Flight Navigation

Clock Error and Pseudo Range

We assumed in the previous discussion that both the satellite and the UE set were generating identical pseudo codes at exactly the same time. Practically speaking, this is not the case. Each satellite carries an atomic clock accurate to 10-9 seconds. Achieving maximum accuracy in synchronizing the codes would require all users to carry atomic clocks with comparable accuracies, significantly increasing both the size and cost of each receiver set. As a compromise, each UE set is equipped with a quartz crystal clock.

Since the accuracy of a quartz crystal clock cannot approach that of an atomic clock, there is a difference between satellite GPS system time and UE set time. As a result, the generation of the two pseudo codes is not perfectly synchronized and a ranging error is induced. Instead of determining actual range, we measure the apparent, or pseudo range, to the satellite. This particular problem area is known as clock bias. Clock bias affects all range measurements equally. The problem is determining the amount of bias error. Using three satellites allows us to determine our position in three dimensions. By using a fourth satellite and comparing pseudo codes, the UE set internally determines the amount of adjustment necessary to make all of the measurements agree.

Satellite Clock Error

It might be safe to assume that since each satellite carries an atomic clock, it would keep extremely accurate time. Since the compact dimensions of an orbiting satellite limit the clock size, its accuracy does not approach that of ground based atomic clocks. As a consequence, there is some error in each satellite’s clock when compared with master GPS system time. The satellite’s generation of the pseudo code is slightly out of synch and some ranging error is induced. This problem is known as satellite clock error. To compensate for this type of error, the GPS control segment comes into play. Monitor stations evaluate the accuracy of the satellite’s clock and its pseudo code generation. This information is then relayed to the master control station where the necessary corrections to the satellite’s transmissions are computed. Updated information is then uploaded to the satellite via the ground antennas.

 

Ephemeris Error

Ephemeris is the ability to determine the location of a celestial body (in this case a satellite) at regular time intervals. Ephemeris error then is caused by the satellite not being exactly where we thought it was. By using estimation theory techniques, the computers at the master control station predict what the satellite’s position should be at a specific time. This predicted position is then compared with the actual position as determined by the monitor stations. Updated information on the satellite’s future position is then uploaded to each satellite on a regular basis via the ground antennas. Each satellite then continuously transmits these corrections to all users. In this way, ranging error caused by uncertainty as to the satellite’s exact position is minimized.

Atmospheric Propagation Error

We assumed that the satellite’s RF signal traveled at the speed of light, as it does in a vacuum like space. But just as light is refracted through a prism, the RF signal is bent and slowed down as it enters the ionosphere. The degree to which the signal is affected depends on the atmospheric conditions between the satellite and receiver and on the signal’s angle as it passes through the ionosphere. Atmospheric propagation error can cause position uncertainties up to 40 meters. By noting the time delay between the two L-Band signals, much of the effect caused by atmospheric propagation can be removed internally by the UE set. Since only the military is capable of simultaneously monitoring both of the frequencies, civilian users are forced to live with this error.

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Space-based GPS, such as the U.S. GPS, the Russian GLONAS, or the forthcoming EU Galileo system all function on the same principles. In fact since the three systems use different frequencies and algorithms, in general receivers of all three systems can be more accurate than a receiver of just one system, since the system errors can be canceled out.

 

Deployment of the NAVSTAR GPS constellation of satellites began with the first launch in 1977. The satellites were launched into precisely controlled orbits, allowing users with GPS equipment to receive data to determine their position. The phenomenal accuracy of GPS was its major selling point, but its many different applications were a close second. GPS determines a position referenced to a common grid known as the World Geodetic System 1984 (WGS 84). The WGS 84 grid is based upon a mathematical model and compensates for the fact that the earth is not a perfect sphere. [Figure 16-9] Derived using precise satellite measurements, it creates an accurate model of the earth’s surface. As a consequence, WGS 84 provides extremely accurate information when compared to older traditional datum references. The value of the WGS 84 grid is that positional data can be standardized worldwide. Many receiver sets are capable of converting WGS 84 data into these other commonly used references.

General System Description

The GPS system is made up of three segments—space, user, and control.

Space Segment

The space segment is deigned to have 24 satellites plus spares in 6 orbital planes. [Figure 16-10] The orbits are arranged precisely such that a minimum of four satellites are in view at all times worldwide.

User Segment

The user segment consists of user equipment (UE) sets, test equipment, and associated support equipment. [Figure 16-11] The UE set, using data transmitted by the satellites, determines the user’s position, altitude, and velocity. Transmissions from the satellites also allow the UE set to evaluate the accuracy of the navigational information being received. This is based on built-in checks of its own performance, the configuration of the satellite constellation in view, and the jamming-to-signal ratios being experienced by the set.

Control Segment

The control segment includes a network of monitor stations and ground antennas placed throughout the world. [Figure 16-12] The monitor stations track all satellites in view and monitor general health of the system. Data from the monitor stations is sent to and processed at the Master Control Station (MCS). This data is then used to refine and update the satellite’s navigational signals. These corrections are transmitted to the individual satellites via ground antennas. The operational master control station is collocated with the Consolidated Space Operations Center at Peterson Field, Colorado. Three ground antenna stations are located at Diego Garcia, Ascension Island, and Kwajalein. Five monitor stations are positioned in Hawaii, Colorado, and at the three ground antenna locations.

Theory of Operation

A UE set is capable of determining position, velocity, and time information by receiving ranging signals from a number of satellites. By measuring the difference between signal transmission and reception times and multiplying that time interval (Dt) by the speed of light, range to the satellite can be determined. In a general sense, this is very similar to the way TACAN DME functions, with one important difference. TACAN DME is an active system in that a signal must be sent from the aircraft to the selected TACAN ground station. The ground station in turn sends a reply signal back to the aircraft. The TACAN set then measures the Dt and then computes and displays the range to the station. GPS is a passive system; no signal is transmitted by the UE set to the satellite. How does the user or receiver determine when the signal was transmitted by the satellite? The solution is to encode the satellite signal so the receiver knows when it was transmitted.

 

In order to encode the signal with its transmission time, the satellite generates what is known as a pseudorandom noise (PRN) sequence or code. This code is broadcast continuously from each satellite. At the same time, the UE set simultaneously generates an identical code. When the set receives the satellite’s signal, it compares it with the code that it has been generating. If a signal arrives at the receiver with the same code generated two seconds ago, we know that the satellite’s signal took 2 seconds to reach us. [Figure 16-13]

If we know the satellite’s location in space and our distance from it, we know we are somewhere on the sphere having the satellite as its center. With two satellites in view, the user’s position is somewhere on the circle representing the intersection of two spheres. A third satellite provides an additional sphere of position whose intersection with the other two defines a three-dimensional navigation fix with timing errors. Figure 16-14 illustrates this concept in two dimensions for clarity. A fourth allows us to eliminate most of the timing errors. The accuracy of the navigation fix would be dependent on the accuracy of the measurement process (how accurately is the digital signal processed), the accuracy of the satellite positions, and the accuracy and stability of the satellite’s clocks and the receiver clock. The user equipment should be able to track the satellite’s signal to within 3 nanoseconds (3 × 10–9 seconds). This is equivalent to a 1-meter error in position. If navigational accuracy on the order of 10 meters is desired, we must be able to establish satellite position at a particular time to within at least 10 meters. This is not a trivial problem. Since the satellite is moving and is subject to complex gravitational attractions and solar winds, measuring and predicting its position within 10 meters as a function of time is quite difficult. Fortunately, this ephemeris (orbital) data is transmitted to the receiver in the form of almanac data.

Ground stations continuously monitor each satellite so its position can be corrected and passed to the other satellites in the network. Each satellite in reception range transmits its coordinates, a time factor correction, and other data. The receiver solves simultaneous equations for the unknown receiver coordinates and the correct time. Four channel GPS receivers can use one channel per satellite to maintain continuous lock on and update. Receivers with less than four channels must continuously switch frequencies and hunt for new satellites that can limit the system responsiveness.

The satellites transmit two code signals: the precision (P) code on 1227.6 MHz and the coarse acquisition (C/A) code on 1575.42 MHz. [Figure 16-15] Both codes carry the same types of information. The C/A code is transmitted with intentional errors to deny the highly accurate position from unauthorized users. The P code, like its relative, the encrypted Y code, does not include these intentional errors. To circumvent the Y code encryption, differential GPS (DGPS) receivers are being designed that receive general GPS signals, as well as a fifth signal from a ground-based transmitter. These differential transmitters can easily determine the intentional error by comparing their GPS position to the surveyed coordinates of the transmitter. The difference is the intentional error. The ground transmitters compute and relay the amount of intentional bias in the C/A code so receivers can remove the position error without use of the P or Y codes.

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Measuring Horizontal Acceleration

The key to a successful inertial system is absolute accuracy in measuring horizontal accelerations. A slight tilt of the stable platform introduces a component of earth’s gravity as acceleration on the aircraft and results in incorrect distances and velocities. [Figure 16-4] Keeping the accelerometers level is the job of the feedback circuit. The computer calculates distance traveled along the surface of the earth and moves the accelerometer through an equivalent arc.

 

Several factors affect aligning the accelerometer using this method. The earth is not a sphere, but an oblate spheroid or geoid. Because the earth is not a smooth surface, there are local deviations in the direction of gravity. The feedback circuit operates on the premise that the arc traversed is proportional to distance traveled. Actually, the arc varies considerably because of the earth’s shape; the variation is greatest at the poles. The computer must solve for this irregularity in converting distance to arc.

The accelerometers are kept level relative to astronomical rather than geocentric latitude. Using the astronomical latitude, the accelerometers are kept aligned with the local horizon and also with the earth’s gravitational field. Feedback from the computer keeps the accelerometers level, correcting for two types of apparent precession. If the inertial unit were stationary at the equator, it would be necessary to rotate the accelerometers to maintain them level because of the earth’s angular rotation of 15° per hour. Also, movement of the stabilized platform would require corrections to keep the accelerometers level. When using a local horizontal system, in which the accelerometers are maintained directly on the gyro platform, the gyro platform must be torqued by a signal from the computer to keep the platform horizontal. Apparent precession is illustrated in Figure 16-5.

A slight error in maintaining the horizontal would induce a major error in distance computation. If an accelerometer picked up an error signal of 1/100 of the G-force, the error on a 1-hour flight would be 208,000 feet (over 34 nautical miles (NM)). In 1923, Dr. Maxmillian Schuler showed a pendulum with a period of 84.4 minutes could solve the problem of eliminating inadvertent acceleration errors.

If a pendulum has a period of 84.4 minutes, it indicates the vertical, regardless of acceleration of the vehicle. He demonstrated that a device with a period of 84.4 minutes would remain vertical to the horizon despite any acceleration on the device. The fundamental principle of the 84.4-minute theorem is that if a pendulum had an arm equal in length to the radius of the earth, gravity would have no effect on the bob. This is because the center of the bob would be at the center of gravity (CG) of the earth, and the pendulum arm would always remain vertical for all motions of the pivot point. While it would be impossible to construct this pendulum, devices with an 84.4-minute cycle can be constructed using gyroscopes. The Schuler pendulum phenomenon prevents the accumulation of errors that would be caused by platform tilt and treating gravity as an acceleration. It does not compensate for errors in azimuth resulting from the precession of the steering gyro. The amplitude of the Schuler cycle depends upon the overall accuracy of the system. Figure 16-6 shows the Schuler-tuned system.

A spinning, untorqued gyro is space-oriented and appears to move as the earth rotates underneath it. This is undesirable for older systems because the accelerometers are not kept perpendicular to the local vertical. To earth-orient the gyro, we control apparent precession. If a force is applied to the axis of a spinning gyro wheel that is free to move in a gimballing structure, the wheel moves in a direction at right angles to the applied force. This is called torquing a gyro and can be considered as mechanized or induced precession. A continuous torque, applied to the appropriate axis by electromagnetic elements called torques, reorients the gyro wheel to maintain the stable element level with respect to the earth and keeps it pointed north. An analog or digital computer determines the torque to be applied to the gyros through a loop that is tuned using the Schuler pendulum principle. The necessary correction for earth rate depends on the position of the aircraft; the correction to be applied about the vertical axis depends on the velocity of the aircraft.

It is important that the stable element be leveled accurately with respect to the local vertical and aligned in azimuth with respect to true north. Precise leveling of the stable element is accomplished prior to flight by the accelerometers that measure acceleration in the horizontal plane. The stable element is moved until the output of the X and Y accelerometers is zero, indicating that they are not measuring any component of gravity and that the platform is level. Azimuth alignment to true north is accomplished before flight by starting with the magnetic compass output and applying variation to roughly come up with true north reference. From this point, gyrocompassing is performed. This process makes use of the ability of the gyros to sense the rotation of the earth. If the stable element is misaligned in azimuth, the east gyro sees the wrong earth rate and causes a precession about the east axis. This precession causes the north accelerometer to tilt. The output of this accelerometer is then used to torque the azimuth and east gyro to ensure a true north alignment and a level condition.

In the more modern gyroscopes, the gyro cannot be physically torqued, because the gyro is either not moving or the gyro is electronically suspended. In these systems, the stable platform is leveled mathematically using gyro data. The precise orientation of the X and Y accelerometers on the stable platform is less critical since the computer can mathematically correct for any orientation. The next generation of INS may work without a stable platform, with orientation and stability maintained mathematically from accelerometer inputs.

 

Integrator

Simply stated, the processing of acceleration is done with an integrator. An integrator integrates the input to produce an output: it multiplies the input signal by the time it was present. Accurate navigation demands extremely accurate integration of both acceleration and velocity. One of the most used analog integrators is the DC amplifier, which uses a charging current stabilized to a specific value proportional to an input voltage. Another analog integrator is the AC tachometer-generator that uses an input to turn a motor, which physically turns the tachometer-generator, producing an output voltage. The rotation of the motor is proportional to an integral of acceleration.

Computer

The computer changes the integrator’s outputs into useful navigation information. To do this, one accelerometer is mounted aligned to north and another is mounted 90° to the first, to sense east-west accelerations. Any movement of this system indicates distance traveled east-west and north-south. The INS maintains a local vertical reference and measures distance traveled over a reference spheroid perpendicular to the local vertical. On this spheroid, the latitude and longitude of the present position are continuously measured by the integration of velocity. In Figure 16-7, Ө represents latitude and λ represents longitude. The axes are designated X, Y, and Z, corresponding to east, north, and local vertical. This defines their positive directions. References to velocities, attitude angles, and rotation rates are about the X, Y, and Z axes. The local vertical (Z) is established by platform leveling. This is the most fundamental reference direction. To complete platform alignment, the INS uses gyrocompassing to establish true north (Y). Gyrocompassing establishes platform alignment to the earth’s axis of revolution or North Pole. The INS is capable of doing this to an accuracy of 10 minutes of arc or less. After alignment, the platform remains oriented to true north and the local vertical, regardless of the maneuvers of the aircraft.

Groundspeed components of velocity in track (V) are measured by the system along the X and Y axes. [Figure 16-8] These components, V_X and V_Y, include all effects on the aircraft, such as wind, thermals, engine accelerations, and speed brake decelerations. Some form of digital readout usually displays the groundspeed (V).

The angles between the aircraft attitude and the platform reference attitude are continuously measured by synchros. The aircraft yaws, rolls, and pitches about the platform in a set of gimbals, each gimbal being rotated through some component of attitude. TH is measured as the horizontal angle between the aircraft’s longitudinal axis and platform north. Roll and pitch angles are measured by synchro transmitters on the platform roll and pitch gimbals.

INS technology has advanced very rapidly within the past few years. Advanced navigation systems are commonly designed with INS as an essential component. INS reliability is exceptional and INS accuracies are second only to GPS. Traditional INS design has capitalized on the advances of the digital computer to increase system responsiveness.

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Inertial navigation is accepted as an ideal navigation system because it meets all the criteria of an ideal system. INS provides worldwide ground plot information regardless of flightpath and aircraft performance. An INS can measure groundspeed independently of wind and independently of the operating environment. INS is completely independent of ground transmissions and passive in operation. It is selfcontained and portable; most units weigh less than 100 pounds. Some ring laser gyro systems weigh as little as 20 pounds. The need for a system with these properties has spurred development to the point where INS is superior to almost every other navigation system. INS provides accurate velocity information instantaneously for all maneuvers, as well as an accurate attitude and heading reference. INS accuracy decreases as the time between position updates increases. INS maintains its accuracy for short flights without position updates; however, longer flights may require periodic inflight updates.

 

Types of Inertial Systems

In the last several years, inertial technology has taken several leaps forward. Early inertials were bulky devices weighing several hundred pounds, whose installation had to be precise and whose operation had to be planned in great detail. Today, there are compact systems that fit in a briefcase and can be bolted to an aircraft in any space available. While some inertial systems still have mechanical gyroscopes, pendulous linear accelerometers, and space stable platforms, most have evolved to keep pace with the advances in technology. Acoustic gyros, ring laser gyros, and electronically suspended gyros have replaced the gimbaled gyroscope. Laser and acoustic accelerometers are replacing the pendulous linear accelerometer. Highly accurate computers and precision sensors have led to modifications of the space stable platform so the INS housing does not need to be accurately aligned with the aircraft. Eventually software will perform all the functions of the space stable platform. Despite all these modern advances, we can better learn about and understand inertial systems by studying the original systems.

The basic principle behind inertial navigation is straightforward. Starting from a known point, you calculate your present position (a continuously running DR) from the direction and speed traveled since starting navigation. The difference between other navigation systems and INS is how it determines direction, distances, and velocities. Accelerations are detected by the three linear accelerometers. These accelerations are integrated over time to determine changes in velocity. Velocity is integrated a second time to determine distance traveled. Changes in vector direction are detected with angular accelerometers. As sensors detect changes in gyroscope orientation, correction signals are generated to reorient the stable platform to the original position and determine new vector direction. INS requires no other inputs. It avoids all environmental inputs, such as indicated or true airspeed, magnetic heading, drift, and winds that are necessary for dead reckoning.

Components

The five basic components of an INS are:

1. Three linear accelerometers arranged orthogonally to supply X, Y, and Z axis components of acceleration.
2. Gyroscopes to measure and use changes in aircraft vector to maintain and orient the stable platform.
3. A stable platform oriented to keep the X and Y axis linear accelerometers oriented north-south and eastwest to provide azimuth orientation and to keep the Z axis aligned with the local gravity vector. The stable platform is necessary to prevent either the X or Y axis accelerometer from picking up the force of gravity and interpreting it as an acceleration on the aircraft.
4. Integrators to convert raw acceleration data into velocity and distance data.
5. A computer to continuously calculate position information.

Linear Accelerometers

Acceleration-measuring devices are the heart of all inertial systems. It is important that they function reliably for all maneuvers within the capability of the aircraft and that all possible sources of error are minimized. Very slight accelerations and changes in heading in all directions must be detected. Changes in temperature and pressure must not affect INS operation. To do this, INS requires two types of accelerometers: linear and angular.

The simplest type of linear accelerometer consists of a pendulous mass that is free to rotate about a pivot axis in the instrument. There is an electrical pickoff that converts the rotation of the pendulous mass about its pivot axis into an output signal. This output signal is used to torque the pendulum to hold it in the original position and, since the signal is proportional to the measured acceleration, it is sent to the navigation computer as an acceleration output signal. [Figure 16-1]

To obtain acceleration in all directions, three accelerometers are mounted mutually perpendicular in a fixed orientation. To convert acceleration into useful information, the acceleration signals must be integrated to produce velocity and then the velocity information is integrated to get the distance traveled. One of the forces measured by the linear accelerometers is gravity. This acceleration may be incorrectly interpreted as an acceleration of the aircraft if the stabilized platform is tilted relative to the local gravity vector. The accelerometers cannot distinguish between actual acceleration and the force of gravity. This means that the linear accelerometers on the stable platform must be kept level relative to the earth’s surface (perpendicular to the local gravity vector). The gyroscopes keep the stabilized platform and the accelerometers level and oriented in a north-south and east-west direction.

Gyroscopes

Gyroscopes are used in inertial systems to measure angular acceleration and changes in orientation and heading. While the types of gyros are briefly discussed here, the function of the gyro is discussed in great detail in the next section on the stable platform. The original gimbaled gyroscope has been replaced by newer designs.

Electronically Suspended Gyros

These gimbal-less gyros consist of a ball that is suspended in a magnetic field and spun electronically. Evacuating the air in the gyro cavity further reduces friction. The result is a near frictionless gyro with precession rates measured in years. Optical sensors measure the ball’s orientation from symbols etched on the surface of the ball.

Ring Laser Gyro (RLG)

Accuracy and dependability of first generation systems have greatly improved with the introduction of the ring laser gyro (RLG) INS. The RLG INS replaces the three pendulous mass accelerometers with three RLG accelerometers. Technically, the RLG is not a gyroscope since it has no moving parts, but it gives the same information as a gyro. A RLG is made from a single block of glass with three holes drilled through the glass to form a triangular path. Two of the openings are plugged with mirrors and the triangular tube is filled with helium neon or other lazing gas. When the gas is charged, the lazing gas produces two counter-rotating laser beams that are reflected around the path by the mirrors. Both laser beams emerge through the third hole in the glass and are superimposed upon each other to produce an interference pattern. As the RLG moves, one beam has a longer path to travel; the other a shorter path. This causes changes in the interference pattern that are detected by photocells. The angular rate and direction of motion are computed as accelerations.

Acoustic Gyros

Another recent development is the inertial sensor based on vibrating quartz crystal technology. Like the RLG, these are not true gyros. Acoustic gyros are manufactured from a single piece of microminiature quartz rate sensor. Angular accelerations affect the patterns produced by a vibrating tuning fork and result in torque on the fork proportional to the angular acceleration. These gyros appeared in inertial units in the late 1990s.

 

Stable Platform

Autopilots and attitude indicators use gyrostabilized platforms. Inertial navigation simply requires a stable platform with higher specifications of accuracy. A gyrostabilized platform on which accelerometers are mounted is called a stable element. It is isolated from the aircraft’s angular motions by three concentric gimbals. The stable element is the mounting for the linear accelerometers, gyroscopes, and other supporting equipment. The supporting equipment includes torque motors, servo motors, pickoffs, amplifiers, and wiring. The effectiveness of the stable platform is determined by all parts of the platform, not just the accelerometers and gyros.

The linear accelerometers measure acceleration in all directions and the gyros control the orientation of the platform. The platform must contain at least two gyros with two degrees of freedom. A simple diagram of a two-degreesof- freedom gyro mounted on a single-axis platform is shown in Figure 16-2. If one-degree-of-freedom rate gyros are used, three units are needed, each gyro having its own independent feedback and control loop. The original gimbaled gyro was not very accurate by today’s standards, producing sizeable amounts of gyroscopic precession. Recent developments such as the air-bearing gyro and the electronically suspended gyro have only 1/10,000,000 the friction of a standard gyro and negligible real precession. Today’s gyros have real precession rates of less than 360° in 30 years.

The desired property of a gyro that we want to capitalize on is its stability in space. A spinning gyro tends to remain in its original position. A free spinning gyro aligned in space tends to remain pointed in the same direction unless a force acts on it. On a stable platform, any displacement of the stable element from its frame of reference is sensed by the electrical pickoffs in the gyroscopes. These electrical signals are amplified and used to drive the platform gimbals to realign the stable element in the original position. More advanced INS have a four-gimbal platform in a three-axis configuration. [Figure 16-3]

The four-gimbal mounting provides a full 360° freedom of rotation about the stable element, thus allowing it to remain level with respect to local gravity and to remain oriented to true north. This is north as established by the gyros and accelerometers, regardless of the inflight attitude of the aircraft. The azimuth, pitch, and outer roll gimbals have a 360° freedom of rotation about their own individual axis. The fourth, or inner roll, gimbal has stops limiting its rotation about its axis. This gimbal is provided to prevent gimbal lock, which is a condition that causes the stable element to tumble. Gimbal lock can occur during flight maneuvers, such as a loop, when two of the gimbal axes become aligned parallel to each other, causing the stable element to lose one of its degrees of freedom.

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NAVAIDS

NAVAIDS are easily added to a computer system. Very high frequency (VHF) omnidirectional range (VOR) or tactical air navigation system (TACAN) bearings and distance measuring equipment (DME) provide the same information as a radar fix. The computer needs the location and frequency of the transmitter, which can be programmed into the computer before the flight begins. Some corrections must be applied to bearing data. The computer must correct for magnetic variation and slant range from the station to the aircraft.

 

Pressure Altimeter

Pressure altimeter data is an input to the true airspeed computations. Additionally, it can be used with temperature data to compute true altitude.

Radar

When a ground mapping radar is incorporated into the navigation system, present position can be corrected based on the measurements to surveyed radar returns. The operator identifies radar returns on his radar scope and measures the range and bearing to the return. The operator determines the aircraft position relative to the return and updates the aircraft position. Automatic systems allows the operator to pre-load the coordinates of radar returns in a database, place a movable electronic cursor (or crosshairs) on the return, and push a button to update the system. The computer determines the distance and bearing from the aircraft to the set coordinates. The computer then generates the cursor on the radarscope at the calculated range and bearing. If there is any error in the navigation system position, the cursor will not fall on the radar return. The operator adjusts the cursor or crosshairs onto the radar return. The operator pushes a button to automatically update the system.

Temperature Sensors

The air data computer uses the information collected by temperature sensors. Temperature gradients can be used with pressure altimeter data to compute true altitude.

True Airspeed

True airspeed can be calculated from indicated airspeed, temperature, and pressure. True airspeed and winds can be used as a backup for cross-checking groundspeed.

Independent Systems

INS and GPS can also act as sensors for a navigation system. They are discussed in greater detail later in this category.

Determining Position

The ever-present problem facing the navigator is determining aircraft position. With a navigation system, this problem is solved because the computer converts input data into a constantly updated present position for the aircraft. Advanced systems provide altitude, attitude, heading, and velocity information.

The mathematics of navigation over the surface of a sphere has been known for several centuries. Starting from an initial position, the computer determines the distance and direction traveled since starting navigation. Aircraft direction, or track, may be supplied by INS, GPS, or the heading reference system in combination with doppler drift. Groundspeed may come from INS, GPS, doppler groundspeed, or may be determined from any NAVAID capable of range and bearing fixes. The computer multiplies speed against time interval to determine distance traveled. Distance is projected along the aircraft track to obtain the new position. Track and speed are sampled and present position is updated many times per second. Waypoint navigation is a simple addition to the navigation computer. A database of coordinates can be added to the system to determine distance to go and estimated time of arrival (ETA). If the aircraft changes speed, the ETA is automatically updated using the new groundspeed.

Decision Algorithm

Simple navigation systems determine position as described above. The operator updates the position for errors that will eventually occur. More complex systems have additional problems. When a system has a variety of sources that provides redundant information, how does the computer decide which source to use? What if the sensors are subject to errors? What if the operator inputs an inaccurate update to the system? How can we get a computer to make simple decisions once left to the navigator? Can we program a computer to analyze and correct for the predictable and unpredictable errors in sensor data? Bias in the accuracy and variability of data are two types of error that navigation systems actually experience and can be solved with the use of statistical software called decision algorithms.

To compensate for these predictable and unpredictable errors in sensor data, we can include statistical measuring software that weigh the accuracy of each data source and the accuracy of the data itself. These programs determine the most likely value for track and velocity in order to compute the most likely present position.

One type of program used to determine the most likely sensor values is called a Kalman filter. Kalman filters are used extensively in computer controlled communications, electronics, and equipment. When used as part of a navigation system, a Kalman filter computes the most likely position of the aircraft and updates the weighing factors with each new position update. The Kalman filter compares the actual sensor data used prior to the update with the data from the update. By comparing the first position with the second position, actual distance and heading can be determined. It then determines the amount of error in the original data and estimates a correction to the data for the next time period. The Kalman filter is an iterative program requiring several updates prior to achieving completely reliable data. If used, the Kalman filter will also be used to evaluate the reliability of operator inputs and weigh how much of each position update to accept. Kalman filtering provides increased reliability in navigation systems so an operator can trust that the information used is valid. Kalman filters protect the operator from inaccurate sensor data and even operator error.

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In the same way an autopilot frees a pilot from the manual operations of flying, a navigation system relieves you of many manual operations required to direct the aircraft. When sensors are tied into a navigation system, the system automatically uses their data to compute present position for the navigator. During a flight from the United States to a foreign country, the aircraft may pass over areas of land, water, and icecaps. You may have to deal with conditions of overcast, undercast, day, night, altitude changes, turn points, and air traffic requirements. To handle these conditions at high speeds more effectively, the navigator uses a navigation system.

 

Types of Systems

Navigation systems can be classified according to many criteria. Systems can be classified by capability, such as visual flight rules (VFR)-only or all-weather. They can be classified as either self-contained or externally-referenced. Each system has advantages and disadvantages, but this discussion is confined to self-contained and externally referenced systems.

Self-Contained Navigation Systems

Self-contained systems (radar, celestial, INS, etc.) are complete in that they do not depend upon externally transmitted data.

Externally-Referenced Navigation Systems

Externally-referenced aids (GPS, NAVAIDs, etc.) include all aids that depend upon transmission of energy or information from an external source to the aircraft. While externally referenced aids have enormous installation and operating costs to the system administrator, they have much lower equipment and maintenance costs to the user.

The Ideal System

Every navigation system has certain advantages and disadvantages. A particular navigation system is selected for use in an aircraft when its advantages outweigh its disadvantages. In some cases, several components are included in a system to provide adequate, redundant information for all possible flight situations. The ultimate navigation system should have the following characteristics:

• Groundplot DR information—the system must indicate the position and velocity relative to the ground.
• Global coverage—capable of positioning and steering the aircraft accurately and reliably any place in the world.
• Self-contained—must not rely on ground or space transmissions of any kind.
• Flexible—works well despite unplanned deviations. The system must work well at all altitudes and speeds.

Components

The navigational system consists of three parts:

1. The computer or central processing unit (CPU)
2. Data-gathering sensors, such as astrotrackers, GPS, ground-mapping radar, or NAVAIDS
3. An operator input/output (I/O) interface

The CPU takes in all available data and converts it into usable navigation information. Control panels or computer keyboards allow the operator to control and make inputs to the computer. Data is displayed for the operator on display panels, radar screens, or computer screens. Additional hardware components could include terrain following radar or television cameras.

Computer Unit

Most navigation systems are hybrids of the two basic computer types: analog and digital.

Analog

Analog computers are more specific in design and function than digital computers. While analog computers process vast amounts of similar data, they are not very flexible and cannot be used for multiple purposes. Radar scan converters efficiently process collected radar signals into video images. Video processors collect and process images into video displays. Other examples of analog computers are terrainavoidance computers and terrain-following computers.

Digital

Digital computers are lighter and more compact than analog computers. Hand-held calculators and laptop computers are two examples of the miniaturization possible with digital computers. You can put a great deal of computing power and capability into a small box; the biggest limitation is increased cost. An analog radar scan converter is very efficient at processing radar data, but it cannot be used for other applications. On the other hand, digital computers can be loaded with navigation software, aerial delivery software, and diagnostic programs. These computers can mathematically manipulate data in any way imaginable, because they deal strictly with digital information. The output from digital computers may need to be converted into an analog format for most efficient use by the navigator; however, the digital computer cannot do that. It can display the digital data in an approximation of analog data. While a digital computer can perform any mathematical function, it must first be programmed for that function. Inflight reprogramming is not generally possible.

 

Sensors

Many types of sensors are used for inputs to navigation systems.

Astrotracker

The use of astrotrackers has decreased; however, they are still excellent sources of position information. They automatically track celestial bodies and compute position information using celestial techniques. They are passive but require clear skies.

Doppler

The Doppler radar measures groundspeed and drift. These two data inputs can be put to several uses in the computer system. Doppler groundspeed is used to determine distance to update the aircraft position. Drift can be used to compute winds and aircraft track. Doppler outputs can be used in platform leveling and verifying inertial groundspeed in an INS. Doppler radar is an essential part of many navigation computer systems.

Heading System

The gyro-stabilized magnetic heading source is corrected to true heading with the local magnetic variation. This can be applied manually or automatically from a database in the computer. Magnetic or true course can be calculated by applying doppler or inertial drift.

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