frequency depends on the aim of the avionic instrument. Radiations from the VLF to MF bands are characterised by long ranges because they are partially transmitted via ground and reflected by the ionosphere, but on the other hand they are affected by meteorological noise and magnetism phenomena; higher


11.9


frequencies do not suffer significantly from meteorological disturbances and allow more coherent and narrow beams, thus being also preferred for directional transmissions, but on the other hand they have optical range and are significantly absorbed by atmosphere.The carrier wave in itself does not contain any audio information. This is added with a modulation of the carrier, i.e. by changing its frequency or amplitude according to the audio signal that must carry; when the carrier signal is received, it is demodulated and the audio content is split from the carrier.

11.12 Radar As anybody knows, the radar is based on the detection of the echo of radio signal

pulses transmitted by the station and reflected by a target (fig. 11.11). The time lapsed from pulse transmission to echo sensing is proportional to the distance from the target, since the signal travels at the constant speed of the electromagnetic radiation. This means that between two consecutive transmissions, the system must switch to a receiving mode for a time suitable to allow the signal to come back. An oscillator (magnetron) generates the radio frequency, then modulated and sent to the aerial; this transmits a high burst of energy in

narrow directed beams, for a short period of time (order of 1 µs); the signal intensity decreases proportionally to the second power of the distance; a considerable amount of power is absorbed by the atmosphere; limited part of the signal striking the aircraft is reflected in direction of the transmitter/receiver antenna; this weak reflected signal received from the radio station is then amplified, detected and filtered.

Fig. 11.11 – Transmitted and reflected radio beam

The used frequencies are considerably high, in the spectrum of UHF and SHF, allowing for sharp beam widths. Many kinds of radar are used in aviation: primary and secondary surveillance, Doppler, altimeter and weather. Primary radar The primary radar is a ground based system used for the air traffic control; the directional antenna rotates at constant speed, then scanning all the 360° of surrounding region at each rotation; the angular position of the antenna is transduced together with the measure of target distance; the display (PPI, planned position indicator) can then indicate the target position in terms of radial and distance with respect to the radio station. The order of magnitude of the transmitted power of a long-range surveillance radar is some Mw, frequency in the UHF band, detection range higher than 250 nm for aircraft flying at high altitudes. Secondary radar The secondary radar completes the task performed by the primary one, with additional information of altitude and aircraft identification. In fact this device sends a radio signal by a directional antenna coupled with the primary radar antenna, that trigger a transponder on the aircraft, interfaced with the altimeter (more precisely with the central air data computer) and connected to an omni-directional antenna located These lecture notes are available for the students of the Polytechnic of Milan for free download. No commercialisation allowed. Queste dispense possono essere gratuitamente scaricate da Internet dagli studenti del Politecnico di Milano. E’ vietata la commercializzazione.

11.10


in the airplane belly; the transponder will reply with a coded radio signal including identification and altitude, and more recently range and bearing. The ground operator will have the full information visualised on the PPI, allowing coordination of the air traffic. Doppler radar A Doppler radar is able to evaluate the velocity of the target, by measuring the shift between the frequency of the radiation transmitted and that one of the echo; due to the Doppler effect the frequency shift is proportional to the velocity. One possible application of this system is the MTI (Moving Target Indication): only the shifted frequency echo signals are processed, then discriminating moving target from stationary objects and erasing these ones on the display. A second application of this system is the Doppler navigation, mentioned later in this chapter. Altimeter A radar altimeter is able to measure the ground distance by sending a vertical radio signal to ground and receiving its echo. Due to the size of the target (the earth surface) and its favourable inclination (orthogonal to the transmission), this type of radar needs a few w of power; radio band is usually SHF and is used for low altitude flight (less than 2500 ft). A second type of radar altimeter works on continuous wave instead of pulses. The frequency is modulated in alternation (up and down), so that the indication of height is obtained by the measure of the difference between the transmitted signal and that of its echo. This allows long-range measures, and is commonly used on satellite as well as on airplanes. Weather radar Moisture and droplets can be visualised by radar operating typically in the SHF band. The antenna is located in the nose of the aircraft, heading forward and protected by the radome, a radio-transparent fairing. Normal order of magnitude of power transmission is 50 kW, allowing for a good detection up to the range of 300 nm.

11.13 ADF The automatic direction finding system is a short-range navigation system, based on the determination of the bearing of the aircraft to a ground radio beacon. The radio beacon transmits a carrier signal in the MF band, then with some possibility of out-of-sight range; the signal contains a Morse identification of the beacon. The radio wave is received by a sense (non directional) antenna and a loop (directional) antenna on the airplane. Originally the determination of the bearing was manually operated by the crew with a radio-goniometer, i.e. a radio receiver integrated with compass, sense and loop antenna; the minimum signal intensity from the loop antenna was obtained when it was aligned with the radio wave. Next development of the ADF consisted in an electric motor that automatically oriented the loop antenna on the basis of its signal intensity with respect of that received by the sense antenna, and then transducing the orientation of the antenna structure to the cockpit indicator. Latest digital generation uses a sense antenna integrated with a loop antenna made of a rose of coils, with no moving parts.


11.11


ADF is the simplest navigation aid and, even if more sophisticated systems are now available, it is still in use. Its main and easiest use is to follow the signal to the radio beacon, which often is in an airport area.

11.14 VOR/DME and TACAN VOR is the acronym of VHF omni-directional range. It is a short-range navigation system that, like the ADF, supplies the bearing of the aircraft to a ground radio beacon. The ground station is made of a group of antennas able of generating two radio signals: a reference phase signal and a variable phase signal. The difference in phase between the two signals changes with the radial position around the beacon, in such a way to represent the radial with respect to the North: for example, a receiver exactly located East of the beacon will read a phase difference of 90°, an so on. From the practical point of view, it is as if the variable phase signal is a directional rotating signal and the reference phase signal is emitted every time the directional one passes through the North (fig. 11.11). The ground station aerials are usually fixed, so that the variable phase signal is electronically generated; older versions used a rotating antenna for the variable phase signal. The airborne system is a receiver where the radio signal is split into three sub-signals, obtaining one signal for radio beacon identification and two signals for phase difference evaluation.

A VOR, like an ADF, enable a bearing evaluation. Two beacons should be at least necessary to evaluate the position, but this would require an intensive VOR distribution on a territory. For this reason a VOR is usually associated to a DME, or distance measuring equipment, that works in the UHF band. The ground station of a DME is a transponder that is interrogated by a radio signal transmitted from the airborne DME system. This signal is made of pulses sequenced in a random fashion, within

certain limits. The signal is received by a ground transponder that, after a 50 ms delay, re-sends the sequence. The airborne receiver reads the signal, easily discriminated from all those concerning other aircraft and, from the time between transmission and reception, the distance is measured.

Fig. 11.11 – VOR functioning principle

A VOR/DME system is then able to find a polar position with respect to the ground station, in terms of bearing and distance, and then to fully localise the aircraft. TACAN, or tactical air navigation system, is still a UHF system for polar localisation, based on a DME-transponder for distance measurement and on comparison between an omni-directional reference signal and a rotating directional signal. The rotating signal is a cardiod pattern revolving at 15 cycles per seconds and, every time its maximum passes through the East, the reference signal is transmitted. The phase difference between the two signals depends on aircraft bearing. VOR and TACAN may also be coupled in a VORTAC system.


11.12


11.15 Hyperbolic navigation systems Hyperbolic navigation systems were the long-range systems available before the GPS started its operating life; at present they are considered a back-up system in case of GPS black out. The three systems used are, in historical order: Decca, Loran and Omega. The ground system consists, in all cases, in a series of non-directional radio beacons; the airborne system consists in a non-directional receiver. The name “hyperbolic” is due to the fact that the positioning is based on the determination of the difference of distance from couples of radio beacons (foci of the hyperbola), as shown in fig. 11.12. Decca was developed during the Second World War for ships positioning, then extended to aircraft and finally discontinued at the end of the 80’s. The beacons transmit synchronised continuous radio signals in the LF band; from the phase difference it is possible to evaluate the difference of distance between the two ground stations, then generating one hyperbola. With a third station is possible to define, by combination with the other two stations, one additional hyperbola, then allowing for position determination. A Decca Navigator consists actually of 4 beacons, one master and three slaves, and phase comparison is made between the master and

each slave. Accuracy vary from 100 m for a receiver close to the radio beacons and good weather conditions, to several nm when the receiver is around 400 nm from the beacons, that is considered the range of this navigator, because of interference with the radio waves transmitted through ground and those reflected by the ionosphere. Loran (long range navigation) is still a LF navigation system. The one still in use is the Loran-C, based on radio pulses transmission. A master station sends a pulse series; this is received by a first slave station that, after a fixed delay, sends another pulse series. Later a second slave station will do the same. The receiver circuit

on the aircraft evaluates the difference of distance with respect to the two master-slave couples from the difference of time when the corresponding signals are received. The range for an acceptable radio beacon signal is within 1000 nm and present coverage is guaranteed on Europe, North America and part of Asia; accuracy varies from 0.1 to 0.25 nm, depending on distance from the beacon and weather conditions.

Fig. 11.12 – Hyperbolic navigation principle

The Loran-C is at present the real back-up in case of GPS failure or intentional signal degradation. Omega is a global navigation system, fully operating since the early 80’s and terminated in 1997. It is based on a network of 8 non-directional VLF-band radio beacons and, like Decca, the difference of distance between the receiver and pairs of beacons is evaluated as phase difference between continuous radio signals.

11.16 GPS The global positioning system, and in particular the differential GPS, is by far the most accurate method of positioning. The system is based on the reception of radio These lecture notes are available for the students of the Polytechnic of Milan for free download. No commercialisation allowed. Queste dispense possono essere gratuitamente scaricate da Internet dagli studenti del Politecnico di Milano. E’ vietata la commercializzazione.

11.13


signals from a constellation of satellites with known positions, whose launches began in 1978 and that became fully operational in 1995 with two types of signals: one restricted for US military use (the system is property of the US Ministry of Defence) and one allowed for civil use, containing an intentional degradation. Since year 2000 the signal degradation is cancelled and the system has levelled the accuracy to all users. The 24 satellites, called the space segment, are uniformly grouped in 6 different orbits in such a way that from any position on the Earth surface at least 4 satellites are above the horizon. A group of US Air Force ground stations, called the ground segment, monitors constantly the satellites, generating a prediction of their future orbit parameters and supplying this information to the satellites themselves. The receivers are called the user segment. Any satellite transmits a non directional signal, in the UHF band, that contains identifier, satellite position and starting time of transmission. The receiver reads the signal and compute distance from lapsed time. Three distances from known points enable for positioning, by solving the simple system of equations:

( ) ( ) ( ) 222222iiiii tcdzzyyxx ⋅==−+−+− i = 1,2,3

where: x,y,z = position of receiver; xi,yi,zi = position of i-th satellite; di = distance of receiver from i-th satellite; c = speed of electromagnetic radiation; ti = time lapsed between transmission from i-th satellite and reception. As a matter of fact, 3 satellites give a poor positioning because ti evaluation contains an important error. In fact the satellite are equipped with atomic clocks, while the receiver have a clock with a precision that is not comparable, even if continuously updated by the satellite. Therefore a timing error te is introduced, that can be reasonably considered constant for the signals received contemporarily from satellites. Then with a fourth satellite an additional equation is generated that increases significantly the fixing accuracy. Nevertheless, the precision of the GPS is affected by other phenomena that can hardly be corrected, as follows: • accuracy in the determination of satellite position; • reflection and refraction of the radio signal in the atmosphere, that changes the

beam velocity; • noise; • position of the satellite above the horizon, that may be geometrically non

favourable for a fixing. All the above-mentioned problems make the overall accuracy of current GPS around 10 m. A significant improvement is obtained by a differential system, or DGPS; this technique makes use of an additional ground station, of known coordinates and equipped with a GPS receiver; the station is then able to estimate the local GPS error, which is considered constant in an area of 100 nm around the station; the ground station broadcasts this parameter that, when included in the fixing evaluation, increases accuracy to less than 1 m. A new technique that reduces noise can increase accuracy to 10-20 cm. These lecture notes are available for the students of the Polytechnic of Milan for free download. No commercialisation allowed. Queste dispense possono essere gratuitamente scaricate da Internet dagli studenti del Politecnico di Milano. E’ vietata la commercializzazione.

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11.17 Airborne navigation systems Airborne navigation systems allow positioning with no external aid. They can be a Doppler or inertial system. A Doppler system is made of a series of Doppler radars in the SHF band, located on

the airplane belly and transmitting beams to ground; beam orientations are such that the velocity vector of the airplane can be computed. Time integration of the velocity allows the construction of the trajectory and then estimation of the current position starting from a previously evaluated position. A “λ” or “X” configuration is possible, depending on the fact that the beams are 3 or 4; in the last case the system is redundant and more accurate. The signals, integrated with information from a gyrocompass, are fed to a computer that evaluate the velocity components along track, drift and height and then performs the integration to determine the

displacement from an initial point, converted to latitude and longitude. Radar accuracy is within 0.2% of velocity, but the overall system accuracy is significantly deteriorated by the heading indication signal; a normal accuracy of the Doppler navigator is 10 nm for a 300 nm flight.

Fig. 11.13 - λ-Doppler configuration

An inertial system is made of a series of gyroscopes and accelerometers. The present solution is called strap down inertial system: three rate gyros supply angular velocities around the three aircraft axes and enable to compute the aircraft reference system orientation starting from a known configuration; three accelerometers supply accelerations along the aircraft axes, are converted to inertial reference system and integrated to estimate position, starting from a known point. Accuracy of present systems is within 0.3 nm/h. The need of airborne systems on civil aircraft is due to a possible failure, or intentional black-out or signal degradation, of other direct positioning systems: an airborne navigator allows a reliable flight tracking for a reasonably long time between two position updates or between a position update and an approach to airport. These systems have major interest on military aircraft, that are likely to fly with no radio aids in combat operations; in this case the inertial system has the advantage of being a completely self contained system, with no radio emissions that could be easily detected from other stations.

11.18 Landing radio aids Radio aids allow safe landing with reduced visibility and are commonly installed in airports and used by civil and military aircraft enabled for instrumental flight. The most common technique is the ILS, or instrument landing system, based on an electronic generation of the approach path to runway. The aircraft instrument is able to detect the shift from path and the pilot or automatic pilot corrects the trajectory. The ground system is made of two transmitters with aerials near the runway touch down point (localiser and glidepath transmitters) and some vertical directional radio beacons along the approach line.


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The localiser is the lateral guidance and is obtained by generating a VHF directional carrier signal with a different predominant modulation at the left and right of the approach path (90 and 150 Hz respectively).

The glidepath is the vertical guidance and is obtained in the same way with a UHF directional signal with different predominant modulation above and under the approach path (90 and 150 Hz respectively). The glidepath slope is in the range 3-6°. Fig. 11.14 shows schematically the principle. When the airplane is in the path, the receiver will read the

4 modulations with same depths, otherwise there will be an unbalance proportional to the deviation from the path, which is pictorially shown on a cockpit indicator.

Fig. 11.14 – ILS localiser and glidepath

The directional radio beacons, called markers, are 2 or 3 transmitters aligned at different distances from the runway end: the outer marker around 4 nm, the middle marker around 0.5 nm and the inner marker around 0.1 nm. They transmit a VHF signal, vertically directed; when the aircraft passes on the marker area, the receiver reads the signal and the pilot is informed. Many ILS are equipped with a DME for an accurate reading of distance from runway.

The ILS are divided into different categories depending on their accuracy, and then allowing landings in different weather conditions; the highest category allows a zero visibility landing and no decision altitude specified for the pilot, which is the altitude where he must decide whether continuing landing or execute a missed approach; this

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Fig. 11.15 – MLS functioning principle
means also that the
anding can be fully operated by the autopilot, that is also interfaced to the ILS eceiver.

he ILS indicator on board is at full scale when the aircraft is more than 2° off the ath. This means that, in these conditions, the pilot has no information of the istance from runway alignment path; therefore aircraft coupling to the ILS approach ath must be done fairly in advance before landing. Such procedure, common for ivil aircraft, is too restrictive for military aircraft. The microwave landing system, or LS, is a SHF band radio aid able to generate a broader azimuth and elevation uidance for aircraft with respect to ILS, and moreover is less affected by isturbances. For this reason MLS is gradually substituting ILS in civil airports.

hese lecture notes are available for the students of the Polytechnic of Milan for free download. o commercialisation allowed. ueste dispense possono essere gratuitamente scaricate da Internet dagli studenti del Politecnico di Milano. ’ vietata la commercializzazione.

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In MLS a narrow directional radio beam sweeps across a wide coverage area, both in azimuth and elevation, at a fixed scan rate. The angular position of the aircraft in that area can be evaluated by measuring the time lapsed by two consecutive receptions of the radio beam. The sector covered is in the range of 40-60° to the left and right of the runway centreline and 15° in elevation, and is reliable up to 20 nm from the runway end. This allows the pilot to chose a convenient approach path, which can also be segmented or curved. Moreover back azimuth guidance is provided for missed approaches and departure navigation. A PDME (precision distance measuring equipment) is usually integrated in the MLS.


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