516A. § 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.
517. § 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 676 Par 516A
normally used in the flight stage for which the cooling tests are conducted. The 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 oll 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 Par 517 677
thermal limit correction factor. This factor is sometimes used if, at test day conditions, 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.
(i) 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.
(її) 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 GE cooling test point at sea level.
Corresponding Power (SHP) MGI (°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 850 600 678 Par 517
(i) 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 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.
(i¡) 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.
518. § 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 Par 517 679
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 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.
518A. § 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 518 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.
519.-530. RESERVED.
680 (thru 692) Par 518
SECTION 30. INDUCTION SYSTEM
531. § 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.
(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 operation, it is essential that the airflow be smooth, uniform, clean, and unrestricted throughout the very wide range of horsepower expected from the engine.
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531A. § 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.
532. § 27.1093 (Amendment 27-20) INDUCTION SYSTEM ICING PROTECTION.
a. Reciprocating Engines.
(1) Explanation.
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) 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 and relative humidity of 30 percent. At low cruise power, ice can occur at outside air temperatures as high as 62° F 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 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.
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(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 OAT is 30° F and a power setting of 75 percent maximum continuous power can be maintained. If this combination cannot be achieved, satisfactory interpolation of data from other test conditions can be achieved using the methodology of AC 23-8A, Flight Test Guide for Certification of Part 23 Airplanes.
In addition to the preheat requirements of § 27.1093(a)(1), (a)(3), and
(a)(4), the design should consider the possibility of impact ice (supercooled droplets below freezing) on engine air inlet components opening into the airstream. However, normal practice is to provide a crew selectable, sheltered, alternate engine air intake arrangement which, by inspection can be determined to be free from impact ice accumulations. Typically, a sheltered alternate air source would be acceptable if the opening is located inside the cowling out of the free airstream. However, precautions should be taken to ensure that backfire flames to be expected do not constitute a fire or explosion hazard.
(ii) For further information review AC 20-113, Pilot Precautions and Procedures to be Taken in Preventing Aircraft Reciprocating Engine Induction System and Fuel System Icing Problems, NACA TN 1790, NTSB AAS-72-1, DOT/FAA/CT 84/44 (1982) and NACA TN 1993 (1949).
b. Turbine Engines - Ice Protection.
(1) Explanation.
This rule requires turbine engines and turbine-engine inlets to perform satisfactorily in atmospheric icing conditions defined in Appendix C of Part 25. On an equivalent safety basis, the limited icing envelopes described in AC 29-2A, Paragraph 386 may be used to show compliance with the intent of the regulation if the rotorcraft is limited to not greater than a 10,000-foot pressure altitude for all operations.
If operations are permitted above 10,000 feet, the Appendix C, Part 25, envelope must be used from 10,000 feet to the service ceiling or 22,000 feet. These possible equivalent safety approaches are not discussed herein. Compliance with the induction system icing protection rule is required regardless of flight manual limitations or restrictions against flight into atmospheric icing conditions.
(il) In showing compliance with § 27.1093(b)(1)(i), the FAA/AUTHORITY has accepted the concept of limited exposure associated with escape from inadvertent ice encounters.
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(A) Under the concept of limited exposure associated with escape from inadvertent ice encounters, it is presumed that there will be a flight manual limitation against flight into known icing, and that the engine induction system will be reevaluated if total aircraft ice protection certification is requested. Under this concept, the rotorcraft is assumed to fly directly through the icing environment; i.e., direct sequential penetration and straight line exit from both the continuous maximum and intermittent maximum icing clouds. Thus, the duration of exposure to the icing environment could be calculated by knowing the aircraft flight speed and cloud horizontal extent. A range of engine power and rotorcraft airspeeds should be evaluated to encompass the operating envelope of the rotorcraft. Note that aircraft speed has a pronounced effect (Ludlam effect) on ice accretion on small surface areas (inlet screens). A review of this phenomena may be found in National Research Council (Canada) Letter LT-92.
(B) When this limited exposure concept is used, the aircraft type certificate data sheet should clearly specify that the engine induction system must be reevaluated if certification to the general ice protection regulation, § 27.877 or § 27.1419, is requested. This direct penetration and exit approach is inappropriate for aircraft for which full icing clearance is requested (reference § 27.1419).
(iii) Engine induction system continuous icing protection would be necessary for aircraft for which full-icing clearance is requested (reference § 27.1419(d)). The approach is much preferred for all programs in order to reduce the scope of any eventual total aircraft icing program effort and to increase the safety level in conducting the rotorcraft natural icing tests. Since at least one rotorcraft has been FAA/AUTHORITY certificated to operate in known icing conditions and others have active development programs to this end, applicants should anticipate eventual full-icing clearance and consider that the engine induction system may be required to operate routinely in a continuous icing environment.
(iv) It is noted in AC 29-2A, Paragraph 386 that some natural icing tests are required to show compliance with the overall rotorcraft ice protection requirements.
It is not required that the engine induction system be evaluated as a part of that natural icing test if adequate verification has been shown by tunnel testing, analysis, or other means to ensure satisfactory operation in an extended continuous icing environment.
If, however, subsequent rotorcraft natural icing testing shows unanticipated detrimental engine inlet effects, the inlet ice protection system should be reexamined.
(v) The regulation specifies the examination of flight idling conditions.
This requirement is normally associated with a low-power letdown at the minimum practical forward airspeed. Alternatively, evaluation of the minimum power and minimum airspeed combination specified in the RFM for operation in visible moisture when below 40° F will accomplish the intent of the idling requirement.
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(vi) An acceptable approach to a finding of compliance would be a combination of analysis of the performance of the ice protection system which covers the range of the applicable icing flight envelope (maximum altitude, minimum temperature, etc., of the basic rotorcraft) supported and validated by tests. Ideally, these tests would be conducted in natural atmospheric ice with special instrumentation for droplet size and liquid water content. In practice, however, natural icing testing may pose unacceptably severe problems since rotorcraft may not have the range and speed to reasonably find icing clouds and may not be equipped with the airtrame and rotor ice protection needed for safety during the testing.
(vii) Problems with analysis emerge if engine inlets incorporate screens, turning vanes, sideward or upward openings, and edge or lip configurations which deviate from the airfoil shapes assumed in most of the analytical procedures described in current technical literature. The applicant should recognize that if meaningful analytical methods are not available, extensive testing with significant conservatism or possibly design changes may be required. Inlet screens in particular, if not adequately heated, fall in this category and can only be accepted if shown by very conservative ice testing to not significantly impede airflow to the engine.
(viii) The icing evaluation should definitely include some test points or other adequate evaluation of flight at ambient temperatures several degrees Fahrenheit above freezing and with very high water content. This condition has actually produced multiengine flameouts in in-service aircraft. The actual icing phenomena involved is not fully understood and in some instances efforts to duplicate the phenomena in icing tunnels were unsuccessful. Usually this condition does not produce rotor icing;
therefore, actual flight testing using special precautions to ensure safe autorotation landings or engine relight capability may be needed to identify this condition.
(2) Procedures.
Review AC 29-2A, Paragraph 386, ADS-4, Report No. FAA-RD-77-767; Aircraft Icing Handbook, FAA Technical Report No. DOT/FAA/CT-88/8, and Advisory Circular 20-73, Aircraft Ice Protection. (The comparative concept described under Item 34 of AC 20-73 is obsolete and should not be considered. These data provide extensive description and methodology for evaluation of ice protection systems; however, as noted above, these data generally apply to near straight line droplet trajectory with impingement onto conventional airfoil shaped inlets. As such, the applicability of these data to rotorcraft engine inlet ducts is limited and may require extensive adjustment to accommodate the different inflow trajectories and shapes of rotorcraft.
(ii) An analysis, appropriate to the configuration; i.e., heated or unheated impingement surfaces, should be prepared. To be acceptable, this analysis should show the inlet to be adequately protected by heat, or if unheated, to show that the inlet Par 532 697
with ice accretions as predicted, will provide adequate airflow to the engine throughout the flight envelope of the rotorcraft.
(A) For heated surfaces, ADS-4 and Report No. FAA-RD-77-76 provide detailed suggestions on heat transfer analysis particularly applicable to bleed air heated inlet lips formed in airfoil shapes. These data are limited in applicability and may not be useful for analyzing engine inlet water droplet trajectories to be expected at low airspeed and high engine airflow. Actual icing tests may be needed to derive the impingement patterns for these conditions.
(1) Acceptability criteria for heated inlet ducts usually require sufficient heat to evaporate the water to be expected in a "continuous maximum" icing cloud and to anti-ice the duct during flight in "intermittent maximum" icing clouds, providing the run-back and refreeze to be expected does not cause additional airflow disruption or damage to the engine. Full-scale inlet icing tests with the engine installed and operating should be conducted to verify the analysis. Engine power changes which may be expected in service should be included in the testing. Wind tunnels equipped for icing tests probably are the most useful means of conducting these tests if natural icing tests are impractical. The rotor downwash effect should be considered to the extent possible by adjusting the inflow angle in the tunnel.
(2) The power loss (bleed air, generator load, etc.) attributable to the heating requirements will affect the performance of the rotorcraft. Normally, this may be accounted for by specifying a gross weight incremental deduction from the flight manual performance data for flight into visible moisture below 40° F.
(3) Special evaluation of the possibility of ice ingestion damage to the engine should be made for heated systems which considers the ice ingestion to be expected when the anti-ice system is actuated after a delay of 1 minute for the pilot to recognize that the rotorcraft has encountered ice. This time delay may be reduced if the crew is provided adequate distinctive cues to alert them that the rotorcraft has encountered icing conditions.
(B) For unheated inlets, an acceptable method for showing compliance would include an extensive, detailed analysis (which shows that ice accretions on and in the inlet do not seriously obstruct adequate airflow to the engine) and tests as necessary to validate the analysis. The analysis of ice accretion becomes even more questionable since the unheated inlet involves ice buildups which themselves progressively change shape during icing exposure.
(1) Flight testing with an instrumented rotorcraft in natural ice to verify the analysis is desirable; however, wind tunnel tests as discussed above may be used. Since unheated inlets typically continue to accrete ice as a function of exposure, both the analysis and the test should realistically consider the actual exposure to be expected in service. This should not be less than penetration of the continuous 698 Par 532
maximum icing cloud followed immediately by exposure to the intermittent maximum cloud for rotorcraft not certified for icing. Engine power changes which may be expected in service should be included in the testing, and a warm-up period at the conclusion of the icing exposure should be shown for some selected test points to evaluate potential ice breakaway and ingestion.
(2) For the nonicing certified rotorcraft using the limited icing exposure concept for inlet certification, some conservatism should be applied to account for the fact that inlet icing may occur without airframe icing, and that the escape procedure from this unapproved operating condition is not defined. A demonstration of 30-minute hold capability in the continuous maximum cloud would be acceptable. Alternatively, if positive cues (perhaps a carefully located ice detector) of potential inlet icing are provided to the crew, the time increment could be reduced to recognition plus 15 minutes (15-minute escape time after recognition is consistent with the single ice protection system failure recognition and escape guidance for aircraft ice protection systems in AC 29-2A, Paragraph 386). It should not be assumed that airframe icing will always be available as a cue to potential inlet icing. The main rotor, for example, may not show icing indications above 25° F, whereas some inlets may ice critically near 32° F ambient. A reduction of the acceptable 30-minute exposure should not be based on observation of ice accretions on protruding components which are likely to be changed. For example, a limited exposure inlet icing program which reduces the inlet icing exposure time based on crew recognition of icing on the windshield wipers may be invalidated at a later date if a new windscreen deletes the wipers.
(Ш) Inlet capability during IGE hover in icing conditions has not generally been considered for rotorcraft not certified for icing. However, the FAA/AUTHORITY is recently aware that some inlets may ice at zero airspeed near 32° F with no indications of airframe icing in the field of view of the crew. This special concern of operating within RFM limitations, and yet placing the induction system in jeopardy, may be addressed in several ways. If the induction system ice protection scheme is not dependent on airspeed for proper function, the issue may be addressed by tunnel testing with inlet airflows approximating hover with no particular attention to tunnel windspeed. For protection schemes which may be sensitive to airspeed (external screens have shown this tendency), actual hover demonstration at or near zero speed tunnel conditions may be appropriate. Icing detectors located to indicate induction system icing in hover may be an option to a hover icing protection demonstration. Recently, on an external screened configuration, the FAA/AUTHORITY has accepted a satisfactory IGE hover demonstration of 30 minutes at the critical ambient temperature (i.e., ambient consistent with no airframe icing but potential inlet icing), 0.6 grams/meter LWC, and 40-micron droplet size as an adequate response to this concern.
(iv) For aircraft requesting full icing approval, or for those electing to show continuous induction system icing protection, the forward flight icing exposure would not be less than that time required to stabilize any ice accretions observed during repeated cycles of the continuous maximum followed by intermittent maximum cloud exposure.
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