CHAPTER 27. POWERPLANT: COOLING
27.1 Cooling Test Procedures—§ 25.1045.
27.1.1 Explanation.
The following guidance applies to cooling test procedures for turbine engine powered airplanes.
27.1.1.1 Purpose.
In accordance with § 25.1041, cooling tests must be 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 engine type certificate data sheet (TCDS), qualification specification sheet for the component, and/or in the approved engine installation handbook.
27.1.1.2 Scope.
Cooling tests should be conducted under (or data corrected to) critical ground and flight operating conditions to the maximum altitude for which approval is requested.
27.1.2 Flight Test Procedures.
27.1.2.1 Moisture.
The tests should be conducted in air free of visible moisture.
27.1.2.2 Weight and CG.
Forward CG at maximum gross weight is usually the most critical condition as this results in the lowest airspeed and/or vertical speed.
However, a lighter initial gross weight or a critical step climb profile may be necessary to get to maximum altitude. In any case, the most critical climb profile must be established by the applicant and agreed to by the FAA prior to commencing testing.
27.1.2.3 Test Conditions.
The critical flight profile(s) should be tested. It may be necessary to fly multiple flights with different flight profiles to assure all components under test are exposed to their most critical anticipated sequence of test conditions. Flight profiles with periods of relatively high heat rejection and relatively low cooling, such as can occur during step climbs and at the top of descent, may be more critical for some components, for example, those in or affected by the engine oil system. The period after engine shutdown on the ground may also be critical for some components. The applicant should identify and obtain the FAA’s approval of the flight test 27-1
profile(s) prior to beginning certification testing. The following sequence of test conditions is usually adequate to cover the critical case for most components:
1. Initiate the flight test once critical temperatures have stabilized after engine start;
2. Use the test engine to perform a 1-mile single engine taxi.
3. Hold at idle power or thrust in a 10 knot or greater crosswind for 20 minutes or until temperatures stabilize (use ground test definition of stabilization).
4. Operate at least the test engine at the rated takeoff power or thrust for the maximum approved period (usually either 5 or 10 minutes) during a maximum gross weight/forward CG takeoff.
5. Perform simulated one-engine-inoperative and all-engines-operating climbs, operating the test engine at maximum continuous power or thrust until the engine temperatures stabilize or the airplane reaches maximum operating altitude (or until the airplane is essentially unable to climb further as indicated by a very low climb rate, e.g., 200 feet/minute).
6. Cruise with the test engine at maximum continuous power or thrust (but a speed no higher than VMO/MMO) at maximum operating altitude until temperatures stabilize.
7. Conduct a normal descent at VMO/MMO to a typical holding altitude and hold until temperatures stabilize.
8. Conduct a normal approach to landing, but from not less than 200 feet above the ground:
a. Perform a simulated engine out go-around;
b. Climb to pattern altitude; and then
c. Perform a normal approach and landing.
9. Taxi back to the ramp and shut down the engines.
10. Allow the test engine’s heat-soak to peak.
27.1.2.4 Oil Quantity.
The critical condition should be tested.
27.1.2.5 Thermostat.
Airplanes that incorporate 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 thermostat 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 exercised when operating an airplane 27-2
with the thermostat removed or blocked during cold weather to prevent failure of the lubricating system components.
27.1.2.6 Instrumentation.
The applicant must identify all critical components (electronic components, actuators, etc.), including structural elements that have temperature limits. The limits are typically based on component qualification or certification testing. Each limit may be expressed in terms of a surface temperature or environmental temperature and may have associated time limits. Instrumentation should be installed to provide data needed to show that each component corrected temperature remains below the identified limit. Accurate and calibrated temperature-measuring devices should be used, along with acceptable thermocouples or temperature pickup devices. The temperature pickup should be located at critical engine positions.
27.1.2.7 Generator.
The alternator/generator should be electrically loaded to the rated capacity for the engine/accessory cooling tests.
27.1.2.8 Maximum Ambient Atmospheric Temperature.
Section 25.1043(b) establishes 100 °F (38 °C) at sea level as the lowest maximum ambient temperature for cooling tests, except for winterization installations. (See paragraph 27.1.2.12 for guidance on certifying winterization equipment.) Applicants may establish a higher temperature limit if desired. In accordance with § 25.1041, applicants must show that cooling provisions can maintain the temperatures of powerplant components, engine fluids, and auxiliary power unit components and fluids within the established temperature limit. The assumed temperature lapse rate is -3.6 °F (-2 °C) per thousand feet of altitude above sea level until a temperature of -69.7 °F (-56.5 °C) is reached, above which altitude the temperature is considered constant at -69.7 °F (-56.5 °C). The compliance demonstration flight test should be conducted with an ambient temperature as close to the desired maximum ambient atmospheric temperature as practical. If testing is accomplished at lower ambient temperatures, then the test data must be corrected to that which would have resulted from testing on a day with the maximum ambient atmospheric temperature. (See paragraph 27.1.5.2.) The maximum ambient temperature selected and demonstrated satisfactorily, taking account of correction factors, shall not be less than the minimum hot day conditions prescribed by § 25.1043(b) and shall be an airplane operating limitation per the requirements of § 25.1521(d). The applicant should correct the engine temperatures to as high a value as possible in order to minimize the impact of this limitation.
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27.1.2.9 Temperature Stabilization.
For the cooling flight tests, a temperature is usually considered stabilized when its observed rate of change is less than 2 °F per minute. However, regardless of the rate of temperature rise, if a component or fluid temperature is still rising and is near the limit for that component or fluid, sufficient test data must be gathered to show that the limit will not be exceeded in that steady state operating condition. A combination of test data and rational analysis may be used to determine the maximum temperature expected for a component or fluid rather than continuing a test condition for a long period of time after the rate of temperature rise has fallen below 2 °F per minute.
27.1.2.10 Temperature inversion.
During an inversion, where a layer of colder air is at ground level, component temperatures at the beginning of the cooling test climb will be lower relative to the ambient air temperature encountered during the climb. For components that require significant time to adjust to ambient temperature changes, the artificially low starting temperature may result in erroneous results. If cooling tests are conducted when an inversion exists, the applicant should present a more rational correction method than either the “degree for degree” method described in paragraph 27.1.5.2 or the method defined in § 25.1043. The FAA has accepted a correction method that used the difference between the test day ambient and the hot day atmospheric temperature taken at the beginning of the flight being used as the correction factor for all flight data. Although conservative, this method accounts for the effects of an inversion.
27.1.2.11 Airport Altitude.
The cooling tests should be conducted from the lowest practical airport altitude, usually below 3,000 feet mean sea level (MSL), to provide test data reasonably close to sea level.
27.1.2.12 Winterization Equipment Procedures.
The following procedures should be applied when certificating winterization equipment:
27.1.2.12.1 Maximum Ambient Sea Level Atmospheric Temperature Less Than 100 °F (38 °C).
Cooling test results for winterization installations may be corrected to any temperature desired by the applicant rather than the conventional 100 °F (38 °C) hot day. For example, an applicant may choose to demonstrate cooling to comply with requirements for a 50 °F (10 °C) or 60 °F
(15.5 °C) 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 27-4
elected by the applicant with a rate of temperature drop of 3.6 °F (2 °C) per thousand feet above sea level.
27.1.2.12.2 Tests.
Cooling tests and temperature correction methods should be the same as for conventional cooling tests.
27.1.2.12.3 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.
27.1.2.12.4 Equipment Marking.
If practical, winterization equipment, 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.
27.1.2.12.5 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.
27.1.3 Ground Test Procedures.
27.1.3.1 General.
The flight testing guidance in paragraphs 27.1.2.1, 27.1.2.4, 27.1.2.5,
27.1.2.6, and 27.1.2.11 are equally applicable to ground testing.
27.1.3.2 Test Conditions.
Ground testing should be conducted with the engine operated at idle power or thrust until all critical temperatures have stabilized, followed by operation of the engine at rated takeoff power or thrust for 5 minutes, and then idle power or thrust until all critical temperatures have stabilized.
27.1.3.3 Temperature Stabilization.
During the ground operation portion of the compliance demonstration, the definition of stabilized temperatures defined for the flight test (rate of change less than 2 °F per minute (1 °C)) should not be used for determining the maximum component temperatures, unless it can be shown that ground operation of the engine is limited to the conditions tested. The reason for using a different definition of stabilized temperatures for ground operation is that during sustained ground operations at a particular condition, such as at idle power or thrust for using the engines as a pneumatic source, may result in temperatures that 27-5
gradually exceed the defined temperature limits. To address this concern, some manufacturers have used engine operating limitations for ground operations that limit engine operation for discrete time periods. For example, engine operation at takeoff power or thrust may be limited to 2 minutes at ambient temperatures above 110 °F (43 °C), or engine operation at idle power or thrust is limited to 2 hours at ambient temperatures above 110 °F.
27.1.3.4 Temperature Correction.
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.
27.1.4 Data Acquisition.
All of 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:
Outside air temperature (OAT).
Altitude.
Airspeed (knots).
Gas generator RPM.
Engine torque.
Time.
Fan RPM.
Engine oil temperature.
Pertinent engine temperature.
Pertinent nacelle and component temperatures.
27.1.5 Data Reduction.
27.1.5.1 Purpose.
Seldom is testing actually accomplished at the maximum required ambient temperature of at least 100 °F at sea level lapsed 3.6 °F (2 °C) per 1,000 feet pressure altitude. Component and fluid temperatures must therefore be corrected to derive the item temperature that would have been reached if the test day had matched exactly the maximum ambient temperature day. The applicant may select a higher maximum ambient temperature for cooling certification than the 100 °F (38 °C) sea level hot 27-6
day prescribed. Provisions are also made for selecting a maximum ambient temperature less than the 100 °F sea level hot day for winterization installations not intended to function at the hot day conditions.
27.1.5.2 Correction Factors for Ambient Conditions.
Unless a more rational method applies, a correction factor of 1.0 is applied to the temperature data as follows: co rrect ed temperature = true temperature + 1.0 [100 - 0.0036 (Hp) - true OAT]. A correction factor other than “degree-for-degree” should be based on engineering test data.
The corrected temperature is then compared with the maximum permissible tempe rature to determine compliance with the cooling requirements. No corrected temperatures may exceed establis hed limits.
Sample Calculation:
27.1.5.3 Correction Factor for Minimum Engine.
An important correction factor that is not discussed in the regulations but is frequently necessary to show the cooling adequacy required by § 25.1041, is the minimum engine (i.e., the thermal limit) correction factor. This factor is sometimes required if, at test day conditions, the engine measured temperature does not correspond to the engine temperature that would have occurred on a minimum specification engine in hot day conditions. The correction factor would not apply to those components not affected by changes in exhaust gas temperature (EGT) at a constant power or thrust. Typical items expected to be affected by changes in the EGT at constant power or thrust would be engine oil temperature, thermocouple harnesses, or other fluid, component, or ambient temperatures in the vicinity of the engine hot-section or exhaust gases. Other items remote from the hot section, like the starter-generator or fuel control, would not be expected to be influenced by EGT variations;
however, the items affected and the magnitude of the factor to be applied should be established by testing. There are several acceptable methods for establishing the appropriate correction factor during development testing.
The general idea is to establish a stabilized flight condition, typically during ground runs, and vary the measured EGT at approximately fixed power or thrust and OAT conditions. This may be accomplished by using engine anti-ice bleed air, customer bleed air, or by ingesting warmer than ambient air (either an external source or the engine bleed air) into the engine inlet. Care should be used when ingesting warmer than ambient air to assure that the warm air is diffused in order to avoid possible engine 27-7
surge. If it is not possible to attain an adequate variation in EGT by these methods, an acceptable, but conservative correction may be obtained by allowing both power or thrust and EGT to vary at stabilized engine operating conditions and OAT. The component temperature is plotted as a function of EGT, and the correction from test EGT for any flight condition, to the EGT that would have existed with minimum specification engines on a hot day, is then applied to derive the corrected component temperature. Both of these methods assume the inlet air and cooling air sources are essentially independent. Where they are not independent, such as those designs that take cooling air from the fan exit stream (e.g., via a fan box in bifurcation) rather than via a free stream scoop, care needs to be taken in selecting a technique to assure the results are neither overly conservative as may result from use of the hot air technique nor overly optimistic as may result from use of the higher power or thrust technique.
27.2 [Reserved.] 27-8
AC 25-7D CHG 1 CHAPTER 28. POWERPLANT: INDUCTION SYSTEM
28.1 Air Induction—§ 25.1091.
28.1.1 Explanation—§ 25.1091(d)(2).
The turbine engines of transport category airplanes are susceptible to surge, stall, and flameout when excessive amounts of water are ingested. The certification requirements for turbine engines include the demonstration of a capability to operate in simulated rainfall with no adverse operating effects. The quantities of water spray that may be directed toward the engine inlets, resulting from the airplane passing through standing water on taxi and runways, may exceed that used in the simulated rainfall ingestion testing. This becomes particularly important during takeoff, where the engines are operated at high power or thrust settings and the airplane will experience a wide range of speeds. During takeoff and landing ground rolls, the airplane’s tires generate bow waves, side spray, and “rooster tails” (spray that is thrown off the tires as they rotate), which can collect into concentrated streams of water that, if ingested into the engines, APU, or air conditioning systems, could affect operation sufficiently to cause an unsafe condition. Similarly, water ingestion into the airspeed system during ground operations may cause errant cockpit indications of airspeed. It is often practical to investigate these effects concurrently with the engine water ingestion evaluation.
28.1.2 References.
AC 20-124.
AC 91-79B.
28.1.3 Procedures.
28.1.3.1 Method of Compliance.
The applicant may show compliance with the requirements of § 25.1091(d)(2) by test or by reasonable analysis. The analytical approach may be acceptable in cases where an airplane and/or its engines were modified, but the overall geometry and configuration remained unchanged.
28.1.3.2 Water Depth.
Takeoffs should not be attempted when the depth of standing water, slush, or wet snow is greater than one-half inch over an appreciable part of the runway. (See AC 91-79A). Therefore, one-half inch of standing water is the accepted criteria for conducting tests to demonstrate compliance with the water ingestion requirements of § 25.1091(d)(2). Testing may be conducted in specially constructed water test beds, where the selected water depth is maintained over not less than 90 percent of the test bed area. If the airplane successfully completes testing with a one-half inch water depth, no operating limitations will be imposed for wet runway operations.
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28.1.3.3 Airplane Configuration.
Testing should be conducted to demonstrate all critical phases of taxi, takeoff, and landing operations. High lift devices and landing gear doors should be in the position appropriate to each phase of operation tested. All portions of the airplane that may affect water spray patterns should be in the production configuration desired for approval (e.g., chine tires, fenders). The most critical powerplant operating configuration (bleed air, electrical load, ignition system, etc.) that is consistent with the applicant’s recommended procedures for operation on wet runways should be used.
28.1.3.4 Test Facility.
The test facility needs to have adequate space for simulated takeoff and landing runs. Intermediate dams may be used in the water test bed to maintain the selected water depth over not less than 90 percent of the test bed area. The test bed length should not be less than that required to produce a spray pattern of one second duration at critical test speeds. The test bed width need only accommodate one landing gear, if acceptable data are presented to the FAA that show the combined nose and main gear spray is not more hazardous than the spray produced by either gear separately.
28.1.3.5 Test Procedures.
The tests should be conducted with the appropriate power or thrust setting.
To establish the critical speed ranges, test runs should be made through the water test bed in speed increments of not more than 20 knots. If the rotated airplane attitude is suspected to be critical, at least one run at VR should be conducted with the airplane rotated to the normal rotation pitch attitude. If thrust reversers or reversing propellers are provided, testing should also be conducted with those systems operating.
28.1.3.6 Test Data.
The applicant is expected to provide personnel and equipment to collect the following data:
28.1.3.6.1 Test Site Data.
The outside air temperature, wind velocity, and water depth should be recorded before each test run.
28.1.3.6.2 Airplane Performance Data.
The airplane velocity should be recorded when it enters and exits the water test bed. Engine rotor speeds and interstage turbine temperature (ITT) or EGT, as applicable) should be recorded and any engine abnormal engine sounds noted.
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28.1.3.6.3 Water Spray Pattern.
Appropriate areas of the airplane should be coated with agents that will permit identification of water spray impingement. Suitable high-speed photography or video equipment should be used to record the origin, trajectory, and configuration of the water spray.
28.1.3.6.4 Pilot Comments.
Pilot comments should be noted to determine if the takeoff would have been affected by abnormal audible engine sounds, airplane instrument anomalies, etc.
28.1.3.7 Test Results.
No hazardous quantities of water are considered to have been ingested when review of the collected test data shows that:
28.1.3.7.1 No engine flameout, performance degradation, distress, or airspeed fluctuations occurred that would create a safety hazard; and
28.1.3.7.2 No abnormal engine sounds occurred, such as pops or bangs, or cockpit instrument indications of incipient engine surge or stall, or they were not of sufficient magnitude to cause a pilot in service to abort the planned operation.
28.2 Induction System Icing Protection—§ 25.1093.
28.2.1 References.
AC 20-73A.
AC 20-147A.
28.2.2 Reciprocating Engines—§ 25.1093(a).
28.2.2.1 Explanation.
28.2.2.1.1 Conditions for Tests.
The carburetor air temperature measurement has been found to provide satisfactory average readings through the use of a minimum of three thermocouples so arranged as to give an average air temperature. This indicator should be calibrated prior to the test. Operationally, it has been determined that the tests should be conducted at an altitude where the free air temperature is 30 °F (-1 °C), or at two altitudes of different temperatures, one of which is near 30 °F (-1 °C).
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28.2.2.1.2 Configuration.
The test should be conducted in the configuration that follows:
Weight—optional.
CG Position—optional.
Wing Flap Position—optional.
Landing Gear Position—optional.
Engines—60 percent maximum continuous power.
Cowl Flaps—appropriate for flight condition.
Mixture Setting—normal cruising position.
28.2.2.2 Procedures.
Test procedures and required data.
28.2.2.2.1 After all temperatures have stabilized (i.e., when the rate of temperature change is less than 2 °F (1 °C) per minute), and with the airplane in level flight and full cold carburetor at 60 percent maximum continuous power, the following data should be recorded:
Pressure altitude.
Ambient air temperature.
Indicated airspeed.
Carburetor air temperature.
Engines’ RPM and manifold pressure.
Torque pressure.
Mixture setting.
Cowl flap setting.
28.2.2.2.2 Preheat should then be applied slowly (power may be restored to 60 percent maximum continuous at the applicant’s option) and the above data recorded again after the temperatures have stabilized. The carburetor heat rise is determined by comparing the results of the data obtained with and without preheat.
28.2.3 Turbine Powered Airplanes—§ 25.1093(b).
28.2.3.1 Explanation.
28.2.3.1.1 Section 25.1093(b) requires that each turbine engine operate throughout its flight power or thrust range without adverse effect on engine operation or serious loss of power or thrust in the icing conditions specified, including 28-4