(VI) There should be a means identified or provided for determining the formation of ice on critical parts of the rotorcraft which can be met by a reliable and safe natural warning or an ice detection system. A system utilizing OAT should include an accurate OAT measurement since the onset of icing can occur in a very narrow temperature band requiring sensitive and accurate OAT measurement. OAT accuracy should be relative to the true temperature of the air mass. Total system accuracy should be t0.5° C in the -5.0° to +5.0° C range and t1° C throughout the remaining temperature range. The location of the sensor has been shown to be very critical and, in effect, there can be a position error or other errors induced by ice formations or solar radiation. If the system measures liquid water content, consideration should be given to the fact that the actual LWC fluctuates considerably as the rotorcraft passes through an icing environment. A warning system displaying or utilizing a peak or average LWC value (rather than an instantaneous readout) should include sufficient conservatism to provide a margin of safety. The value of an LWC detecting system lies in its utility as a warning that ice is being encountered. The actual magnitude of LWC in combination with OAT and MVD can be used to indicate the icing severity level. The U.S. Army is developing an advanced ice detection system (1990) for potential application to rotorcraft.
(2) In-flight Ice Detection Sensing Systems With the advent and development of In-flight Ice Detection Sensing Systems (IIDSS) technology designed to warn flight crews of potential ice accumulation on critical helicopter components, standardized guidelines for certification have been established. These guidelines will permit applicants for new, amended and supplemental type certificates under FAR 27 to present a rational compliance plan to the respective Authority. Currently there are two types of lIDSS; advisory and primary.
The advisory system enunciates the presence of icing conditions. The flight crew is responsible for monitoring the icing conditions as defined in the Rotorcraft Flight Manual (RFM) and activation by the flight crew of the anti-icing or de-icing systems) remains a requirement.
(ili) The primary system has automatic control of anti-icing or de-icing systems when the flight crew has selected the automatic switch position. The automatic feature can be de-selected and the system reverts to advisory, where the crew is responsible for monitoring the icing conditions and activating the anti-icing or de-icing systems). Neither the advisory nor the primary IDSS are designed to operate on the The following factors should be considered during the design and certification of an IIDSS:
(A) The IDSS displays) status lights and/or crew alerting messages must be located so that they are within the seated flight crew's forward vision scan area Page F - 83
AC 27-1B, Chg 1 2/12/03 while performing their normal duties. The IDSS display must also meet the applicable requirements of § 27.1322. Fixed probes must be located in areas easily scanned by the crew, and must be visible under normal daytime and nighttime flying conditions.
(B) cing conditions can exist when visible moisture in any form is present and the Outside Air Temperature (OAT) is 41°F (or 5° Celsius) or lower or the Total Air Temperature (TAT) is 50°F or lower. It should be noted that icing conditions may also exist when the OAT is 50°F (10°C) or lower while operating on the ground where surface snow, water, slush, etc., may be ingested by engines or freeze on engines, nacelles, engine sensor probes, rotors, or other critical surtaces.
(C) The core of any IDSS is the ice detector device and its location.
Ice detectors should be installed in carefully determined locations to avoid interference from air data sensors, external protuberances (including aircraft options such as rescue hoists, flotation devices, and radar units), rotor downwash induced water impingement, wheel splash during ground operation, etc.
(D) From the standpoint of powerplant icing protection, airframe visual icing cues are not an acceptable means to advise the flight crew to activate the engine ice protection system. Delaying the use of the engine ice protection system until ice build-up is visible from the cockpit may result in severe engine damage and/or flameout due to shed ice ingestion, and is therefore unacceptable. The engine ice protection system is to be activated by the flight crew in accordance with approved Rotorcraft Flight Manual (RFM) procedures when icing conditions exist, if using an advisory system, or automatically if using a primary system. Current engine induction ice protection systems are operated as anti-ice systems, as required by the RFM limitations and procedures. These requirements provide the necessary margin between system activation and the ambient conditions. In-service experience has shown that adherence to these RFM limitations and procedures has provided satisfactory engine operation.
(E) An IIDSS that is intended as the prime means of alerting the flight crew or operating the de-icing/anti-icing systems should contain certain features:
(1) An IIDSS system hazard analysis should be completed in order to choose the IIDSS architecture.
(2) The effectiveness of the IIDSS must be demonstrated during the icing flight tests of §§ 27.1093 and 27.1419.
(3) The threshold level chosen to activate the ice detection and annunciation system should be guided by:
The assurance that when the amount of ice that is accreted on the critical surfaces is shed, there will be no damage to the helicopter or engines, and Page F - 84
The assurance that the amount of ice accreted can be safely eliminated by the ice protection system.
Ш An advisory system should not be overly sensitive and enunciate frequent changes from "on" to "off" and thereby induce the pilot to ignore detector indications. However, the system must be sensitive enough to readily detect sudden exposure to icing conditions throughout the complete approved icing envelope.
(4) If overheat of structure (such as engine inlet cowl or rotorblade) can result from the anti-ice/de-ice systems being "on" during any operations, then a means should be provided to alert the flight crew or an automatic means included that will prevent such a condition.
(5) The operation of an anti-ice/de-ice system should be examined for the combined effect of an undetected failure of the annunciation system together with a time delay before the flight crew manually activates these systems. Specific considerations that warrant investigation include:
(i) The amount of ice that can be accreted on critical areas.
(ii) The effect of the ice shedding on the aircraft and propulsion system.
(iii) The capability of the anti-ice/de-ice systems.
(6) The RFM should address the following:
(i) Normal operational use of the IIDSS and any limitations.
(ii) Procedures to use in case of disagreement between the dual ice detectors, if applicable.
(ii) Failure mode indications, and appropriate crew procedures.
(Z) An IIDSS must meet the applicable requirements of FAR 27.1309. Multiple systems, automatic fault monitoring, Built-in Test Equipment (BITE), pre-flight status tests, etc., may be used to support the design reliability, depending on IIDSS system hazard analysis.
8) To ensure the continued airworthiness of IIDSS, it will be necessary to develop maintenance procedures.
(F) Compliance with the regulations may be demonstrated by tests, analyses, models, similarity with approved systems, or a combination thereof as outlined in the compliance plan and approved by the Authority.
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AC 27-1B, Chg 1 2/12/03
(3) Instrumentation and Data Collection.
Instrumentation proposed for certification tests, including flight strain surveys, should be reviewed as early as possible in the program to establish that it will provide the necessary data. The need for accurate OAT measurement previously noted for operation in icing also applies to the certificated configuration. Mechanical devices such as the rotating multicylinder and rotating disc have been used for measuring the ice accretion rate which is related by calibration to average WC and MVD. More recently, hybrid mechanical/electronic LWC measuring devices have been used.
Devices that rely on ice accretion as a signal source are subject to the Ludlam limit (the limits whereby latent heat of fusion is not totally absorbed, thus resulting in incomplete freezing of the moisture and some inaccuracy in the indication). The Ludlam limit is a function of various parameters including OAT, airspeed, LWC, and MVD. The Ludlam limit may vary from one device to another. (See References AC 27.1419d(8) and AC 27.1419d(9)(i) for further information). Gelatin slides, soot and oil slides, and more recently, laser nephelometers have been used to measure droplet size. Other calibrated devices intended for measurement of LWC should be used.
Paragraph AC 27.1419d(16) describes several of these devices. Photographic coverage of critical areas may be necessary to ascertain that ice protection systems are functioning properly and that there are no runback problems. (The term "runback" refers to liquid water that has not been evaporated by surface deice equipment and flows back to an unheated area subject to freezing.) Reference AC 27.1419d(19) highlights use of video techniques and equipment for this purpose. Some systems will require acceptable calibration techniques and data.
(її) Gelatin, soot, and oil slides provide data that can be used to estimate MVD at discrete intervals while laser nephelometer data can provide time histories of MVD droplet size distributions. Gelatin slide data should be taken frequently during test flights to properly characterize the cloud. Laser nephelometer data have been found to be highly dependent upon knowledge of the equipment and calibration. Proper calibration, maintenance, and data processing techniques should be utilized and demonstrated. Additional information on the subject may be found in Reference AC 27.1419d(18).
(iii) Structural instrumentation requirements should also be established as early as possible in the program. Flight strain measurements are strongly recommended in assessing the ice imposed stress on the rotorcraft. The flight strain measurements should determine the effect on fatigue life due to ice accumulation for such items as main rotor blades, main rotor hub components, rotating and fixed controls, horizontal stabilizer, tail rotor, etc. The subsequent proper operation of retractable devices such as landing gear should be demonstrated with representative ice accretion. In addition, the static and fatigue strength of the blade with heater mat should be substantiated. Any effect of the heater mat on fatigue strength of the blades should be considered.
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(4) Additional Considerations. The following are items to consider in an icing certification program. They are not intended to be all-inclusive, and the possibility of widely differing characteristics and critical areas among various rotorcraft in icing should be considered.
The rotorcraft should be shown by analysis and confirmed by either simulated or natural icing tests to be capable of holding for 30 minutes in the design conditions of the continuous maximum icing envelope at the most critical weight, CG, and altitude with a fully functional ice protection system.
A single ice protection system and power source may be considered acceptable provided that after any single failure of the ice protection system, the rotorcraft can be shown by analysis and/or test to be capable of safe operation (no hazard) for 15 minutes following failure recognition in the continuous icing envelope used as the basis for certification within the same icing limits used for the 30-minute hold criteria. During this 15-minute period the rotorcraft may exhibit degraded characteristics. Pilot controllable operating limitations such as airspeed may be used to satisfy this continued safe flight criteria. For purposes of determining performance and handling qualities degradation, ice protection system failure need not be considered to occur simultaneously with engine failure unless ice protection system operation is dependent upon engine operation.
(iii) Although current airborne weather radar technology systems may be useful in avoiding potential icing conditions by detecting precipitation, the use of weather radar is not an FAA/AUTHORITY requirement for icing certification.
(iv) Section 27.1419(e) says there must be a means to advise the crew when the rotorcraft is in icing conditions in order that the system may be activated.
No autorotational performance data is required for rotorcraft which have Category A powerplant installations. The rotorcraft must be capable of full autorotational landings with the ice protection system operating (§ 27.143a(2) (vi).
Autorotational entry, steady state, and flare entry flying qualities and performance should be evaluated with the ice load to be expected with the deice system operating and with the ice load to be expected 15 minutes after failure of the system. Since the en route performance can vary as the ice protection system operates, a mean value of cyclic torque is acceptable provided, at no time the power required drops below that required for level flight. The rotorcraft is assumed to be clear of ice prior to takeoff, and, therefore, the takeoff performance is not degraded. The landing performance can be based on the in-flight assessment of overall performance degradation. Items such as fuel burns can be used as part of the in-flight performance degradation determination.
Regardless of the methods used to determine performance degradation, they must be easily used by the crew. The hover performance should be addressed for the termination of a flight after an icing encounter. The engines should be protected from the adverse effects of ice. When ice does accumulate on the inlets, screens, etc., it Page F - 87
must be accounted for in performance, engine operating characteristics, and inlet distortion.
(vi) The handling qualities of the rotorcraft must be substantiated if ice can accumulate on any surface. When ice can accumulate on unprotected surfaces, the rotorcraft must exhibit satisfactory IFR handling qualities. In addition, following the failure of the deice system, the rotorcraft must be safely controllable for 15 minutes, i.e., the rotorcraft must be free from excessive and rapid divergence. Artificial ice shapes may be acceptable for acquisition of flight test data necessary for handling qualities and performance evaluations and demonstrations.
(vii) Items such as fuel tank vents, cooling vents, antennas, etc., should be substantiated for maximum icing effects.
(vili) The ice protection system should be sufficiently reliable to perform its intended function in accordance with the requirements of § 27.1309. These requirements may in some instances be met by the use of sound engineering judgment during design and compliance demonstrations. In many instances, use of good design practices, failure modes and effects analysis, and similarity analyses combined with good judgment will be adequate. In some instances the need for reliability analyses may be desirable. Additional information pertaining to reliability is contained in paragraph AC 27.1309 (§ 27.1309).
(ix) The subject of lightning should be addressed. The criteria applied on rotorcraft with ice protection systems are that "the rotorcraft should be protected in such a manner to minimize lightning risk." The general rules of § 27.1309(a), (b), and (c) are applicable to ensure adequate lightning protection. (Amendment 27-21, November 6, 1984, added lightning protection requirements in § 27.610.)
(x) Ice protection of pitot-static sources, windshields, inlets, exposed control linkages, etc., should be considered.
(xi) The impact of ice protection system failure, complete and partial, and achieving adequate warning thereof should be assessed.
(xii) The impact of delayed application of ice protection systems should be assessed. Hazardous conditions should not be apparent. Any rotorcraft characteristic changes resulting should be covered in cautionary material in the rotorcraft flight manual.
(xiii) Possible droop stop malfunction with ice accumulation and its potential hazard to the rotorcraft, its occupants, and ground personnel should be assessed.
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(xiv) Possible ice shedding hazards to ground personnel or equipment in proximity to turning rotors following flight in icing conditions should be given consideration.
(5) Flight Manual. Areas of the flight manual which may require input are:
(•) Operating limitations including approved types of operation and prohibiting operation in freezing rain or freezing drizzle conditions. Avoidance procedures may also be useful.
(ii) Normal Operating Procedures. Information on the ice detection means or system and ice protection system and their capabilities.
(ili) Emergency Operating Procedures. Operating procedures containing essential information particularly with system failure.
(iv) Caution Notes. These caution notes should advise or address:
(A) Against inducing asymmetric shedding with rapid control inputs or rotor speed changes, except possibly as a last resort. Rotor speed changes appear to be more effective than control inputs in removing ice from the rotor blades of some rotorcraft.
(B) Loss in range, climb rate, and hover capability following prolonged operation in icing.
(C) The need for clean blade surfaces and use of approved cleaning solvents or ground deicing/anti-icing agents prior to start of rotors turning.
(D) Changes in autorotational characteristics resulting from ice formations.
(E) Although the rotorcraft has been certificated for flight in supercooled clouds and falling and blowing snow, flight in other conditions such as freezing rain, freezing drizzle, sleet, hail, and combinations of these conditions with supercooled clouds must be avoided.
(F) The potential hazards to ground personnel, passengers deplaning, and equipment in proximity to turning rotors following flight in icing conditions.
d. Icing Reterences.
(1) FAA Technical Report ADS-4, Engineering Summary of Airframe Icing Technical Data, December 1963.
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(2) Advisory Circular 20-73, Aircraft Ice Protection, 21 April 71.
(3) Advisory Circular 91-51, Airplane Deice and Anti-ice Systems, 9/15/77.
(4) FAA Report RD-77-76, Engineering Summary of Powerplant Icing Technical Data, July 1977.
(5) United States Army Aviation Engineering Flight Activity Reports:
(i) Natural Icing Tests, UH-1H Helicopter, Final Report, June 1974, USAASTA Project No. 74-31.
(1) Artificial Icing Tests, UH-1H Helicopter, Part I, Final Report, January 1974, USAASTA Project No. 73-04-4.
(iii) Artificial Icing Tests, UH-1H Helicopter, Part II, Heated Glass Windshield, Final Report, USAASTA Project No. 73-04-4.
(iv) Artificial Icing Tests, Lockheed Advanced Ice Protection System Installed on a UH-1H Helicopter, Final Report, June 1975, USAAEFA Project No. 74-13.
(v) Artificial and Natural Icing Tests for Qualification of the UH-1H, Kit A Aircraft, Letter Report, USAAEFA Project No. 78-21-1.
(vi) Microphysical Properties of Artificial and Natural Clouds and Their Effects on UH-1H Helicopter Icing, Report USAAEFA Project No. 78-21-2.
(vii) Helicopter Icing Spray System (HISS) Nozzle Improvement Evaluation, Final Report, September 1981, USAAEFA Project No. 79-002-2.
(viii) Artificial and Natural Icing Tests of the YCH-4TD, Final Report, May 1981, USAAEFA Project No. 79-07.
(ix) Limited Artificial Icing Tests of the OV-ID, Letter Report, July 1981, USAAEFA Project No. 80-16, (Limited Distribution).
(x) JUH-IH Ice Phobic Coating Tests, Final Report, July 1980, USAAEFA Project No. 79-02.
(xi) Artificial and Natural Icing Tests, Production UH-60A Helicopter, Final Report, June 1980, USAAEFA Project No. 79-19.
(xii) Helicopter Icing Spray System (HISS) Evaluation and Improvements, Letter Report, June 1981, USAAEFA Project No. 80-04.
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(xili) Artificial Icing Test of CH-47C Helicopter with Fiberglass Rotor Blades, Final Report, July 1979, USAAEFA Project No. 78-18.
(xiv) Limited Artificial and Natural Icing Tests, Production UH-60A Helicopter (Reevaluation), Final Report, August 1981, USAAEFA Project No. 80-14.
(6) Further Icing Experiments on an Unheated Nonrotating Cylinder, National Research Council, Canada Report LTR-LT-105, dated November 1979, by J.R. Stallabrass and P.F. Hearty.
(7) Ludlam, F.H., Heat Economy of a Rimed Cylinder, Quarterly Journal, Royal Meteorological Society, Vol. 77, 1951.
(8) U.S. Army AMRDL Reports:
USAAMRDL TR 73-38, Ice Protection Investigation For Advanced Rotary Wing Aircraft, J.B. Werner, August 1973, AD 7711182.
(ii) Werner, J.B., The Development of an Advanced Anti-Icing/Deicing Capability for U.S. Army Helicopters, Volume 1, Design Criteria and Technology Considerations, USAAMRDL - TR-75-34A, Eustis Directorate, U.S. Army Air Mobility R&D Laboratory, November 1975, AD A019044.
(iii) Werner, J.B., The Development of an Advanced Anti-Icing/Deicing Capability for U.S. Army Helicopters, Volume 2, Ice Protection System Application to the UH-1H Helicopter, USAAMRDL - TR-75-34B, Eustis Directorate, U.S. Army Air Mobility Research and Development Laboratory, November 1975, AD A019049.
(iv) USAAMRDL-TR-76-32, Ottawa Spray Rig Tests of an Ice Protection System Applied to the UH-1H Helicopter, November 1976, AD A0034458.
(v) USARTL-TR-78-48, Icing Tests of a UH-1H Helicopter with an Electrothermal Ice Protection System Under Simulated and Natural Icing Conditions, April 1979.
(vi) USAAMRDL-TR77-36, Final Report, Natural Icing Flights and Additional Simulated Icing Tests of a UH-IH Helicopter Incorporating an Electrothermal Ice Protection System, July 1978, AD A059704.
(9) Technical Feasibility Test of Ice Phobic Coatings for Rain Erosion in Simulated Flight Conditions, U.S. Army Test and Evaluation Command, Final Report, 4-AI-192-IPS-001, August 1980.
(10) Technical Feasibility Test of Ice Phobic Coatings in Simulated Icing Flight Conditions, U.S. Army TECOM, Final Report, 4-CO-160-000-048, September 1980.
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(11) Aircraft Icing, NASA Conference Publication 2086, FAA-RD-78-109, July 1978.
(12) Helicopter Icing Review, FAA Technical Center, Final Report, FAA-CT-80-210, September 1980.
(13) National Icing Facilities Requirements Investigation, Final Report, FAA Technical Center, FAA-CT-81-35, March 1981.
(14) Aircraft Icing, AGARD Advisory Report No. 127, November 1978.
(15) Rotorcraft Icing - Review and Prospects, AGARD Advisory Report, AR-166, September 1981.
(16) Advisory Circular 20-117, Hazards Following Ground Deicing and Ground Operations in Conditions Conducive to Aircraft Icing, Dec. 17, 1982.
(17) Olson, W., Experimental Comparison of Icing Cloud Instruments, January 1983, NASA TM 83340.
(18) JUH-1H Redesigned Pneumatic Boot Deicing System Flight Test Evaluation. Hayworth, L., Graham, M., August 1987. USAAEFA Edwards AFB, California. Project No. 83-13.
(19) An Appraisal of the Single Rotating Cylinder Method of Liquid Water Content Measurement, National Research Council Canada Report LTR-LT-92, dated November 1978, by J.R. Stallabrass.
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30 - 20 10 2000 4000 6000 8000 10, ã00
PRESSURE ALTITUDE - FT FIGURE 4C 27 1419-1 CONTINUOUS ICING - TEMPERATURE VS ALTITUDE LIMITE Figures 27.1419-1 through 4 represent the approach to a 10,000-foot altitude limit See Paragraph b(5)ții) for a discussion on this approachi.
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АС 27-1B 9/30/99 30 - AMBIENT TEMPERATURE: DEG F 20 — 10- 2000 4000 6000 8000 là, 000
PRESSURE ALTITUDE - FT FIGURE AC 27 1419-2 INTERMITTENT ICING - TEMPERATURE VS ALTITUDE LIMITE Figures AC 27.1419-1 through 4 represent the approach to a 10,000-foot altitude limit.
See Paragraph b(5)(iil) for a discussion on this approach.
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