* 501. § 29.1017 (through Amendment 29-19) OIL LINES AND FITTINGS.
a - Explanation. This regulation outlines the certification requirements for oil lines and fittings.
b. Procedures. The oil system lines and fittings are required to meet the requirements of § 29.993; therefore, the routing and clamping described in Par 709, Chapter 14, Section 2, of AC 43.13-1A may be utilized as guidance for the system design. An evaluation carried out through the development and certification test period will usually surface any problems of interference and/or vibration.
(1) When flexible hoses are used in the lubrication system they must be substantiated. Hoses listed in ISO C53a may be used which would preclude certain substantiation requirements.
(2) Location of the breather lines and discharge should be carefully evaluated to determine that the requirements of this paragraph are followed.
(3) The routing of fluid lines should be such that drooping lines and fluid traps which are undrainable are avoided.
502. § 29.1019 (through Amendment 29-19) OIL STRAINER OR FILTER.
a • Explanation. This regulation defines the requirements for the engine oil system strainer or filter. If a strainer or filter which meets the requirements of this paragraph is incorporated as part of the type certificated engine, an additional airframe filter is not required.
b. Procedures. This paragraph requires an oil strainer or filter through which all of the oil flows for each turbine engine installation. The strainer or filter should be sized to allow oil flow at the flow rates and within the pressure limits as specified in the engine requirements. The effect of oil at the minimum temperature for which certification is sought should be accounted for.
(1) For each oil strainer or filter required by § 29.1019(a) which has a bypaas, the bypass should be sized to allow oil flow at the normal rate through the oil system with the filtration means completely blocked.
(2) For each oil strainer or filter installed per this rule, the capacity must be such that the oil flow and pressure are within the operating limits established for the engine. The mesh requirements are determined by the engine specification for the filtration of particle size and density.
Chap 2 Par 501 795
AC 29-2, CHG 1 11/19/84
(3) Section 29.1019(a) (3) requires an indicator that will show when the contaminant level of the filtration system, as specified in § 29.1019(a) (2), has been reached. The indicator should signal a contaminant level which has not caused the filter to go into a bypass condition. Consideration should also be given so that the contaminant level at which the indicator is activated is such that the filter would not bypass during a flight time based on full fuel at a cruise condition with the lubricant contaminated to the degree used to show compliance with $ 29.1019(a) (2).
An evaluation of the construction and location of the bypass associated with the strainer or filter should be accomplished. The appropriate installation of the filter based on this evaluation would preclude the release of the collected contaminants in the bypass oil flow.
(5) If an oil strainer or filter installed in compliance with this regulation does not have a bypass, there must be a means to connect it to the warning system required in § 29.1305(a) (18). This warning should indicate to the pilot the contamination before it reaches the capacity established in § 29.1019(a)(2). Section 29.1019(b) covers the blocked oil filter requirements associated with reciprocating engine installations. The lubrication system should be such that the normal oil flow will occur with the filter completely blocked.
503. § 29.1021 (through Amendment 29-19) 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 ocours 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 (through Amendment 29-19) 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.
Chap 2 796 Par 502
AC 29-2, CHG 1 11/19/84 * 499. § 29.1013 (through Amendment 29-19) OIL TANKS.
a. Explanation This regulation identifies the requirements that each oil tank must meet. It also specifies that the oil tank installation must meet the installation requirements of § 29.967.
b. Procedures.
(1) The oil tanks usually are constructed of aluminum, aluminum alloy, or stainless steel and are of such a design to permit installation in the aircraft as close to the engine as the design allows. The choice of materials will generally be determined by the selected location of the tank. The tank envelope or outline will generally be determined by the location within the structure of the helicopter.
(2) The design of the tank is required to meet the expansion space requirements as specified in the regulation for the particular installation.
This is generally accomplished by locating the filler cap in such a manner that the expansion space cannot be inadvertently filled with the rotorcraft in normal ground attitude.
(3) The tank is required to be properly vented and the vent requirements are identified in the regulation.
(4) Unless alternate means are provided, it is good design practice to locate the oil tank with respect to the engine so that when the helicopter is in its normal ground attitude, a positive head to the oil pump inlet is provided.
(5) Sections of the regulation address specific requirements when Category A certification is requested.
(6) The designer should be aware of the requirements associated with the location of the oil tank outlet and the marking requirements specified in $ 29,1557(0) (2).
(7) Flexible oil tank liners may be used; however, they must be approved or shown to be suitable for the particular installation.
(8) An "external oil system" which is defined as being those components, lines, etc., of an oil system which are outside the engine and not supplied as part of a certificated engine. The components of such a system which are within the fire zone and required to be fire resistant. Those outside the fire zone need not be fire resistant.
500. § 29.1015 OIL TANK TESTS. (RESERVED) Chap 2 790 (thru 794) Par 499
SECTION 28. OIL SYSTEM * 498. § 29.1011 (through Amendment 29-19) GENERAL.
a • Explanation.
(1) This regulation defines the oil system requirements for engines and the rotor drive systems which require continuous lubrication.
(2) Each engine must have an independent oil system. This lubrication system performs two functions. It provides an adequate oil supply and it incorporates a means of cooling the hot oil discharged from the engine.
Acceptable oil quantities must be determined and the system designed to accommodate this quantity.
(3) The adequacy of the oil systems to maintain the oil temperature at or below the specified limits must be shown under the applicable requirements of $§ 29.1041 through 29.1049.
b. Procedures.
(1) To provide engine isolation, meeting the regulations requires completely independent lubrication systems in multiengine rotorcraft. This includes separate oil tanks, oil coolers, oil cooler blowers, and the associated plumbing. Alternate designs providing an equivalent level of independence may be used if substantiated. A single failure of any one system may not--
(1) Prevent continued safe operations of the remaining engine(s); or
(11) Require immediate action by any crewmember for continued safe operations.
(2) Any rotor drive system which requires continuous lubrication must be sufficiently independent to ensure operation with any engine inoperative. The oil cooling provisions of this system should be sufficient to provide adequate cooling at maximum continuous power of the drive system and critical conditions associated with any single engine failure. The system must also provide for a safe autorotation.
(3) The usable oil quantity for the oil systems may be determined in several ways and the required quantities will vary between piston engine installation and turbine engine installation. Several methods of determining the usable oil capacity are provided in the regulations. However, when oil-fuel ratios other than those prescribed are used, they must be substantiated by data on the oil comsumption of the engine.
Chap 2 Par 498 789
* (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.
MULTIENGINE POWER AVAILABLE 20K 15K HOT DAY XMSN LIMIT PRESSURE ALTITUDE 15°C 10°C 10K 5K 25°C 30°c POWER POINT "A" FIGURE 516-1 ADDITIONAL COOLING TEST POINT Chap 2 Par 516 815
(3) The component and fluid temperature oorrection factor to be applied when test ambients do not correspond to the hot day conditions is commonly called the "degree-for-degree correotion." It may be possible to juatify, 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.
o. 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 ohanges 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 diffuged, in order to avoid possible engine surge.
Chap 2 Par 517 817
AC 29-2, CHG 1 11/19/84 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 obtained by allowing both shaft horsepower and MCT to vary at a stabilized flight condition and OAT.
(i1) The component temperature is plotted as a function of MGI, 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 Twin Engine Hot Day Power Available 1,050 750 Hot Day Power Required at Maximum G.W. 650 Engine Maximum Allowable MGT (Instrument Marking) 765 Test Day (90° F OAT) Parameters 850 600
(1) Notice that the installed hot day power available MGT from the engine performance program, 1s 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.
(i1) 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.
Chap 2 818 Par 517
518. § 29.1045 (through Amendment 29-19) 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 OEI 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), 1.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 f.p.m., 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.
(1) 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 or other limitations, then 10,000 feet is the all-engine-critical altitude. Point "A" in figure 518-l illustrates the all-engine critical altitude.
(11) The 150 f.p.m. climb oriteria should be based on multiengine operation at maximum continuous power available at hot day conditions at maximum gross weight. * Chap 2 Par 518 819
AC 29-2, CHG 1 11/19/84 (1i1) 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 OLl 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 OFI power approval is not requested, the power to be utilized would be maximum continuous hot day power available.
(1) Rotorcraft for which approval of a continuous OEI power rating is requested would use the power available on a hot day at the maximum continuous DEI rating following the 30-minute OEI climb phase (or for the entire climb if approval of 30-minute OEI power is not requested).
(11) If the maximum continuous OEI approval is not requested, then the highest hot day power available approved for continuous usage from the remaining engine(s) under OEI conditions would be used following the 30-minute DEI climb phase (or for the entire climb 1f approval of 30-minute OEI power 1s
(3) In order to achieve representative test results, the helicopter climb rate and airspeed should approximate those which would occur on a hot day.
This is accomplished by adjusting helicopter 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 oenter-of-gravity of the test airoraft.
(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 OEI 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.
(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 OEI powers available to be utilized during the descent would be the same as those prescribed previously for OEI olimb cooling. OEI operation should continue until component and fluid temperatures stabilize.
Chap 2 820 Par 518
11/19/84 SECTION 31. EXHAUST SYSTEM
548. 29.1121 GENERAL. (RESERVED)
549. § 29.1123 EXHAUST PIPING. (RESERVED)
550. $ 29.1125 EXHAUST HEAT EXCHANGERS. (RESERVED)
551.-560. RESERVED.
SECTION 32. POWERPLANT CONTROLS AND ACCESSORIES
561. § 29.1141 POWERPLANT CONTROLS: GENERAL. (RESERVED)
562. § 29.1142 AUXILIARY POWER UNIT CONTROLS. (RESERVED
563. § 29.1143 ENGINE CONTROLS. (RESERVED)
564. § 29.1145 IGNITION SWITCHES. (RESERVED)
565. § 29.1147 MIXTURE CONTROLS. (RESERVED)
566. § 29.1151 ROTOR BRAKE CONTROLS.
567. § 29.1157 CARBURETOR AIR TEMPERATURE CONTROLS.
568. § 29.1159 SUPERCHARGER CONTROLS. (RESERVED)
569. § 29.1163 POWERPLANT ACCESSORIES. (RESERVED) 570- 0 29.1165 ENGINE IGNITION SYSTEMS. (RESERVED)
571.-583. RESERVED.
SECTION 33. POWERPLANT FIRE PROTECTION
584. { 29.1181 DESIGNATED FIRE ZONES: REGIONS INCLUDED. (RESERVED)
585. $ 29.1183 FLAMMABLE FLUID-CARRYING COMPONENTS. (RESERVED)
586. § 29.1185 FLAMMABLE FLUIDS. (RESERVED) 587- § 29.1187 DRAINAGE AND VENTILATION OF FIRE ZONES. (RESERVED)
588. § 29.1189 SHUTOFF MEANS. (RESERVED).
589. § 29.1191 FIREWALLS. (RESERVED)
590. § 29.1193 COWLING AND ENGINE COMPARTMENT COVERING. (RESERVED) Chap 2 Par 548 873
AC 29-2, CHG 1 11/19/84
591. § 29.1194 OTHER SURFACES. (RESERVED)
592. § 29.1195 FIRE EXTINGUISHING SYSTEMS. (RESERVED) 593• 29.1197 FIRE EXTINGUISHING AGENTS. (RESERVED)
594. § 29.1199 EXTINGUISHING AGENT CONTAINERS. (RESERVED)
595. § 29.1201 FIRE EXTINGUISHING SYSTEM MATERIALS. (RESERVED)
596. § 29.1203 FIRE DETECTOR SYSTEMS. (RESERVED)
597.-616. RESERVED.
SECTION 34. EQUIPMENT - GENERAL
617. § 29.1301 FUNCTION AND INSTALLATION. (RESERVED)
618. § 29.1303 FLIGHT AND NAVIGATION INSTRUMENTS. (RESERVED)
619. 29.1305 POWERPLANT INSTRUMENTS. (RESERVED)
620. § 29.1307 MISCELLANEOUS EQUIPMENT. (RESERVED) Chap 2 874 (thru 990) Par 591
(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).
MULTIENGINE POWER AVAILABLE 20K
HOT DAY 15K 10°C 10K CRITICAL ALTITUDE _ - 15°G 20°C POINT "A" SK 25°C 30°% 35°C POWER FIGURE 518-1 ALL-ENGINE-CRITICAL ALTITUDE Chap 2 821 Par 518