can damage the engine. This is not an issue for reciprocating engines equipped with FADEC because the system should make the throttle speed irrelevant.
9. § 23.943 Negative Acceleration.
a. Explanation. Tests should be conducted to show that no hazardous malfunction occurs under negative accelerations within the flight envelope. A hazardous malfunction in this case usually is considered to be one which causes a loss or sustained malfunction of the engine, or improper operation of the engine accessories or systems.
b. Procedures.
(1) Tests. Critical points of negative acceleration may be determined through tests.
Consideration should be given to the possibility of critical level of fuel and oil.
(2) Normal, Utility and Acrobatic Category Airplanes. With engines operating at maximum continuous power, the airplane is flown at a critical negative acceleration within the prescribed flight envelope. Normally, a duration of the negative acceleration in separate tests of -0.2g for five seconds, -0.3g for four seconds, -0.4g for three seconds, and -0.5g for two seconds should reveal any existing hazardous malfunctioning of the engine. Alternately, -0.5g for five seconds may be used.
(3) Acrobatic Category Airplanes. In addition, for acrobatic category airplanes, for which certification is requested for inverted flight or for negative g maneuvers, the airplane should be subjected to the maximum value and time of negative acceleration for which approval is requested.
(4) Commuter Category Airplanes. For commuter category airplanes, one continuous period of at least five seconds at -0·5g, and separately a period containing at least two excursions to 0·5g in rapid succession, in which the total time at less than zero g is at least five seconds has to be shown without any existing hazardous malfunctioning of the engine.
(5) Accelerations. In addition, it may be necessary to consider other points within the flight envelope at other levels of fuel with shorter applications of accelerations. In all cases, the accelerations are measured as near as practicable to the c.g. of the airplane.
Section 2. Fuel System
1. § 23.959 Unusable Fuel Supply. Refer to AC 23-16 for guidance.
2. § 23.961 Fuel System Hot Weather Operation. Refer to AC 23-16 for guidance.
Section 3. Fuel System Components
1. § 23.1001 Fuel Jettisoning System.
a. Explanation. The basic purpose of these tests is to determine that the required amount of fuel may be safely jettisoned under reasonably anticipated operating conditions within the prescribed time limit without danger from fire, explosion, or adverse effects on the flying qualities. The applicant should ha ve made sufficient jettisoning tests to prove the safety of the jettisoning system.
b. Procedures.
(1) Fire Hazard.
(i) Fuel in liquid or vapor form should not impinge upon any external surface of the airplane during or after jettisoning. Colored fuel, or surfaces so treated that liquid or vaporous fuel changes the appearance of the airplane surface, may be used for detection purposes. Other equivalent methods for detection may be acceptable.
(ii) Fuel in liquid or vapor form should not enter any portion of the airplane during or after jettisoning. The fuel may be det ected by its scent, combustible mixture detector, or by visual inspection. In pre ssurized airplanes, the presence of liquid or vaporous fuel should be checked with the airplane unpressurized.
(iii) There should be no evidence of fuel valve leakage after it is closed.
(iv) If there is any evidence that wing flap (slats/slots) positions other than those used for the test may adversely affect the flow pattern, the airplane should be placarded "Fuel should not be jettisoned except when flaps (slats/slots) are set at ___ degrees."
(v) The applicant should select, for demonstration, the tanks or tank combinations that are critical for demonstrating the flow rate during jettisoning.
(vi) Fuel jettisoning flow pattern should be demonstrated from all normally used tank or tank combinations on both sides of the airplane whether or not both sides are symmetrical.
(vii) Fuel jettisoning rate may be demonstrated from only one side of symmetrical tank or tank combinations that are critical for flow rate.
(viii) Fuel jettisoning rate and flow pattern should be demonstrated when jettisoning from full tanks using fuel.
(2) Control.
(i) Changes in the airplane control forces during the fuel jettisoning tests should be noted.
(ii) The capability to shut off the fuel jettisoning system should be demonstrated in flight.
(3) Residual Fuel. The residual fuel should be measured by draining the tanks from which fuel has been jettisoned in flight, measuring the total drained fuel, and subtracting from the total the unusable fuel quantity for each tank to determine if there is sufficient reserve fuel after jettisoning to meet the requirements of this section. This may be a ground test.
Section 4. Oil System
1. § 23.1027 Propeller Feathering System. Refer to paragraph 23.903 of this AC for guidance.
Section 5. Cooling
1. § 23.1041 General. Refer to paragraphs 23.1043, 23.1045, and 23.1047 of this AC for guidance.
2. § 23.1043 Cooling Tests.
a. Explanation. Paragraphs 23.1045 and 23.1047 of this AC provide details on reciprocating engine and turbine engine cooling tests. Additional procedures for certification of winterization equipment are given below.
b. Weight and C.G. Forward c.g. at maximum gross weight is usually the most critical condition. For reciprocating engine-powered airplanes of more than 6,000 pounds maximum weight and for turbine engine-powered airplanes, the takeoff weight need not exceed that at which compliance with § 23.63(c)(1) has been shown. If engine cooling is critical at high altitude, it may not be possible to achieve the critical point with the maximum weight, in which case a lower weight may represent the most critical weight condition.
c. Winterization Equipment Procedures. The following procedures should be applied when certificating winterization equipment:
(1) Other Than a 38 °C (100 °F). Cooling test results for winterization installations may be corrected to any temperature desired by the applicant rather than the conventional 38 °C (100 °F) hot day. For example, an applicant may choose to demonstrate cooling to comply with requirements for a 10 C (50 °F) or 15.5 C (60 °F) day with winterization equipment installed.
This temperature becomes a limitation to be shown in the AFM. In such a case, the sea level temperature for correction purposes should be considered to be the value elected by the applicant with a rate of temperature drop of 2 C (3.6 °F) per thousand feet above sea level.
(2) Tests. Cooling tests and temperature corr ection methods should be the same as for conventional cooling tests.
(3) Limit Temperature. The AFM should clearly indicate that winterization equipment should be removed whenever the temperature reaches the limit for which adequate cooling has been demonstrated. The cockpit should be placarded accordingly.
(4) Equipment Marking. If practical, winterization equi pment, such as baffles for oil radiators or for engine cooling air openings, should be marked clearly to indicate the limiting temperature at which this equipment should be removed.
(5) Installation Instructions. Since winterization equipment is often supplied in kit form and accompanied by instructions for its installation, manufacturers should provide suitable information regarding temperature limitations in the installation instructions.
3. § 23.1045 Cooling Test Procedures for Turbine Engine-Powered Airplanes.
a. Explanation.
(1) Purpose. Cooling tests are conducted to determine the ability of the powerplant cooling provisions to maintain the temperatures of powerplant components and engine fluids within the temperature limits for which they have been certificated. These limits will normally be specified on the TC data sheet.
(2) Components With Time/Temperature Limits. The conventional method of approving engine components is to establish a temperature limit that will ensure satisfactory operation during the overhaul life of the engine. However, a component that exceeds the temperature limit can be approved at the elevated temperature for a specific period of time. To ensure that a component having a time/temperature limit will operate within the established limitation, a means should be provided to record the time and temperature of an excessive temperature and warn the pilot accordingly. The method of recording elapsed time and temperature should be automatic or activated by the pilot with a simple operation. Operating limitations requiring the pilot to detect a critical airplane operating condition and record the elapsed time in the airplane logs would not be acceptable due to the other pilot duties during the critical airplane operating condition.
(3) Altitude. Cooling tests should be conducte d under critical ground and flight operating conditions to the maximum altitude for which approval is requested.
b. Test Procedures Applicable to Both Single-Engine and Multiengine Airplanes.
(1) Performance and Configuration. Refer to § 23.45, which has performance requirements related to engine cooling.
(2) Moisture. The tests should be conducted in air free of visible moisture.
(3) Weight and C.G. Forward c.g. at maximum gro ss weight is usually the most critical condition.
(4) Oil Quantity. The critical condi tion should be tested.
(5) Thermostat. Airplanes that incor porate a thermostat in the engine oil system may have the thermostat retained, removed, or blocked in such a manner as to pass all engine oil through the oil cooler. If the ther mostat is retained, the oil temperature readings obtained on a cooler day corrected to hot-day conditions may, therefore, be greater than those obtained under actual hot-day conditions. Caution should be ex ercised when operating an airplane with the thermostat removed or blocked during cold weather to prevent failure of the lubricating system components.
(6) Instrumentation. Accurate and calibrated temper ature-measuring devices should be used, along with acceptable thermocouples or temperature-pickup devices. The proper pickup should be located at critical engine positions.
(7) Generator. The alternator/generator should be electrically loaded to the rated capacity for the engine/accessory cooling tests.
(8) Temperature Limitations. For cooling tests, a maximum anticipated temperature (hot-day conditions) of at least 38 C (100 °F) at sea level must be used. Temperatures at higher altitudes assume a change at 2 C (3.6 °F) per 1,000 feet of altitude, up to -56.5 C (-69.7 °F).
The maximum ambient temperature selected and demonstrated satisfactorily becomes an airplane operating limitation per the requirements of § 23.1521(e).
(9) Temperature Stabilization. For the cooling tests, a temperature is considered stabilized when its observed rate of change is less than 2 °F per minute.
(10) Altitude. The cooling tests should be starte d at the lowest practical altitude, usually below 3,000 feet MSL, to provide a test data point reasonably close to sea level.
(11) Temperature Correction for Ground Operation. Recorded ground temperatures should be corrected to the maximum ambient temperature selected, without consideration of the altitude temperature lapse rate. For example, if an auxiliary power unit is being tested for ground cooling margins, the cooling margin should be determined from the recorded ground temperature without regard to the test site altitude.
c. Test Procedures for Single-Engine, Turbine-Powered Airplanes.
(1) Normal engine start. A normal engine start should be made and all systems checked out. The engine should be run at ground idle and temperatures and other pertinent data should be recorded.
(2) Taxi. Taxi airplane for approximately one mile to simulate normal taxi operations. Record cooling da ta at one-minute intervals.
(3) Seaplanes. For hull-type seaplanes operating on water, taxi tests should be conducted such that spray characteristics do not bias the cooling characteristics. Engine cooling during water taxiing should be checked by taxiing downwind at a speed approximately five knots above the step speed for a minimum of ten minutes continuously. Record cooling data at oneminute intervals.
(4) Pre-takeoff. Establish a pre-takeoff holding c ondition on the taxiway (crosswind) for 20 minutes minimum or until temperatures stabilize. Record cooling data at five-minute intervals.
(5) Runway. On the runway, set takeoff power and record cooling data.
(6) Takeoff. Takeoff as prescribed in § 23.53 and climb to pattern altitude. Record cooling data upon reaching pattern altitude or at one-minute intervals if it takes more than oneminute to reach pattern altitude.
(7) Flaps and climb. Retract flaps and continue climb with maximum continuous power at the speed selected to meet the requirements of § 23.65(b). Climb to the maximum approved altitude, recording cooling data at one-minute intervals.
(8) Cruise. Cruise at maximum continuous power (or V MO/MMO, if limiting) at maximum operating altitude until temperatures stabilize. Record cooling data at one-minute intervals. For many components, this will be the critical temperat ure operating condition.
(9) Descent. Conduct a normal descent at V MO/MMO to holding altitude and hold until temperatures stabilize. Record cooling data at one-minute intervals.
(10) Approach. Conduct a normal approach to landing. Record cooling data at oneminute intervals.
(11) Balked landing. From not less than 200 feet above the ground, perform a balked landing go-around in accordance with § 23.77. Reco rd cooling data at one-minute intervals during a traffic pattern circuit.
(12) Climb to pattern. Climb to pattern altitude, perform a normal approach and landing in accordance with the applicable portion of § 23.75. Record cooling data at one-minute intervals.
(13) Taxi back to ramp. Shut down engine. Allow e ngine to heat-soak. Record temperature data at one-minute intervals until five minutes after temperatures peak.
d. Test Procedures for Multiengine, Turbine-Powered Airplanes. A multiengine airplane should conduct the same profile as the single-engine airplane, in an all-engine configuration. On completion of the all-engine profile, conduct the applicable one-engine inoperative cooling climb test recording data at one-minute intervals. Shut down critical engine and, with its propeller (if applicable) in the minimum drag position, the remaining engine(s) at not more than maximum continuous power or thrust, landing gear retracted, and wing flaps in the most favorable position. Climb at the speed used to show compliance with § 23.67. Continue until five minutes after temperatures peak.
e. Data Acquisition. The following data should be recorded at the time intervals specified in the particular test program. The data may be manually recorded unless the quantity and frequency necessitate automatic or semi-automatic means:
(1) Outside air temperature (OAT);
(2) Altitude;
(3) Airspeed (knots);
(4) Gas generator r.p.m.;
(5) Engine torque;
(6) Time;
(7) Propeller r.p.m.;
(8) Engine oil temperature;
(9) Pertinent engine temperature; and
(10) Pertinent nacelle and component temperatures.
f. Data Reduction.
(1) Limitations. A maximum anticipated temperature (hot-day conditions) of at least 38 °C (100 °F) at sea level must be used. The assumed temperature lapse rate is 2 °C (3.6 °F) per 1,000 feet altitude up to the altitude at which a temperature of -56.5 C (-69.7 oF is reached, above which altitude the temperature is constant at -56.5 oC (-69.7 oF). On turbine-powered airplanes, the maximum ambient temperature selected becomes an airplane operating limitation in accordance with the requirements of § 23.1521(e). On turbine-powered airplanes, the applicant should correct the engine temperatures to as high a value as possible in order to not be limited.
(2) Correction Factors. Unless a more rational method applies, a correction factor of
1.0 is applied to the temperature data as follows and as shown in the sample calculation below:
corrected temperature = true temperature + 1.0 [100 - 0.0036 (Hp) - true OAT].
Figure 24 – Sample Calculation 128
Sample Calculation:
True Temperature -- 300 °F True OAT -- 15 °F Hp -- 5,000 feet The corrected temperature = 300 + 1.0 [100 - 0.0036 (5,000) - 15] = 367 °F.
The corrected temperature is then compared with the maximum permissible temperature to determine compliance with the cooling requirements.
4. § 23.1047 Cooling Test Procedures for Reciprocating Engine-Powered Airplanes.
a. Procedures.
(1) Additional Procedures. The procedures of paragraph 23.1045b(1) through
23.1045b(6) of this AC also apply to reciprocating engines.
(2) Altitude. Engine cooling tests for reciproc ating engine airplanes are normally initiated below 2,000 feet pressure altitude. Service experience indicates that engine cooling tests started above 5,000 feet may not assure adequate cooling margins when the airplane is operated at sea level. If an appl icant elects not to take the airplane to a low altitude test site, additional cooling margins have been found acceptable. If engine cooling tests cannot be commenced below 2,000 feet pressure altitude, the temperature margin should be increased. It should be increased by 17 C (30 °F) at 7,000 feet for cylinder heads and 33 C (60 °F) for both engine oil and cylinder barrels with a straight line variation from sea level to 7,000 feet, unless the applicant demonstrates that some other correction margin is more applicable.
(3) Hull-Type Seaplanes. Cooling tests on hull-type seaplanes should include, after temperatures stabilize, a downwind taxi for ten minutes at five knots above the step speed, recording cooling data at one-minute intervals.
(4) Test Termination. If at any time during the test temperatures exceed the manufacturer's specified limits, the test is to be terminated.
(5) Climb Transition. At the beginning of the cooli ng climb, caution should be used in depleting the kinetic energy of the airplane while establishing the climb speed. The climb should not be started by "zooming" into the climb. The power may be momentarily reduced provided that the stabilized temperatures are not allowed to drop excessively. This means that a minimum of time should be used in slowing the airplane from the high cruise speed to the selected cooling climb speed. This may be accomplished by maneuver loading the airplane or any other means that provides minimum slow-down time.
(6) Component Cooling. Accessories or components on the engine or in the engine compartment that have temperature limits should be tested and should be at the maximum anticipated operating conditions during the cooling tests; for example, generators should be at maximum anticipated loads.
(7) Superchargers. Superchargers and turbo-superchargers should be used as described in the AFM. Engine cooling should be evaluated in the cruise condition at the maximum operating altitude, since this may be a more critical point than in climb. Also, 129
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.