turbocharged engines sometimes give a false peak and the climb should be continued long enough to be sure that the temperatures do not begin to increase again.
(8) Single-Engine Airplanes. The cooling tests for singl e-engine airplanes should be conducted as follows:
(i) At the lowest practical altitude, establish a level flight condition at not less than 75 percent maximum continuous power until temperatures stabilize. Record cooling data.
(ii) Increase engine power to takeoff rating and climb at a speed corresponding to the applicable performance data given in the AFM/POH, which are criteria relative to cooling.
Maintain takeoff power for one minute. Record cooling data.
(iii) At the end of one minute, reduce engine power to maximum continuous and continue climb for at least five minutes after temperatures peak or the maximum operating altitude is reached. Record c ooling data at one-minute intervals. If a leaning schedule is furnished to the pilot, it should be used.
(9) Multiengine Airplanes. For multiengine-powered airplanes that meet the minimum one-engine-inoperative climb performance specified in § 23.67 with the airplane in the configuration used in establishing critical one-engine-inoperative climb performance, the cooling tests should be conducted as follows:
(i) At the lower altitude of 1,000 feet below engine critical altitude or 1,000 feet below the altitude at which the minimum one-engine-inoperative climb gradient is 1.5 percent, or at the lowest practical altitude (when applicable), stabilize temperatures of the test engine in level flight at not less than 75 percent maximum continuous power. Record cooling data.
(ii) After temperatures stabilize, initiate a climb at a speed not more than the highest speed at which compliance with the climb requirement of § 23.67 is shown. With the test engine at maximum continuous power (or full throttle), continue climb until five minutes after temperatures peak or the maximum operating altitude is reached. Record cooling data at one-minute intervals.
(10) Performance Limited Multiengine Airplanes. For multiengine airplanes that cannot meet the minimum, one-engine-inoperative performance specified in § 23.67 is shown below:
(i) Set zero thrust on the planned "inoperative" engine and determine an approximate rate of sink (or climb). A minimum safe test altitude should then be established.
(ii) Stabilize temperatures in level flight with engines operating at no less than 75 percent maximum continuous power and as near sea level as practicable or the minimum safe test altitude.
(iii) After temperatures stabilize, initiate a climb at a speed not more than the highest speed at which compliance with the climb requirements of § 23.67 is shown with one engine inoperative and remaining engine(s) at maximum continuous power. Continue for at least five minutes after temperatures peak. Record cooling data at one-minute intervals.
b. Data Acquisition. The following data should be recorded at the time intervals specified in the applicable test programs and may be manually recorded unless the quantity and frequency necessitate automatic or semi-automatic means:
(1) Time;
(2) Hottest cylinder head temperature;
(3) Hottest cylinder barrel temperature (only if a limitation);
(4) Engine oil inlet temperature;
(5) Outside air temperature;
(6) Indicated airspeed (knots);
(7) Pressure altitude;
(8) Engine r.p.m.;
(9) Propeller r.p.m.;
(10) Manifold pressure;
(11) Carburetor air temperature;
(12) Mixture setting;
(13) Throttle setting; and
(14) Temperatures of components or accessories that have established limits that may be affected by powerplant heat generation.
c. To Correct Cylinder Barrel Temperature to Anticipated Hot-Day Conditions.
(1) Cooling requirements. To determine compliance with cooling requirements and find the corrected cylinder barrel temperature, follow the formulas shown in figure 25 below:
Figure 25 – Corrected Cylinder Barrel Temperature Corrected cylinder barrel temperature = true observed cylinder barrel temperature +
0.7 [100 - 0.0036 (pressure altitude) - true OAT].
For example:
True observed maximum cylinder barrel temperature-- 244 °F.
Pressure Altitude --8330 feet True OAT --+55 °F.
Corrected cylinder barrel temperature (°F) = 244 + 0.7 [100 - 0.0036 (8330) - 55] = 255°F.
(2) Corrected temperatures. The corrected temperatures are then compared with the maximum permissible temperatures to determine compliance with cooling requirements.
d. To Correct Cylinder Head or Other Temperatures to Anticipated Hot-Day Conditions.
(1) Cooling requirements. To determine compliance with cooling requirements, find the correct cylinder head temperature and follow the formula shown in figure 26 below:
Figure 26 – Correct Cylinder Head Temperature 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].
For example (using metric units):
True maximum cylinder head temperature-- 208 °C.
Pressure Altitude -- 8330 feet True OAT -- +13 °C.
Corrected cylinder head temperature = 208 + 1.0 [38 - 0.002 (8330) -13] = 216 °C.
(2) Corrected temperatures. The corrected temperatures are then compared with the maximum permissible temperatures to determine compliance with cooling requirements.
e. Liquid Cooled Engines. Reserved.
Section 6. Induction System
1. § 23.1091 Air Induction. AC 20-124 covers the turbine engine water ingestion aspects of this requirement.
2. § 23.1093 Induction System Icing Protection.
a. Explanation.
(1) Purpose. Tests of engine induction syst em 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 fu el injected engines. See
23.1093(a)(3)
(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 he at 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 assu re 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 place d 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) Stabilize. At low altitude, stabilize airplane w ith 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 heat. Apply carburetor heat and, after condition stabilizes, record data.
(3) Reduce airspeed. 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 heat. Apply carburetor heat and, after condition stabilizes, record data.
(5) Reduce airspeed. 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 heat. Apply carburetor heat and, after condition stabilizes, record data.
(7) Repeat. At the intermediate altitude , repeat steps (1) through (6).
(8) Repeat. 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 27 and 28 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.
:
Hay be fl ovu HIHI HUM ALTI TUD E
INTERMEDIATE
ALTITUDE atooly one altitude Full Throttle 90% IAS of 80% IAS of Full Thro ttle 90% IAS 80% IAS if O.
A.T.
o.~ 30°F i1 or H C Coluon #1 Column #l of HC of of vailable Power* Power* Column #1 Column fl :a rbure tor Air Heat Control Position C H C
H C H C 11 C H C B Q :Q - Pres 1ure Al ti tude (ft.)
......__,, (a3) m m O.A.T.
( F) 83 83 83 - .A.T.
( F) 84 215 84 205 84 200 jleat Ri se (in) t8
(ii) - - - I.
A.
S.
(H.P.H.)
105 99 95 92 84 82 l!..P .H.
2850 2730 2690 2590 2430 2310 H.P.
(I
n.
R1t.)
26.4
25.7
24.0
23.5
22.0
21.
3 Indicated B.H.P .
144 132 120 11 2 105 99 ~td.
Temperature fo r - IPreuure Altitud e , ( P) 54 Temperature Cor rection Factor (See note
l) .972 .8 72 .972 .879 .972 .882 !lctu al B.H.P.
140 11 5 117
98.4 102
87.4
• Ra te d B.H.P.
~ w (,;)
(Su note
2)
82.2
70.2
62.4 - --- '-,,I' lrhrottle Pos it io n FT FT p p p p *Super charge d Engines On ly !!Q!!_!_ :
temperature Correction Factor• ~:
Rated BHP• 140 atd tem p (°F) + 460 CAT (Of) + 46 0 ~:
Cir cled nu mbers indicate data plotted on figure 256-2.
'-,,I'
m) ml
m) 72 72 72 - - 73 201 73 189 73 184 ~ Q
(m)I - - - 96 88 87 78 77 70 2800 2640 2555 2400 241 0 2280 23 .5
22.8
19.6
19.3 19.0 18.5 125 114 92 85 76 72 -, 41 .970 .810 · .970 .879 .9 70 .88 2 121 99 .2 89
74.7 74
63.5 Q
71.0 0
53.4
62.
~
45.3 ......__,, ......__,, - FT FT p p p p
MAXIMUM
ALTITUDE (75%) Full Throttle 90 % IAS 80% IAS or MC of of Pow er * Colummn fl Column fl C R C B C B Q - '-,,I'
(60) /6ol (60' 60 60 60 - - 61 190 61 18 5 61 176 ~ (i'2Sl 616 - - - 90 80 82 75 72 67 2770 2525 2665 2480 2525 2310
21.2 20 .4
19.9
19.
18.0 17.2 113 100 101 90 73 65 -, 30 .970 .868 .970 .871 .97 0 .1178 110
86.8 98 78 .4 71
57.1 bs~
62.1 no)
56.0 ~
40.8 --- --- Ft FT p p p p NOTE 11/16/2011 AC 23-8C Figure 27 - Carburetor Air Heat Rise Calculations 135
- i, ' ~-~ , I/ / I ~ I II' / / / I/ .I,'\.. I
I -I/ I I ,ii ) / / , rt) ~-~ / I/ ~-~ / ' , r,< I/ I ~ ... / ,, -.I.
I J \JJ V / I/ / / / ,N J «t-:,,,. / / / ) II' / I/ I ) I'-~ r-1( I/ "'-' / '\.
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\.
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"' 11/16/2011 AC 23-8C Figure 28 - Carburetor Air Heat Rise Plots 100
PERCENT
OF RATED
BHP
OUTSIDE AIR
TEMPERATURE
- 0F 50
110 120 130 140 150 CARBURETOR AIR HEAT RISE - 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 Section 7. Powerplant Controls and Accessories 136
1. § 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. (Refer to § 23.777). Controls are required to maintain any position set by the pilot without tendency to creep due to vibration or control loads.
Electronic engine control system installations must meet the requirements of § 23.1309.
b. Procedures. None.
2. § 23.1145 Ignition Switches. Reserved.
3. § 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.
Section 8. Powerplant Fire Protection
1. § 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.
Chapter 5. Subpart F--Equipment Section 1. General
1. § 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. The flight test pilot should evaluate all installations of this nature to verify that the equipment functions properly when installed.
b. Avionics Test.
(1) Very High Frequency VHF Communication Systems. Refer to 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.
(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 severa l 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 th e 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 switch es 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 (prop-modulation). 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 Aeronautical 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 two 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. Contac t the nearest FAA Airway Facilities Sector Office to locate a Doppler VOR.
(iii) Low-Angle Reception. Perform a 360 degree right a nd 360 degree left turn at a bank angle of at least ten 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. [note to myself, do a search]
(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. In VMC, incur fault condition to test activation of failure flag on the display.
(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, paragraph (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 AIM. This signal should be received for 360 degrees of the airplane heading at all bank angles up to ten degrees left or right at all normal pitch attitudes and at an altitude of approximately 2,000 feet (refer to RTCA guidance).
(B) Bank Angles. Satisfactory results should al so 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 configurati on 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 five miles before reaching the outer marker, change the heading of the airplane to obtain full needle deflection. Then, fly the airplane to esta blish 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. In VMC, incur fault condition to test activation of failure flag on the display.
(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 rece iver should be of sufficient strength to keep the warning flags out of view at all distances to ten n.m. from the facility. This performance should be demonstrated at all airplane headings between 30 degrees right and left of the localizer course (refer to 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 ten 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 the table below (table 10) and time the marker beacon light duration.
Table 10 – 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 (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 installe d in the airplane should provide operation with errors not exceeding five degrees, and the aural signal should be clearly audible up to the distance listed for any one of the following types of radio beacons: