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.
Par 532 699
Typically, any ice accretions resulting from these repeated cycles would be expected to stabilize in less than 30 minutes. The 30-minute hold capability in the continuous maximum icing environment could thus be ensured without special testing by careful selection of the test points for this repeated cycle.
(v) A rotorcraft requesting full icing approval should also have hover capability in the icing environment. Intermittent maximum icing conditions are not likely to exist near ground level and a satisfactory demonstration could involve the ability to hover indefinitely in the continuous maximum icing environment. Alternatively, carefully worded RFM limitations to restrict hover time may be acceptable if the system is not capable of indefinite exposure. Hover capability verification may not involve zero airspeed demonstration if the inlet protection system is insensitive to rotorcraft airspeed.
(vi) The engine(s) must be installed or protected to avoid engine damage from ice ingestion due to ice accretion in the inlet or on other parts of the rotorcraft, including the rotors, which may break away to enter the inlet. If screens or bypass arrangements are provided for these purposes, they should be included in the icing tests and shown by test or rational analysis to effectively protect the engine.
(vii) For unheated inlets, significant ice accumulations to be expected on the inlet may adversely affect the engine stall margin, acceleration characteristics, duct loss, etc. Dry air flight tests to evaluate these aspects can be accomplished by affixing ice shapes to the inlet. These shapes should closely match the actual ice shapes defined by test or analysis. In addition, it should be determined that ice shedding into the engine inlet either during continued flight into icing conditions or after emerging from the icing environment does not damage engine compressor on other inlet components.
c. Turbine Engines - Snow Protection.
(1) Explanation.
(• Section 27.1093(b)(1)(ii) provides that the turbine engine and its air inlet system operate satisfactorily within the limitations established for the rotorcraft, in both falling and blowing snow. The section does not provide the definition of falling and blowing snow.
(il) Since the regulation provides for certification "within the limitations established for the rotorcraft," the FAA/AUTHORITY can accept a restriction against snow operations in the limitations section of the RFM in lieu of demonstration of compliance. If no restriction on snow operations appears in the RFM, it is presumed that the aircraft may operate in snow at the pilot's discretion.
(2) Acceptance Criteria.
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The FAA/AUTHORITY has accepted that engine induction system operation in falling and blowing snow can be approved without restriction if normal operations under the following conditions are demonstrated:
Visibility: one-quarter mile or less as limited by snow.
Temperature: 25° F to 34° F (28° F to 34° F desired), unless other temperatures are deemed critical.
Operations: Ground operations - 20 minutes IGE hover - 5 minutes Level flight - 1 hour Descent and landing (її) RFM visibility restrictions for falling and blowing snow operations are not appropriate.
(iii) Time limitations, other than possibly for ground and hover operations, are not appropriate.
(iv) Artificially produced snow should not be used as the sole means of showing compliance.
(3) Rationale.
The test conditions specified--visibility, temperature, and operations-- are based on previous certification programs, previous FAA/AUTHORITY guidance, and on research by the FAA technical center and others.
(A) Visibility. The test visibility defined, 1/4-mile visibility or less as limited by snow, represents a heavy snowstorm and is the maximum likely to be encountered in service. Rotorcraft which have been certified to the 1/4-mile visibility test criteria have not shown engine inlet snow-related service difficulties. It is important to note that the visibility specified is a test parameter rather than an operational limitation to be imposed on the rotorcraft after the tests are completed.
B) Temperature.
(1) The ambient temperature specified is conducive to wet snow conditions. Wet snow tends to accumulate on unheated surfaces subject to impingement.
(2) Colder ambients, more conducive to dry snow conditions, may be critical for some induction systems. Colder exterior surfaces may be bypassed, and Par 532 701
the snow crystals may stick to partially heated interior surfaces where partial melting and refreezing may occur.
(3) Company development testing or experience with very similar type induction systems may be adequate to determine the critical ambient conditions for certification testing.
(C) Operations.
(1) Ground running, taxiing, and IGE hover operations are generally the most critical since the rotorcraft may be operating in recirculating snow. Twenty-five minutes under these extreme conditions would seem a reasonable maximum, both from the view of pilot stress and the maximum expected taxi time prior to takeoff in bad weather.
One hour of level flight operation under ¼-mile visibility snow conditions should provide ample opportunity for hazardous accumulations to begin to build.
(3) The descent and landing will provide an engine power change, an induction system airflow change, and a variation in the external airflow pattern near the induction system entrance. The initiation of the descent and final flare for landing may also produce additional airframe vibration transmitted to the induction system.
These power, airflow, and vibration changes may provide an opportunity for any level flight accumulations to be ingested into the engine. Hazardous accumulations are not acceptable during or after any test phase.
(ii) Visibility may fluctuate rapidly in snowstorms. It is affected by the presence of fog or ice crystals, is not crew measured or controlled, and is difficult to estimate. A visibility operational limitation based on snow, therefore, is not appropriate.
(ili) Since during cruise in snow conditions the aircraft is likely to be in and out of heavy snowfall, it is not practical for the crew to account for the time spent in snow in level flight conditions. Thus, it is not appropriate to include time limitations in the RFM for level flight snow operations.
(iv) A practical ground and IGE hover time limitation of less than 25 minutes in recirculating snow may be considered. The expected action at the expiration of this specified time period would be shut down and inspection of the inlet system or transition to a safe flight condition where demonstration has shown that moisture accumulations will not intensify or shed and cause engine operational problems.
(v) Artificially produced snow is an excellent development tool and has been successfully used to indicate potential problem areas in induction systems. These 702 Par 532
7/30/97 АС 27-1A devices are usually restricted to use for hover and ground evaluations, and the snow pellets produced by these machines are not sufficiently similar to natural snowflakes to justify the use of artificial snow as the sole basis of certification.
(4) Procedures.
(i) Satisfactory demonstration of the test conditions requires that the engine, induction system, and proximate cowling surfaces remain free of excessive snow, ice, or water accumulation. Excessive accumulation is defined as accumulation that may cause engine instability, damage, or significant loss of engine power. If a questionable amount of snow or moisture accumulates in the inlet, the applicant may elect to demonstrate that this amount in the form of snow or water and ice, as appropriate, can be ingested by the engine without incurring surge, flameout, or (її) The conditions specified assume actual flight demonstration in natural snow. The ground operations and IGE hover test conditions assume operation in recirculating snow. Blowing snow, resulting from rotor airflow recirculation, can be expected to be more severe than natural blowing snow if the rotorcraft continues to move slowly over freshly fallen snow. Thus, the blowing snow operational capability is usually demonstrated by the taxi and hover operations in recirculating snow.
For VFR rotorcraft, the airspeeds for the level flight test condition should include the maximum consistent with the visibility conditions. For IFR operations, the airspeed should be the maximum cruise speed or the maximum speed specified for snow operations in the flight manual limitations, unless other airspeeds are deemed more critical. It is recognized that many rotorcraft initially VFR certified are later IF certified with a resulting possible increase in airspeed in snow conditions.
This factor should be considered if IFR certification is anticipated.
(iv) The visibility specified assumes that visual measurements are made in falling snow in the absence of fog or recirculating snow by an observer at the test site outside the tests rotorcraft's area of influence. An accepted equation for relating this measured visibility to snow concentration is V = 374.9/C°7734 where C is the snow concentration (grams/meter*) and V is the visibility (meters).
(A) This equation can be reasonably applied to all snowflake type classifications and is credited to J.R. Stallabrass, National Research Council of Canada.
(B) Other equations may be applied if they are shown to be accurate for the particular snowflake types for the test program.
(v) The snow concentration corresponding to the 1/4-mile or less visibility prescribed will be extremely difficult to locate in nature. Data from Ottawa, Canada, Par 532 703
research indicate that fewer than 4 percent of the snowstorms encountered there meet the 0.91 grams/m' concentration associated with the 1/4-mile visibility. Furthermore, the likelihood that the desired concentration will exist for the duration of the testing is even more remote. Because of these testing realities, it is very likely that exact target test conditions will not be achieved. Those involved in certification must exercise good judgment in accepting alternate approaches.
(vi) For some engine induction systems, it may become apparent by inspecting for moisture accumulations that ground and IGE hover operations in recirculating snow are much more severe than the level flight test. In this instance, it is reasonable to accept prolonged IGE operations in recirculating snow and to accept durations of less than 1-hour level flight in 1/4-mile or less visibility. Best efforts should be made to ensure that at least some level flight time is accomplished at 1/4-mile or less visibility to ensure that the spectrum is covered.
(vil) It should be determined that the visibility established at the test sight is limited by snow and not by fog or poor lighting (twilight) conditions.
(viii) The concentration of snow approaching the inlet in severe recirculation will far exceed the quantity encountered in the natural snowfall.
Recirculation is necessarily a qualitative judgment by the test pilot. The snow concentration at the inlets during recirculation would vary for different rotorcraft types and would be dependent on rotor characteristics, power setting, and inlet location. For test purposes, recirculation should be the highest snow concentration attainable in the maneuver, or that corresponding to the lowest visibility at which (in the pilot's judgment) control of the rotorcraft is possible in the IGE condition. The 1/4-mile or less visibility specification outside of the recirculation influence becomes inconsequential provided that fresh, loose snow is continually experienced during the ground operation and GE hover testing phase. However, since it is intended that the test phases be accomplished sequentially to ensure that transition to takeoff and other transients are considered, the conditions at takeoff, level flight, and descent and landing should approximate the 1/4-mile visibility criteria.
d. Turbine Engines - Ground cing.
(1) Explanation. This requirement addresses the situation where extended ground operation in icing exposes the rotorcraft and its engine inlet to icing (ground fog) conditions which may have different droplet impingement patterns and involve different and/or less effective means of ice protection. Note that the requirement is effective in Amendment 10 and is applicable regardless of any desire to prohibit dispatch into known icing conditions.
(2) Procedure. Since this condition assumes zero airspeed, wind tunnel testing may be inappropriate unless conservative extrapolation of low speed tunnel data can be determined to be valid. For protection schemes which are dependent primarily on 704 Par 532
airspeed for proper functions (external screens have shown this tendency), it may be necessary to verify adequate ground operation protection capability by very low speed tunnels or by the use of outside facilities such as the Canadian National Research Council's spray rig at Ottawa, Canada. For heated systems or for internal bypass schemes, tunnel speed may not be important, and adequate demonstration may be accomplished at higher tunnel speeds provided that internal inlet airflows and heat available are properly considered. Testing should proximate the regulatory test conditions and be continued for 30 minutes using engine power and control manipulation as normally accepted during taxiway operations, followed by an acceleration to takeoff power. The test time may be shortened if deice/anti-ice protection is adequate or if stabilization of ice build-up is affirmed. The induction system should be in condition for safe flight at the conclusion of the test.
e. § 27.1093(c) Supercharged Reciprocating Engines.
(1) Explanation. This rule authorizes the designer to take credit for the heat-of-compression available downstream of an engine air inlet supercharger to meet the induction system heat rise requirements of § 27.1093(a) (3) or (a)(4), provided the heat rise is automatically available for the applicable altitude and operating condition.
(2) Procedures. Since a wide variety of superchargers and supercharger controls (waste-gate controls) have been devised, it is impracticable to outline specific instructions for determining compliance. However, the certification engineer can properly evaluate the arrangement by analyzing the system for trends (in heat rise available) and conducting measurements to verify these trends and quantify the actual values of heat rise available. Some factors to be considered are:
(i) Mechanically driven superchargers for rotorcraft usually operate over a very narrow speed range, thus the heat of compression (4 temperature) may remain constant over the altitude range. Conversely, turbosuperchargers usually are controlled (via waste-gate position modulation) to gradually increase the compression with increase in altitude, thus the critical (or lowest) 4 temperature may be available at very low altitudes.
(il) Other waste-gate controllers sense carburetor deck pressure and respond by modulating the waste gate to maintain a constant carburetor deck pressure within the capabilities of the turbo unit. Heat of compression may then vary with altitude and engine power in a complicated fashion such as to require experimental temperature measurements across a wide range of operating conditions to determine compliance.
(iii) Turbosuperchargers which are not controlled (by waste-gate modulation) but respond to an orificed exhaust generally will (at constant power) produce more heat rise at altitude than at sea level; however, size matching between Par 532 705
engine and turbo unit may affect this. Instrumented flight tests should be used as a final compliance verification method.
532A. § 27.1093 (Amendment 27-23) INDUCTION SYSTEM ICING PROTECTION.
a. Explanation. Amendment 27-23 clarifies that the phrase, "within the limitations established for the rotorcraft" applies only to the requirement in § 27.1093(b)(1) (i) for demonstrating flight in falling and blowing snow.
b. Procedures. All of the policy material for this section remains in effect with the update that turbine engines and turbine engine inlets should perform satisfactorily in atmospheric icing conditions defined in Appendix C of FAR 29 instead of FAR 25. In addition to the procedures of Paragraph 532, the following procedures should be followed:
(1) A "serious loss of power" in this section has been interpreted to be any power loss that requires immediate pilot action. In addition, the term "adverse effect on engine operation" in § 27.1093(b)(1) (ii) has been interpreted to be an effect that would prevent the engine from achieving rated aircraft flight manual performance (takeoff/climb/etc.). This term also includes effects on the engine induction system characteristics to an acceptable level established by the engine manufacturer (inlet distortion, etc.).
(2) It should be shown that rotorcraft that are prohibited from flight into falling and blowing snow can exit inadvertent entrance into those conditions without adverse effect upon the operating characteristics of the engine or the rotorcraft.
(3) For full flight capability into snow, both falling and blowing, it should be shown that each engine, and its inlet system, will operate satisfactorily throughout the flight power range of the engine and the operating limitations of the rotorcraft. It should be shown that any build-up or accumulation of snow will not reduce or block the flow of inlet air to the engine. Any accumulations that become dislodged should not affect engine operation.
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533.-547. RESERVED.
Intentionally Left Blank Par 533 707 (thru 721)
SECTION 31. EXHAUST SYSTEM
548. § 27.1121 (Amendment 27-12) GENERAL.
a. Explanation.
(1) This section addresses the arrangement of exhaust components and the protection against hazardous conditions which exist with hot exhaust gases.
(2) The objective is to allow for thermal expansion of manifolds and pipes, prevent local hot spots, and eliminate the possibility of igniting flammable fluids or vapors.
b. Procedures.
(1) Sufficient clearance of hot exhaust components must be maintained from structure, fuel cells, flammable fluid lines, and electrical components to compensate for thermal growth under normal and most extreme operating temperatures. Verify that adequate clearance exists between the exhaust system components and the surrounding structure, and that no interference occurs under the most adverse temperature excursions.
(2) Hot spots that can occur on fuselage or rotor blade skin as a result of impingement or in compartments due to an accumulation of hot gases should be eliminated with deflectors or by providing adequate flow-through ventilation.
Compliance may be shown by demonstration or analysis.
(3) It should not be possible to ingest sufficient quantities of exhaust gases which will produce engine surges, stalls, or flameouts during normal and emergency operation within the range of operating limitations of the aircraft and of the engine.
Analysis and/or flight testing may be required to demonstrate compliance. If flight testing is required, particular attention should be placed upon critical azimuths and wind conditions.
(4) Exhaust system surfaces hot enough to ignite flammable fluids or vapors must meet the isolation or shielding requirements of this section in addition to the requirements of §§ 27.1183 and 27.1185. Good design practice suggests that the isolation and shielding features incorporated would continue to be effective under the emergency landing conditions specified in § 27.561.
(5) It should be demonstrated that exhaust gases are discharged in such a manner that they do not cause distortion or glare which seriously affects pilot visibility at night. One method of compliance would be a night flight evaluation at critical azimuths and variable wind conditions to verify that no degradation exists.
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(6) Compliance with § 27.1121(f) can be accomplished by ensuring that the drain will discharge positively and is a minimum of 0.25 inches in diameter. No drain may discharge where it might cause a fire hazard. This can be demonstrated by discharging a colored liquid through the drain system in flight and on the ground. The dye should not impinge on any ignition source.
(7) Section 27.1121(g) is self-explanatory in specifying that a means must be provided to prevent blockage of the exhaust port after any internal heat exchanger failure. Compliance can be shown by demonstration or by analysis. In either case, it must be shown that any internal failure will not result in a significant power loss from the engine.
549. § 27.1123 (Amendment 27-11) EXHAUST PIPING.
a. Explanation. This section contains the following requirements that must be met for proper certification of exhaust piping on engines, auxiliary propulsion units (APU), and other similar devices.
(1) § 27.1123(a) requires that the piping be heat and corrosion resistant so that it performs its intended function during its operational life (either the life of the rotorcraft or a specified limited life) without significant metal corrosion, metal erosion, or creation of hazardous hot spots. The piping system should be designed, have an installation design, or a combination that allows performance of its function without thermal expansion (thermal strain) induced structural failures such as ruptures caused by operating temperature excursions and overpressurization during its operational life.
(2) § 27.1123(b) requires that the piping be supported to withstand the vibration and loading environment (including inertia loads) to which it will be subjected in service.
(3) § 27.1123(c) requires that piping that connects to components between which relative motion exists in service must have the necessary flexibility and structural integrity to withstand the relative motion without exceeding limit load (at the maximum operating temperature) of the piping, or creating unintended loads (or load paths) on the components to which the piping connects.
b. Procedures. Exhaust piping is typically certified by analysis and installation tests conducted during the basic certification process, including flight tests, as follows:
(1) For compliance with § 27.1123(a), because of its durability in the hot exhaust environment, exhaust piping is typically made from stainless steel or alloy steel of the appropriate structurally and thermally derived wall thickness. Hot aircraft exhaust gases are very corrosive; thus, proper material selection and corrosion protective design should be performed and validated during certification. Advisory Circular Par 548 723