airplane configuration, airspeed, angle of attack, altitude) and the icing conditions of part 25, Appendix C (temperature, liquid water content, mean effective drop diameter). The applicant should substantiate the critical mean effective drop diameter, Liquid Water Content (LWC), and temperature that result in the formation of an ice accretion that is critical to the airplane's operational characteristics. For deicing systems, intercycle and residual ice accretions need to be considered.
(3) The 45-Minute Hold Condition. The 45-minute hold criterion should be evaluated when determining critical ice shapes for which the operational characteristics of the overall airplane are to be analyzed. The airplane's tolerance to continuous ice accumulation on the unprotected surfaces should be evaluated in accordance with the information contained in AC
20-73. The applicant should determine the effect of the 45-minute hold in continuous maximum icing conditions.
(a) A mean effective droplet diameter of 22 microns and a liquid water content of
0.5 gm/m3 with no horizontal extent correction are normally used for this analysis; however, the applicant should substantiate the specific values used, including temperature, which represent the critical conditions for the airplane's performance and handling qualities. The analysis should consider that the airplane would remain in an icing cloud based on a rectangular course with leg lengths not exceeding the cloud horizontal extent and all turns being made within the icing cloud.
Critical flight conditions should be considered such as weight and speed for critical angle of attack, and airspeed and altitude for maximum water catch.
(b) The applicant may elect to use more severe liquid water contents that are more representative of expected holding altitudes. The critical ice shapes derived from this analysis should be compared to critical shapes derived from other analyses (climb, cruise, and descent) to establish the most critical simulated ice shapes to be used during dry air flight tests.
c. Flutter Analysis. AC 23.629-1A, “Means of Compliance With Section 23.629, Flutter,” provides guidance for showing compliance with § 23.629. The flutter analyses should reflect any mass accumulations on unprotected and protected surfaces from exposure to part 25, Appendix C icing conditions, including any accretions that could develop on control surfaces.
The 45-minute hold should also be considered. Ice accretions due to failure of the ice protection system should also be addressed.
d. Power Sources.
(1) Electrical Load Analysis. The applicants should evaluate the power sources in their ice protection system design. Electrical, bleed air, and pneumatic sources are normally used. A load analysis or test should be conducted on each power source to determine that the power source is adequate to supply the ice protection system, plus all other essential loads throughout the airplane flight envelope under conditions requiring operation of the ice protection system.
(2) Effect on Essential Systems. The effect of an ice protection system component failure on power availability to other essential loads should be evaluated, and any resultant hazard should be prevented on multi-engine designs and minimized on single-engine designs.
The applicant should show that there is no hazard to the airplane in the event of any power source failure during flight in icing conditions. Two separate power sources (installed so that the failure of one source does not affect the ability of the remaining source to provide system power) are adequate if the single source can carry all the essential loads.
(a) Tw o-Engine Airplanes require two sources in accordance with § 23.1309(c).
If a single source system is planned, additional reliability evaluation of the power source under system loads and environmental conditions may be required. All power sources that affect engine or engine ice protection systems for multi-engine airplanes must comply with the engine isolation requirements of § 23.903(c).
(b) Single-Engine Airplanes. Section 23.1309(a) requires that the ice protection system be designed to minimize hazards to the airplane in the event of a probable malfunction or failure. Failure condition classifications of “major”, “hazardous” or “catastrophic” are considered hazards. Complete loss of the airframe ice protection system has been considered at least “major” on past certification programs. Since experience has shown that the failure of generators currently in service is probable, for example, systems that utilize a generator, such as an alternator, would require two sources of electrical power. This is also consistent with past project specific guidance on interpretation of § 23.1309, which stated that the level of safety in a single engine airplane is established by engine reliability and the ice protection system should not compromise it.
e. Failure Analysis. AC 23.1309-1C provides guidance and information for an acceptable means, but not the only means, for showing compliance with the requirements of § 23.1309(a) and (b) (amendment 23-49) for equipment, systems, and installations in 14 CFR, part 23 airplanes. The regulatory requirements are in § 23.1309. Substantiation of the hazard classification of ice protection system failure conditions is typically accomplished through analyses and/or testing. It has been standard industry practice to assign a probability of encountering icing conditions of “one” for an airplane certificated for flight in icing.
(1) During the analyses, each identifiable failure within the system should be examined for its effect on the airplane and its occupants. Examples of failures that need to be examined include:
(a) (a) Those that allow ice to accumulate beyond design levels; or
(b) Those that allow asymmetric ice accumulation.
(c) System failures such as loss of pneumatic boot vacuum or overheat of a thermal system
(2) A probable malfunction or failure is any single malfunction or failure that is expected to occur during the life of any single airplane of a specific type. This may be determined on the basis of past service experience with similar components in comparable airplane applications. This definition should be extended to multiple malfunctions or failure when:
(a) The first malfunction or failure would not be detected during normal operation of the system, including periodic checks established at intervals that are consistent with the degree of hazard involved; or
(b) The first malfunction would inevitably lead to other malfunctions or failures.
A procedure requiring a pilot to exit icing conditions would not be acceptable after any failure condition that would become catastrophic within the average exposure time it takes to exit icing conditions.
f. Similarity Analyses.
(1) In the case of certification based on similarities to other type certificated airplanes previously approved for flight in icing conditions, the applicant should specify the reference airplane model and the component to which the reference applies. Specific similarities should be shown for physical, functional, thermodynamic, pneumatic, aerodynamic, and environmental areas. Analyses should be conducted to show that the component installation and operation is equivalent to the previously approved installation.
(2) Similarity requires an evaluation of both the system and installation differences that may adversely affect the system performance. An assessment of a new installation should consider differences affecting the aircraft and the system. Similarity may be used as the basis for certification without the need for additional tests provided:
(a) Only minimal differences exist between the previously certificated system and installation, and the system and installation to be certificated; and
(b) The previously certificated system and installation have no unresolved icing related service history problems.
(3) FAA Order 8110.4B should be consulted regarding the use of previously approved FAA data.
If there is uncertainty about the effects of the differences, additional tests or analyses, or both, should be conducted as necessary and appropriate to resolve the open issues.
g. Induction Air System Protection. The induction air system for airplanes is certificated for ice protection in accordance with § 23.1093. These requirements are for all airplanes even those not certificated for flight in icing conditions. Thus ice protection systems installed on 13
previously type certificated airplanes to protect the engine induction air system should be adequate and need not be re-examined, unless the inlet is being modified, the original certification basis is inadequate, or an in-flight AFM limitation was used to comply with the falling and blowing snow regulation. When natural icing flight tests are conducted to show compliance to 23.1419, engine operation, engine ice protection system procedures, and engine inlet ice accretion should be evaluated if flight tests in natural icing conditions were not previously accomplished on the engine/inlet configuration.
h. Pitot Probe Ice Protection. Compliance to the TSO qualification standard for electrically heated pitot probes (TSO-C16) is not sufficient by itself in demonstrating compliance to the installation requirements of § 23.1309(b)(1) and § 23.1419. Section 23.1309(b)(1) requires that the system must perform its intended function under any foreseeable operating condition. Section 23.1419 requires that an airplane certificated for flight in icing must be able to safely operate in part 25, Appendix C icing conditions. It is unlikely that the conditions of part 25, Appendix C that are critical to the air data system equipment will be encountered during flight tests. Consequently certification programs should supplement the icing flight tests with an icing tunnel test and/or airborne icing tanker test data, as necessary, or reference testing that has been accomplished by the pitot probe manufacturer. In-service experience during severe atmospheric conditions has shown that pitot tubes qualified to the older standards have resulted in airspeed fluctuations and even loss of indicated airspeed. As these components should perform as intended in all expected atmospheric environments, it is reasonable to require that they be qualified to the Continuous and intermittent maximum icing conditions defined in part 25, Appendix C. Additionally, more recent standards indicate that these components should be qualified for operation in atmospheric conditions beyond the environment described by part 25, Appendix C, specifically during ice crystal and mixed phase icing conditions.
(1) Conditions Within the Part 25, Appendix C Icing Envelope. TSO C16, Airspeed tubes (electrically heated) require compliance to the performance specifications of SAE Aeronautical Standard AS393. This standard and AS393A (out-of-date), are non-current.
SAE AS393A includes a test to demonstrate deicing and anti-icing capability, but only temperature and airspeed are specified. Liquid water content is not specified but it influences heat requirements. Although functioning of pitot probes are evaluated in natural icing conditions during certification test programs, there is no requirement to flight test at the part 25, Appendix C icing limits because the low probability of finding those conditions imposes a burden. The airframe manufacturer is responsible for showing the pitot heat is adequate throughout the part 25, Appendix C icing envelope. If not obtained in flight test, analysis or icing tunnel test data should be submitted.
(2) Conditions Outside of the Part 25, Appendix C Icing Envelope. Although part 25, Appendix C only considers the liquid water content of icing conditions, recent cloud characterization research has indicated that approximately 40 percent of icing condition events consist of liquid water drops and ice crystals (mixed-phase icing conditions). Also ice crystal atmospheric conditions are encountered during aircraft operations. The ice crystal environment may be more critical than liquid water for thermal systems since more energy is required to melt 14
the ice crystals. Recently, some aircraft manufacturers and foreign certification authorities have required pitot and pitot-static probes to be tested in ice crystal and mixed phase icing conditions along with supercooled liquid water conditions. As a result, some pitot tube manufacturers now use the icing environment of British Specification (BS) 2G.135 “Specification for Electrically- Heated Pitot and Pitot-Static Pressure Heads,” as modified by the maximum rate that the icing tunnel facility can produce ice crystal, in addition to the requirements of the TSO. Even though the part 23 and part 25 regulations only address liquid water and testing in the mixed phase or ice crystal conditions are not required for FAA approval, it is good design practice to assure the pitot heat is sufficient for the ice crystal and mixed phase conditions of BS 2G.135.
i. Stall Warning Ice Protection. Compliance to the TSO qualification standard for stall warning instruments (TSO-C54) is not sufficient by itself in demonstrating compliance to the installation requirements of § 23.1309(b)(1) and § 23.1419.
(1) Conditions Within the Part 25, Appendix C Icing Envelope. TSO-C54, “Stall Warning Instruments”, requires compliance to the performance specifications of SAE Aeronautical Standard AS403A with some exceptions and additions. This standard is noncurrent. As in AS393A, the requirements include a test to demonstrate deicing and anti-icing capability, but only temperature and airspeed are specified. The precipitation test conditions of AS403A include moderate icing conditions for Type II instruments. However, "moderate" is not defined. The same comments from 9i(1) apply. The airframe manufacturer is responsible for showing that stall warning heat is adequate throughout the part 25, Appendix C icing envelope.
(2) Conditions Outside of the Part 25, Appendix C Icing Envelope. The same comments from paragraph 9i(2) above apply to stall warning ice protection systems.
j. Ice Detector. Besides the pilot's appraisal of icing conditions (i.e., defined by temperature and visible moisture or visual detection of ice accretions on wiper blades, window frames or propeller spinner, etc.), some airplanes use In-Flight Ice Detection Systems (IIDS).
IIDS may either directly detect the presence of ice on an airplane reference surface or detect that the airplane is in icing conditions. There are basically two classes of IIDS:
(1) Advisory In-flight Ice Detection System. An advisory IIDS provides information to advise the flightcrew of the presence of ice accretion or icing conditions. The flight crew has primary responsibility for detecting icing conditions or ice accretions, using the means defined in the AFM, and activating the Ice Protection Systems (IPSs). An advisory IIDS can automatically activate the IPS. However, the AFM must state the flight crew has primary responsibility for detecting icing conditions or ice accretions.
(2) Primary In-flight Ice Detection System. A Primary IIDS (PIIDS) is considered the sole means used to determine when the IPS must be activated. The IPS may be automatically activated by the PIIDS or the PIIDS may provide a flight deck signal that will direct the crew to manually activate the IPS. Installation of an ice detection system is considered a safety enhancement since the icing conditions may be identified or confirmed at an early stage and appropriate actions can be initiated in a timely manner. However, recent investigations indicate 15
that previously certificated ice detection systems may not detect airframe and engine icing in some part 25, Appendix C conditions. It has also been demonstrated in an icing wind tunnel that atmospheric moisture may fail to freeze on ice detector probes even though ice may be accreting on other airplane surfaces. With the continuing development of ice detection systems and due to recent in-service incidents, the FAA has determined there is a need to define specific criteria to certify ice detection systems that are used as the sole means of determining when the ice protection systems are activated.
(a) Applicable regulations. When certificating a PIIDS, compliance with §§ 23.929, 23.1093, 23.1301, 23.1309 and 23.1419 should be substantiated with the ice accretions that form prior to activation of the ice protection system and ice accretion during the system response time (e.g., for a thermal ice protection system, the time to heat the surface and remove the ice).
(b) System Safety Assessment. A PIIDS and associated components considered separately and in relation to other systems, shall be designed so that failure conditions classification and effects must be in compliance with § 23.1309. In particular, reliability of the PIIDS hardware and software should be established using means for showing compliance provided by AC 23.1309-1C.
1. The probability of encountering part 25, Appendix C icing conditions is considered to be 1.
2. An undetected failure of the PIIDS should be considered as catastrophic unless a lower failure condition classification is substantiated and agreed to by the FAA.
3. Information concerning unsafe PIIDS operating conditions must be provided to the flight crew. For example, loss of PIIDS capability should be annunciated.
(c) System Operation. To be compliant with § 23.1301 a PIIDS must be capable of detecting the presence of icing conditions or actual ice accretion under all atmospheric conditions defined in part 25, Appendix C, or the airplane must be capable of safely operating without restriction in the icing conditions that are not detected by the PIIDS. Certification data must be provided that substantiates that the PIIDS complies with applicable 14 CFR, part 23 requirements and the specific tests identified in this appendix. Compliance should be demonstrated by analysis that is verified using flight test data gathered and measured in natural icing conditions, and, as found necessary, by icing wind tunnel or tanker tests.
(d) Freezing Fraction Consideration. Certain icing conditions defined by part 25, Appendix C may result in failure of the liquid water to freeze on ice detector probes even though ice may be accreting on other airplane surfaces. In these cases the ice detector accretion may be insufficient to trigger the ice detection signal, or the response time for the detector to detect icing conditions may be extended such that significant ice can accrete on protected airplane or propulsion components. This phenomenon occurs when the freezing fraction is well below one. A specific assessment of this phenomenon should be accomplished for the PIIDS. Icing wind tunnel evaluations used to evaluate these conditions should include the use of actual size leading edge 16
models of the airfoils used on the airplane. Additionally the evaluation should include engine inlets, vanes and blades to assure that these conditions will not adversely affect the propulsion system unless specific AFM provisions are provided that assure engine ice protection independent of the PIIDS detection of icing conditions.
(e) Location of the Sensors. Performance of the PIIDS is affected by the physical installation and can only be verified after installation. The ice detecting devices should be installed at locations that assure safe flight by detection of icing prior to a hazardous build up of ice on the airframe or propulsion components. Proper location of ice detectors may require various computational fluid dynamic analyses and icing wind tunnel tests to assure that the ice detector is capable of adequately sampling the icing environment. It should be shown by analysis and/or flight test that the detection system sensors are located properly for all flight phases and airplane configurations and all combinations of icing conditions parameters as defined in part 25, Appendix C (liquid water content, mean effective drop diameter, temperature, altitude and horizontal extent).
Additionally, the airframe manufacturer should show that ice accretions occurring forward of the sensor, such as on the radome, do not interfere with the airflow ice sensing, and that shed ice will not damage the detectors. Also other probes forward of the ice detector should not interfere with the proper functioning of the detector.
(f) Annunciation. The PIIDS display(s) flight deck lights and crew alerting messages must be located so that they are within the seated flight crew's seat forward vision scan area while performing their normal duties.
(g) Icing Conditions Outside of Part 25, Appendix C. Ice detectors typically have limitations with certain atmospheric phenomena, e.g. ice crystal conditions, outside of part 25, Appendix C, which the airframe manufacturer must understand and not rely on the PIIDS to detect these conditions. There have been cases (an anomaly that has affected some engine models) where atmospheric ice crystals ingested in the initial fan stages of jet engines have melted, and re-accreted on subsequent low pressure compressor stages or initial stages to the high pressure compressor resulting in engine core ice. Subsequent shedding of this type of ice accretion has resulted in at least one unrecoverable engine surge. Therefore, airframe manufacturers must be aware of the potential limitations of engines during ice crystal conditions without reliance on the ice detection system. Current policy considers this potential condition in § 23.1093(b)(1)(ii) which requires demonstration of adequate engine performance in ice crystal conditions during falling or blowing snow. 14 CFR part 23 and part 33 regulations and compliance methods intend to provide unrestricted engine operation throughout the atmospheric environmental envelope.
k. Fluid (Freezing Point Depressant) Systems. Freezing point depressant fluid systems have been successfully certificated on part 23 airplanes. However, the system is not as common as other ice protection systems and that prompted the FAA in 1986 to publish information on certification of these systems in DOT/FAA/CT-TN86/11. Certification highlights from this publication are repeated below.
(1) Analyses. The two critical analyses required are the fluid flow rate required and an evaluation of the operational angles of attack, which will dictate chordwise coverage.
(2) Fluid Capacity. The fluid capacity does not have to exceed the maximum endurance of the airplane but the minimum should be as follows in Table 2.
TABLE 2. FLUID CAPACITY
Airplane Type Minimum Fluid Capacity is the greater of:
Turbojet powered airplanes 90 minutes or 15 percent of the maximum endurance based on the flow rate required in continuous maximum icing conditions Turbopropeller airplanes with maximum operating altitude above 30,000 feet Turbopropeller airplanes with maximum operating altitude 30,000 feet and below 150 minutes or 20 percent of the maximum endurance based on the flow rate required in continuous maximum icing conditions Reciprocating engine powered airplanes
(3) Fluid Quantity.
(a) There should be a fluid quantity indicator to allow the crew to determine how much longer the fluid will last. The fluid quantity indicator should be evaluated to determine that it is plainly visible to the pilot and that the indicator provided can be effectively read.
(b) If it is desired that the airplane be dispatched with less than full fluid in the reservoir, the AFM should contain a chart or table to allow the crew to determine the minimum fluid level. The AFM limitations should state a minimum dispatch fluid level of 90 minutes based on the flow rate required in continuous maximum icing conditions.
(4) Fluid Characteristics.
(a) The freezing point depressant fluids used become a gel at very cold temperatures and the temperature extremes to which the airplane will be operated should be considered when evaluating the reservoir, pump and plumbing installations.
(b) An accessible shutoff should be provided in systems using flammable fluids.
The alcohol used on propeller systems is flammable and should be addressed. The common fluid used in freezing point depressant systems in the U.S., AL5, has been shown to have low flammability properties.
(c) The effect of fluid ingestion into engines, Auxiliary Power Units (APUs), and accessories should be evaluated.
(d) The effect of fluid compatibility with electrical contacts and with composite materials should be evaluated.
(e) There should be sufficient AFM and maintenance manual warnings on handling fluids. AL5 is 85 percent mono-ethylene glycol, which is considered extremely toxic.
(f) Another freezing point depressant fluid, TKS80, is available. Icing tunnel tests of TKS80 and AL5 fluids have shown that they perform differently. The fluid reservoir should be placarded to permit only the fluid(s) that have been tested during certification.
(g) To avoid confusion, the fluid filler tank cap should be distinguishable from other caps such as fuel and a placard stating approved fluid should be located near the filler cap.
(h) There should be adequate drainage in the areas that hold or can receive spillage from leaks.
(5) Windshield Visibility. The effect of the fluid on windshield visibility should be evaluated to show compliance to § 23.775(f) if either the windshield or propellers are protected with a freezing point depressant system. On many of the approved installations the windshield system is turned off just prior to landing.
10. FLIGHT TEST PLANNING. When operating any airplane in an icing environment, degradation in performance and flying qualities may be expected. The primary purposes for flight testing an airplane equipped for flight in icing conditions is to evaluate such degradation, determining that the flying qualities remain adequate, and that performance levels are acceptable for this flight environment. For airplanes with a certification basis of 23-43 or higher, § 23.1419 requires that an airplane comply with the performance, stability, controllability and maneuverability regulations of part 23, Subpart B.
a. The flight tests and analyses of flight tests should:
(1) Evaluate normal operation of the airplane with the ice protection system installed in non-icing flight.
(2) Evaluate operation of the airplane with anticipated in-flight accumulations of ice.
(3) Verify the analyses conducted to show adequacy of the ice protection system throughout the icing envelope of part 25, Appendix C.
(4) Develop procedures and limitations for the use of the ice protection system in normal, abnormal, and emergency conditions.
b. Flight tests are generally conducted in three stages:
(1) Initial Dry Air Tests With Ice Protection Equipment Installed And Operating.
Initial dry air tests are usually the first steps conducted to extend the basic airplane certification to evaluate the airplane with the ice protection system installed and operating. The initial dry air tests are conducted to verify that the ice protection system does not affect the flying qualities of the basic airplane in clear air.
(2) Dry Air Tests With Predicted Simulated Ice Shapes Installed. Dry air tests with predicted, simulated critical ice shapes installed are usually the second step for certification for flight in icing. Airplane performance and handling characteristics are evaluated with these simulated ice shapes. Often, it is more economical to verify specific analyses by ground tests such as icing tunnel tests where the design variables can be controlled to some extent.
(3) Flight Tests in Icing Conditions. Flight tests in icing conditions, including artificial icing tests such as flight behind an airborne icing tanker aircraft, are normally employed to demonstrate that the ice protection system performs under flight conditions as the analysis or ground test indicated. These demonstrations should be made at various points (targeting critical points) in the icing envelope of part 25, Appendix C, to verify the airplane's ability to safely operate throughout that icing envelope.
11. FLIGHT TESTS. The following sections cover the major flight tests and/or analyses normally performed to substantiate the flight aspects of an ice protection system:
a. Initial Dry Air Tests with Ice Protection Equipment Installed. Depending upon the detailed design of the ice protection system, some preliminary ground tests of the equipment may be warranted to verify the basic function of each item. Quantitative data on such items as temperatures of thermal devices, fluid flow rates and flow patterns on liquid devices, or operating pressures of pneumatic components may be obtained as necessary to verify the system designs. The airplane should be shown to comply with the certification requirements when all ice protection system components are installed and functioning. This can normally be accomplished by performing tests at those conditions found to be most critical to basic airplane aerodynamics, ice protection system design, and powerplant functions. Section 23.1419 requires that the adequacy of the airplane's ice protection system be established based on operational needs. Tests must be conducted to demonstrate that the airplane is capable of operating safely in continuous maximum and intermittent maximum icing conditions as described in part 25, Appendix C. Several commonly used ice protection system components are discussed below to illustrate typical flight test practices. Other types of equipment should be evaluated as their specific design dictates.
(1) Pneumatic Leading Edge Boots.
(a) O peration. Boot inflation rate and inflated boot shape are important parameters in ice accretion removal. Tests should demonstrate a rapid rise and decrease in operating pressures 20
for effective ice removal. This pressure rise time, as well as the maximum operating pressure for each boot, should be evaluated throughout the altitude band of part 25, Appendix C—Mean Sea Level (MSL) to 22,000 feet above MSL—unless performance constraints or aircraft limitations unrelated to the ice protection system in the AFM restrict the airplane to a lesser altitude range.
(b) Minimum Operating Temperature. Boots should be operated in flight at the minimum envelope temperature (-22 degrees Fahrenheit (F) of part 25, Appendix C, to demonstrate adequate inflation/deflation and throughout the entire flight envelope. Boots should be operated near the proposed AFM operating temperature limit, if below -22 degrees F, to demonstrate that no damage occurs. The appropriate speed and temperature (if any) limitation on activation of boots should be included in the AFM.
(c) Effect of Inflated Boots. The operation of the boots (inflation) should have no hazardous affect on airplane performance and handling qualities. An example of an unacceptable hazardous effect is that some boot inflation sequencing schemes result in abnormal pitch attitude changes. If there are anomalous pitch changes in any configuration, appropriate information or limitation should be in the AFM. This can be shown by inflating the boots at several speeds in the flight envelope from stall speed to (VNE + VD)/2 or (VMO + VD)/2 and observing the reaction of the airplane. This test can be hazardous and should be approached in a build-up manner by inflating the boots at incremental airspeeds starting from the middle of the flight envelope.
(d) Cockpit Annunciations. Section 23.1309(b)(3) states that warning information must be provided to alert the crew to unsafe system operating conditions, and to enable them to take appropriate corrective action. Systems, controls, and associated monitoring and warning means must be designed to minimize crew errors, which could create additional hazards. Section 23.1416(c) requires “that means to indicate to the flight crew that the pneumatic deicer boots system is receiving adequate pressure and is functioning normally must be provided.” Stall speed can increase dramatically with inflated boots and failures in the vacuum system need to be addressed. If the indicating lights for the pneumatic deicing system (boots) illuminate at a lower pressure than required for proper boot operation, misleading information is furnished to the flight crew, which can, in turn, lead to an unsafe operation. This could mislead the crew to believe the boots are operating normally when, in fact, the boots might not be inflating sufficiently to shed ice.
(e) Water Contamination. Consideration should be given to the potential for accumulation of liquid water inside the pneumatic deicing boots, which could freeze within the system and prevent proper operation. The pneumatic and boot arrangement should be examined for low points, which may collect water, and consideration should be given to the installation of water drainage points. Periodic inspection and drainage procedure instructions should be provided in the appropriate manual. Similarly, placement of the pressure sensor(s) should be evaluated to prevent misleading boot inflation indications. An evaluation of the effectiveness of water/air separators and/or drainage holes should be accomplished by flying through precipitation followed by a demonstration of system operation at freezing altitudes.
(2) Electric Propeller Boots.
(a) Dry Air Function Flight Test.
1. System Function. When flying in dry air, the systems should be monitored to confirm proper function. It is suggested that system current, brush block voltage (between each input brush and the ground brush), and system duty-cycles be monitored to assure that proper power is applied to the deicers.
2. Temperatu re Measurements. If not furnished by the manufacturer, surface temperature measurements may be made during dry air tests. These surface temperature measurements are useful for correlating analytically predicted dry air temperatures with measured temperatures or as a general indicator that the system is functioning and that each deicer is heating.
3. Vibration. The system operation should be checked throughout the full certificated Revolutions Per Minute (RPM) and propeller cyclic pitch range expected during icing flights. Any significant vibrations should be investigated.
(b) Maximum Temperature. Consideration should be given to the maximum temperatures that a composite propeller blade may be subjected to when the deicers are energized. It may be useful to monitor deicer bond-side temperatures. When performing this evaluation, the most critical conditions should be investigated. For example, this may occur on the ground (propellers non-rotating) on a hot day with the system inadvertently energized.
Service difficulty reports have indicated propeller damage during de-ice/anti-ice system activation during maintenance without the engines running.
(c) Precipitatio n. The system should be monitored to confirm proper function in precipitation. There have been designs that allowed water to reach the electrical brush blocks and prevent their operation.
(3) Electric Windshield Anti-Ice.
(a) Thermal Analysis Validation. Dry air flight tests should be conducted in support of the systems design, as required. Inner and outer windshield surface temperature evaluations of the protected area may be needed to support thermal analyses. Thermal analysis should substantiate that the surface temperature is sufficient to maintain anti-icing capability without causing structural damage to the windshield. In the case of add-on plates, temperatures of the basic airplane windshield, inside and out, may also be needed, particularly with pressurized airplanes.
(b) Cockpit Visibility. An evaluation of the visibility, including distortion effects through the protected area, should be made in both day and night operations to show compliance to § 23.775(f).