Civil Aviation Order 100.5 (General requirements in respect of maintenance of Australian aircraft) 2011
21Glidescope in I.F.R. — periodic checking
Start here
Get a plain-English read of 21
Turn the raw legal text into a practical explanation grounded in Civil Aviation Order 100.5 (General requirements in respect of maintenance of Australian aircraft) 2011.
21.1This clause applies only for the following navigation system of an Australian aircraft for which the holder of the certificate of registration has elected to use the CASA maintenance schedule: a glidescope system installed in an I.F.R. aircraft.
21.2Check and ensure that each of the following requirements is the case:
(a) the deviation indicator must centre within less than ½ (± 12 microamperes) when the tone ratio is 0 dB (ddm = 0);
(b) the sensitivity must be such that the deviation indicator deflection is 4.3 dots 0.7 dots (110 to 150 microamperes) for both up and down deflections when using a signal with 3.3 dB tone ratios or 2.7 dots ± 0.3 dot (68 to 93 microamperes), when using a 2 dB tone ratio, and the deflection must be in the correct sense;
(c) the flag must remain concealed during each of the tests mentioned in paragraphs (a) and (b), however, when the signal input level is reduced so as to cause the deviation indicator deflection to fall by 2 dots, the flag must be at least partly visible.
Note 1 A simulator may be used for the checking mentioned in paragraphs 21.2 (a) to (c).
Note 2 The sensitivities referred to in this subclause are related to a “standard” 5 dot-0-5 dot, 150 microamperes-0-150 microamperes indicator. Proportional deflections apply to other than “standard” indicators.
21.3The checks mentioned in this clause must be carried out each time a periodic inspection is carried out in accordance with paragraphs 2.4 and 2.5 in Part 2 of Schedule 5 of CAR.
Attachment 1 to Appendix 1
Note See subclauses 3.3, 4.2 and 4.6 in Appendix 1.
Testing procedures for:
Pitot-static systems
Pressure altimeter systems
Air data computers
Automatic pressure altitude encoders
Other transponder devices
Clause 1 Pitot-static system test
(1)Static pressure systems
Performance of the test procedures set out below, with all static instruments connected, must ensure that any leakage present is within the tolerance specified for the procedure.
(a) Visually inspect the ports, plumbing, accessories and instruments connected to the static system. Repair or replace those parts which are defective, for example, broken “B” nuts, cracked flare sleeves, deteriorated flexible tubing and quick disconnects, bad valves etc. If purging is necessary, use compressed air or nitrogen to remove foreign matter which may have accumulated in the tubing. Ensure that all static instruments are disconnected before commencing to purge.
Section 21 - Civil Aviation Order 100.5 (General requirements in respect of maintenance of Australian aircraft) 2011 - Glidescope in I.F.R. — periodic checking | Zoe
(b) Ensure that no alterations or deformations of the airframe surface are present that would affect static air sensing. This is of particular importance for RVSM aircraft.
(c) Check any static port heaters to assure proper operation.
(d) If an aircraft has more than 1 static system, test each system separately to assure its independence and that the leak rate for each system is within tolerance.
(e) Connect the test equipment directly to the static ports, if practicable. If not practicable, connect to a static system drain or tee connection and seal off the static ports. If the test equipment is connected to the static system at any point other than the static port, it must be made at a point where the connection may be readily inspected for system integrity after the system is returned to its normal configuration.
(f) Determine that any leakage is within the tolerances mentioned in paragraph (g) or (h) (as the case requires).
(g) For unpressurised airplanes — evacuate the static pressure system to a pressure differential of approximately 33 hPa or to a reading on the altimeter that is 1 000 feet above the aircraft’s elevation at the time of the test. Without additional pumping for a period of 1 minute, the loss of indicated altitude must not exceed 100 feet on the altimeter.
(h) For pressurised airplanes — evacuate the static pressure system until a pressure differential equivalent to the maximum cabin pressure differential for which the airplane is type certificated is achieved. Without additional pumping for a period of 1 minute, the loss of indicated altitude must not exceed 2 per cent of the equivalent altitude of the maximum cabin differential pressure or 100 feet, whichever is greater.
(i) On completion of the static pressure system test, ensure that all static port seals are removed.
(2)Pitot-systems
(a) The pitot system is tested for leaks by applying a pressure at the pitot head sufficient to cause the airspeed indicator to read 120 knots, or the maximum indicated speed, whichever is the greater.
(b) There must be no discernible lag in the movement of the airspeed indicator pointer with the application of the pressure, as such a lag indicates restrictions in the piping.
(c) There must be no decrease in the reading when the system is sealed for at least 10 seconds.
Clause 2 Tests for altimeters and air data computers
Note For testing of air data computers, see subclause 2 (8) below.
(1)Environmental conditions test
(a) Vibration (intended to minimise the effects of friction). If suitable test equipment is available, each test for performance may be conducted with the instrument installed in the aircraft. If suitable test equipment for an installed test is not available, or if the instrument fails the installed test, the instrument must be removed from the aircraft and tested or retested with vibration applied.
(b) Temperature. When tests are conducted with the temperature substantially different from ambient temperature of approximately 25°C, allowance must be made for that temperature difference.
(2)Scale error test
(a) With the barometric pressure scale at 1 013 hPa, the altimeter must be successively subjected to pressures corresponding to the altitude specified in Table 1 up to the maximum, normally expected, operating altitude of the aircraft in which the altimeter is, or is to be, installed.
(b) The reduction in pressure must be made at a rate not in excess of 20 000 feet per minute to within approximately 2 000 feet of the test point.
(c) The test point must be approached at a rate compatible with the test equipment.
(d) The altimeter must be kept at the pressure corresponding to each test point for at least 1 minute, but not more than 10 minutes, before a reading is taken.
(e) The error at all test points must not exceed the tolerances specified in Table 1.
(a) The hysteresis test must begin within 15 minutes of the altimeter’s initial exposure to the pressure corresponding to the upper limit of the scale error test in subclause (2). While the altimeter is at this pressure, the hysteresis test is to commence.
(b) Pressure must be increased at a rate simulating a descent in altitude at the rate of 5 000 to 20 000 feet per minute until within 3 000 feet of the first test point (50 per cent of maximum altitude).
(c) The test point is then to be approached at a rate of approximately 3 000 feet per minute. The altimeter must be kept at this pressure for at least 5 minutes, but not more than 15 minutes, before the test reading is taken.
(d) After the reading has been taken, the pressure must be increased further, in the same manner as before, until the pressure corresponding to the second test point (40 per cent of maximum altitude) is reached. The altimeter must be kept at this pressure for at least 1 minute, but not more than 10 minutes, before the test reading is taken.
(e) After the reading has been taken, the pressure must be increased further, in the same manner as before, until atmospheric pressure is reached.
(f) The reading of the altimeter at either of the 2 test points must not differ by more than the tolerance specified in Table 2 in subclause 2 (4) from the reading of the altimeter for the corresponding altitude recorded during the scale error test prescribed in subclause (2).
(4)After effect test
Within 5 minutes following the completion of the hysteresis test set out in subclause (3), the reading of the altimeter (corrected for any change in atmospheric pressure) must not differ from the original atmospheric pressure reading by more than the tolerance specified in Table 2.
Table 2 Test tolerances
Test
Tolerance (feet)
Case leak test
±100
Hysteresis test:
First test point (50 per cent of maximum altitude)
75
Second test point (40 per cent of maximum altitude)
75
After effect test
30
(5)Friction test
(a) The altimeter is to be subjected to a steady rate of decrease of pressure approximating 750 feet per minute.
(b) At each altitude listed in Table 3, the change in reading of the pointers after vibration (using a light tapping of the instrument panel adjacent to the altimeter if the altimeter does not have an integral vibrator) must not exceed the corresponding tolerance listed in Table 3.
(c) If the altimeter fails the friction test while installed on the aircraft, the altimeter must be removed and retested.
Table 3 Friction
Altitude (feet)
Tolerance(feet)
1 000 -
±70
2 000 -
70
3 000
70
5 000
70
10 000
80
15 000
90
20 000
100
25 000
120
30 000
140
35 000
160
40 000
180
50 000
250
(6)Case leak test
The leakage of the altimeter case, when the pressure within it corresponds to an altitude of 18 000 feet, must not change the altimeter reading by more than the tolerance shown in Table 2 in subclause 2 (4) during an interval of 1 minute.
(7)Barometric scale error test
At constant atmospheric pressure, the barometric pressure scale must be set at each of the pressures (falling within its range of adjustment) that are listed in Table 4, and this must cause the pointer to indicate the equivalent altitude difference shown in Table 4 within a tolerance of plus or minus 25 feet.
Table 4 Pressure-altitude difference
Pressure(hectopascal)
Altitude difference(feet)
952
-1 727
965
-1 340
982
-863
999
-392
1013
0
1033
+531
1046
+893
1049
+974
(8)Air data computers test
(a) The tests set out in subclauses (1) to (7) do not apply for air data computers or for systems similar to air data computers (a similar type).
(b) Paragraph (c) sets out the tests for the following altimeters:
(i) an altimeter that is an air data computer or similar type with associated computing systems;
(ii) an altimeter that incorporates air data correction internally.
(c) An altimeter mentioned in paragraph (b), must be tested as follows:
(i) in the manner, and to the specifications, provided by the manufacturer of the equipment or aircraft in which the altimeter is installed;
(ii) in accordance with the instructions for continuing airworthiness incorporated in a modification approval for the equipment or aircraft in which the altimeter is installed.
Clause 3 Automatic pressure altitude encoders and ATC transponder system integration test
Automatic pressure altitude encoder test and other transponder devices
Measure the automatic pressure altitude value at the output of the installed ATC transponder when interrogated on Mode C at a sufficient number of test points to ensure that the altitude reporting equipment, altimeters, and ATC transponders perform their intended functions as installed in the aircraft. The difference between the automatic reporting output and the altitude displayed at each altimeter must not exceed 125 feet.
Appendix 2
Prescribed single-engine aeroplanes (PSEA)
Note Paragraph 10.1 of this CAO requires an aeroplane operated as a PSEA to comply with each of the following requirements. Aircraft systems and equipment mentioned within these requirements must be approved under regulation 21.305 of CASR.