before the filter reaches a bypass condition. The indicator may be a pop-out button or other maintenance cue that is checked on each preflight inspection.
AC 27.1021. § 27.1021 OIL SYSTEM DRAINS.
a. Explanation. This regulation requires provisions be provided for safe drainage of the entire oil system with the rotorcraft at normal ground attitude and defines certain requirements for assuring that no inadvertent oil flow occurs from the system provided.
b. Procedures.
(1) The design of the oil system must provide a means for safe drainage of the entire oil system. This may require one or more drains depending on the design of the system. The routing of fluid lines should be such that drooping lines and fluid traps which are undrainable are avoided.
(2) The drain(s) must provide a means for a positive lock in the closed position.
The method by which the lock is accomplished may be manual or automatic.
AC 27.1027. § 27.1027 (Amendment 27-23) TRANSMISSION AND GEARBOXES:
GENERAL.
a. Explanation. Amendment 27-23 adds a new § 27.1027. This new section provides the regulations for rotorcraft transmission and gearbox lubrication systems. It incorporates lubrication system requirements that were derived from existing engine oil system requirements. These additional requirements have been adjusted or modified to reflect the needs of transmissions and gearboxes. Transmission and gearbox lubrication system regulations are similar to those for engines; therefore, reference is made to the engine lubrication sections as applicable.
b. Procedures.
(1) The pressure lubrication systems for rotorcraft transmissions and gearboxes should comply with the same requirements as the engine lubrication systems stipulated in §§ 27.1013 (except §§ 27.1013(c)), 27.1015, 27.1017, 27.1021, and 27.1337(d).
These sections provide the requirements for oil tanks, tank tests, oil lines and fittings, and oil system drains.
(2) Each pressure lubrication system for rotorcraft transmissions and gearboxes should have an oil strainer or filter. The strainer or filter should:
(i) Remove any contaminants from the lubricant which may damage the transmission, gearbox, or other drive system component and any contaminants that may impede the lubricant flow to a hazardous degree.
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Be equipped with a means to indicate that the bypass system (required by § 27.1027 (b)) is at the point of opening due to the collection of contaminants on the strainer or filter, and;
(iii) Be equipped with a bypass system that will permit lubricant to continue to flow at the normal rate if the strainer or filter is completely blocked. In addition, the bypass system should be designed so that contaminants that have collected on the filter will not enter the bypass flow path when the system is in the bypass mode.
(3) Section 27.1027(c) requires a screen at the outlet of each lubricant tank or sump that supplies lubrication to rotor drive systems and rotor drive system components. The screen should remove any object that might obstruct the flow of lubricant to the filter required by § 27.1027 (b). The requirements of § 27.1027(b) do not apply to the tank outlet screen.
(4) Splash-type lubrication systems for rotor drive system gearboxes should comply with §§ 27.1021 and 27.1337(d).
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SUBPART E - POWERPLANT
COOLING AC 27.1041. § 27.1041 (Amendment 27-2) COOLING--GENERAL.
a. Explanation. The rotorcraft design should provide for cooling to maintain the temperature of all powerplant and power transmission components and fluids within the limitations established for the items. Cooling provisions should be adequate for shutdown and for water, ground, and flight operating conditions. The adequacy of the cooling provisions should be demonstrated by flight testing.
b. Procedures.
(1) Test conditions and procedures necessary to demonstrate adequate cooling for water, ground, flight, and shutdown conditions should be agreed upon between the applicant and the FAA/AUTHORITY certification engineer. A cooling test proposal which defines the agreed test points and procedures should be prepared well in advance of the official certification testing.
(2) The test conditions selected would typically include climb, cruise, hover, and shutdown after a prolonged hover. Hover OGE should be evaluated if sling load operation is envisioned for the rotorcraft. One test condition which should be examined, particularly with regard to transmission cooling, is the point of highest multiengine mechanical power at the maximum ambient temperature. This is identified as test point "A" in figure AC 27.1041-1. The selection of test points should be tempered with engineering judgment and based on results from similar aircraft if such data are available. In showing compliance with the cooling requirements, the applicant should not be required to exceed rotorcraft established limits (gross weight, drive system torque, measured gas temperature, etc.), aircraft power required, or power available.
The applicant may elect, however, to exceed these limits in order to minimize test points by conservative testing, or to anticipate future growth (increased gross weight, etc.).
(3) The need for a comprehensive cooling test plan prior to certification testing cannot be overemphasized. Highly derated engine installations, the relationship of power required to power available, the use of bleed air devices which would increase the measured gas temperature while aircraft power required remains the same, auxiliary cooling provisions, and the increase in engine temperatures with engine deterioration are factors which could affect the selection of cooling demonstration test points.
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"MULTIENGINE POWER AVAILABLE 20K г
HOT DAY 15K- XMSN LIMIT PRESSURE ALTITUDE OAT 10 C ЛОК 15° C 20° C 25° C 30° C 35° C POWER POINT "A" FIGURE AC 27 1041-1 ADDITIONAL COOLING TEST POINT Page E - 93
AC 27.1041A. § 27.1041 (Amendment 27-23) GENERAL.
a. Explanation. Amendment 27-23 provides clarification and definition of powerplant components required to be considered when evaluating the performance of the powerplant cooling systems and arrangements.
b. Procedures. The rotorcraft design should provide for cooling to maintain the temperature of all powerplant and power transmission components and fluids within the limitations established for the items. Components to be considered include, but are not limited to, engines, rotor drive system components, auxiliary power units, and the cooling or lubricating fluids used with these components.
AC 27.1043. § 27.1043 (Amendment 27-14) COOLING TESTS.
a. Explanation.
(1) This section defines the requirements for accomplishing the required cooling tests. Section 27.1043(a)(1) requires that certain ambient temperature correction factors be applied unless testing is accomplished at the maximum ambient atmospheric temperature prescribed. No corrected temperature may exceed specified limits. The requirement in § 27.1043(a)(4) that test procedures be in accordance with § 27.1045 does not limit testing to the conditions prescribed in that section.
Section 27.1041(a) provides the basis for examination of other possible critical operating and shutdown conditions.
(2) This section establishes the hot-day condition as 100° F at sea level, decreasing 3.6° F per 1,000 feet of altitude above sea level. The applicant may select a lower maximum ambient atmospheric temperature for winterization installations. If the cooling tests are conducted under conditions deviating from the maximum anticipated air temperature, then the following correction factors are required unless another FAA/AUTHORITY-approved method is applicable.
(3) The temperature of engine fluids and powerplant components (except cylinder barrels) which have established limits must be corrected by adding to them the difference between the maximum anticipated air temperature and the ambient air temperature at the time of the first occurrence of the maximum component or fluid temperatures recorded during the cooling tests.
(4) Cylinder barrel temperatures must be corrected by adding 0.7 of the difference between the maximum anticipated air temperature and the ambient air temperature at the time of the first account of the maximum cylinder barrel temperature recorded during the cooling tests.
(5) During the cooling tests for reciprocating engines, the fuel used must be of the minimum grade approved for the engine and the mixture settings should be those normally used in the flight stage for which the cooling tests are conducted. The Page E - 94
carburetor parts list used during these tests becomes a requirement in the definition of the engine/carburetor configuration.
b. Procedures.
(1) Seldom is testing actually accomplished at the maximum required ambient temperature of at least 100° F at sea level lapsed 3.6° F 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 sea level hot 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.
(2) When cooling test ambient conditions are cooler than the selected or prescribed hot day conditions, the applicant may take advantage of cooling air or fluid flows that would exist at hot day conditions. For example, thermostatically controlled oil cooler flow could be set for hot day conditions.
(3) The component and fluid temperature correction factor to be applied when test ambients do not correspond to the hot day conditions is commonly called the "degree-for-degree correction." It may be possible to justify, and the regulation allows the application of a more refined, less conservative correction factor. A correction factor other than degree-for-degree should be based on engineering test data.
(4) No corrected temperatures may exceed established limits. In order to maintain temperatures within established limits, the applicant may be willing to accept lesser performance than the full capability of a device. For example, a starter/generator capable of cooling under test cell conditions to 200 amperes continuous load may be limited to a lesser value, perhaps to 150 amperes, when installed in the aircraft due to cooling considerations. This continuous load for cooling must be equal to or greater than the allowable continuous load designated on aircraft instruments.
(5) If the engine or transmission cooling system heat load is increased in any way by rotorcraft configuration changes (affecting airflow, etc.), by other systems, by accessories (alternators, generators, etc.), or by any other heat source or potential heat source, then the maximum cumulative heat load from the worst-case combination of all these sources which is possible in service must be present during the cooling tests.
c. Thermal Limit Correction.
(1) An important correction factor which is not discussed in the regulations, but is frequently necessary to show the cooling adequacy required by § 27.1041, is the thermal limit correction factor. This factor is sometimes used if, at test day conditions, Page E - 95
the engine measured gas temperature does not correspond to that which would have occurred on a minimum specification engine at hot day conditions.
(2) The correction factor would not apply to those components not affected by changes in measured gas temperature (MGT) at a constant power. Typical items expected to be affected by changes in the MGT at constant power 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, perhaps the starter-generator or fuel control, would not be expected to be influenced by MGT variations; however, the items affected and the magnitude of the factor to be applied should be established by testing.
(3) There are several acceptable methods for establishing the appropriate thermal limit correction factor during development testing. The general idea is to establish a stabilized flight condition, typically ground-run or IGE hover, and to vary the measured gas temperature at approximately fixed power and OAT conditions. This may be accomplished by utilizing 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 in ingesting warmer than ambient air to ensure that the warm air is diffused in order to avoid possible engine surge.
If it is not possible to attain a suitable variation in MGT by these methods, an acceptable, but more conservative, thermal limit correction may be obtained by allowing both shaft horsepower and MGT to vary at a stabilized flight condition and OAT.
(ii) The component temperature is plotted as a function of MGT, and the thermal limit correction from any test day MGT for any flight condition, to the MGT that would have existed with minimum specification engines on a hot day, is then applied to derive the final measured component temperature.
(4) In certain rare instances, it may not be required that the correction factor be applied to the full thermal limit capability of the engine. Consider the following example for the hot day hover IGE cooling test point at sea level.
Corresponding Power (SHP) MGT (°C) Drive System Limit 900 Twin-Engine Hot Day Power Available 1050 750 Hot Day Power Required at Maximum G.W. 850 650 Engine Maximum Allowable MGT (Instrument Marking) 765 Test Day (90° F OAT) Parameters 850 600
(1) Notice that the installed hot day power available MGT from the engine performance program is 15° C cooler than the limit MGT (750 vs. 765° C), thus the Page E - 96
engine has 15° C "field margin" which would allow the engine temperature to gradually increase 15° C to maintain a given power as engine life is utilized. Secondly, the measured gas temperature corresponding to hot day power required at maximum gross weight is less than that corresponding to either the drive system limit or twin-engine hot day power available. Thus, the thermal limit correction could be applied from the test day MGT, 600° C, to the power required MGT plus the field margin, 650° C plus 15° C, rather than applying the correction factor to the full thermal capability of the engine, 765° C.
(її) Care should be used in applying this relieving method, because as the hover altitude changes, the maximum gross weight and power required (and the associated MGT) will vary. The data must be corrected to at least the maximum MGT for a minimum specification engine that can occur in service at the flight condition under investigation.
AC 27.1045. § 27.1045 COOLING TEST PROCEDURES.
a. Explanation.
(1) Section 27.1045(a) requires that cooling tests be conducted for the rotorcraft in the configurations and under the conditions most critical for cooling.
(2) Section 27.1045(b) requires that a temperature be stabilized prior to start of a cooling test for any test rotorcraft and any test stage. This is to ensure that the system reaches the maximum temperature from which it must be cooled. Temperature stabilization is achieved when the rate of change is less than 2° F per minute.
Therefore, for each test rotorcraft in each test phase, the temperature must be stabilized prior to entry into flight test. If temperature stabilization cannot be achieved as a normal result of the entry condition, then operation through the full entry condition must be accomplished prior to entry into the flight test segment being conducted. This allows the temperatures to reach their maximum natural levels prior to test initiation. Also, for each test rotorcraft during the takeoff stage of flight, the climb at takeoff power must be preceded by a hover during which temperature stabilization is achieved.
(3) Section 27.1045(c) requires that a test must be conducted for each flight stage until either the temperatures stabilize, the flight test stage is completed, or an operating limitation is reached.
b. Procedures.
(1) To comply with § 27.1045(a), an applicant typically submits a cooling test proposal to the FAA/AUTHORITY for approval. The proposal should encompass detailed procedures to demonstrate cooling capability for each critical rotorcraft configuration and test condition (test point). If a single-most critical test configuration and condition is not readily identifiable (which is usually the case), then a series of cooling tests must be conducted. Typical cooling test segments are climb, takeoff and Page E - 97
climb, various cruise speeds and altitudes, hover, shutdown after prolonged hover, and sling load cooling if a sling is used. Any other appropriate test conditions and procedures necessary to demonstrate adequate cooling for water, ground, flight, emergency, and shutdown conditions should be addressed in the test proposal. For multiengine rotorcraft (particularly in regard to transmission cooling), one test point that should be investigated is the point of highest multiengine mechanical power at the maximum ambient temperature. Other significant test conditions to be considered for multiengine rotorcraft are the OEI test conditions. The selection of all test points should be tempered with engineering judgment and should consider test points and procedures used on previous, similar rotorcraft certification work, if available.
(2) Compliance with § 27.1045(b) is typically shown by use of existing cockpit instrumentation or add-on test instrumentation from which temperature data are read prior to and during test segments. Test plans should clearly identify what is to be used and who is authorized to make and record readings and the accuracy and current calibration requirements for test instrumentation.
(3) Compliance with § 27.1045(c) is typically shown during FAA/AUTHORITY authorized testing by conducting each test segment until at least one of the three criteria (temperature stabilization, flight test segment completion, or an operating limitation) is reached. If an adverse operating limitation is reached, such as overheating, or the test cannot otherwise be successfully completed, then compliance has not been shown and a reevaluation is required.
AC 27.1045A. § 27.1045 (Amendment 27-23) COOLING TEST PROCEDURES.
a. Explanation. Amendment 27-23 clarifies acceptance criteria for the powerplant cooling tests that are appropriate if, during the cooling test, component temperatures peak and then decline rather than stabilize. In these instances, the previous requirement to continue the test until "stabilization" occurred was unduly restrictive and was eliminated.
b. Procedures. All of the policy material pertaining to this section remains in effect except that the engine fluid temperatures do not have to stabilize.
Paragraph AC 27.1045 currently lists three criteria for test completion: temperature stabilization, flight test segment completion, or an operation limitation. With Amendment 27-23, a fourth criteria for test completion is: 5 minutes after the peak temperature is reached, the test can be considered complete.
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SUBPART E - POWERPLANT
INDUCTION SYSTEM AC 27.1091. § 27.1091 (Amendment 27-2) AIR INDUCTION.
a. Explanation.
(1) The air induction system for each engine should be of a configuration to supply the air required under the operating conditions for which certification is required.
(2) The intake system shall be designed such that if a backfire flame occurs, it will emerge outside the engine compartment cowling.
(3) Where required in the induction system, drains must be provided which discharge clear of the rotorcraft and out of the path of exhaust flames.
(4) For rotorcraft powered by a turbine, the inlets should be located or protected to minimize foreign object ingestion as defined in the regulation. The inlets must be protected during takeoff, landing, and taxiing. There must also be means to prevent leakage of hazardous amounts of flammable fluids from entering the engine intake system.
b. Procedures.
(1) For turbine-engine installation, the induction system should supply air of suitable quality to meet the installation requirements of the engine manufacturers. The installation requirements should be met throughout the operating envelope of the rotorcraft. In addition, the design and location of the air induction system should prevent accumulations of rain or hail, either external or internal to the induction system, that could adversely affect engine operation.
(2) The inlet design should account for the prevention of hazardous fluids entering the engine. Some designs will have inlet ducts which are free from any fluid lines; however, other designs may route the engine inlet air through a compartment which has flammable fluid lines. When the condition exists, test demonstrations of critical leakage during operations have been used to substantiate the installation. The fluid leakage may not have an adverse effect on engine operation.
(3) The air induction system design should also account for and minimize the possibility of foreign matter ingestion during takeoff, landing, and taxiing.
(4) For reciprocating engine installations, the induction system should supply air of suitable quality and quantity to the carburetor inlet of the engine. The condition of the air at the entering face of the carburetor is extremely important. For proper Page E - 99
AC 27-1B, Chg 4 3/31/2014 operation, it is essential that the airflow be smooth, uniform, clean, and unrestricted throughout the very wide range of horsepower expected from the engine.
AC 27.1091A. § 27.1091 (Amendment 27-23) AIR INDUCTION.
a. Explanation. Amendment 27-23 removed § 27.1091(d) since the specific test defined by this paragraph was not critical for certain rotorcraft. The turbine inlet foreignobject-ingestion protection provided by § 27.1091(d) is adequately evaluated by existing requirements in § 27.1091(e)(2).
b. Procedures. This rule change did not change the current suggested methods of compliance.
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AC 27.1093. § 27.1093 (Amendment 27-20) INDUCTION SYSTEM ICING PROTECTION.
a. Reciprocating Engines.
(1) Explanation.
(i) Atmospheric moisture, even in clean air and temperatures above freezing, can result in ice accumulations in induction systems to a degree which can easily cause engine failure.
(A) Impact Ice. This forms as supercooled water droplets impact on engine induction system components. Particularly heavy accumulations must be expected where bends or turns in the induction system force changes in the airflow direction thus centrifuging the droplets out of the air stream where they freeze on impact with induction system components. A serious form of impact ice is the collection of ice on fuel metering elements of the carburetor, the alternate (preheat) valve, and any screens in the system.
(B) Throttle Ice. This type of ice forms at or near the throttle in a partly closed position [up to 30°F (-1°C)] due to cooling effect resulting from the increase in kinetic energy (increased velocity) of the air in the restricted flow area.
(C) Refrigeration Ice. This forms as a result of the cooling effect of the fuel evaporating after the fuel is introduced into the airstream. For some float type carburetors, it is possible in rare instances to accumulate serious ice during a closed throttle glide with ambient air temperatures as high as 93°F (+34°C), and relative humidity of 30 percent. At low cruise power, ice can occur at outside air temperatures as high as 62°F (+17°C) and relative humidities as low as 60 percent. Most of the heat necessary to evaporate fuel is supplied from the air as it drops in temperature. Fuel evaporation ice can affect airflow by blocking the throat of the manifold riser; it can fect the fuel al ratio br mixture affect the fuel-air-ratio by interfering with the fuel flow; and it can affect mixture distribution or quantity of mixture flowing to individual cylinders by upsetting the fuel flow distribution, or quantity of mixture flowing to individual cylinders by upsetting the fuel flow distribution at the fuel nozzle or airflow distribution in the manifold throat. This refrigeration phenomenon is the most serious of all factors causing carburetor ice.
(2) Procedures. Normally, flight tests with carburetor air temperature instrumentation are required. Unless otherwise justified, conduct all tests at maximum gross weight, a median center of gravity, in level flight, and at the engine speed which, considering cooling fan effects, if any, produces the minimum heat to the carburetor muff or engine component area utilized to provide the carburetor air heat. The optimum test condition is flight at the altitude at which the measured outside air temperature (OAT) is 30°F (-1°C) and when a power setting of 75 percent maximum continuous power can be maintained. If this combination cannot be achieved, satisfactory Page E - 101