* b. Procedures.
(1) Provide the specified aisle minimum width where a longitudinal main aisle is provided in the type design.
(2) Historically, certain rotorcraft with short, wide cabins were initially designed without a longitudinal main aisle for military and cargo use, but were later fitted and approved for civil passenger configuration. These craft generally have 19 or less passenger seats and have either (1) outboard seats/seat backs, or (3) a combination of lateral and longitudinal rows with and * racing passenger seats, (2) a limited number of lateral rows with fold down without main aisles to facilitate entrance and egress.
345. § 29.831 VENTILATION. (RESERVED)
346. $ 29.833 HEATERS. (RESERVED)
347.-356. RESERVED.
SECTION 21. FIRE PROTECTION
357. § 29.851 FIRE EXTINGUISHERS. (RESERVED) Chap 2 Par 344 575 (thru 596)
AC 29-2 5/20/83 SECTION 27. FUEL SYSTEM COMPONENTS
483. § 29.991 (through Amendment 29-19) FUEL PUMPS.
a • Explanation.
(1) Section 29.991, paragraph (a), provides a definition of the main pump(s) and § 29.991, paragraph (b), requires an "emergency pump(s)." The main pump (s) that is certified as part of the engine does not fall under § 29.991 requirements. The main pump(s) discussed under § 29.991 should therefore be considered "main airoraft pump(s)."
(2) The main aircraft pump(s) consists of whatever pump(s) is required to meet engine or fuel system operation throughout the range of ambient temperature, fuel temperature, fuel pressure, altitude, and fuel types intended for the rotorcraft. If the main aircraft pump(s) is required to meet the above criteria, then an emergency pump(s) is required.
b. Procedures.
(1) Each pump classified as a main aircraft pump, which is also a positive displacement pump, must have provisions for a fuel bypass. An exception is made for fuel injection pumps used on certain reciprocating engines and for the positive displacement, high pressure, fuel pumps routinely used in turbine engines. The bypass may be accomplished via internal spring check valve and fuel passage, or by external plumbing and a check valve. High capacity positive displacement pumps with internal pressure relief and recirculation passages should be checked for overheating if they may be expected to operate continuously at or near 100 percent recirculation.
(2) Section 29.991, paragraph (b) specifies a requirement for "emergency" pumps to provide the necessary fuel after failure of any (one) main aircraft pump. (Injection pumps and high pressure pumps used on turbine engines are exempt.) As stated in this rule, the "emergency" pump must be operated continuously or started automatically to assure continued normal operation of the engine. For some multiengine rotorcraft, another main aircraft pump may possibly be used as the required "emergenoy" pump. In this case, the dual role of this pump requires it to have capacity to feed two engines at the critical pressure/flow condition. Availability of fuel flow from this backup pump must be automatic and this funotion should be verified in the preflight check procedure.
For Category A rotorcraft, a comprehensive fault analysis of the fuel system is mandatory to assure compliance with § 29.903, paragraph (b).
Chap 2 Par 483 771
AC 29-2, CHG 1 11/19/84
(3) Section 29.991, paragraphs (o)(1)(1) and (11) address the situation, usually associated with superoharged reciprocating engines, where fuel pressure must be modulated with respect to carburetor deck pressure. This is aocomplished with interconnecting air lines from the carburetor intake (after the supercharger) to the pressure relief connection on the fuel pump(s). A similar connection from the carburetor intake to the vented side of the fuel pressure gauge is needed to obtain correct fuel pressure reading. These systems may require orifices and/or surge chambers to operate correctly.
(4) Section 29.991, paragraphs (c) (2) and (3) requires seal drains which drain safely. A drain impingement test 1s normally required to verify safe drainage. Use of a colored dye to simulate fuel discharge at the drain line exit or a fluid sensitive coating (Bon Ami) on the aircraft skins will facilitate evaluation of the safety aspects of drain impingement. Pump seal drain requirements would not be applicable for tank immersed pumps.
484. § 29.993 (through Amendment 29-19) FUEL SYSTEM LINES AND FITTINGS.
Explanation. This rule outlines design requirements for fuel aystem lines.
b. Procedures.
(1) Compliance is usually obtained by employing routing and clamping as described in paragraph 709, Chapter 14, Section 2 of AC 43.13-1A and by monitoring the arrangement throughout the developmental and certification test period. Requirements for approved flexible lines may be resolved by utilizing lines listed as TSO C53a approved for installation in either normal or high temperature areas as appropriate.
(2) Verify adequate clearance exists between lines and elements of the rotorcraft control system at extremes of control travel, including control deflections and, for flexible lines (hoses), possible variations in routing.
(3) Flexible lines inside fuel or oil tanks require special evaluation to assure that the external surfaces of these lines are compatible with the fluids involved and that fluid sloshing will not cause line failure. Lines inside tanks should be routed to avoid impingement by fuel or oil filler nozzles.
(4) Good design praotice suggests that all flammable fluid lines should be routed to minimize the possibility of rupture in the event of a crash or from engine rotor diso failure.
485. § 29.995 (through Amendment 29-19) FUEL VALVES.
a. Explanation. This regulation requires that fuel valves be supported so that no loads resulting from their operation or from accelerated flight conditions are transmitted to the lines attached to the valve.
b. Procedures. Compliance with this rule is usually aocomplished by designing the installation of the fuel valve so that the valve is supported by either primary or secondary airframe structure.
Chap 2 772 (thru 774) Par 483
486. § 29.997 (through Amendment 29-19) FUEL STRAINER OR FILTER.
a • Explanation. This rule provides for a main in-line fuel filter designed 00523 to collect all fuel impurities whioh could adversely affect fuel system and engine components downstream of the filter. The rule also requires a sediment bowl and drain (or that the bowl be removable for drain purposes) to facilitate separation of contaminations, both solid and liquid, from the fuel.
b. Procedures•
(1) The filter should be mounted in a horizontal segment of the fuel line to facilitate proper action of the sediment bowl. If the filter is located above the fuel tank, it becomes necessary to activitate a fuel boost pump to achieve positive drainage of the filter bowl. Without pump pressure, air may enter the fuel system during the filter draining operation and, for turbine engines, result in transient power surges or engine failure during subsequent engine operation. A flight manual note to require pump(a) to be "on" during filter draining would be appropriate.
(2) Section 29.997(d) sets forth a requirement for filter capacity and for filter mesh. The capacity requirement may be substantiated by showing that the filter, when partially blooked by fuel contaminates (to a degree corresponding to the indicator marking or setting required by $29.1305(a) (17)), does not impair the ability of the fuel system to deliver fuel at pressure and flow values established as minimum limitations for the engine. The filter mesh must be sized to prevent passage of particulate whioh cannot be tolerated by the engine. FAR Part 33 requires that the degree and type of filtration be established. This information should be the base for selecting the filter mesh. Although a test may be devised and conducted, data from the filter manufacturer usually are acceptable to verify compliance. Note that when the filter capacity is reached, continued flow of contaminated fuel may result in engine failure. A flight manual note.
regarding precautionary procedures is appropriate.
(3) FAR Part 33 (through Amendment 33-6) has an identioal requirement for a fuel filter for engine fuel systems; however, it is not intended that two filters should be required.
Chap 2 775 Par 486
AC 29-2, CHG 1 11/19/84 * 487. § 29.999 (through Amendment 29-19) FUEL SYSTEM DRAINS.
a. Explanation. This regulation provides for fuel system drains and defines the requirements which the system must meet.
b. Procedures.
(1) The location and function of the fuel gystem drains are an integral part of any fuel system. There may be several drains required dependent upon the fuel system design. Each fuel tank sump and certain types of fuel strainers or filters require a means to drain (ref. $§ 29.971 and 29.997).
(2) Selection of the location and orientation of the drain discharge in the design phase is important to assure that there is no impingement on any part of the rotorcraft. To show compliance with the requirement may require tests dependent upon whether the applicant has a previously approved design which is similiar, or if the system is a new design for which no previous experience is available.
(3) The location of the drain valve should be selected so that the requirements for accessibility, ease of operation, and protection are met.
(4) Advisory Circular 20-119 provides an acceptable means, but not the only means, of compliance with the requirement for positive looking of fuel drain valves in the closed position.
(5) The fuel drain installation on aircraft with retractable landing gear will be satisfactory if recessed within the outside surface of the aircraft.*
488.-497. RESERVED.
Chap 2 776 (thru 788) Par 487
* 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