(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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LIQUID WATER CONTENT - GRAMS/CUBIC METER 08 - SEA LEVEL TO 10,000 FT PRESSURE ALTITUDE 0,6 AMBIENT TEMP OF 0,4 02 10 20 30 40
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)(i) 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 Õ‚8 Õ4 10 20 30 50
MEAN EFFECTIVE DROP DIAMETER - MICRONS FIGURE AC 27, 1419-4. INTERMITTENT IONG-LIQUID
WATER CONTENT VS DROP DIAMETER Figures AC 27.1419-1 through 4 represent one approach to a 10,000-foot altitude limi See paragraph b(5)(ii) 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
(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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(2) Cockpit Area Microphone (CAM). The third channel of recorded information is specified to be from a cockpit area microphone or from voice activated lip microphones at the first and second pilot stations. It should be noted that a continuously recording or "hot" microphone at both the first and second pilot stations would satisfy this CAM requirement. Due to the ambient noise level in rotorcraft, the use of "hot" microphone results in objectionable constant hissing in the pilot's headsets.
Therefore, it is recommended that "hot" microphones not be used on rotorcraft.
(3) CVR Mechanical Installation. The CVR or the portion thereof which contains the recording should be physically located to enhance the probability of the recording surviving a crash. Normally, such a location would be in the lower portion of the rotorcraft as far aft as possible.
(4) Intelligibility of Recordings. Tests should be accomplished to determine that the recording is intelligible enough to make a positive identification of the speaker and the words or phrases spoken. This is usually accomplished by flight operations to produce the maximum cockpit background noise. The operation should provide for the normal speech of all crew members to be recorded on the pertinent channels. Then, during playback, preferably using a different listener, the listener should be able to identify the different crew members, the words and phrases spoken by the crew, and the radio communications made by and to the crew. The use of special filters and multiple playbacks to improve intelligibility is acceptable.
(5) Electrical Power Supply. The rule requires that the CVR should be supplied with power from the most reliable source that does not jeopardize essential or emergency loads. Since the functioning of the CVR is required by operating rules for some operations, it should be given priority over other nonessential loads.
(6) Self-Test Function. The CVR should be provided with a means in the cockpit that will allow a test to ensure the CVR is functioning properly. This may be accomplished by a manual playback feature.
(7) Bulk erasure. If this function is provided, the installation should be as follows:
Any probable malfunction will not cause erasure of the recording medium.
(ii) The crash impact forces will not cause activation of the bulk erasure function.
(ili) Inadvertent actuation of the bulk erasure function is minimized.
Usually, this is accomplished by requiring two separate actions to operate the bulk erasure.
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AC 27.1459. § 27.1459 (Amendment 27-22) FLIGHT RECORDERS.
a. Explanation. The function of the flight recorder, sometimes referred to as a flight data recorder, is to provide a record of various aircraft and air data parameters during the operation of the rotorcraft. This data is utilized by accident investigators to aid in determination of the probable cause of an accident. The problems associated with acquisition of this data in aircraft not equipped with flight recorders have been complicated by the use of advanced instrument systems such as EFIS, FADEC, EICAS, and IDS. The very nature of the operation of these systems precludes the deduction of post accident data, as was possible with mechanical and electromechanical instruments, annunciators, hydromechanical engine controls, and switches. The National Transportation Safety Board (NTSB) therefore made a recommendation to the FAA that aircraft should be required to have flight recorders. Subsequently Congress mandated that flight recorders be required on many rotorcraft involved in passenger-carrying operations in accordance with FAR 91 and FAR 135.
b. Procedures. The following areas are of particular consideration in the approval of a flight data recorder installation.
(1) Equipment Qualification. The recommended procedure to obtain an approval for the flight recorder (and associated control panel, if appropriate) is to qualify the flight recorder to TSO C-124. The required underwater locating device should be qualified to the provisions of TSO C-121.
(2) Recorded Parameters and Accuracy.
Airspeed. The installed flight recorder should record the airspeed with an accuracy of 3 percent or 5 knots (whichever is greater) from a speed of 20 knots to a speed of 80 percent more than Vy.
(ii) Flight Recorder. The flight recorder should be capable of recording the pressure altitude of the rotorcraft with a range of -1,000 feet to the maximum certified altitude. The error of this recording at sea level should not exceed +50 feet.
(¡¡i) Direction. The flight recorder should be capable of recording the magnetic heading of the rotorcraft within at least 10° for any heading. Larger deviations caused by the temporary operation of high current electrical devices such as heated windshields are acceptable.
(iv) Vertical Acceleration. The flight recorder should be capable of recording the normal acceleration of the center of gravity of the rotorcraft. The recommended range of this recording is an envelope of -3 to +6 G with an accuracy of at least 10.2 G.
(v) Time Correlation. The flight recorder should provide a time scaled correlation between the data recorded and the time at which this information was Page F - 101
presented to the first pilot via the required flight instruments. This correlation should normally be established before flight, and should have an accuracy rate that does not diverge by more than 4 minutes and 4 seconds in 8 hours.
(vi) Caveat. It should be noted that even though the requirements outlined above provide for compliance with the specific provisions of § 27.1459 regarding the acquired data and its accuracy, a flight recorder certified to these minimum standards will not meet the requirements of Appendix F of FAR 91 or Appendix C of FAR 135. If the flight recorder is to be used to comply with these operating rules, it is recommended that the appropriate appendix be consulted prior to requesting certification. The approved configuration may then be certified as meeting the requirements of the appropriate appendix.
(3) Flight Recorder Mechanical Installation. The non-ejectable flight recorder or the portion thereof which contains the recorded data should be physically located to enhance the probability of the recording surviving a crash. Normally, such a location would be in the lower portion of the rotorcraft as far aft as possible. However, other locations in the rotorcraft may be suitable to meet the requirement to "minimize the probability of container rupture resulting from crash impact and subsequent damage to the record from firei The normal accelerometer should be located within the most restrictive center of gravity of the rotorcraft. The required underwater locator is usually mounted to the case of the flight recorder.
(4) Electrical Power Supply. The rule requires that the flight recorder should be supplied with power from the most reliable source that does not jeopardize essential or emergency loads. Since the functioning of the flight recorder is required by operating rules for some operations, it should be given priority over other nonessential loads.
(5) Self-Test Function. The flight recorder should be provided with a preflight test that will provide confirmation that the recorder and its recording medium are functioning properly.
(6) Data Erasure Feature. If this function is provided and the flight recorder is not powered solely by an engine or transmission driven generator, the installation should provide the following features:
(i) Any probable malfunction will not cause erasure of the recording medium.
(ii) The crash impact forces will not cause activation of the data erasure function.
(iii) Inadvertent actuation of the data erasure function is minimized.
Usually, this is accomplished by requiring two separate actions to operate the data erasure.
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