Control and Display Unit (CDU)

Control and Display Unit (CDU)

The CDU provides the major input link to the system and allows the flight crew to make inputs to obtain EPR and airspeed displays and can also be used for obtaining decision-making data in relation to an aircraft's flight profile. The CRT has a 2 in x 3 in screen and enables data to be displayed over a 13 (column) x 6 (row) matrix.

The selection of EPR and airspeed data for various phases of flight is accomplished by a flight mode select switch, the modes and associated displays being as follows:

      

• TO : Take-off EPR limits for the outside air temperature entered by the flight crew

• CLB: EPR and speeds for the desired climb profile; best economy, maximum   climb rate, or crew-selected speeds

• CRZ:EPR and speeds for the desired cruise schedule; best economy, long-range cruise, or crew-selected speeds

            

• DES: Descent speed, time and distance for best economy

• HOLD:            EPR, speed and endurance for holding

• CON:  Maximum continuous EPR limits for existing altitude, temperature and speed

• GA: Go-around EPR limit for existing altitude, temperature and speed

The standby (STBY) position of the select switch is used for data entry and for an automatic check-out of the system.

The function of the `ENGAGE' key is to couple the target command `bugs' of the Mach/airspeed indicator and EPR indicators to computer command signals which drive the bugs to indicate the speed and EPR values corresponding to those displayed on the CRT screen. If the data is verified by the computer to be valid, engageable and different from the data presently engaged, the engage key illuminates and is extinguished after engagement takes place; at the same time the appropriate light of the mode annunciator is illuminated.

The key marked `TURB' is for use only in cruise and when turbulent flight conditions are to be encountered. When pressed it causes the CRT to display the appropriate turbulence penetration data, i.e., airspeed in knots (also Mach number at high altitudes), pitch attitude and the N<sub>l</sub> percentage rpm. In the turbulence mode, the target command speed and EPR `bugs' engage automatically. This mode is disengaged by pressing the key a second time or else engaging another flight mode.

In order that the flight crew may load keyboard-selected data into the system, three push-button switches are provided above the keyboard for Selecting, Clearing and entering data. In connection with the selection and entering of data, question marks and two symbols are displayed at the right-hand end of a data line; a caret (<) and an asterisk (*). The caret signifies that the computer is ready to accept data, while the asterisk signifies that the data next to it may be entered or changed if necessary.

The keyboard primarily serves a dual function in that it (i) permits the flight crew to enter pure numeric data into the computer and (ii) permits desired performance function data to be called up from the computer for display. The data appropriate to the keys is given in Table 2.4 and is displayed in the form of pages, each page being numbered in the top right-hand corner. For example, the page shown on the CDU in Figure 2.4 is page 1 of a set of four relating to `economy fuel' in the cruise mode. In order to call up each of the remaining pages the PAGE key is successively pressed. Similarly, the PAGE key permits cycling of the pages in reverse order. When a flight mode or performance function is first selected, the first page of a set is always automatically displayed. The RCL key is used whenever a performance function is being displayed and if it is required to recall a display corresponding to a selected flight mode.

The two switches in the upper right-hand corner of the CDU are associated with auto-throttle system operation. When the A/T Annunciator switch is pressed, an internal light is illuminated to indicate connection of the auto - throttle system and at the same time an `EPR' light in the mode Annunciator is illuminated. The PDCS then adjusts the throttles to track the EPR target values displayed on the CDU and by the command bugs of the associated EPR indicators. In order for the auto-throttle system to adjust engine power in relation to indicated airspeed, the second switch JAS SEL Annunciator' is operated; the system then drives the throttles so as to track the speed target values displayed on the CDU and by the command bug of the Mach/airspeed indicator.

2.5 Computer

The computer is of the hybrid type, and the inputs, outputs and unit interfaces are as shown in Figure 2.4. Program storage is by means of a PROM and an additional non-volatile memory for retaining all entered data during any interruption of the power supply. Built-in test equipment circuits and software operate continuously to check all critical circuits of the system. The fuel summation unit that is a component of the Performance Data Computer System (PDCS,) develops an a.c. voltage signal that is proportional to the total fuel on board the aircraft; the signal being a combination of those produced by the fuel-quantity-indicating system sensing probes which are located in the fuel tanks.

Failure lights on the front of the computer indicate whether a fault is in the computer, CDU or input signals. The INDEX NUMBER switches, which are of the rotary type, are used for programming a flight index number from 0 to 200 into the computer so that maximum economy flight modes are modified according to time-related costs compared to fuel costs. The switches are guarded to eliminate the possibility of inadvertent changing of the index number.

PERFORMANCE DATA COMPUTER SYSTEM (TYPICAL SYSTEMS)

Performance Data Computer System (Typical Systems)

This system provides advisory data in alphanumeric format on a CRT display, in addition to the positioning of target command `bugs' on a Mach/airspeed indicator and EPR indicators, such indicators operating on electrical servomechanism principles. Provision is also made for interfacing the system with autothrottle and automatic flight control systems. A schematic diagram of the system .

Abbreviations are extensively used for the display of data by the control and display units of this and other flight management computer systems, and these abbreviations/ acronyms and their definitions are given in Table 2.2.

PERFORMANCE ADVISORY AND FLIGHT MANAGEMENT SYSTEMS

PERFORMANCE Advisory and Flight Management Systems

 Systems designed in various forms to carry out performance advisory or comprehensive flight management functions are now an essential feature of a number of types of commercial transport aircraft, their development having stemmed from the need to ensure the most efficient use of fuel, the need to reduce workload and the need to reduce operating costs overall. Fuel usage and other economic factors associated with aircraft operations have always been ones attracting the attention of the manufacturing and operating sectors of the industry, but in about the early 1970s when certain of the oil-producing states were creating sharp increases in the costs of crude oil and for political reasons were imposing oil embargos against some Western nations, the industry was forced to pay even greater attention to the above factors. As a result, many research and development programmes were instituted and were centred on the fuel efficiency of engines, improvements in LID ratios of airframes (e.g. by such means as use of supercritical wing sections) and on reductions in structural weight by use of composite materials.

Computer technology, although limited at the time in its application to aircraft systems, was nevertheless more advanced in overall concepts, and so by the production of software, which took into account the many operational variables, computers offered an additional and quicker route to the attainment of economic flight operations by performing automatic adjustments to relevant control systems. In conjunction with developments in the areas noted earlier, a system of computerized flight management has currently become the `elite' of avionic systems and it is probably not unfair to say that it is the most developed, besides being the one most readily retro-fitted to aircraft.

FMS development has, of course, resulted in a number of variations on the original theme of controlling engine power and flight operations consistent with the most efficient use of fuel at all times and, consequently, a variety of system designations has been applied by the manufacturers of” systems. These designations, some of which are interchangeable, while others indicate distinctly different capabilities, are given in Table 2.1 although Table 2.1 has not been compiled to a rigid scale of evolution, it does provide some indication of development of system functions which may be advisory only, or a combination of advisory and control

 Table 2.1 System designations

System                                                    Function

Performance advisory system (PAS)

           

- Advises of best altitude and speed to fly at to save fuel. Flight crew has to transfer values into automatic flight control system and throttle settings

*Performance data computer system

                                                                       

- Similar to PAS but typically linked (PDCS) to provide automatic pitch and throttle            control

*Performance management system (PMS)

- Similar to PDCS but with additional lateral navigation capability

*Automatic performance management            Similar to PAS

System (APMS)

*Flight management computer system (FMCS)        

- Full performance and navigation capabilities, flight planning and operation in a three-dimensional capacity

*Flight management system                               - Similar to FMCS (FMS)

In performing an advisory function a system merely advises the flight crew of the optimum settings of various control parameters, such as engine pressure ratio (EPR) and climb rate under varying flight conditions, in order to achieve the most economical use of the available fuel. Such systems require adjustments of controls on the part of the flight crew if they are to be utilized to maximum advantage. Examples of advisory systems are the PAS and PDC systems noted in Table 2.1

A system performing a combined function is one in which the sensing computer and display units are interfaced with an auto-throttle control system and pitch channel of an automatic flight control system; thus, in removing the flight crew from the control loop, an integrated automatic FMS is formed to provide greater precision of engine power and vertical flight path control. Early forms of flight management systems, whether purely advisory or combined function, were limited to supervising control parameters affecting the vertical flight path. In order to ensure maximum fuel economy it is, however, also necessary to integrate this optimized flight path management with the lateral flight path; in other words, a system must also be provided with a navigation capability. This requires interfacing the computer with such navigation systems as Doppler, inertial reference system, DME and VOR. The inputs from these systems permit continuous monitoring of an aircraft's track in relation to a flight plan, which may be pre-stored in the computer memory and an immediate identification of deviations. Furthermore, it allows flight plan variants to be constructed and evaluated. It is thus apparent that by combining these inputs with those controlling the vertical flight path parameters mentioned earlier, an FMS can integrate the functions of navigation, performance management, flight planning and three-dimensional guidance and control along a pre-planned flight path.

 Inputs: In addition to changing data inputs from such systems as those mentioned above, an FMS system also requires data bases for storing bulk navigation data, and the characteristics of an aircraft and its engines, in order that the system will operate in a full three-dimensional capacity. The navigation data base is capable of storing the necessary flight environmental data associated with a typical airline's entire route structure, including pertinent navigation aids and waypoints, airports and runways, published terminal area procedures, etc. The memory bank also contains flight profile data for a variety of situation modes, such as take-off, climb, cruise, descent, holding, go-around and 'engine -out'. The cruise mode is also sub-divided into sub-mode variants such as economy, long-range, manual and thrust-limited. The integration of all the foregoing data, plus other variable inputs such as wind speeds and air traffic control clearances, permit the automatic generation or modification of flight plans to meet the needs of any specific flight operation

MONITORING TECHNIQUES

Performance and Failure Assessment Monitor (PAFAM) System

This system is also one which uses a digital computer and a colour CRT display, its purpose being to operate in conjunction with an automatic flight guidance system (AFGS) to provide a flight crew with a prediction of the quality of an automatic approach and landing manoeuvre being carried out in low visibility. It monitors aircraft attitude, heading, and performance of the AFGS and makes a continual assessment of whether or not a successful automatic landing will result. In the event that the progress of the manoeuvre is unsuccessful, a `TAKEOVER' command is displayed; if the aircraft is being flown manually with commands from the flight director system, and the approach path is unacceptable, the legend NO TRACK is displayed. A block diagram of the system  necessary for proper operation of the AFGS and auto throttle/speed control system. Electrical power is applied when the AFGS LAND ARM mode or flight director ILS modes of operation are selected, and the system is automatically switched to its operational condition when the ILS localizer and glideslope are being tracked.

The signal inputs to the computer are a.c. and d.c. analog and are multiplexed into an A/D converter which is under programmed  memory control by one of two control processors in the computer; this processor performs most of the landing performance and prediction computations. Discrete signal inputs are multiplexed directly into the second processor, which provides display drive commands, landing system failure assessment, and controls signals for discrete outputs. Interconnection between the two processors is through two 18-bit storage registers.

 Analog signals from the computer are applied to the display electronics unit, and they provide commands for blanking out a portion of two raster-scanned CRT display units (one for each pilot) as well as commands which determine the location of desired characters in the display. The location of a display unit is shown in  the viewing area of the CRT is 38 mm x 76 mm. Discrete signal outputs are supplied to the AFGS and auto throttle/speed control system.

The digital signal outputs from the computer are applied to timing and logic circuits in the display electronics unit for the development of analog character signals via fixed memory circuits in a symbol generator. The character signals are amplified by horizontal and vertical summing amplifiers, and then fed to deflection amplifier and blanking circuits so that desired symbols and words are `painted' on the CRT screen. A colour control logic circuit supplies the CRT with a command signal which varies the level of a high - voltage supply so as to vary the colour. As the voltage is increased in selected steps, the colour of the character or raster being generated changes from red to red-orange, to amber, to yellow, and finally to green.

 Operation

When the LAND or ILS mode logic is available from the AFGS computers, a TEST mode display first comes into view

 . Then, after a very short time period, a raster pattern is displayed on the CRTs representing the airport runway over which is superimposed a cross depicting the predicted touchdown point, . The runway symbol has a yellow border within which there is a green area representing the center touchdown zone. The area corresponds approximately to a zone ±18 m (60 ft) laterally and ±300 m (1000 ft) longitudinally about the nominal or ideal touchdown point, which is on the centreline and approximately 150 m (500 ft) beyond the glideslope transmitter location. The horizontal line closest to the bottom of the display corresponds to the runway threshold.

The symbol `expands' by moving downwards as a function of computed range in a manner corresponding to the same rate of expansion that would be apparent if the real runway were visible out the flight deck windscreens. As the aircraft crosses the runway threshold, the bottom line of the display moves out of view, and on the basis of range computation, the lower edge of the green area reaches the bottom of the display when the aircraft passes over the nominal touchdown point. The green area continues to move down until touchdown. At touchdown, a green downward-pointed triangle is displayed on the CRT screen (Figure 2.3(c)) and the display then remains static for about three seconds, after which the whole system is de - energized.

The orange-coloured cross symbol is positioned to show the predicted touchdown point on the runway as determined by simplified dynamic models of the aircraft/AFGS combination operated in an accelerated time scale. In an ideal landing situation, the cross will be superimposed over the touchdown area, with the intersection of the arms coincident with the centre of the area. The arms of the cross represent the uncertainty of the basic touchdown prediction (e.g., the effect of winds, ILS beam anomalies, and acceptability of AFGS responses), the uncertainty being reflected by variations in the lengths of the arms, and deflection of the cross from the centre of the displayed runway area. All parts of the cross should remain in this area for an acceptable landing.

If the system detects such uncertainty of performance that an approach could be seriously impaired, then the message TAKEOVER or NO TRACK is displayed in red letters on each pilot's display unit, together with a yellow arrow symbol in one of the four corners to indicate which of the two pilots should perform a manual go-around procedure, or a continued manual landing. The arrow appears in the left or right corners depending on which pilot the more valid information (Figure 2.3(d)) has.

In the event of failure of either of the ILS ground transmitters (localizer or glideslope), the loss of the desired guidance signal information is detected by the PAFAM system and the message NO ILS is displayed in red (and flashing) on the indicator units (Figure 2.3(f)). Normally under these circumstances the AFGS would automatically disengage, but as the PAFAM system uses other references to supply equivalent ILS deviation for assessing performance, then if no deterioration in performance is detected, disengagement of the AFGS is inhibited for up to 5s. If, during this interval, ILS signals are restored, the PAFAM system reverts to the normal display mode and the landing continues uninterrupted. If the signals are not restored, or if the performance is affected, the AFGS would then disengage and the NO ILS message in the display units would automatically change to TAKEOVER.

 An automatic self-testing function is built into the system and checks the whole of its operation twice each second. Any malfunction causes latching-type annunciators to trip, and gross distortions or blanking out of symbols on the display units.

ON BOARD MAINTENANCE SYSTEM (OBMS)

ON BOARD MAINTENANCE SYSTEM (OBMS)

This is a system found in modern larger commercial aircraft that integrates all required information forming the basis of generating maintenance guidance, trouble shooting information on faults.

System has a panel, called Maintenance Panel, normally located behind the first officer (also called Lateral Panel). The complete lateral/maintenance panel is divided into different system panels. Each system maintenance panel has dedicated bite/test knobs, system instruments, anunciator lights.

OBMS computers centralizes data from system sensors, processes those data to find faults/failure ranges and keep the fault in non-volatile memory. Centrally there displays a BITE light whenever there is a FAULT light in the local panel of any system. This display of BITE light in the OBMS panel, in conjunction with FAULT light in the local panel means that the fault has been codified by the OBMS and stored in the BITE software. Now, when BITE switches in the corresponding maintenance panel is to be positioned to the test and BITE position. DBMS computer will then send codes on the dedicated CRTs or ECAM CRTs. The codes are to be taken out to decode and follow the trouble shooting manual.

This panel is primarily intended to be used by maintenance personnel as an assistance in the performance of system maintenance. Normally only component condition memory indicator lights are installed. The memory indicators are illuminated if the maintenance panel switch ANN LT is in the READ position and if a signal has been stored.

For example, in a typical engine with FADEC system, there is a FADEC BITE panel in the OBMS panel (Figure 1.10). When ever there is a FADEC fault, OBMS BITE light comes on, meaning that a fault on FADEC has been stored in the NVM. To trouble shoot the fault (including normal test), FADEC BITE test is carried out.

Placing the BITE switch of any engine to test position, status data received by the EEC of FADEC system is transmitted to the ECAM CRTs. Placing the BITE switch to the BITE Display position causes maintenance words in hexadecimal code on the L CRT (Figure 1.11, 1.12).

Hexadecimal words are taken to convert into label-bit through conversion chart and ultimately fault code will be obtained that helps tracing a trouble shooting tree in the trouble shooting manual.

ENGINE INDICATING AND CREW ALERTING SYSTEM (EICAS)

ENGINE INDICATING AND CREW ALERTING SYSTEM (EICAS)

Basic system of EICAS is the same as ECAM with the same philosophy of system designs and system architecture. It covers engine area and computers are installed for data acquisition, processing, warning generation, normal display and failure display along with procedures.

1.7 ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS)

This is centralized system for flight path and navigational area and electronic display systems are designed appropriate to it.

The EFIS enables display for the crew of flight path control and navigation data.

In a typical aircraft, EFIS system provides two CRTs at disposal of the Captain, two CRTs for the pilot on which is displayed all the information that is usually displayed by the following instruments:

- Attitude Director Indicator

- Horizontal situation indicator

- Mach-Airspeed Indicator

- Radio Altimeter Indicator

- Weather Radar Indicator

- Flight Mode Annunciator.

Thanks to the high level of flexibility of the CRTs, the EFIS makes it possible to display, at a given moment, only the information required by the present flight phase.

The upper instrument displays the information necessary for short-term aircraft control, hence the name: Primary Flight Display (PFD), that is, aircraft attitude, computed airspeed (VC) and Mach number, altitude and course deviations etc.

The lower instrument Display (ND), displays the navigation information necessary for the flight phase: rose, flight plan, radar image.

Like ECAM system, EFIS system display is generated by SGUs. SGUs receive input from SDAC and FWC. These SGUs, SDC and FWCs are also used by the ECAM. 

LINE MAINTENANCE HELP

The system offers facilities for fault diagnosis from the monitored data into its BITE circuit. A Maintenance Test Panel (MTP) is common in many aircraft for this purpose. A typical MTP is illustrated in Figure 1.2.

 SYSTEM INTEGRATION

The avionics system monitoring computers interface with many other computers and peripherals in the aircraft for exchange of avionics data. The aircraft AIDS system is interfaced with the system. The AIDS data is used for aircraft reliability monitoring, engine trend monitoring and so on.

Self-test of the AFS and peripherical systems

• Self-test of the AFS and peripherical systems:

- Most of the systems or units have their own built-in self-test,

- For the AFS they consist in power-up and AFS tests,

- The purpose of the self-tests is to check the integrity of the hardware within each computer and the integrity of most of the LRUs without BITE but in direct connection with the computer (sensors, actuators, switches...),

- These tests do not rely on the control laws and logics of the avionics program but are solely turned towards the components integrity.

For the AFS when power up and AFS test are successful it is considered that all AFS safety devices are operative and that nearly all components within the AFS are healthy.

1.3 IN FLIGHT HELP

Performance and Failure Assessment Monitor (PAFAM) System uses a digital computer and a colour CRT display, its purpose being to operate in conjunction with an automatic flight guidance system (AFGS) to provide a flight crew with a prediction of the quality of an automatic approach and landing manoeuvre being carried out in low visibility. It monitors aircraft attitude, heading, and performance of the AFGS and makes a continual assessment of whether or not a successful automatic landing will result. In the event that the progress of the manoeuvre is unsuccessful, a `TAKEOVER' command is displayed; if the aircraft is being flown manually with commands from the flight director system, and the approach path is unacceptable, the legend NO TRACK is displayed.

The computer accepts electrical input signals from those sensors and sub-systems necessary for proper operation of the AFGS and auto throttle/speed control system. Electrical power is applied when the AFGS LAND ARM mode or flight director ILS modes of operation are selected, and the system is automatically switched to its operational condition when the ILS localizer and glide slope are being tracked.

The signal inputs to the computer are a.c. and d.c. analog and are multiplexed into an A/D converter which is under programmed memory control by one of two control processors in the computer; this processor performs most of the landing performance and prediction computations. Discrete signal inputs are multiplexed directly into the second processor, which provides display drive commands, landing system failure assessment, and controls signals for discrete outputs. Interconnection between the two processors is through two 18-bit storage registers.

Analog signals from the computer are applied to the display electronics unit, and they provide commands for blanking out a portion of two raster-scanned CRT display units (one for each pilot) as well as commands which determine the location of desired characters in the display. 

SCOPE OF MONITORING AND FAILURE ANALYSIS

SCOPE OF MONITORING AND FAILURE ANALYSIS

Monitoring and Failure analysis system has the scope of doing the following:

• AFS Trouble Shooting: AFS monitors system and memorizes fault into its non-volatile memory which can be retrieved and decoded step by step. This is a great help for the flight line maintenance of avionics systems. Principle of the AFS Fault Isolation and Detection System (FIDS) is:

 Whenever a basic monitoring condition becomes invalid the associated AFS lever trips,

 The FIDS then takes a snapshot of all parameters and starts its analysis to identify the invalid signal

From the invalid signal, and using its own cross-reference table, the FIDS designates and memorizes of its own the system or unit generating this signal.

- In-service experience has shown that the FIDS of the AFS were able to correctly identify nearly all systems when associated levers have tripped (invalid signal generator).

- In some cases and based on experience or the surrounding context, tripping is considered by the crew as normal and therefore not reported in the log book.

- However the FIDS do not ignore it and a message will be memorized by the MTP when no action is in fact required from the maintenance team.

BENEFITS OF MONITORING

 Benefits OF monitoring

Monitoring system of the AFS fulfills the following functions:

• Fault isolation: Data are accessible in flight and on the ground by pressing the DISPLAY pushbutton switch on the Maintenance Test Panel (MTP). It is to be noted that this pushbutton switch is normally active in flight in many aircraft.

• Test after replacement of a unit: This test is performed on the ground by ground maintenance personnel in order to check correct operation of a system after replacement of a unit. This test called AFS TEST. This test is inhibited in flight.

• LAND mode test: This test is performed on the ground only (inhibited in flight), to check availability of CAT 3 capability. If aircraft is operated in actual CAT 3, this test is recommended after removal/installation of the many AFS components.

• Complex fault isolation: This function is activated on the ground in the event of complex failure detection.