b. Procedures. Unless the filter is located at the oil tank outlet, § 29.1019(a)(3) requires that the oil strainer or filter have the means to indicate when the contaminant level of the filtration system, as specified in § 29.1019(a) (2), has been reached. If the indicator is installed, it should signal a contaminant level that will allow completion of the flight before the filter reaches a bypass condition. The indicator may be a pop-out button or other maintenance cue that is checked on each preflight inspection.
503. § 29.1021 OIL SYSTEM DRAINS.
a. Explanation. This regulation requires provisions be provided for safe drainage of the entire oil systems and defines certain requirements for assuring that no inadvertent oil flow occurs from the system provided.
b. Procedures. The design of the oil system must provide a means for safe drainage of the entire oil system. This may require one or more drains dependent upon the design of the system. If a valve is used for this function, it must provide a means for a positive lock in the closed position. The method by which the lock is accomplished may be manual or automatic.
504. § 29.1023 OIL RADIATORS.
a. Explanation. This regulation defines the installation requirements to be considered for oil system radiators.
b. Procedures.
(1) The primary concern with respect to oil radiators is that they are sized to provide the required heat rejection and to provide adequate fluid flow within the prescribed pressure limits.
(2) The structural design of the radiator must consider the system oil pressure requirements and the service involvement of the intended application. The selection of the location of the radiator can have a significant bearing on its ability to withstand the vibration and inertia loads.
(3) If the system design incorporates an air duct to direct the airflow, the effects of a fire as defined in this regulation must be considered.
505. § 29.1025 OIL VALVES.
a. Explanation. This regulation identifies the requirements which oil system valves must meet. In addition to the items specified in this rule, this regulation specifies compliance with the requirements of § 29.1189.
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b. Procedures. The closing of the oil shutoffs may not preclude a safe autorotation. Compliance with this requirement is best accomplished in the design phase. This can be accommodated by proper orientation of the valve and/or system plumbing routing. Another means is to design adequate entrapment of lubricants to provide for the autorotation state. The design of the oil shutoff valve must consider the stop or index provisions of this rule. The installation must be such that the loads specified in the rule are addressed.
506. § 29.1027 (Amendment 29-26) TRANSMISSION AND GEARBOXES: GENERAL.
a. Explanation. Amendment 29-26 adds a new § 29.1027. This new section provides the regulations for rotorcraft transmission and gearbox lubrication systems. It incorporates lubrication system requirements that were removed from § 29.1011 and adds additional lubrication system requirements that were derived from existing engine-oil system requirements. These additional requirements have been adjusted or modified to reflect the needs of transmissions and gearboxes. Transmission and gearbox lubrication system regulations are similar to those for engines; therefore, reference is made to the engine lubrication sections as applicable.
b. Procedures.
(1) The pressure lubrication systems for rotorcraft transmissions and gearboxes should comply with the same requirements as the engine lubrication systems stipulated in § 29.1013 (except §§ 29.1013(b)(1), 29.1015, 29.1017, 29.1021, and 29.1337(d)). These sections provide the requirements for oil tanks, tank tests, oil lines and fittings, and oil system drains.
(2) Each pressure lubrication system for rotorcraft transmissions and gearboxes should have an oil strainer or filter. The strainer or filter should:
(i> Remove any contaminants from the lubricant that may damage the transmission, gearbox, or other drive system component and any contaminants that may impede the lubricant flow to a hazardous degree.
(¡i) Be equipped with a means to indicate that the bypass system (required by § 29.1027(b)) is at the point of opening, due to the collection of contaminants on the strainer or filter; and,
(iii) Be equipped with a bypass system that will permit lubricant to continue to flow at the normal rate if the strainer or filter is completely blocked. In addition, the bypass system should be designed so that contaminants, that have collected on the filter, will not enter the bypass flow path when the system is in the bypass mode.
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(3) Section 29.1027(b)(2) requires a screen at the outlet of each lubricant tank or sump that supplies lubrication to rotor drive systems and rotor drive system components. The screen should remove any object that might obstruct the flow of lubricant to the filter required by § 29.1027(b)(1). The requirements of § 29.1027 (b)(1) do not apply to the tank outlet screen.
(4) Splash-type lubrication systems for rotor drive system gearboxes should comply with §§ 29.1021 and 29.1337(d).
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507.-515. RESERVED.
Intentionally Left Blank 798 (thru 806) Par 507
SECTION 29. COOLING
516. § 29.1041 GENERAL.
a. Background.
(1) Few substantive changes have been made to the cooling provision requirements, §§ 29.1041 through 29.1049, since the rules were defined in the Civil Air Regulations, Part 7, effective August 1, 1956. Testing procedures utilized have not precisely followed those rigorously set forth in §§ 29.1045 through 29.1049 as industry and the FAA/AUTHORITY have recognized the need to vary procedures slightly to accomplish the practical test objectives.
(2) In the paragraphs which follow, the cooling regulations will be explained, and in some instances where the regulations provide specific procedures, "alternative procedures" which have been found acceptable in achieving the rule objectives will be presented. The intent of providing those alternative procedures is not to promulgate new regulations, but rather to provide recognized, accepted procedures for compliance with the objective of the current standards.
b. Explanation.
(1) The rotorcraft design should provide for cooling to maintain the temperatures of all powerplant, auxiliary power unit, and power transmission components and fluids within the limitations established for these items.
(2) Cooling provisions should be adequate for shutdown and for water, ground, and flight operating conditions.
(3) The adequacy of the cooling provisions should be demonstrated by flight testing.
c. Procedures.
(1) Detailed procedures for the demonstration of climb, takeoff and climb, and hover cooling are given in §§ 29.1045 through 29.1049. Other test conditions and procedures necessary to demonstrate adequate cooling for water, ground, flight, and shutdown conditions must be negotiated between the applicant and the FAA/AUTHORITY certification engineer. A cooling test proposal which defines the agreed test points and procedures should be prepared well in advance of the official certification testing.
(2) The test conditions selected, in addition to those in §§ 29.1045 through
29.1049, would typically include cruise at various airspeeds and altitudes, shutdown Par 516 807
after prolonged hover, and sling load cooling if applicable. One test condition which should be examined, particularly with regard to transmission cooling, is the point of highest multiengine mechanical power at the maximum ambient temperature. This is identified as test point "A" in Figure 516-1. The selection of test points should be tempered with engineering judgment and based on results from similar aircraft, if such data are available.
(3) In showing compliance with the cooling requirements, the applicant should not be required to exceed rotorcraft established limits (gross weight, drive system torque, measured gas temperature, etc.), aircraft power required, or power available.
The applicant may elect, however, to exceed these limits in order to minimize test points by conservative testing, or to anticipate future growth (increased gross weight, etc.).
(4) The need for a comprehensive cooling test plan prior to certification testing cannot be overemphasized. Highly derated engine installations, the relationship of power required to power available, the use of bleed air devices which would increase the measured gas temperature while aircraft power required remains the same, auxiliary cooling provisions, and the increase in engine temperatures with engine deterioration are factors which could affect the selection of cooling demonstration test points. The following paragraphs will provide some general guidance, but the cooling test plan is the key to a successful program.
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MULTIENGINE POWER AVAILABLE 20K 15K XMSN LIMIT PRESSURE ALTITUDE OAT 10° C 10K 15° C 20° C 25° C 30° C 35° C ol POWER POINT "A" FIGURE 516-1 ADDITIONAL COOLING TEST POINT Par 516 809
517. § 29.1043 (Amendment 29-15) COOLING TESTS.
a. Explanation.
(1) Section 29.1043(a) requires that certain ambient temperature correction factors be applied unless testing is accomplished at the maximum ambient atmospheric temperature prescribed.
(2) No corrected temperatures may exceed established limits.
(3) The statement of § 29.1043(a)(4) which requires that test procedures be in accordance with §§ 29.1045 through 29.1049 does not limit testing to the conditions prescribed in those sections. Section 29.1041(a) and (b) provide the basis for examination of other operating and shutdown conditions.
(4) The maximum ambient atmospheric temperature must be at least 100° Fat sea level, lapsed to altitude at a rate of 3.6° F per 1,000 feet pressure altitude. The applicant may select a lower maximum ambient atmospheric temperature for winterization installations.
(5) Unless a more rational correction applies, the temperature data (except for cylinder barrels) are to be corrected by adding the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum component or fluid temperature recorded during the cooling test.
(6) Cylinder barrel temperature data are corrected in a similar manner to other components except 0.7 times the difference between the maximum ambient atmospheric temperature and the ambient temperature at the first occurrence of the maximum cylinder barrel temperature is applied.
b. Procedures.
(1) Seldom is testing actually accomplished at the maximum required ambient temperature of at least 100° F at sea level lapsed 3.6° F per 1,000 feet pressure altitude. Component and fluid temperatures must therefore be corrected to derive the item temperature that would have been reached if the test day had matched exactly the maximum ambient temperature day. The applicant may select a higher maximum ambient temperature for cooling certification than the 100° F sea level hot day prescribed. Provisions are also made for selecting a maximum ambient temperature less than the 100° F sea level hot day for winterization installations not intended to function at the hot day conditions.
(2) When cooling test ambient conditions are cooler than the selected or prescribed hot day conditions, the applicant may take advantage of cooling air or fluid 810 Par 517
flows that would exist at hot day conditions. For example, thermostatically controlled oil cooler flow could be set for hot day conditions.
(3) The component and fluid temperature correction factor to be applied when test ambients do not correspond to the hot day conditions is commonly called the "degree-for-degree correction." It may be possible to justify, and the regulation allows the application of a more rational, less conservative correction factor. A correction factor other than degree-for-degree should be based on engineering test data.
(4) No corrected temperatures may exceed established limits. In order to maintain temperatures within established limits, the applicant may be willing to accept lesser performance than the full capability of a device. For example, a starter/generator capable of cooling under test cell conditions to 200 amperes continuous load may be limited to a lesser value, perhaps to 150 amperes, when installed in the aircraft due to cooling considerations. This continuous load for cooling must be equal to or greater than the allowable continuous load designated on aircraft instruments.
c. Thermal Limit Correction.
(1) An important correction factor which is not discussed in the regulations, but is frequently necessary to show the cooling adequacy required by § 29.1041, is the thermal limit correction factor. This factor is sometimes required if, at test day conditions, the engine measured gas temperature does not correspond to that which would have occurred on a minimum specification engine at hot day conditions.
(2) The correction factor would not apply to those components not affected by changes in measured gas temperature (MGT) at a constant power. Typical items expected to be affected by changes in the MGT at constant power would be engine oil temperature, thermocouple harnesses, or other fluid, component, or ambient temperatures in the vicinity of the engine hot-section or exhaust gases. Other items remote from the hot-section, perhaps the starter-generator or fuel control, would not be expected to be influenced by MGT variations; however, the items affected and the magnitude of the factor to be applied should be established by testing.
(3) There are several acceptable methods for establishing the appropriate thermal limit correction factor during development testing. The general idea is to establish a stabilized flight condition, typically ground-run or IGE hover, and to vary the measured gas temperature at approximately fixed power and OAT conditions. This may be accomplished by utilizing engine anti-ice bleed air, customer bleed air, or by ingesting warmer than ambient air (either an external source or the engine bleed air) into the engine inlet. Care should be used in ingesting warmer than ambient air to assure that the warm air is diffused in order to avoid possible engine surge.
If it is not possible to attain a suitable variation in MGT by these methods, an acceptable, but more conservative thermal limit correction may be Par 517 811
obtained by allowing both shaft horsepower and MGT to vary at a stabilized flight condition and OAT.
The component temperature is plotted as a function of MGT, and the thermal limit correction from any test day MGT for any flight condition, to the MGT that would have existed with minimum specification engines on a hot day, is then applied to derive the final measured component temperature.
(4) In certain rare instances, it may not be required that the correction factor be applied to the full thermal limit capability of the engine. Consider the following example for the hot day hover IGE cooling test point at sea level.
Corresponding Power (SHP) MGT (°C) Drive System Limit 900 Twin-Engine Hot Day Power Available 1,050 750 Hot Day Power Required at Maximum G.W. 850 650 Engine Maximum Allowable MGT (Instrument Marking) 765 Test Day (90° F OAT) Parameters 850 600
(i) Notice that the installed hot day power available MGT from the engine performance program, is 15° C cooler than the limit MGT (750° vs. 765° C), thus the engine has 15° C "field margin" which would allow the engine temperature to gradually increase 15° C to maintain a given power as engine life is utilized. Secondly, the measured gas temperature corresponding to hot day power required at maximum gross weight, is less than that corresponding to either the drive system limit or twin-engine hot day power available. Thus, the thermal limit correction could be applied from the test day MGT, 600° C, to the power required MGT plus the field margin, 650° C plus 15° C, rather than applying the correction factor to the full thermal capability of the engine, 765° C..
(ii) Care should be used in applying this relieving method, because as the hover altitude changes, the maximum gross weight and power required (and the associated MGT) will vary. The data must be corrected to at least the maximum MGT for a minimum specification engine that can occur in service at the flight condition under investigation.
517A. § 29.1043 (Amendment 29-26) COOLING TESTS.
a. Explanation. Amendment 29-26 adds a new paragraph. to § 29.1043(a)(5), to define "stabilization" as it pertains to powerplant systems cooling tests.
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b. Procedures. All of the policy material pertaining to this section remains in effect with additional information that "stabilized temperatures" are achieved when the rate of change is less than 2° F per minute.
518. § 29.1045 CLIMB COOLING TEST PROCEDURES.
a. Objective. The objective of the regulation is to verify, for Category A and for Category B rotorcraft described, that cooling provisions are adequate for a one-engine-inoperative (OEI) climb or descent initiated from a multiengine cruise at the critical altitude with stabilized component temperatures. The specific flight conditions and powers are described in the regulation.
b. Explanation.
(1) This regulation specifies climb or descent cooling with OEl for Category A rotorcraft and for Category B rotorcraft with Category A powerplant isolation and fireproof or isolated structure, controls, etc., which are essential for controlled flight and landing. For the Category B machine described, the testing should be accomplished at the steady rate of climb or descent established under § 29.67(b), i.e., at the best OEI rate of climb (or descent) and the remaining engine at maximum continuous power or 30-minute power, whichever is applicable.
(2) The engine whose shutdown has the most adverse effect on the cooling conditions for the remaining engine(s) and powerplant components should be inoperative.
(3) The regulation provides that the climb cooling test may be conducted in conjunction with the takeoff cooling test of § 29.1047. This possible combining of tests applies only to § 29.1047(a), since § 29.1047(b) is a multiengine climb and not related to the OEI climb procedures of § 29.1045.
c. Procedures.
(1) The OEI climb cooling test point begins from a multiengine cruise, with stabilized fluid and component temperatures, 1,000 feet below either the all-engine-critical altitude or the maximum altitude at which the rate of climb is 150 FPM, whichever is the lowest altitude. If the minimum altitude derived is less than sea level, the climb should begin from a twin engine cruise with stabilized fluid and component temperatures at the minimum practical altitude.
(i) The all-engine-critical altitude is the maximum altitude at which, for the ambient conditions prescribed, it is possible to maintain the multiengine specified power. For example, if for multiengine operations, the transmission maximum continuous torque can be maintained on the hot day to a maximum altitude of 10,000 feet above which power would have to be reduced because of gas temperature Par 517A 813
or other limitations, then 10,000 feet is the all-engine-critical altitude. Point "A" in Figure 518-1 illustrates the all-engine-critical altitude.
(ii) The 150 FPM climb criteria should be based on multiengine operation at maximum continuous power available at hot day conditions at maximum gross weight.
(ill) Fluid and component temperatures are considered stabilized when the rate of change is less than 2° F per minute.
(2) The OEI climb power to be utilized is 30-minute OEl hot day power available (if approval of 30-minute power on the aircraft is requested), followed by maximum continuous hot day power available. If 30-minute OEI power approval is not requested, the power to be utilized would be maximum continuous hot day power available.
Rotorcraft for which approval of a continuous OEl power rating is requested would use the power available on a hot day at the maximum continuous OEI rating following the 30-minute OEI climb phase (or for the entire climb if approval of 30-minute OEI power is not requested).
(ii) If the maximum continuous OEl approval is not requested, then the highest hot day power available approved for continuous usage from the remaining engines) under OE conditions would be used following the 30-minute OEI climb phase (or for the entire climb if approval of 30-minute OE power is not requested).
(3) In order to achieve representative test results, the rotorcraft climb rate and airspeed should approximate those which would occur on a hot day. This is accomplished by adjusting rotorcraft gross weight as required to produce the desired climb rate based on published or predicted climb performance data. The possible adverse effects of climb fuselage attitude on cooling air duct entrances should be considered in the selection of center-of-gravity of the test aircraft.
(4) The OEI climb should be continued for at least 5 minutes after the occurrence of the highest temperature recorded or until the maximum certification altitude is reached. Generally, temperatures would be expected to peak a short time after the climb begins since component and fluid temperatures are stabilized prior to entry to the climb phase.
(5) For Category B rotorcraft, defined in § 29.1045(a)(2) without a positive OEl rate of climb, the descent should begin from a hot day maximum continuous power multiengine cruise, with stabilized fluid and component temperatures, at the all-engine-critical altitude.
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