7/30/97 АС 29-2B
(6) The descent should conclude at either the maximum altitude at which level flight can be maintained with one engine inoperative or at the minimum practical altitude, whichever is higher.
(7) The OE powers available to be utilized during the descent would be the same as those prescribed previously for OEI climb cooling. OE operation should continue until component and fluid temperatures stabilize.
(8) The airspeeds utilized in the climb and descents should be representative of normal speeds unless cooling provisions are sensitive to rotorcraft airspeed, in which case the airspeeds most critical for cooling should be used. In no case, however, should it be required that the selected airspeeds exceed the speeds established under §§ 29.67(a)(2) and 29.67(b).
Par 518 815
MULTIENGINE POWER AVAILABLE 20K
HOT DAY- 15K PRESSURE ALTITUDE OAT 10° C CRITICAL ALTITUDE 15° C 10K - - POINT "A"- 20° C 5K 25° C 30° C POWER FIGURE 518-1 ALL-ENGINE CRITICAL ALTITUDE
ADDITIONAL COOLING TEST POINT 816 Par 518
519. § 29.1047 (Amendment 29-1) TAKEOFF COOLING TEST PROCEDURES.
a. Objective.
(1) For Category A rotorcraft, the objective is to verify satisfactory takeoff and OEl climb cooling for the Category A takeoff profile defined in aircraft performance §§ 29.59(c) and 29.67(a) following a prolonged hover.
(2) For Category B rotorcraft, the objective is to verify satisfactory cooling for the takeoff and subsequent climb for the Category B takeoff defined in performance §§ 29.63 and 29.65(a) following a prolonged hover.
b. Procedure - Category A.
(1) The rotorcraft is hovered in-ground-effect (IGE) at the power required to hover on the test day at the maximum Category A takeoff gross weight for the hot day, until temperatures stabilize.
Alternate Procedure. If the test day OAT is high, it may not be possible to hover IGE at the prescribed gross weight without entering the takeoff range on the measured gas temperature (MGT) indicator. Since operations in the takeoff range are allowed only for 5 minutes and the typical stabilization time is 20 to 35 minutes, it is permissible to reduce the initial aircraft gross weight so the initial MGT will be at least at the MCP limit, but will not be in the takeoff range for more than 5 minutes; and (її) The fuel burn during the anticipated 20 to 35 minute stabilization period may cause the aircraft to leave the prescribed hover IGE condition unless power is reduced or additional weight is added by fluid transfer or other methods. It is permissible to reduce power to maintain the IGE hover for this phase of testing rather than attempt special weight control procedures.
(2) After temperatures have stabilized, an OE climb is initiated from the lowest practicable altitude.
Multiengine power may be used from the stabilized IGE hover to the CDP before OE operations for cooling verification begin.
(ii) Actual shutdown of the simulated failed engine may not be necessary if the applicant can show that cooling of the remaining engine, fluids, and components is not affected by operation of the "failed" engine at idle power.
(ill) The power utilized at the initiation of the OEI climb would be the same as for establishing the takeoff climbout path of § 29.59, typically 2.5-minute OEl hot-day power available.
Par 519 817
(3) After the time period for which the power is used in establishing the takeoff climbout path has expired, OEI power is changed to that used in meeting the steady rate of climb (150 FPM, 1,000 feet above the takeoff surface of § 29.67 (a)(2)).
The power to be used for this phase is 30-minute OEl hot-day power available, if approval of this power rating for performance is requested.
(ii) If 30-minute OEl approval is not requested, the highest hot-day power available approved for continuous usage under OEI conditions would be utilized.
(4) Climb at the OEl power used in meeting § 29.67(a)(2) would continue for at least--
(i) Thirty minutes if 30-minute OE power is used; or
(ii) Five minutes after the occurrence of the highest temperature recorded, if other than 30-minute OEl is used.
(5) Unlike § 29.1045, the procedure set forth in § 29.1047 for Category A rotorcraft does not specifically require continuation of the OEI climb beyond the 30-minute duration allotted for 30-minute OEl power usage.
c. Procedure - Category B.
(1) The rotorcraft is hovered IGE until temperatures stabilize at the power required on the test day to hover IGE at the maximum Category B takeoff gross weight for the hot day.
(i> Alternate Procedure. If the test day OAT is high, it may not be possible to hover IGE at the prescribed gross weight without entering the takeoff range on the MGT indicator. Since operation in the takeoff range is allowed only for 5 minutes and the typical stabilization time is 20 to 35 minutes, it is permissible to reduce the initial aircraft gross weight so the initial MGT will be at least at the MCP limit, but will not be in the takeoff range for more than 5 minutes; and
(ii) The fuel burn during the anticipated 20 to 35 minute stabilization period may cause the aircraft to leave the prescribed hover IGE condition unless power is reduced or additional weight is added by fluid transfer or other methods. It is permissible to reduce power to maintain the IGE hover for this phase of testing rather than attempt special weight control procedures.
(2) After temperatures have stabilized in hover IGE, a multiengine climb is initiated at hot-day takeoff power available from the lowest practicable altitude.
Section 29.1047(b)(3) requires only that takeoff power be maintained for the same time 818 Par 519
interval as used in determining the takeoff flight path under § 29.63. This time interval could be less than the 5 minutes for which takeoff power is approved. Unless the applicant can show that the time interval used in § 29.63 provides more conservative results, or unless additional testing is proposed, the full 5 minutes allowed for takeoff power should be used to assure that the most critical condition has been surveyed.
(3) After the use of takeoff power for the appropriate time interval, the power should be reduced to multiengine maximum continuous hot-day power available and the climb continued until at least 5 minutes after the occurrence of the highest temperature recorded.
(4) The airspeeds utilized in the climb should be representative of normal speeds unless cooling provisions are sensitive to rotorcraft airspeed, in which case the airspeed most critical for cooling should be used. The airspeed need not exceed the speed for best rate of climb with maximum continuous power available.
520. § 29.1049 HOVERING COOLING TEST PROCEDURES.
a. Objective. The objective is to verify satisfactory hover IGE cooling at sea level and at the hover ceiling for hot-day conditions.
b. Explanation. The rule provides for a hover IGE cooling check in still air at sea level and at the hover ceiling at maximum continuous power. Still air is interpreted as a wind speed of 5 knots or less.
c. Procedures.
(1) The aircraft should be hovered IGE at the maximum certificated hover weight or at the IGE hover weight corresponding to hot-day maximum continuous power available, whichever is less.
(1) The power utilized would normally be hot-day maximum continuous power available and the initial gross weight would be selected as required to achieve hover IGE on the test day.
(ii) After initiation of the hover, special weight control procedures need not be implemented in attempting to maintain hover IGE as fuel burn-off occurs. The power may be gradually reduced to maintain the IGE hover condition.
(2) The hover test is to continue until at least 5 minutes after the occurrence of the highest temperature recorded.
(3) Section 29.1049 also requires a hover IGE at the maximum continuous power available at the altitude resulting in zero rate of climb.
Par 519 819
Often, compliance is illustrated by extrapolating component cooling margins from sea level test results and from selected altitude test site results to the altitude resulting in zero rate of climb.
(ii) Considerable engineering judgment must be exercised in utilizing the extrapolation method described. In general, if test data is extrapolated more than 2,000 feet to the hover ceiling from the highest altitude site selected and the resulting component margin is less than 5° F, additional verification at altitude may be required.
521.-530. RESERVED.
820 (thru 830) Par 520
SECTION 30. INDUCTION SYSTEM
531. § 29.1091 (Amendment 29-17) AIR INDUCTION.
a. Explanation.
(1) The air induction system for each engine and auxiliary power unit must supply the air required under the operating conditions for which certification is requested. For reciprocating engine installations, the system must provide air that is suitable for proper fuel metering and mixture distribution. This should be shown with the induction system valves in any position.
(2) The intake system shall be designed such that a backfire flame will not constitute a fire hazard within the engine accessory compartment or within other areas of the powerplant compartment.
(3) Each reciprocating engine must have an alternate air source which must be located to prevent entrance of rain, ice, or other foreign matter.
(4) For rotorcraft powered by turbine engines and rotorcraft incorporating auxiliary power units, there must be means to prevent leakage of hazardous amounts of flammable fluids from entering the engine or auxiliary power unit intake system.
(5) Also, the air ducts must be located or protected to minimize the ingestion of foreign matter during takeoff, landing, and taxiing.
b. Procedures.
(1) For turbine-engine installation, the induction system should supply air of suitable quality to meet the installation requirements of the engine manufacturer. The installation requirements should be met throughout the operating envelope of the rotorcraft.
(2) The inlet design should account for the prevention of hazardous fluids entering the engine. Some designs will have inlet ducts which are free from any fluid lines; however, other designs may route the engine inlet air through a compartment which has flammable fluid lines. When this condition exists, test demonstrations of critical leakage during operation have been used to substantiate the installation. The fluid leakage may not have an adverse effect on engine operation.
(3) The air induction system design should also account for and minimize the possibility of foreign matter ingestion during takeoff, landing, and taxiing.
Par 531 831
(4) For reciprocating engine installations, the induction system should supply air of suitable quality and quantity to the combustion system of the engine. The condition of this air at the entering face of the carburetor is extremely important. For proper operation, it is essential that the airflow be smooth and uniform, clean, and unrestricted throughout the very wide range of horsepower expected from the engine.
532. § 29.1093 (Amendment 29-22) INDUCTION SYSTEM ICING PROTECTION.
a. Reciprocating Engines. No advisory material is presented here for reciprocating engines since it is unlikely that these types will be used in transport rotorcraft.
b. Turbine Engines - Ice Protection.
(1) Explanation.
(i) This rule requires turbine engines and turbine-engine inlets to perform satisfactorily in atmospheric icing conditions defined in Appendix C of Part 25. On an equivalent safety basis, the limited icing envelopes described in Paragraph 386 herein may be used to show compliance with the intent of the regulation if the rotorcraft is limited to not greater than a 10,000-foot pressure altitude for all operations. If operations are permitted above 10,000 feet, the Appendix C, Part 25, envelope must be used from 10,000 feet to the service ceiling or 22,000 feet. These possible equivalent safety approaches are not discussed herein. Compliance with the induction system icing protection rule is required regardless of flight manual limitations or restrictions against flight into atmospheric icing conditions.
In showing compliance with § 29.1093(b)(1)(i), the FAA/AUTHORITY has accepted the concept of limited exposure associated with escape from inadvertent ice encounters.
Note: Although this approach (whether deliberate or by oversight in selection of test conditions) has been used in past certifications, the Office of Chief Counsel (AGC) of the FAA/AUTHORITY will be asked to provide the interpretation of § 29.1093(b)(1)(i) regarding acceptability of the limited exposure concept. Guidance material herein which addresses this limited exposure concept will be withdrawn if the AGC decision finds the limited exposure concept unacceptable.
(A) It is presumed that there will be a flight manual limitation against flight into known icing, and that the engine induction system will be reevaluated if total aircraft ice protection certification is requested. Under this concept, the rotorcraft is assumed to fly directly through the icing environment; i.e., direct sequential penetration and straight line exit from both the continuous maximum and intermittent maximum icing clouds.
Thus, the duration of exposure to the icing environment could be calculated by knowing the aircraft flight speed and cloud horizontal extent. A range of engine power and 832 Par 531
rotorcraft airspeeds should be evaluated to encompass the operating envelope of the rotorcraft.
(B) When this limited exposure concept is used, the aircraft type certificate data sheet snould clearly specity that the engine induction system must be reevaluated if certification to the general ice protection regulation, § 29.877 or § 29.1419, is requested. This direct penetration and exit approach is inappropriate for aircraft for which full icing clearance is requested (reference § 29.1419).
(lI) Engine induction system continuous icing protection would be necessary for aircraft for which full-icing clearance is requested (reference § 29.1419(d)). The approach is much preferred for all programs in order to reduce the scope of any eventual total aircraft icing program effort and to increase the safety level in conducting the rotorcraft natural icing tests. Since at least one rotorcraft has been FAA/AUTHORITY certificated to operate in known icing conditions and others have active development programs to this end, applicants should anticipate eventual full-icing clearance and consider that the engine induction system may be required to operate routinely in a continuous icing environment.
(iv) It is noted in Paragraph 386 that some natural icing tests are required to show compliance with the overall rotorcraft ice protection requirements. It is not required that the engine induction system be evaluated as a part of that natural icing test if adequate verification has been shown by tunnel testing, analysis, or other means to assure satisfactory operation in an extended continuous icing environment. If, however, subsequent rotorcraft natural icing testing shows unanticipated detrimental engine inlet effects, the inlet ice protection system should be reexamined.
(v) The regulation specifies the examination of flight idling conditions. This requirement is normally associated with a low-power letdown at the minimum practical forward airspeed. Alternatively, evaluation of the minimum power and minimum airspeed combination specified in the RFM for operation in visible moisture when below 40° F will accomplish the intent of the idling requirement.
(vi) An acceptable approach to a finding of compliance would be a combination of analysis of the performance of the ice protection system which covers the range of the applicable icing flight envelope (maximum altitude, minimum temperature, etc., of the basic rotorcraft) supported and validated by tests. Ideally, these tests would be conducted in natural atmospheric ice with special instrumentation for droplet size and liquid water content. In practice, however, natural icing testing may pose unacceptably severe problems since rotorcraft may not have the range and speed to reasonably find icing clouds and may not be equipped with the airframe and rotor ice protection needed for safety during the testing.
(vii) Problems with analysis emerge if engine inlets incorporate screens, turning vanes, sideward or upward openings, and edge or lip configurations which Par 532 833
АС 29-28 7/30/97 deviate from the airfoil shapes assumed in most of the analytical procedures described in current technical literature. The applicant should recognize that if meaningful analytical methods are not available, extensive testing with significant conservatism or possibly design changes may be required. Inlet screens in particular, if not adequately heated, fall in this category and can only be accepted if shown by very conservative ice testing to not significantly impede airflow to the engine.
(2) Procedures.
Review Paragraph 386 of this AC, ADS-4, Report No. FAA-RD-77-767, and Advisory Circular 20-73. (The comparative concept described under Item 34 of AC 20-73 is obsolete and should not be considered.) These data provide extensive description and methodology for evaluation of ice protection systems, however, as noted above, these data generally apply to near straight line droplet trajectory with impingement onto conventional airtoil shaped inlets. As such, the applicability of these data to rotorcraft engine inlet ducts is limited and may require extensive adjustment to accommodate the different inflow trajectories and shapes of rotorcraft.
(її) An analysis, appropriate to the configuration; i.e., heated or unheated impingement surfaces, should be prepared. To be acceptable, this analysis should show the inlet to be adequately protected by heat, or if unheated, to show that the inlet with ice accretions as predicted, will provide adequate airflow to the engine throughout the flight envelope of the rotorcraft.
(A) For heated surfaces, ADS-4 and Report No. FAA-RD-77-76 provide detailed suggestions on heat transfer analysis particularly applicable to bleed air heated inlet lips formed in airfoil shapes. These data are limited in applicability and may not be useful for analyzing engine inlet water droplet trajectories to be expected at low airspeed and high engine airflow. Actual icing tests may be needed to derive the impingement patterns for these conditions.
(1) Acceptability criteria for heated inlet ducts usually require sufficient heat to evaporate the water to be expected in a "continuous maximum" icing cloud and to anti-ice the duct during flight in "intermittent maximum" icing clouds, providing the run-back and refreeze to be expected does not cause additional airflow disruption or damage to the engine. Full-scale inlet icing tests with the engine installed and operating should be conducted to verify the analysis. Engine power changes which may be expected in service should be included in the testing. Wind tunnels equipped for icing tests probably are the most useful means of conducting these tests if natural icing tests are impractical. The rotor downwash effect should be considered to the extent possible by adjusting the inflow angle in the tunnel.
(2) The power loss (bleed air, generator load, etc.) attributable to the heating requirements will affect the performance of the rotorcraft. Normally, this may be 834 Par 532
accounted for by specifying a gross weight incremental deduction from the flight manual performance data for flight into visible moisture below 40° F.
(3) Special evaluation of the possibility of ice ingestion damage to the engine should be made for heated systems which considers the ice ingestion to be expected when the anti-ice system is actuated after a delay of 1 minute for the pilot to recognize that the rotorcraft has encountered ice. This time delay may be reduced if the crew is provided adequate distinctive cues to alert them that the rotorcraft has encountered icing conditions.
(B) For unheated inlets, an acceptable method for showing compliance would include an extensive, detailed analysis (which shows that ice accretions on and in the inlet do not obstruct adequate airflow to the engine) and tests as necessary to validate the analysis. The analysis of ice accretion becomes even more questionable since the unheated inlet involves ice buildups which themselves progressively change shape during icing exposure.
(1) Flight testing with an instrumented rotorcraft in natural ice to verify the analysis is desirable; however, wind tunnel tests as discussed above may be used. Since unheated inlets typically continue to accrete ice as a function of exposure, both the analysis and the test should realistically consider the actual exposure to be expected in service. This should not be less than penetration of the continuous maximum icing cloud followed immediately by exposure to the intermittent maximum cloud for rotorcraft not certified for icing. Engine power changes which may be expected in service should be included in the testing, and a warm-up period at the conclusion of the icing exposure should be shown for some selected test points to evaluate potential ice breakaway and ingestion.
(2) For the nonicing certified rotorcraft using the limited icing exposure concept for inlet certification, some conservatism should be applied to account for the fact that inlet icing may occur without airtrame icing, and that the escape procedure from this unapproved operating condition is not defined. A demonstration of 30-minute hold capability in the continuous maximum cloud would be acceptable. Alternatively, if positive cues (perhaps a carefully located ice detector) of potential inlet icing are provided to the crew, the time increment could be reduced to recognition plus 15 minutes (15-minute escape time after recognition is consistent with the single ice protection system failure recognition and escape guidance for aircraft ice protection systems in Paragraph 386). It should not be assumed that airframe icing will always be available as a cue to potential inlet icing. The main rotor, for example, may not show icing indications above 25° F, whereas some inlets may ice critically near 32° F ambient. A reduction of the acceptable 30-minute exposure should not be based on observation of ice accretions on protruding components which are likely to be changed. For example, a limited exposure inlet icing program which reduces the inlet icing exposure time based on crew recognition of icing on the windshield wipers may be invalidated at a later date if a new windscreen deletes the wipers.
Par 532 835
AC 29-2B 7130/97 (Hi) Inlet capability during IGE hover in icing conditions has not generally been considered for rotorcraft not certified for icing. Recently, however, the FAA/AUTHORITY is aware that some inlets may ice at zero airspeed near 32° F with no indications of airframe icing in the field of view of the crew. This special concern of operating within RFM limitations, and yet placing the induction system in jeopardy, may be addressed in several ways. If the induction system ice protection scheme is not dependent on airspeed for proper function, the issue may be addressed by tunnel testing with inlet airflows approximating hover with no particular attention to tunnel windspeed. For protection schemes which may be sensitive to airspeed (external screens have shown this tendency), actual hover demonstration at or near zero speed tunnel conditions may be appropriate. Icing detectors located to indicate induction system icing in hover may be an option to a hover icing protection demonstration.
Recently, on an external screened configuration, the FAA/AUTHORITY has accepted a satisfactory IGE hover demonstration of 30 minutes at the critical ambient temperature
(i.e., ambient consistent with no airframe icing but potential inlet icing),
0.6 grams/meter LWC and 40 micron droplet size as an adequate response to this concern.
(iv) For aircraft requesting full icing approval, or for those electing to show continuous induction system icing protection, the forward flight icing exposure would not be less than that time required to stabilize any ice accretions observed during repeated cycles of the continuous maximum followed by intermittent maximum cloud exposure.
Typically, any ice accretions resulting from these repeated cycles would be expected to stabilize in less than 30 minutes. The 30-minute hold capability in the continuous maximum icing environment could thus be assured without special testing by careful selection of the test points for this repeated cycle.
(V) A rotorcraft requesting full icing approval should also have hover capability in the icing environment. Intermittent maximum icing conditions are not likely to exist near ground level and a satisfactory demonstration could involve the ability to hover indefinitely in the continuous maximum icing environment. Alternatively, carefully worded RFM limitations to restrict hover time may be acceptable if the system is not capable of indefinite exposure. Hover capability verification may not involve zero airspeed demonstration if the inlet protection system is insensitive to rotorcraft airspeed.
(vi) The engine(s) must be installed or protected to avoid engine damage from ice ingestion due to ice accretion in the inlet or on other parts of the rotorcraft, including the rotors, which may break away to enter the inlet. If screens or bypass arrangements are provided for these purposes, they should be included in the icing tests and shown by test or rational analysis to effectively protect the engine.
(vii) For unheated inlets, significant ice accumulations to be expected on the inlet may adversely affect the engine stall margin, acceleration characteristics, duct loss, etc. Dry air flight tests to evaluate these aspects can be accomplished by affixing 836 Par 532