d. To Correct Cylinder Head or Other Temperatures to Anticipated Hot-Day Conditions.
(1) Corrected temperature (°F) = true temperature + 1.0 [100 - 0.0036 (pressure altitude) - true outside air temperature].
Corrected temperature (°C) = true temperature + 1.0[38-0.002(pressure altitude, ft) – true outside air temperature, °C].
(2) For example (using metric units):
True maximum cylinder head temperature -- 208 °C.
Pressure Altitude -- 8330 feet True OAT -- +13 °C.
(3) Corrected cylinder head temperature = 208 + 1.0 [38 - 0.002 (8330) -13] = 216 °C.
(4) The corrected temperatures are then compared with the maximum permissible temperatures to determine compliance with cooling requirements.
e. Liquid Cooled Engines. (RESERVED)
249.-254. RESERVED.
Section 6. INDUCTION SYSTEM
255. SECTION 23.1091 AIR INDUCTION.
AC 20-124 covers the turbine engine water ingestion aspects of this requirement.
256. SECTION 23.1093 INDUCTION SYSTEM ICING PROTECTION.
a. Explanation.
(1) Purpose. Tests of engine induction system icing protection provisions are conducted to ensure that the engine is able to operate throughout its flight power range without any adverse effect on engine operation. Reciprocating engines utilize a preheater or a sheltered alternate air source to provide adequate heat rise to prevent or eliminate ice formation in the engine induction system. The adequacy of this heat rise is evaluated during the test. The amount of heat available is determined by measuring the intake heat rise by temperature measurements of the air before it enters the carburetor. Turbine engine inlet ducts must be protected to prevent the accumulation of ice as specified in § 23.1093(b)(1).
(2) Reciprocating Sea Level Engine Configurations.
(i) Venturi Carburetor. Section 23.1093(a)(1) requires a 50 °C (90 °F) heat rise at 75 percent maximum continuous power at –1 °C (30 °F) OAT.
(ii) Single-Engine Airplanes With a Carburetor Tending to Prevent Icing (Pressure Carburetor). Section 23.1093(a)(5) requires an alternate air source with a temperature equal to that of the air downstream of the cylinders.
(iii) Multiengine Airplane With Carburetor Tending to Prevent Icing (Pressure Carburetor). Section 23.1093(a)(5) requires a 50 °C (90 °F) heat rise at 75 percent maximum continuous power at –1 °C (30 °F) OAT.
(iv) Fuel Injection With Ram Air Tubes. A heat rise of 50 °C (90 °F) at 75 percent maximum continuous power is recommended.
(v) Fuel Injection Without Projections Into the Induction Air Flow. An alternate air source with a temperature not less than the cylinder downstream air is recommended.
(3) Reciprocating Altitude Engine Configurations.
(i) Venturi Carburetor. Section 23.1093(a)(2) requires a 67 °C (120 °F) heat rise at 75 percent maximum continuous power at –1 °C (30 °F) OAT.
(ii) Carburetors Tending to Prevent Icing (Pressure Carburetor). Section 23.1093(a)(3) requires a heat rise of 100 °F at 60 percent maximum continuous power at –1 °C (30 °F) OAT or 22 °C (40 °F) heat rise if an approved fluid de-icing system is used.
(iii) Fuel Injection. Same as for sea level fuel injected engines.
(4) Turbine Engines. Section 23.1093(b) requires turbine engines to be capable of operating without adverse effects on operation or serious loss of power or thrust under the icing conditions specified in Part 25, Appendix C. The powerplant should be protected from ice at all times, whether or not the airplane is certificated for flight into known icing conditions.
b. Reciprocating Engine Test Considerations.
(1) Visible Moisture. The tests should be conducted in air free of visible moisture.
(2) Instrumentation. All instruments used during the test should be calibrated and all calibration curves made part of the Type Inspection Report. Calibrations should be made of complete systems as installed and shall cover the temperature range expected during the tests.
(3) Heat Rise. All carburetor air heat rise requirements should be met at an outside air temperature of –1 °C (30 °F). If the test cannot be conducted in an atmosphere with an ambient air temperature of –1 °C (30 °F), it will normally be flown at low, intermediate, and high altitudes. If a –1 °C (30 °F) day exists at an altitude where 75 percent of rated power is available, only one test is necessary.
(4) Intake Air. Care should be exercised to assure that the method of measuring the temperature of the air will give an indication of the average temperature of the airflow through the intake and not just a stratum of air. This may be accomplished by temperature measurements of the intake air at several points. Usually, the temperate probe is placed at the carburetor deck. Other temperature probe locations may be acceptable, provided that the temperatures measured can be correlated with the air temperature at the carburetor deck.
Care should be taken not to measure the temperature of a stratum of air, but the average temperature of the air entering the carburetor. Temperature probes placed in the carburetor throat, or venturi, are not acceptable for the carburetor air heat rise test. Temperature probe locations should be specified in any test plan for determining induction system ice protection.
c. Test Procedures for Reciprocating Engine Airplanes.
(1) At low altitude, stabilize airplane with full throttle or, if the engine is supercharged, with maximum continuous power on the test engine. With carburetor air heat control in the "cold" position, record data. Manually operated turbochargers should be off. For integrally turbocharged engines, heat rise data should be taken at lowest altitude conditions, where the turbo is providing minimum output.
(2) Apply carburetor heat and, after condition stabilizes, record data.
(3) Reduce airspeed to 90 percent of that attained under item (1). With carburetor air heat control in the "cold" position and condition stabilized, record data.
(4) Apply carburetor heat and, after condition stabilizes, record data.
(5) Reduce airspeed to 80 percent of that attained under item (1). With carburetor air heat control in the "cold" position and condition stabilized, record data.
(6) Apply carburetor heat and, after condition stabilizes, record data.
(7) At the intermediate altitude, repeat steps (1) through (6).
(8) At high altitude, repeat steps (1) through (6). Data to be recorded:
(i) Altitude (feet);
(ii) Airspeed (IAS) (Knots);
(iii) Ambient air temperature °F;
(iv) Carburetor air temperature °F;
(v) Carburetor heat control position;
(vi) Engine r.p.m.;
(vii) Engine manifold pressure (in Hg); and
(viii) Throttle position.
d. Data Reduction. Figures 256-1 and 256-2 show sample carburetor air heat rise determinations.
e. Test Procedures for Turbine Engine-Powered Airplanes. Tests to determine the capability of the turbine engine to operate throughout its flight power range without adverse effect on engine operation or serious loss of power or thrust should be conducted to encompass the icing conditions specified in Part 25, Appendix C. Each airplane should be evaluated for compliance.
Thermodynamic exercises and dry air tests alone are not usually adequate, and actual icing encounters or wind tunnel testing are necessary.
(752) 80% IAS = Column 2525 2310 17.2 18.0 - 878 40.8 P MAXIMUM ALTITUDE 90% IAS of Colummn 185 2480 2665 19.4 19.9 101 .871 98 78.4 56.0 Full Throttle Power* 2525 20.4 100 - 868 62.1 or MC 2770 21.2 113 .970 ALTITUDE 80% IAS of Column #1 73 77 2280 2410 18.5 19.0 72 76 . 882 . 970 63.5 45.3 INTERMEDIATE 90% TAS of Column #l 72 73 68T 78 2555 2400 19.6 85 .879 970 74.7 53.4 201 2640 22.8 114 . 870* 99.2 71.0 Full of MC Power* 5000 72 73 96 2800 23.5 125 970 121 80% IAS of Column #1 002 +78 2430 2310 21.3 22.0 105 882 .972 87.4 102 62.4 (°F) + 460 ate data plotted on figure 256-2.
MINIMUM 90% IAS of Column #1 SOZ 2690 2590 23.5 24.0 112 120 .879 . 972 98.4 70.2 Only Engines Thrott [* SIZ (132) 2730 25.7 132 .872 115 82.2 *Supercharged NOTE 1: Temperature Correction Factor Full 105 2850 26.4 144 .972 140 flown be May only Air tude (ft.) Pressure Alti Heat Rise (м. P.н.) Indicated B.H.P. for Temperature Pressure Altitude Temperature Correc- note 1) tion Factor (See Actual B.H.P. 7 Rated B.H.P. (See note 2) Throttle Position Rated BHP • 140 NOTE 2: NOTE 3: Circled numbers in‹ FIGURE 256-1 - CARBURETOR AIR HEAT RISE CALCULATIONS 125
100 90 80 70 60 50 90 80 70 60 50 40 30 20 110 120 130 140 150 CARBURETOR AIR HEAT RISE - 0F
PERCENT
OF RATED BHP
OUTSIDE AIR
TEMPERATURE
- 0F 140 0F CARB. AIR HEAT RISE AT 30 0F
OUTSIDE AIR TEMP 75% BHP PRESS. ALT.
1500 FT PRESS. ALT.
5000 FT PRESS. ALT.
8000 FT
FIGURE 256-2 - CARBURETOR AIR HEAT RISE PLOTS
257.-265. RESERVED.
Section 7. POWERPLANT CONTROLS AND ACCESSORIES
266. SECTION 23.1141 POWERPLANT CONTROLS: GENERAL.
a. Explanation. Powerplant controls for each powerplant function will be grouped for each engine allowing simultaneous or independent operation as desired. Each control will be clearly marked as to function and control position. (Also see § 23.777). Controls are required to maintain any position set by the pilot without tendency to creep due to vibration or control loads.
b. Procedures. None.
267. SECTION 23.1145 IGNITION SWITCHES. (RESERVED)
268. SECTION 23.1153 PROPELLER FEATHERING CONTROLS.
a. Explanation. If the propeller pitch or speed control lever also controls the propeller feathering control, some means are required to prevent inadvertent movement to the feathering position.
b. Procedures. None.
269.-278. RESERVED.
Section 8. POWERPLANT FIRE PROTECTION
279. SECTION 23.1189 SHUTOFF MEANS.
a. Explanation. The location and operation of any required shutoff means is substantiated by analysis of design data, inspection, or test. The location and guarding of the control (switch), the location and clarity of any required indicators, and the ability to operate the controls with the shoulder harnesses locked (if applicable) should be evaluated.
b. Procedures. Control locations and guarding and indicator effectiveness should be part of the complete cockpit evaluation. Check the shutoff means function by performing an after-flight engine shutdown using the fuel shutoff.
280.-285. RESERVED.
CHAPTER 5. EQUIPMENT
Section 1 GENERAL
286. RESERVED
287. SECTION 23.1301 FUNCTION AND INSTALLATION.
a. Explanation. Section 23.1301 gives specific installation requirements. Particular attention should be given to those installations where an external piece of equipment could affect the flight characteristics. All installations of this nature should be evaluated by the flight test pilot to verify that the equipment functions properly when installed.
b. Avionics Test.
(1) Very High Frequency VHF Communication Systems. See AC 20-67B. AC 20-67B references Radio Technical Commission for Aeronautics (RTCA) document DO-186. DO- 186, paragraph 3.4.2.3 speaks to ground facility coverage area. FAA Order 6050.32, Appendix 2, shows the coverage limits for various facility parameters. Contact the nearest FAA Airway Facilities Sector Office to examine the order.
(2) High Frequency (HF) Communication Systems.
(i) Ground Station Contacts. Acceptable communication should be demonstrated by contacting a ground station on as wide a range of frequencies as HF propagation conditions allow. Distances may vary from 100 to several hundred nautical miles (n.m.).
The system should perform satisfactorily in its design modes.
(ii) Precipitation Static. It should be demonstrated that precipitation static is not excessive when the airplane is flying at cruise speed (in areas of high electrical activity, including clouds and rain if possible). Use the minimum amount of installed dischargers for which approval is sought.
(iii) Electromagnetic Compatibility (EMC). Electromagnetic compatibility tests should be conducted on the ground and in flight at 1.0 MHz intervals. Any electromagnetic interference (EMI) noted on the ground should be repeated in flight at the frequency at which the EMI occurred on the ground. Since squat switches may isolate some systems from operation on the ground (that is, air data system. pressurization, and so forth), EMI should be evaluated with all systems operating in flight to verify that no adverse effects are present in the engine, fuel control computer, brake antiskid, and so forth, systems.
(3) Very High Frequency Omnirange (VOR) Systems.
(i) Antenna Radiation Patterns. These flight tests may be reduced if adequate antenna radiation pattern studies have been made and these studies show the patterns to be without significant holes (with the airplane configuration used in flight; that is, flaps, landing gear, and so forth). Particular note should be made in recognition that certain propeller r.p.m. settings may cause modulation of the course deviation indication (propmodulation). This information should be made a part of the AFM.
(A) Reception. The airborne VOR system should operate normally with warning flags out of view at all headings of the airplane (wings level) throughout the standard service volumes depicted in the Airman's Information Manual (AIM) up to the maximum altitude for which the airplane is certified.
(B) Accuracy. The accuracy determination should be made such that the indicated reciprocal agrees within 2 degrees. Tests should be conducted over at least two known points on the ground such that data are obtained in each quadrant. Data should correlate with the ground calibration, and in no case should the absolute error exceed ±6 degrees. There should be no excessive fluctuation in the course deviation indications.
(ii) En Route Reception. Fly from a VOR facility rated for high altitude along a radial at an altitude of 90 percent of the airplane's maximum certificated altitude to the standard service volume range. The VOR warning flag should not come into view, nor should there be deterioration of the station identification signal. The course width should be 20 degrees ±5 degrees (10 degrees either side at the selected radial). The tests should be flown along published route segments to preclude ground station anomalies. If practical, perform an en route segment on a Doppler VOR station to verify the compatibility of the airborne unit. Large errors have been found when incompatibility exists. Contact the nearest FAA Airway Facilities Sector Office to locate a Doppler VOR.
(iii) Low-Angle Reception. Perform a 360 degree right and 360 degree left turn at a bank angle of at least 10 degrees at an altitude just above the lowest edge of the standard service volume and at the maximum service volume distance. Signal dropout should not occur as evidenced by the warning flag appearance. Dropouts that are relieved by a reduction of bank angle at the same relative heading to the station are satisfactory. The VOR identification should be satisfactory during the left and right turns.
(iv) High-Angle Reception. Repeat the turns described in (c) above, but a distance of 50- 70 n.m. (20-30 n.m. for airplanes not to be operated above 18,000 feet) from the VOR facility and at an altitude of at least 90 percent of the maximum certificated altitude of the airplane.
(v) En Route Station Passage. Verify that the to-from indicator correctly changes as the airplane passes through the cone of confusion above a VOR facility.
(vi) VOR Approach. Conduct VOR approaches with gear and flaps down. With the facility 12-15 n.m. behind the airplane, use sufficient maneuvering in the approach to ensure the signal reception is maintained during beam tracking.
(vii) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(4) Localizer Systems.
(i) Antenna Radiation Patterns. Flight test requirements should be modified to allow for adequate antenna radiation pattern measurements as discussed in VOR systems, subparagraph (3)(a).
(A) Signal Strength. The input to the receiver, presented by the antenna system, should be of sufficient strength to keep the malfunction indicator out of view when the airplane is in the approach configuration (landing gear extended - approach flaps) and within the normal limits of localizer coverage shown in the Airman's Information 129
Manual (AIM). This signal should be received for 360 degrees of the airplane heading at all bank angles up to 10 degrees left or right at all normal pitch attitudes and at an altitude of approximately 2,000 feet (see RTCA Document D-102).
(B) Bank Angles. Satisfactory results should also be obtained at bank angles up to 30 degrees when the airplane heading is within 60 degrees of the inbound localizer course. Satisfactory results should result with bank angles up to 15 degrees on headings from 60 degrees to 90 degrees of the localizer inbound course and up to 10 degrees bank angle on headings for 90 degrees to 180 degrees from the localizer inbound course.
(C) Course Deviation Indicator (CDI). The deviation indicator should properly direct the airplane back to course when the airplane is right or left of course.
(D) Station Identification. The station identification signal should be of adequate strength and sufficiently free from interference to provide positive station identification, and voice signals should be intelligible with all electric equipment operating and pulse equipment transmitting.
(ii) Localizer Intercept. In the approach configuration and at a distance of at least 18
n.m. from the localizer facility, fly toward the localizer front course, inbound, at an angle of at least 50 degrees. Perform this maneuver from both left and right of the localizer beam. No flags should appear during the time the deviation indicator moves from full deflection to on-course.
(iii) Localizer Tracking. While flying the localizer inbound and not more than 5 miles before reaching the outer marker, change the heading of the airplane to obtain full needle deflection. Then, fly the airplane to establish localizer on-course operation. The localizer deviation indicators should direct the airplane to the localizer on-course.
Perform this maneuver with both a left and a right needle deflection. Continue tracking the localizer until over the transmitter. Acceptable front course and back course approaches should be conducted to 200 feet or published minimums.
(iv) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight system.
(5) Glide Slope Systems.
(i) Signal Strength. The signal input to the receiver should be of sufficient strength to keep the warning flags out of view at all distances to 10 n.m. from the facility. This performance should be demonstrated at all airplane headings between 30 degrees right and left of the localizer course (see RTCA Document DO-1010). The deviation indicator should properly direct the airplane back to path when the airplane is above or below the path. Interference with the navigation operation, within 10 n.m. of the facility, should not occur with all airplane equipment operating and all pulse equipment transmitting. There should be no interference with other equipment as a result of glide slope operation.
(ii) Glide Slope Tracking. While tracking the glide slope, maneuver the airplane through normal pitch and roll attitudes. The glide slope deviation indicator should show proper operation with no flags. Acceptable approaches to 200 feet or less above threshold should be conducted.
(iii) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(6) Marker Beacon System.
(i) Flight Test.
(A) In low sensitivity, the marker beacon annunciator light should be illuminated for a distance of 2,000 to 3,000 feet when flying at an altitude of 1,000 feet AGL on the localizer centerline in all flap and gear configurations.
(B) An acceptable test to determine distances of 2,000 to 3,000 feet is to fly at a ground speed listed in table 1 and time the marker beacon light duration.
Altitude = 1,000 feet (AGL)
Ground Speed Light Time (Seconds) Knots 2000 feet 3000 feet 90 13 20 110 11 16 130 9 14 150 8 12
TABLE 287-1 – Light Duration
(C) For ground speeds other than table values, the following formulas may be used:
Upper limit = 1775 (seconds) Ground Speed in Knots
Lower limit = 1183 (seconds) Ground Speed in Knots
(D) In high sensitivity, the marker beacon annunciator light and audio will remain on longer than when in low sensitivity.
(E) The audio signal should be of adequate strength and sufficiently free from interference to provide positive identification.
(F) As an alternate procedure, cross the outer marker at normal ILS approach altitudes and determine adequate marker aural and visual indication.
(ii) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight system.
(7) Automatic Direction Finding (ADF) Equipment.
(i) Range and Accuracy. The ADF system installed in the airplane should provide operation with errors not exceeding 5 degrees, and the aural signal should be clearly audible up to the distance listed for any one of the following types of radio beacons:
(A) 75 n.m. from an HH facility.
(B) 50 n.m. from an H facility. Caution —service ranges of individual facilities may be less than 50 n.m.
(C) 25 n.m. from an MH facility.
(D) 15 n.m. from a compass locator.
(ii) Needle Reversal. The ADF indicator needle should make only one 180 degree reversal when the airplane flies over a radio beacon. This test should be made with and without the landing gear extended.
(iii) Indicator Response. When switching stations with relative bearings differing by 180 degrees ± 5 degrees, the indicator should indicate the new bearing within ± 5 degrees in not more than 10 seconds.
(iv) Antenna Mutual Interaction. For dual installations, there should not be excessive coupling between the antennas.
(v) Technique.
(A) Range and Accuracy. Tune in a number of radio beacons spaced throughout the 190-535 kHz range and located at distances near the maximum range for the beacon.
The identification signals should be understandable and the ADF should indicate the approximate direction to the stations. Beginning at a distance of at least 15 n.m. from a compass locator in the approach configuration (landing gear extended, approach flaps), fly inbound on the localizer front course and make a normal ILS approach.
Evaluate the aural identification signal for strength and clarity and the ADF for proper performance with the receiver in the ADF mode. All electrical equipment on the airplane should be operating and all pulse equipment should be transmitting. Fly over a ground or appropriately established checkpoint with relative bearings to the facility of 0 degrees, 45 degrees, 90 degrees, 135 degrees, 180 degrees, 225 degrees, 270 degrees, and 315 degrees. The indicated bearings to the station should correlate within 5 degrees. The effects of the landing gear on bearing accuracy should be determined. (A calibration placard should be provided, if appropriate.)
(B) Needle Reversal. Fly the airplane over an H, MH, or compass locator facility at an altitude 1,000 to 2,000 feet above ground level. Partial reversals that lead or lag the main reversal are permissible.
(C) Indicator Response. With the ADF indicating station dead ahead, switch to a station having a relative bearing of 175 degrees. The indicator should indicate within ±5 degrees of the bearing in not more than 10 seconds.
(D) Antenna Mutual Interaction.
(1) If the ADF installation being tested is dual, check for coupling between the antenna by using the following procedure.
(2) With #1 ADF receiver tuned to a station near the low end of the ADF band, tune the #2 receiver slowly throughout the frequency range of all bands and determine whether the #1 ADF indicator is adversely affected.
(3) Repeat (2) with the #1 ADF receiver turned to a station near the high end of the ADF band.
(vi) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(8) Distance Measuring Equipment (DME).
(i) Tracking Performances. The DME system should continue to track without dropouts when the airplane is maneuvered throughout the airspace within the standard service volume of the VORTAC/DME station and at altitudes above the lower edge of the standard service volume to the maximum operating altitude. This tracking standard should be met with the airplane:
(A) In cruise configuration;
(B) At bank angle up to 10 degrees;
(C) Climbing and descending at normal maximum climb and descent attitude;
(D) Orbiting a DME facility; and
(E) Providing a clearly readable identification of the DME facility.
(ii) Electromagnetic Compatibility (EMC). With all systems operating in flight, verify, by observation, that no adverse effects are present in the required flight systems.
(iii) Climb and Maximum Distance. Determine that there is no mutual interference between the DME system and other equipment aboard the airplane. Beginning at a distance of at least 10 n.m. from a DME facility and at an altitude of 2,000 feet above the DME facility, fly the airplane on a heading so that the airplane will pass over the facility.
At a distance of 5-10 n.m. beyond the DME facility, operate the airplane at its normal maximum climb attitude up to 90 percent of the maximum operating altitude, maintaining the airplane on a station radial (within 5 degrees). The DME should continue to track with no unlocks to the range of the standard service volume.
(iv) Long-Range Reception.
(A) Perform two 360 degree turns, one to the right and one to the left, at a bank angle of at least 10 degrees at the maximum service volume distance of the DME facility and at an altitude of at least 90 percent of the maximum operating altitude.
(B) Unlocks may occur and are acceptable if they do not interfere with the intended flight path of the airplane or are relieved by a reduction of bank angle at the same relative heading to the station.
(v) High-Angle Reception. Repeat the flight pattern and observations of (iii) above at a distance of 50-70 n.m. (20-30 n.m. for airplanes not to be operated above 18,000 feet) from the DME facility and at an altitude of at least 90 percent of the maximum operating altitude.
(vi) Penetration. From 90 percent of the maximum operating altitude, perform a letdown directly toward the ground station using normal maximum rate of descent procedures to a DME facility so as to reach an altitude of 5,000 feet above the DME facility 5-10 n.m. before reaching the DME facility. The DME should continue to track during the maneuver with no unlocks.