Chapter 4. Subpart E--Powerplant Section 1. General
1. § 23.901 Installation.
2. § 23.903 Engines.
a. Explanation.
(1) Automatic Propeller Feathering Systems. All parts of the feathering device that are integral with the propeller or attached to it in a manner that may affect propeller airworthiness should be considered. The determination of airworthiness should be made on the following basis:
(i) The automatic propeller feathering system should not adversely affect normal propeller operation and should function properly under all temperatures, altitudes, airspeeds, vibrations, accelerations, and other conditions to be expected in formal ground and flight operation.
(ii) The automatic device should be demonstrated to be free from malfunctioning, which may cause feathering under any conditions other than those under which it is intended to operate. For example, it should not cause feathering during the following:
(A) Momentary loss of power; and (B) Approaches with reduced throttle settings.
(iii) The automatic propeller feathering system should be capable of operating in its intended manner whenever the throttle control is in the normal position to provide takeoff power. No special operations at the time of engine failure should be necessary on the part of the crew in order to make the automatic feathering system operative.
(iv) The automatic propeller feathering installation should be such that not more than one propeller will be feathered automatically even if more than one engine fails simultaneously.
(v) The automatic propeller feathering installation should be such that normal operation may be regained after the propeller has begun to feather automatically.
(vi) The automatic propeller feathering installation should incorporate a switch or equivalent means to make the system inoperative. (Also refer to §§ 23.67 and 23.1501.)
(vii) If performance credit is given for the automatic propeller feathering system, there should be means provided to satisfactorily preflight check the system.
(viii) Some turbopropeller airplanes are equipped with some type of engine ignition system intended for use during flight in heavy precipitation conditions and for takeoff/landing on wet or slush-covered runways. The engine ignition system may be either automatic or continuous. The purpose of this system is to prevent or minimize the possibility of an engine flameout due to water ingestion. Compatibility with auto-feather systems should be ensured.
(2) Negative Torque Sensing Systems. Reserved.
b. Procedures.
(1) Automatic and Manual Propeller Feathering System Operational Tests.
(i) Tests should be conducted to determine the time required for the propeller to change from windmilling (with the propeller controls set for takeoff) to the feathered position at the takeoff speed determined in § 23.51.
(ii) The propeller feathering system should be tested at one engine inoperative climb airspeed. The configuration should be the following:
(A) Critical engine inoperative;
(B) Wing flaps retracted;
(C) Landing gear retracted; and (D) Cowl flaps closed.
Note: If the feathered propeller has a residual rotation, this has to be considered for aircraft performance.
(iii) The propeller should be tested in the actual configuration for an emergency descent. A sufficient speed range should be cove red to assure that any propeller rotation is not hazardous. In addition, the propeller should not inadvertently unfeather during these tests.
(iv) In order to demonstrate that the feathering system operates satisfactorily, propeller feather should be demonstrated throughout both the airspeed and the altitude envelope since engine failure may occur at any time. Propeller unfeathering, manually or automatically, need only be demonstrated up to the maximum one-engine-inoperative service ceiling or maximum airstart altitude, whichever is higher. Satisfactory propeller unfeathering should also be demonstrated after a 30-minute cold soak.
(2) Continued Rotation of Turbine Engines.
(i) Means should be provided to completely stop the rotation of turbine engines if continued rotation would cause a hazard to the airplane. Devices such as feathering propellers, brakes, doors, or other means may be used to stop turbine engine rotation.
(ii) If engine induction air duct doors or other types of brakes are provided to control engine rotation, no single fault or failure of the system controlling engine rotation should cause the inadvertent travel of the doors toward the closed position. Also, it should not cause the inadvertent energizing of braking means unless compensating features are provided to ensure that engine failure or a critical operating condition will not occur. Such provisions should be of a high order of reliability, and the probability should be remote that doors or brakes will not function normally on demand.
(3) Engine Operation with Automatic Propeller Control System Installed.
(i) When an automatic control system for simultaneous r.p.m. control of all propellers is installed, it should be shown that no single failure or malfunction in this system or in the engine controlling this system will cause the following:
(A) The allowable engine overspeed for this condition to be exceeded at any time; and (B) A loss of thrust that will cause the airplane to fail to meet the requirements of §§ 23.51 through 23.77 if such system is certificated for use during takeoff and climb. This should be shown for all weights and altitudes for which certification is desired. A period of five seconds should be allowed from the time the malfunction occurs to the initial motion of the cockpit control for corrective action taken by the crew.
(ii) Compliance with this policy may be shown by analysis, flight demonstration, or a combination thereof.
c. Restart Envelope.
(1) Explanation. The applicant should propose a practicable airstart envelope wherein satisfactory inflight engine restarts may be accomplished as required by the regulation.
Airstarts should be accomplished satisfactorily at critical combinations of airspeed and altitude.
During these tests, normally, time history data showing airspeed, altitude, r.p.m., exhaust temperature, and so forth, are obtained for inclusion in the type inspection report.
(i) The airstart envelope should be included in the limitations section of the AFM, and the procedures used to restart the engine(s) should be contained in the emergency or abnormal procedures section of the AFM.
(ii) Results of restart tests completed by the engine manufacturer on the same type of engine in an altitude test facility or flying test bed, if available, are included in the type inspection report. The experience accumulated in other aircraft with the same engine and engine installation, may be taken into account if justified.
(2) Procedures. To establish the required enve lope of altitude and airspeed, sufficient flight tests should be made as follows:
(i) From sea-level to the maximum declared restarting altitude in all appropriate configurations likely to affect restarting, including the emergency descent configuration;
(ii) From the minimum to the maximum declared airspeed at all altitudes up to the maximum declared engine restarting altitude. The airspeed range of the declared restart envelope normally should cover at least 30 knots, but should be adapted to the type of airplane.
(A) The tests should include the effect on engine restarting performance of delay periods between engine shutdown and restarting as follows:
(iii) Up to two minutes; and
(iv) At least until the engine oil temperature is stabilized at its approximate cold soak value.
3. § 23.905 Propellers. Refer to paragraph 23.903 of this AC for guidance.
4. § 23.909 Turbo Superchargers. Refer to AC 23-16A for guidance.
5. § 23.925 Propeller Clearance. Reserved.
6. § 23.929 Engine Installation Ice Protection.
a. Explanation. This regulation requires that propellers and other components of the complete engine installation, such as oil cooling inlets, generator cooling inlets, and so forth, function satisfactorily and operate properly without an appreciable and unacceptable loss of power when the applicant requests approval for flight in icing conditions. An unacceptable loss of power may depend on the kind of aircraft and the power available. For details refer to the latest version of AC 23.1419-2. Refer to § 23.1093 for induction system ice protection requirements.
b. Procedures. Each propeller and other components of the complete installation that is to be approved for flight in icing conditions should be evaluated under the icing conditions specified in part 25, Appendix C. If the propellers are equipped with fluid-type deicers, the flow test should be conducted starting with a full tank of fluid and operated at maximum flow for a time period found operationally suitable. The operation should be checked at all engine speeds and powers.
7. § 23.933 Reversing Systems.
a. Explanation. Self-explanatory.
b. Procedures. Reversing system installations may be approved provided the following is acceptable:
(1) Pilot skill. Exceptional pilot skill should not be required in taxiing or any condition in which reverse thrust is to be used.
(2) Operating procedures. Necessary operating procedures, operating limitations, and placards are established.
(3) Control characteristics. The airplane control characte ristics are satisfactory with regard to control forces encountered, and buffeting should not cause structural damage.
(4) Directional control. The directional control is adequate using normal piloting skill.
(5) Sudden engine failure. A determination is made that no dangerous condition is encountered in the event of sudden failure of one engine in any likely operating condition.
(6) Reasonable safeguards. The operating procedures and airplane configuration are such as to provide reasonable safeguards against serious structural damage to parts of the airplane due to the reverse airflow.
(7) Pilot’s vision. It is determined that the pilo t’s vision is not dangerously obscured under normal operating conditions on dusty or wet runways and where light snow is on the runway.
(8) Pilots vision for seaplanes. It is determined that the pilot’s vision is not dangerously obscured by spray due to reverse airflow under normal water operating conditions with seaplanes.
(9) Reverse idle setting. The procedure and mechan isms for reversing should provide a reverse idle setting such that without requiring exceptional piloting skill at least the following conditions are met:
(i) Sufficient power is maintained to keep the engine running at an adequate speed to prevent engine stalling during and after the propeller reversing operation.
(ii) The propeller does not overspeed during and after the propeller reversing operation.
(10) Engine cooling. The engine cooling characteris tics should be satisfactory in any likely operating condition.
(11) Ground idle position. If using ground idle for disking drag credit on landing distance, the ground idle position of the power levers should be identified with a gate or a detent with satisfactory tactile feel (reference paragraph 23.75a(6)(iv) of this AC).
(12) Thrust reverser position. If compliance with § 23.933(a )(1)(ii) is intended to be shown by flight tests, any possible position of any one thrust reverser has to be assumed.
8. § 23.939 Powerplant Operating Characteristics.
a. Explanation. Self-explanatory.
b. Procedures.
(1) Stall, Surge, Flameout Tests. For turbine engines, te sts should be conducted to determine that stall, surge, and flameout will not occur, to a hazardous degree, on any engine during acceleration and deceleration throughout the normal flight envelope of the airplane. This would include tests throughout the approved altitude range and throughout the airspeed range from VS to VMO/MMO using sideslip angles appropriate to the individual airplane. For normal category multiengine airplanes, an appropriate sideslip angle is generally considered to be approximately one ball width on a standard slip-skid indicator. The low airspeed tests should be accomplished at light weight and with gear and flaps extended to further reduce the stall speed.
Tests need not be accomplished with gear and flaps extended at airspeeds above which extension is prohibited in the AFM. At the conditions ment ioned above, the effects of engine bleed air off and on and engine ice protection systems off and on should be investigated.
(2) Throttle Techniques for Turbine Engines. With the engine stabilized at maximum continuous power, rapidly retard the throttle to the flight idle position. Before the engine reaches idle power or r.p.m., rapidly advance the throttle to maximum continuous power. Repeat this process, except begin with the engine stabilized at flight idle power. Rapid throttle movement is generally defined as one which results in the throttle moving from maximum continuous power to flight idle, or vice versa, in not more than 0.5 seconds.
(3) Throttle Techniques for Reciprocating Engines. The techniques are similar to turbine engines except throttle movements should be slower and smoother to avoid damaging the engine. Most reciprocating engi nes are still manually controlled and rapid throttle movements 122
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