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.
(viii) 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:
(i) 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.
(iii) Emergency Operating Procedures. Operating procedures containing essential information particularly with system failure.
(iv) Caution Notes. These caution notes should advise or addr ess:
(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 References.
(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.
(ii) 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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(xiii) 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:
(i) 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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40 - 30 20 - 10 - 0-
- 10 - T T 2000 4000 6000 8000 10,000
PRESSURE ALTITUDE - FT FIGURE AC 27.1419-1 CONTINUOUSTCING - TEMPERATURE VS ALTITUDE LIMITS Figures 27.1419-1 through 4 represent the approach to a 10,000-foot altitude limit. See Paragraph b(5)(ili) for a discussion on this approach.
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40 30 - AMBIENT TEMPERATURE - DEG F 20 - 10 °
- 10 2000 4000 6000 8000 10,000
PRESSURE ALTITUDE - FT FIGURE AC 27.1419-2 INTERMITTENT ICING - TEMPERATURE VS ALTITUDE LIMITS Figures AC 27.1419-1 through 4 represent the approach to a 10,000-toot altitude limit.
See Paragraph b(5) (iii) for a discussion on this approach.
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1.0 - LIQUID WATER CONTENT - GRAMS/CUBIC METER 0.8 - SEA LEVEL TO 10,000 FT PRESSURE ALTITUDE AMBIENT TEMP °F 10 20 30 40 50
MEAN EFFECTIVE DROP DIAMETER - MICRONS FIGURE 27.1419-3 CONTINUOUS ICING-LIQUID WATER CONTENT VS. DROP DIAMETER Figures AC 27.1419-1 through 4 represent one approach to a 10,000-foot altitude limit.
See Paragraph b(5)(iii) for a discussion of this approach.
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2.0 SEA LEVEL TO 10,000 FT PRESSURE ALTITUDE LIQUID WATER CONTENT - GRAMS/CUBIC METER 1.6 -
AMBIENT TEMP
- 32 °F
1.2 14 °F
0.8
- 0°F
0.4 10 20 30 40 50
MEAN EFFECTIVE DROP DIAMETER - MICRONS FIGURE AC 27.1419-4. INTERMITTENT ICING-LIQUID
WATER CONTENT VS DROP DIAMETER Figures AC 27.1419-1 through 4 represent one approach to a 10,000-toot altitude limit.
See paragraph b(5)(iii) for a discussion of this approach.
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AC 27.1435. § 27.1435 HYDRAULIC SYSTEMS.
a. References. The following sections of Part 27 are either incorporated in the provisions of § 27.1435 or are otherwise applicable to hydraulic system design:
§§ 27.695, 27.861, 27.863, 27.1183, 27.1185, 27.1189, 27.1309, and 27.1322.
b. System Design. It is assumed that the hydraulic system will be utilized to operate utility systems and the primary control system of the rotorcraft.
(1) Section 27.1309(a) and (b) provides for functioning reliably under any foreseeable operating condition and prevention of hazards after any malfunction or failure.
(2) The substantiating data should include a failure analysis that considers every possible system component failure, such as (but not limited to) ruptured lines, pump failure, regulator failure, ruptured seals, clogged filters, broken pilot valve connections, and so forth. Also, consideration of the specific requirements of § 27.1435 should be included.
(3) If the rotorcraft cannot be safely operated without the hydraulic system, the requirements of § 27.1309(a) and (b) are met by dual independent hydraulic systems.
From the reservoir, hydraulic pump, regulator, connecting tubing, and hoses through the actuators, there must be no commonality in the fluid-containing components. A break in one system should not result in fluid loss in the remaining systems. The pumps should be separated as far as practicable; i.e., on opposite sides of the rotor drive transmission, on separate engines, or one pump on an engine and the other on the rotor drive transmission. The tubing and hoses should also be routed with as much physical separation as practicable. The purpose of this separation is to prevent total loss of the hydraulic systems in the event of a malfunction such as fire or rotor burst wherein one projectile could disable both systems.
(4) Dual actuators must be designed to ensure that any single failure, such as a cracked housing, broken interconnecting input, or output link, does not result in loss of total hydraulic system function.
(5) If installed, the pressure-indicating system is normally included as a dial, vertical scale, or digital indicator. The indicator should enable the crew to detect pressure trends. Paragraph AC 27.1322 concerns § 29.1322 regarding proper colors for annunciators if used to supplement the indicating system.
(6) A combination of analysis and tests should be included in the substantiating data file to show compliance with the provisions of § 27.1435.
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(7) Extra caution should be exercised to ensure that control input forces at the mechanical connection to the actuator pilot valves do not exceed their intended value.
Consideration should be given to the most adverse tolerance buildup in parts fabrication and control system rigging.
(8) The substantiating data should show that the hydraulic components will perform their intended function reliably under the most adverse continuous and short-time environmental conditions to which they are exposed. These variables include but are not limited to temperature, humidity, vibration, altitude, and shock.
Paragraph AC 27.1309b(2)(i) contains a method of temperature correction to cover the entire operating temperature envelope being certified.
(9) The system component strength must be sufficient for its material fatigue life to exceed the number of cycles imposed by pump ripple pressure.
c. Installation Precautions and Fire Protection.
(1) All components and tubing routed through fire zones may be designed to comply with the fire protection requirements of §§ 29.1183, 29.1185, and 29.1189. As an alternative, a fireproof shield may be used around the component to be protected.
The component should be sufficiently protected to assure fluid leakage will not occur and fuel the fire.
(2) All hydraulic lines should be sufficiently isolated from the engine, bleed air lines, environmental control unit, oil cooler, or other heat source to ensure expected line life.
(3) If flammable hydraulic fluid is used, the hydraulic components should be isolated from ignition sources to ensure that failure of any of the hydraulic components will not result in a fire or explosion. In the case of electrical ignition sources in the proximity of hydraulic components, the electrical equipment should be hermetically sealed or otherwise substantiated as not being an ignition source (reference paragraph AC 27.1309b(1)(i)).
(4) The installation detail should be thoroughly reviewed for adequacy of line clamping and clearance from sharp edges. As much physical separation as possible should be provided between hydraulic lines and electrical cables.
(5) While the control system is being moved from stop to stop, observation should be made to determine that hose flexing and tube bending is minimized.
d. Testing.
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(1) Individual components should be substantiated by either a vendor’s or a primary manufacturer’s laboratory test reports. These tests should establish performance ratings such as pressures, flow rates, environmental capability, etc., to be approved.
(2) After the total system is installed, ground tests should be conducted to ensure the system performs as intended and that each component is functioning within its design rating. System testing should consider the provisions of § 27.1435(b).
(3) If the total system design permits each combined independent power source and actuator to be disabled by shutoff valves, engine shutdown, etc., each combination should be disabled and the remaining combination verified to perform the necessary control functions. The test should be accomplished again with the functioning combination disabled and the disabled combination functioning. These tests should be accomplished first by ground tests, then repeated in flight.
(4) Temperature and pressure instrumentation should be provided at the critical points in the system. Temperature results should be corrected for hot day conditions.
(Paragraph AC 27.1309b(2)(i) gives a recommended procedure.)
(5) All controls should be cycled throughout their complete range of travel while accomplishing the provisions of paragraph d(2) above.
(6) Satisfactory hydraulic system performance should be verified while the pump drive sources (rotor, engine, etc.) are individually varied throughout their approved operating range.
(7) Flight tests should be conducted throughout all altitudes, maneuvers, and control ranges while the system is instrumented as in paragraphs d(2) and (4) above to determine that component ratings are not exceeded.
AC 27.1457. § 27.1457 (Amendment 27-22) COCKPIT VOICE RECORDER.
a. Explanation. The function of the cockpit voice recorder (CVR) is to provide a record of the crew communications preceding and during rotorcraft accidents. Over the last several years, the National Transportation Safety Board (NTSB) has determined that CVR’s are invaluable in determining probable cause of an accident. Because of this fact and mandates of Congress, the use of CVR’s is required by the operating rules on many rotorcraft involved in passenger-carrying operations.
b. Procedures. The following areas are of particular consideration in the approval of a CVR installation:
(1) Equipment Qualifications. The CVR should be approved. The most common way of obtaining an approval is to qualify the CVR (and associated control panel, if appropriate) to TSO C84 or C123.
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