(viii) Flight testing. It should be shown by flight tests that the device is safely controllable and maneuverable during all requested flight regimes without requiring exceptional piloting skill. The flight tests should entail the PCDS weighted to the most critical weight. Litters and other types of PCDS designs may spin, twist or otherwise respond unacceptably in flight. These designs should be structurally restrained with devices such as a spider, a harness, or an equivalent device to minimize undesirable flight dynamics.
(ix) Medical design considerations. The PCDS should be designed to the maximum practicable extent and placarded to maximize the HECs protection from medical considerations such as blocked air passages induced by improper body configuration and excessive loss of body heat during operations. Injured or water soaked persons may be exposed to high body heat loss from sources such as rotor wash and the airstream. PCDS occupant safety from transit induced medical considerations can be greatly increased by proper design.
(x) Hoist operator safety device. W hen hoisting operations require the presence of a hoist operator on board, appropriate provisions shall be provided to allow the hoist operator to perform his task safely. These provisions shall include an appropriate hoist operator restraint system. This safety device is typically composed of a safety harness and a strap attached to the cabin used to adequately restrain the hoist operator inside the cabin while operating the hoist. For certification approval, t he hoist operator safety device should comply with § 29.561(b)(3) for personnel safety. The applicant should submit a design evaluation that assures the necessary level of personnel safety is provided. The safety harness and strap should meet TSO C22g, TSO C114, or TSO C167 requirements; however, as a minimum, the following should be evaluated:
(A) The strap attaching point on the body harness should not be located in the back of the hoist operator due to safety issues in case of a fall or crash.
(B) The safety device should be designed to be adjustable so that the strap is tightened behind the hoist operator.
(C) The strap might include a QRS including a DAD to allow hoist operator detaching from cabin in emergency conditions (e.g., crash).
(D) The safety device should be easily and readily ingressed or egressed.
(E) It should be placarded for proper capacity and lifetime limitation.
(F) Any fabric used should be durable and should be at least flame resistant.
(11) Section 29.865(c)(4) Intercom Systems for HEC Operations. For all HEC operations, the rotorcraft is required to be equipped for, or otherwise allow direct Page D - 161
AC 29-2C, Chg 7 2/4/2016 intercommunication under any operational conditions among crewmembers and the HEC. An intercommunications system may also be approved after the external load system certification as operators are standardizing communications equipment within their fleet for safety benefits. In lieu of an intercommunication system being provided as part of the external load system, this onus may be moved onto the operator through a limitation may be placed in the RFM or RFMS (see paragraph d.(4) of this guidance).
(12) Section 29.865(c)(6) HEC Operations Requiring Category A Rotorcraft. For jettisonable HEC operations, it may be required by Operations Requirements, § 133.45(e), that the rotorcraft meet the category A engine isolation requirements of part 29 and that the rotorcraft have OEI hover performance capability OGE in its approved, jettisonable HEC weight, altitude, and temperature envelope.
(i) In determining OEI hover performance, dynamic engine failures should be considered. Each hover verification test should begin from a stabilized hover at the maximum OEI hover weight, at the requested IGE or OGE skid or wheel height, and with all engines operating. At this point the critical engine should be failed and the aircraft should remain in a stabilized hover condition without exceeding any rotor limits or engine limits for the operating engine(s). As with all performance testing, engine power should be limited to minimum specification power. Engine failures may be simulated by rapidly moving the throttle to idle provided a ‘needle split’ is obtained between the rotor and engine RPM.
(ii) Pilot reaction times used following the engine failure to maintain the stabilized hover flight condition will depend on the environment in which such operations are planned. For HEC operations planned for day visual meteorological conditions (VMC), normal pilot reaction times are acceptable. Recent certification projects for HEC operations planned to take place overwater at night while in VMC or in instrument meteorological conditions (IMC) required greater pilot reaction times. When hovering OGE or IGE at maximum OEI hover weight, an engine failure should not result in an altitude loss of more than 10 percent or four (4) feet, whichever is greater, of the altitude established at the time of engine failure. In either case, sufficient power margin should be available from the operating engine(s) to regain the altitude lost during the dynamic engine failure and then to transition to forward flight.
(iii) Consideration should also be given to the time required to recover (reel in and bring aboard) the HEC and then to transition to forward flight. This time increment may limit the use of short duration OEI power ratings. For example, for a helicopter that sustains an engine failure at a height of 40 feet, the time required to stabilize the hover at the original hover height, recover the HEC (given the hoist speed limitations), and then transition to forward flight (with minimal altitude loss) would likely preclude the use of 30-second engine ratings and may encroach upon the 2 ½ -minute ratings. Such encroachment into the 2 ½ - ratings is not acceptable. In accordance with the intent of OEI power ratings, recovery and transition procedures shall not utilize more than one iteration of either a 30-second / 2-minute power combination or a 2½-minute rating.
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(iv) For helicopters that incorporate engine driven generators, the hoist should remain operational following an engine or generator failure. A hoist should not be powered from a bus that is automatically shed following the loss of an engine or generator. Maximum two- engine generator loads should be established so that when one engine or generator fails, the remaining generator can assume the entire rotorcraft electrical load (including the maximum hoist electrical load) without exceeding approved limitations.
(v) The RFM or RFMS should contain information that describes the expected altitude loss, any special recovery techniques, and the time increment used for recovery of the external load when establishing maximum weights and wheel or skid heights.
Engine failure procedures that permit safe recovery of the HEC and safe fly-away capability shall be included in the RFM or RFMS The OEI hover chart should be placed in the performance section of the RFM or RFMS. Allowable altitude extrapolation for the hover data should not exceed 2000 feet. The following statement (modified appropriately) should be included to indicate to operators which charts are applicable to these HEC operations.
• For external load equipment installations specifically intended for HEC applications requiring the use of Transport Category A rotorcraft, OEI hover performance data, and procedures may be found in Figures/Charts/Section XX - YY…
(13) Section 29.865(e) External Loads Placards and Markings. Placards and markings should be installed next to the external load attaching means, in a clearly noticeable location, that state the primary operational limitations - specifically including the maximum authorized external load. Not all operational limitations need be stated on the placard (or equivalent markings) only those clearly necessary for immediate reference in operations. Other more detailed operational limitations of lesser immediate importance should be stated either directly in the RFM or in a RFMS.
(14) Other Considerations.
(i) Agricultural Installation (AI). AIs can be approved for either jettisonable or non-jettisonable NHEC or HEC operations as long as they meet relevant certification and operations requirements and follow appropriate compliance methods. However, most current AI designs are external fixtures (see definition) - not external loads.
External fixtures are not approvable as jettisonable external cargo because they do not have a true payload (see definition), true jettison capability (see definition), or a complete QRS. Many AI designs can dump their solid or liquid chemical loads by use of a "purge port" release over a relatively long time period (i.e., greater than 30 seconds).
This is not considered true jettison capability (see definition) since the external load is not released by a QRS and since the release time span is typically greater than 30 seconds (see paragraph c.(21)). Thus, these types of AIs shoul d be approved as a non-jettisonable external load. However, other designs that have the entire AI (or significant portions thereof) attached to the rotorcraft, that have short time frame jettison Page D - 163
AC 29-2C, Chg 7 2/4/2016 (or release) capability provided by a QRS that meets the definitions herein and that have no post-jettison characteristics that would endanger continued safe flight and landing may be approved as a jettisonable external load. For example, if all the relevant criteria are properly met, a jettisonable fluid load can be approved as a NHEC external cargo. Section AC 29 MG 5 of this AC discusses other AI certification methodology.
(ii) External Tanks. External tank configurations that have true payload (see definition) and true jettison capability (see definition) should be approved as jettisonable NHEC. External tank configurations that have a true payload capability but do not have true jettison capability should be approved as non-jettisonable NHEC. An external tank that has neither a true payload capability nor true jettison capability is an external fixture; it should not be approved as an external load under § 29.865. If an external tank is to be jettisoned in flight, it should have a QRS that is approved for the maximum jettisonable external tank payload and is either inoperable or is otherwise rendered reliable to minimize inadvertent jettisons above the maximum jettisonable external tank payload.
(iii) Logging Operations. These operations are very susceptible to low -cycle fatigue because of the large loads and relatively high load cycles that are common to this industry. It is recommended that load-measuring devices (such as load cells) be used to assure that no unrecorded overloads occur and to assure that cycles producing high fatigue damage are properly considered. Cycle counters are recommended to assure acceptable cumulative fatigue damage levels are identifiable and are not exceeded. As either a supplementary method or alternate method, maintenance instructions should be considered to assure proper cycle counting and load recording during operations.
(15) Noise Certification. 14 CFR part 36 is the noise certification standard.
Section 36.1(a)(4) specifically exempts helicopters that are designed exclusively for agricultural work, carrying firefighting materials, or external loads activity from the noise standards. Section 21.93(b)(4) also contains specific information regarding external loads and what configurations do not constitute an acoustical change.
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SUBPART D - DESIGN AND CONSTRUCTION
MISCELLANEOUS AC 29.871. § 29.871 LEVELING MARKS.
a. Explanation. Reference marks are required for leveling the rotorcraft on the ground. These marks are necessary for accu rate determination of weight and balance effects, particularly after modifications to the basic rotorcraft.
b. Procedures.
(1) Reference marks are sometimes provided in pairs, one high in the cabin and one low. The plumb weight is suspended from the high mark by an appropriate mechanical attachment, and the lower mark is used to level the rotorcraft by centering the plumb weight. The lower reference mark should be a raised or depressed target symbol and shall be applied to a permanent structural component or permanently attached plate in a readily accessible location. Seat tracks, floors, or door sills which are attached with permanent fasteners are typical locations.
(2) Horizontal reference marks for support of bubble levels may also be used, particularly for smaller rotorcraft.
(3) Proper reference should be made to identify the leveling marks or points on the rotorcraft. Design provisions should be made to ensure these locations are not obscured by equipment, fairings, repair, or rework.
AC 29.873. § 29.873 BALLAST PROVISIONS.
a. Explanation.
(1) This rule requires that ballast provisions prevent inadvertent ballast shifting while in flight or as a result of a landing. Shifting of the ballast may cause a hazardous change in the center of gravity thereby affecting rotorcraft controllability.
(2) Other rules noted here allow removable and fixed ballast and require markings or placards to prevent overloading the ballast installation.
(i) Section 29.29 specifies that the rotorcraft empty weight will include any fixed ballast. Section 29.31 allows the use of removable ballast to comply with the flight requirements. However, ballast may not be adjusted (moved, reduced or increased) in flight.
(ii) Section 29.1541 requires conspicuous and durable markings or placards. Section 29.1557 requires placards stating allowable maximum weight, distributed loading, if necessary, and other appropriate limitations for ballast installation.
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(3) Section 29.1583(c) concerns Rotorcraft Flight Manual instructions and information about removable ballast or loading information. The instructions must be included in the operating limitations section of the flight manual to allow ready observance of the limitations.
b. Procedures.
(1) The ballast installation may be substantiated by analysis or by static test.
The design ultimate load may be derived from flight, landing, or minor crash conditions load factor specified in the rules. Substantiation by analysis will require use of the fitting factor prescribed by § 29.625 where appropriate. If static tests are to be conducted, a test plan should be prepared, submitted for evaluation and agreed upon prior to the test.
(2) Ballast installations in the aft part of the fuselage and tail boom may be subject to significant landing condition angular inertia load factors as well as the usual linear load factors.
(3) Substantiation methods and procedures acceptable for the airframe substantiation may be used for the ballast installation as well.
(4) Removable ballast will require attention to assure the ballast is secured easily and properly and will remain secured under the appropriate ballast design load factor requirements. The f light manual instructions should be evaluated for compliance with § 29.1583(c) by flight test and airframe personnel.
(5) The installation must be designed and placarded or marked for the maximum allowable ballast load and for other appropriate loading limits. Normally compliance with § 29.1541 is accomplished with a drawing review by airframe personnel along with an EMDO compliance and conformity inspection. An additional compliance inspection by airframe personnel can be conducted if desired.
AC 29.877. § 29.877 ICE PROTECTION.
NOTE: § 29.877 was removed and replaced by § 29.1419 in Amendment 29-21. This material is retained since this is one way to show compliance with § 29.877.
a. Background.
(1) In March 1984, the FAA/AUTHORITY for the first time certificated a rotorcraft for flight into known icing conditions. Several other manufacturers are pursuing designs for icing flight capability with certification planned for 1985 or 1986.
2) Most rotorcraft icing technology has been developed for military rotorcraft.
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liquid water content (LWC) does not exceed 1.0gm/m3 and outside air temperature (OAT) is not below -20° C.
(3) Many rotorcraft operators have voiced a high priority on obtaining rotorcraft approved for operation in icing conditions.
(4) The icing characteristics envelope of FAR Part 25, Appendix C, has served as a satisfactory design criteria for fixed-wing operations for two decades. The envelope, as presented, extends to 22,000 feet with possible extents to 30,000 feet but does not present icing severity as a function of altitude. At the time the envelope was derived, it was assumed that all transport category airplanes would operate to at least 22,000 feet. For present state-of-the-art rotorcraft, this assumption is not valid. As such, an altitude limited icing envelope based on the same data used to derive the Part 25, Appendix C, and the Part 29, Appendix C, envelopes is presented as an alternate to the full icing envelope. In addition, a second icing envelope which effectively characterizes supercooled clouds from ground level to 10,000 feet is presented as a second alternative to the Part 29, Appendix C, envelope. The second altitude limited envelope described in reference 386d(2) was derived from recent additional airborne measurements.
b. Explanation.
(1) General.
(i) The discussion in this paragraph pertains generally to certifications to the full icing envelope of Part 29, Appendix C, within the altitude limitations of the rotorcraft or to one of the altitude limited icing envelopes based on a 10,000-foot pressure altitude limit. The actual icing envelope considered may be further restricted based on the actual pressure attitude envelope for which certification is requested. It envisions certification with full ice protection systems (rotor blades, windshields, engine inlets, stabilizer surfaces, etc.). With the exception of pilot controllable variables such as altitude and airspeed, limited certification (either in terms of icing envelope or protection capability) is not envisaged at this time due to the difficulty in forecasting the severity of icing conditions, relating the effects of the forecasted conditions to the type of aircraft, and relating the effects of reported icing among various types of aircraft, particularly between fixed and rotary-wing aircraft. In addition, with a limited protection capability, viable escape options may not be operationally available if limitations are exceeded.
(ii) The discussion in this paragraph, regarding rotor blade ice protection, is oriented primarily toward electrothermal rotor deicing systems, since these have the most widespread acceptance and projected use within the industry. Also, most of the testing and research into rotorcraft ice protection to date has been conducted with this type of system. Research is continuing with other types of systems such as anti -icing fluid systems, and information will be added to address certification of these as necessary. It should also be noted that most of the rotorcraft icing experience Page D - 167
accumulated to date has been on rotorcraft with symmetrical airfoil sections. The application of this experience to rotorcraft with asymmetrical airfoils should be carefully evaluated. Limited experience has been gained during development and qualification testing of the Army Blackhawk on asymmetrical airfoil icing characteristics. The most prominent difference appears to be a more rapid degradation of airfoil performance.
Rapidity of performance degradation is also dependent upon severity of the icing condition (primarily a function of liquid water content) and ice shape (primarily a function of OAT and median volumetric droplet diameter (MVD)).
(iii) The effects of ice can vary considerably from rotorcraft to rotorcraft.
Experience gained for a rotor system with an identical blade profile could provide valuable information but should be used cautiously when applied to another rotorcraft.
Assumptions cannot necessarily be made based on icing test results from another rotorcraft. Particular care should be exercised when drawing from fixed- wing icing experience as the widely different and varying conditions seen by the rotor blades make many comparisons with fixed-wing results invalid. Likewise, icing effects on rotor blades vary significantly from those on other parts of the rotorcraft. This is due to changing blade velocity as compared with the constant velocity of the remaining parts.
(2) Reference Material. Prior to commencement of efforts to design and certify a rotorcraft, the references listed in paragraph AC 29.877d should be reviewed. FAA Technical Report ADS-4, Engineering Summary of Airframe Icing Technical Data, December 1963, although somewhat dated, is recommended for basic aircraft icing protection system design information.
(3) Objective. The objective of icing certification is to verify that throughout the approved envelope, the rotorcraft can operate safely in icing conditions expected to be encountered in service (i.e., Appendix C of Part 29 or one of the altitude limited icing envelopes presented herein). This will entail determining that no icing limitations exist or defining what the limitations are, as well as establishing the adequacy of the ice warning means (or system) and the ice protection system. A limiting condition may manifest itself in one of several areas such as handling qualities, performance, autorotation, asymmetric shedding from the rotors, visibility through the windshield, etc.
Prior to flight tests in icing conditions, sufficient analyses should have been conducted to determine the design points for the particular item of the rotorcraft being analyzed (windshield, engine inlet, rotor blades, etc.). After the analyses are reviewed and found adequate, tests should be conducted to confirm that the analyses are valid and that the rotorcraft can operate safely in any supercooled cloud icing condition defined by Part 29, Appendix C, or one of the altitude limited icing envelopes. References 386d(1) and (3) may be useful in determining the design points and extrapolation of test data to the desired design points.
(4) Planning. F or best utilization of both the applicant’s and the FAA/AUTHORITY’s resources, the applicant should submit a certification plan at the start of the design and development effort. The certification plan should describe all Page D - 168
efforts intended to lead to certification and should include the following basic information:
Rotorcraft and systems description.
Ice protection systems description.
Certification checklist.
Description of analyses or tests planned to demonstrate compliance.
Projected schedules of design, analyses, testing, and reporting efforts.
Methods of test - artificial vs. natural.
Methods of control of variables.
Data acquisition instrumentation.
Data reduction procedures.
(5) Environment.
(i) Definitions.
(A) Supercooled Clouds. Cloud s containing water droplets (below 32° F) that have remained in the liquid state. Supercooled water droplets will freeze upon impact with another object. Water droplets have been observed in the liquid state at ambient temperatures as low as -60° F. The rate of ice accretion on an aircraft component is dependent upon many factors such as droplet size, cloud liquid water content, ambient temperature, and component size, shape, and velocity.
(B) Ice Crystal Clouds. Glaciated clouds existing usually at very cold temperatures where moisture has frozen to the solid or crystal state.
(C) Mixed Conditions. Partially glaciated clouds at ambient temperatures below 32° F containing a mixture of ice crystals and supercooled water droplets.
(D) Freezing Rain and Freezing Drizzle. Precipitation existing within clouds or below clouds at ambient temperatures below 32° F where rain droplets remain in the supercooled liquid state.
(E) Sleet. Precipitation of transparent or translucent pellets of ice which have a diameter of 5mm or less.
(F) Hail. Solid precipitation in the form of balls or pieces of ice (hail stones) with diameters ranging from 5mm to more than 50mm.
(ii) Appendix C of Part 29 defines the supercooled cloud environment necessary for certification of rotorcraft in icing except that the pressure altitude limitation is that of the rotorcraft or that selected by the applicant, provided the remaining altitude envelope is operationally practical. Due to air traffic system compatibility constraints, approval of a maximum altitude less than 10,000 feet pressure altitude should be discouraged. However, there are operations where a lower maximum altitude has no Page D - 169
effect on the air traffic system and would still be operationally useful. Figures 3 and 6 of Appendix C, Part 29, relate the variation of average LWC as a function of cloud horizontal extent. These relationships should be used for design assessment of the most critical combinations of conditions as a function of en route distance. This, in combination with a capability to hold in icing conditions for 30 minutes at the destination, is commensurate with policies previously established for fixed-wing aircraft. Figures 3 and 6 should be used in conjunction with the altitude limited criteria of figures AC 29.877-1 through AC 29.877-4 herein. The new criteria of figure AC 29.877-5 includes “duration” (horizontal extent) as the third dimension. It is emphasized that LWC extremes expressed in Part 29, Appendix C, criteria and the alternate envelopes represent the maximum average values to be anticipated within an exceedance probability of 99.9 percent. Transient, instantaneous peak values of much higher LWC have been observed. These instantaneous peak values appear to be of little significance to the design of protected and unprotected surfaces; however, these high values, if encountered, may induce shedding of ice from some unprotected surfaces. This is due to radical changes in the rate of release of latent heat and resultant changes in the structural properties and adhesion force of ice.
(iii) A recent analysis performed at the FAA Technical Center concludes that the aircraft icing environment below 10,000 feet is not as severe in terms of LWC and OAT as that depicted in Part 29, Appendix C, envelope. This AC presents two different altitude limited envelopes that may be employed by those applicants who elect to certify with a 10,000-foot pressure altitude limit. One of these altitude limited envelopes is based upon the same data that were used to derive the design criteria of the Part 29, Appendix C (figures AC 29.877-1 thru 4), while the other is based upon a recently established characterization of supercool clouds below 10,000 feet (figure AC 29.877-5). The applicant may select either of the approaches to altitude limitation. At the present time, applicants have not consistently selected one or the other. If experience shows a unanimous preference for one or the other, the one not used will be deleted in a future revision. The data used to derive these limited envelopes cannot be used to further define icing conditions between 10,000 feet and 22,000 feet; hence, above 10,000 feet, the Part 29, Appendix C, envelopes should be used. It should be noted that the engine inlets should still meet the icing requirements of § 29.1093. The limited icing envelopes may be used on an equivalent safety basis to show compliance with the intent of § 29.1093 if the altitude limit established for the rotorcraft is not greater than 10,000 feet.
(iv) Significant effects can result from various combinations of parameters.
For example, most rapid ice accumulations occur at the high values of liquid water content, and the greatest impingement area occurs at the high values of droplet size.
Most critical ice shapes are a function of each of these parameters in addition to airspeed, surface temperature, and surface contour. Care should be taken to explore the entire specified ranges of these parameters during the design, development, and certification efforts.
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(v) Mixed conditions (i.e., a combination of ice crystals and supercooled water droplets) and freezing rain or freezing drizzle are not addressed in the Part 29 environmental criteria but can present more severe icing conditions than those defined.
Although the probability of encountering freezing rain is relatively low, mixed conditions commonly occur in supercooled cloud formations. Little data have been gathered on the effects of encountering mixed conditions (see reference AC 29.877d(7)). There are no criteria for certification in mixed conditions or freezing rain at present. In addition to the hazards of operating any aircraft in icing, certain aspects of rotorcraft icing (relatively low altitude operation, asymmetric shedding with resulting vibration, and ice damage or ingestion) warrant a caution notice in the RFM advising that the rotorcraft is not certified for operation in freezing rain or freezing drizzle. Avoidance procedures (e.g., climb or descent) may also be useful.
(6) Flight Test Prerequisites.
(i) The prototype rotorcraft should be capable of IFR and IMC flight.
(ii) Sufficient analyses should be developed, submitted, and accepted by FAA/AUTHORITY to show that the rotorcraft is capable of safely operating to the selected design points of both the continuous maximum and intermittent maximum conditions of Part 29, Appendix C, or one of the altitude limited icing envelopes. A detailed failure modes and effects analysis (FMEA) should be performed.
(iii) Specific attention should be given to (1) assuring that the selected design condition(s) of atmospheric and rotorcraft flight envelopes have been identified;
(2) qualification and design of ice protection systems and components; and
(3) component installation and ice formation effects upon basic rotorcraft structural properties and handling qualities. These assurances can be established from analyses, bench test, and/or dry air flight tests or simulated icing tests, as appropriate prior to flight tests in natural icing.
(iv) The applicant should assess rotor blade stability with ice deposits to assure that dynamic instability will not occur in icing conditions. This assessment may be accomplished by analysis including consideration of failure of the most critical segment of the rotor blade ice protection system. It also may be accomplished by experimental means such as attaching dummy ice shapes to the blades and using a whirl stand or wind tunnel.
c. Procedures.
(1) Compliance.
(i) In general, compliance can be established when there is reasonable assurance that while operating in the specified icing environment (1) the engine(s) will not flameout or experience significant power losses or damage; (2) stress levels are not reached with ice accumulations that can endanger the rotorcraft or cause serious Page D - 171
reductions in component life; (3) the handling qualities, performance, visibility, and systems operation are defined and are not deteriorated unacceptably; (4) inlet, vent or drain blockage (such as fuel vent, engine, or transmission cooler) is not excessive; and
(5) autorotation characteristics are acceptable with maximum ice accretion between deice cycles. Assessment of performance loss should include not only the drag and weight of the ice itself but electrical or other load demands of the ice protection system and any performance changes resulting from modified rotor blade contours.
(ii) It is emphasized that ice formations (shape, weight, etc.) vary significantly under varying conditions of outside air temperature (OAT), liquid water content (LWC), median volume diameter (MVD), airspeed, attitude, and rotor RPM. The most critical conditions should be defined by means of analyses or test and verified by test. Performance changes under these various conditions should be determined and found acceptable.
(iii) Laboratory, icing tunnel, ground spray rig, and airborne icing tanker tests are all very useful in developing an ice protection capability, but none of these, either individually or collectively, can satisfy the full requirements for certification. None can presently duplicate the combinations of liquid water content, droplet size, flow field, and random shedding patterns found in natural icing conditions. Airborne tankers hold considerable promise of being able to fulfill certification requirements (in addition to the advantage of being able to produce an icing environment on demand rather than having to wait for it to occur in nature), but tankers have not been able to generate droplet sizes that cover the complete envelope for certification. Many improvements have been made in some tankers in recent years; however, large droplet sizes have typically been a problem. Also, the size of existing tanker clouds is not of sufficient cross section to immerse the entire rotorcraft. There are also solar radiation and relative humidity effects to be considered and correlated with natural icing when using a tanker. The tanker should be able to immerse the entire rotor system as a minimum and should have a means of controlling and changing the cloud characteristics uniformly and repeatably. Until an artificial method has been successfully demonstrated and accepted, icing certification should include flight tests in natural icing conditions.
(iv) Flight testing in natural icing conditions also has limitations.
Reference AC 29.877d(16) contains information that may be useful in planning natural icing flight tests. The key limitation of natural icing flight tests is being able to find the combinations of conditions that comprise critical design points. This is especially true of those points falling near the 99.9 percentile of exceedence probability; e.g., high LWC at low OAT with large MVD. It is emphasized that some more severe design points, however, may exist within the atmospheric icing envelope rather than near the edges or corners of the envelope. This does not mean that natural icing tests must be conducted at all the selected design conditions. Natural icing tests should be conducted in conditions as close to design points as possible and sufficient correlation shown with the analyses to assure that the rotorcraft can operate safely throughout the design envelope.
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