SECTION 24. POWERPLANT - GENERAL
397. § 27.901 (through Amendment 27-20) INSTALLATION.
a. Section 27.901 (a)
(1) Explanation. Paragraph (a) provides a definition of parts of rotorcraft for which safety requirements are set forth under the general title, SUBPART E - POWERPLANT. These parts include not only major propulsive elements and power transmissive components but also controls, instruments, safety devices, including fire protection and other devices to protect personnel, and critical flight structure in event of fires.
(2) Procedures. To ensure that no certification aspect is overlooked in establishing compliance, certification engineers should make at least an informal breakdown of all components of the rotorcraft, assigning responsibility to powerplant certification engineers of all items within the above definition. While this procedure is usually straightforward, the following items of FAA/AUTHORITY powerplant responsibility are listed to minimize questions regarding authority and responsibility.
(• Drive system components. All parts of the transmission, clutches, shafting, including the driveshafts (masts) of main and auxiliary rotors, powerplant cooling components, and powerplant instrumentation requirements under §§ 27.1305,
27.1337, 27.1543, 27.1549, 27.1551, 27.1553, 27.1555, and 27.1583.
NOTE: The division of responsibility between FAA/AUTHORITY airframe engineers and FAA/AUTHORITY powerplant engineers (in accordance with FAA/AUTHORITY practice) regarding the driveshaft is at the flange or spline interface between the driveshaft and the rotor hub. Rotor hubs, controls, blades, and associated components are the airframe engineers' responsibility. (Industry practice may not agree with this concept.)
(il) Engines, except for mount structure.
(iii) Auxiliary power units, except for mount structure.
(iv) Combustion heaters, except for downstream ventilation air ducting, mixing, and distribution systems and for electrical aspects of controls and safety devices.
(v) Water/alcohol or other fluid power augmentation systems.
(vi) Engine induction systems including induction icing and snow ingestion, and exhaust systems, including exhaust shrouds and drains.
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(vii) All fuel systems, including those serving engines, auxiliary power units, combustion heaters, power augmentation systems, etc., and vents and drains for those systems.
(vili) Oil systems for engines, auxiliary power units, rotor drive transmissions, and gearboxes, including grease lubrication.
(ix) Cooling aspects of engines, rotordrive transmissions and gearboxes, and auxiliary power units.
NOTE: Electrical generating equipment and hydraulic component cooling may be the responsibility of the systems and equipment engineer provided agreement is established among responsible personnel.
(x) Rotor brakes, except hydraulic and electrical aspects and structural aspects of nonrotating brake components.
(xi) Fire protection, including firewalls, fire extinguisher systems, fire detector systems, flammable fluid lines, fittings, and shutoff valves. The powerplant engineer has responsibility for evaluating compliance with §§ 27.861 and 27.863 as it pertains to fuel and oil systems.
(xii) Engine and transmission cowling and covering, including latches.
(xiii) Powerplant flexible controls.
(xiv) Powerplant accessories.
(xv) Pneumatic systems (engine bleed air) within the engine compartments, including shut-off valves and engine isolation features of bleed systems.
(xvi) Powerplant aspects of instrument markings and powerplant aspects of flight manuals, including limitations, normal and emergency procedures, engine pertormance; powerplant aspects of maintenance manuals, with emphasis on the limitations section of the manual and verification of the limitations established under § 27.1521.
b. Section 27,901(b).
(1) Explanation. Paragraph (b) requires that the various powerplant components and systems be investigated for general airworthiness.
(2) Procedures.
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АС 27-1A 7/30/97 Each item of the powerplant area of responsibility should be shown to be suitable for its intended purpose and installed to operate satisfactorily and safely between normal inspections and overhauls. Accessories mounted on engine or transmission drive pads should be determined to be compatible with the pad limits, including fit and speed range, overhang moment loads, running torque and static torque. This latter term pertains to protection of the engine or transmission which drives the accessory from damage to be expected from malfunction of the accessory. This protection is usually supplied by providing a shear section in the accessory drive shaft designed to fail before exceeding the static torque limit of the engine or transmission driving component. Note that when evaluating the strength of the mechanical shear section, material allowables quoted in materials handbooks should not be used since these are minimum strength values. Shear sections should consider maximum strength values to be expected which are on the order of 130 percent of the minimum strength values. Also, it should be verified that design data for shear sections are dimensioned to limit the maximum diameter as well as the minimum diameter. Installation of starter-generators may also require verification that horsepower extraction limits are not exceeded. Special flightcrew instructions in the flight manual to monitor generator load or to disconnect electrically loaded items to protect accessory or engine-transmission pad limits should be avoided.
(її) Environmental qualification requires consideration or protection against adverse effects of extremes of cold weather, salt and sand/dust atmosphere, altitude effects, etc. Most powerplant components are subjected to many of these aspects during the individual qualification tests; however, satisfactory overall integrated system performance under these adverse conditions must be verified. Cold weather testing should include verification that lubricating oils and greases function properly, and that engine starting procedures are safe and do not impose excessive loads on accessories, engines, or drive system components. Powerplant engineers should coordinate compliance efforts in this area with system engineer's investigations of compliance with §§ 27.1301 and 27.1309. Full-scale rotorcraft operations in cold weather should be required, including at least some exposure in the range of -10° to -20° F if the aircraft is to be certified to these ambients. Cold soak or overnight exposure to cold weather is appropriate followed by starting and pretakeoff procedures in accordance with the flight manual. Attention should be given to the practicality of important mandatory inspection procedures as affected by cold weather.
(ili) Accessibility for maintenance should be reviewed. Typically, some maintenance activities must involve disassembly or removal of adjacent components.
This should be avoided if repetitive activity can jeopardize the performance of critical or safety-related equipment. Verify that easy access exists to items such as oil system sight gauges or dip sticks, filler ports and drain valves for engines, auxiliary propulsion units, transmissions, fuel tanks and filters, etc.
(iv) Electrical interconnections to prevent difference of potential should be provided in the form of grounding straps or wires sized to carry the currents to be 502 Par 397
7/30/97 АС 27-1A expected. Verify that the attachments for these grounding devices are not compromised by paint or zinc chromate which will tend to electrically insulate the engine or component. Note that engine mount structure should not be accepted as a grounding device since electrical current will cause corrosion at attach points.
(v) Axial and radial expansion of turbine engines is usually not a problem unless redundant mount arrangements are used. Special expansion provisions are usually required if engine components other than mounting points are attached to bulkheads, firewalls, other engines, or drive system components. Engine output shaft axial or bending loads due to thermal expansion and to deflection of supports under ground or flight loads should be checked. Other components of concern are compressor inlet flanges, exhaust ducts, and rigid fluid or air lines between aircraft structure and the engine. The engine installation data will provide limit loads to be considered for parts of the engine which normally are attached to airframe components.
c. Section 27.901(c).
(1) Explanation. Paragraph (c), in conjunction with the installation manual requirements of § 33.5, is intended to assure compliance with the detail installation requirements developed by the engine manufacturer to assure safe, continued operation of the engine.
(2) Procedure. Compliance with most of the detail requirements in the engine installation manual can be established by test or by design features and arrangements negotiated between the rotorcraft manufacturer and the FAA/AUTHORITY powerplant engineer. Some aspects, usually involving inlet and/or exhaust distortion limitations, vibration limitations and aircraft/engine interface items may require direct assistance and information from the engine manufacturer to determine that compliance with the installation manual exists. Fuel control/engine/rotor system torsional matching is usually a developmental problem to be worked out before presentation of the rotorcraft to the FAA/AUTHORITY; however, final flight tests for surge or stall, torsional stability, and acceleration/ deceleration schedules may require direct coordination among FAA/AUTHORITY installation engineers, engine manufacturers' representatives, and the FAA/AUTHORITY engine certification engineers. Reciprocating, carburetor equipped engines usually require a particular carburetor configuration to achieve adequate engine cooling. This configuration, identified as a "carburetor parts list," must be approved for the engine under Part 33 and should be listed on the type data sheet for the rotorcraft.
397A. § 27.901 (Amendment 27-23) INSTALLATION.
a. Explanation. Amendment 27-23 changes § 27.901(b)(1) to require a satisfactory determination that rotorcraft can operate safely throughout adverse environmental conditions such as high altitude and temperature extremes. This amendment was needed to provide consistent application of environmental qualification Par 397 503
aspects. This amendment also added a new Paragraph (§ 27.901(b)(5)) to require design precautions to minimize the potential for incorrect assembly of components and equipment essential to safe operation.
b. Procedures. All of the policy material pertaining to this section remains in effect with the addition of design precautions. Design precautions should be taken to minimize the possibility of improper assembly of components essential to the safe operation of the rotorcraft. Fluid lines, electrical connectors, control linkages, etc., should be designed so that they cannot be incorrectly assembled. This can be achieved by incorporating different sizes, lengths, and types of connectors, wires, fluid lines, and mounting methods.
398. § 27.903 (Amendment 27-11) ENGINES.
a. Engine Type Certification.
(1) Explanation. Section 27.903(a) is intended to ensure that engines used in type certified aircraft are properly qualified and that the associated installation requirements are established.
(2) Procedures.
Compliance can be documented by verification that a type certificate data sheet has been issued by the FAA/AUTHORITY for the engine identified by the rotorcraft manufacturer as the engine planned for use in the rotorcraft. Reciprocating engines must have been qualified to a special test plan (§ 33.49(d)) to be eligible in rotorcraft. This eligibility should be verified by a note on the engine type certificate data sheet.
On some occasions, the engine certification program is conducted concurrently with the rotorcraft certification program. This is technically acceptable provided the engine type certificate is issued prior to the rotorcraft type certificate.
However, practical considerations involving the use of unapproved engine installation data and the probability of engine design changes during the engine certification program that impact the rotorcraft certification program dictate that special procedures must be introduced to assure that the final rotorcraft certification program is satisfactory.
If the engine under consideration is merely a minor model change from a previously certificated engine and these changes are unlikely to cause rotorcraft certification problems and do not involve significant installation aspects, the rotorcraft project engineer need only to follow the engine certification program by routine checks with the FAA/AUTHORITY office responsible for engine certification and, as a final pre-type certification item, verify that the engine type certificate has been issued. Rotorcraft Type Board agenda/minutes should reflect the ongoing status of the engine TC program. For rotorcraft certification programs involving new or significantly changed engines, the powerplant certification engineer for the rotorcraft should become as 504 Par 397A
familiar with the engine as practicable with particular attention to engine ratings, limitations, performance, engine/rotorcraft interface aspects, and any Part 27 certification requirement involved in the engine program (fuel/oil filters, fuel heaters, integral firewalls, etc.) and establish an appropriate working arrangement with the FAA/AUTHORITY engine certification office to monitor changes in the engine certification progress which may impact the rotorcraft certification program. In addition, any rotorcraft certification activity such as test plans, analysis, compliance inspections, etc., which involves the engine should be accepted on a conditional basis; i.e., pending confirmation of completion of the engine program without changes pertinent to these aspects of rotorcraft program. The rotorcraft applicant should be advised of any limitations in this procedure, and that normally, the engine certification program should be complete before authorizing formal FAA/AUTHORITY participation in the rotorcraft certification plan; i.e., TIA.
b. Engine cooling fan protection.
(1) Explanation. Section 27.903(b) is intended to provide safety to the rotorcraft in the event of an assumed cooling fan blade failure or to prescribe a test to show that the cooling fan blade retention means is sufficient that blade failure is not a consideration.
(2) Procedures. The applicant may select § 27.903(b)(1), (b)(2), or (b) (3) to show compliance with this section. If § 27.903(b)(1) is selected, a demonstration should be conducted to show that at the maximum fan speed to be expected, a failed blade is contained within a housing or shroud which is included in the proposed type design and designated by the applicant as the containment shield. The rotational speed required may be related to an overspeed limiting device or to the maximum transient speed to be expected from analysis or test of the system or component which drives the fan. For components driven directly by the engine, output shaft disconnect and the subsequent terminal speed of the engine may set the test condition. To conduct an overspeed blade failure containment demonstration, applicants have found it convenient to progressively weaken a blade to induce failure at or above the required demonstration speed. Blade failure may be expected to subsequently fail some or all of the remaining blades. This condition, provided all blades are contained, is acceptable for showing compliance with this rule. However, the corresponding loss of cooling may be unacceptable if it causes the loss of any function essential to a controlled landing.
(3) Section 27.903(b)(2) may be selected; however, without containment, damage to any component or structure in the plane of the fan rotor or any other trajectory to be expected should not cause the loss of any function essential to a controlled landing.
(4) If § 27.903(b)(3) is selected, a spin test at 122.5 percent of the maximum speed associated with either engine terminal speed or an overspeed limiting device would be acceptable to show compliance. No failure should occur and distortion should Par 398 505
not result in fan element contact with housings or other adjacent components. (Note:
150 percent of the centrifugal force is achieved at 122.5 percent of the rotational speed.)
c. Turbine Engine Installation.
(1) Explanation. The certification of turbine engines and particularly, the qualification of turbine rotors, assumes that the limitations established during these certifications will be accurately and rigorously observed during ground and flight operations in an aircraft. This paragraph is intended to promote this concept.
(2) Procedures. Primary engine limitations in the form of time, gas temperature, torque, and rotational speed and their corresponding allowable transient values are defined in the approved engine installation manual. The rotorcraft manufacturer must provide reliable, accurate means to assure that these limitations are not exceeded. These means may be in the form of automatic limiters or by crew monitoring of appropriately marked instruments. The FAA/AUTHORITY powerplant certification engineer and the rotorcraft manufacturer's staff should verify these aspects by:
(• Evaluating all applicable instrument, indicator, or warning devices, including transmitters, and limiting devices, if any, for system tolerances.
(її) Closely reviewing the component qualification reports of items in 398c(2)(i) above to verify that these devices are properly qualified and that any deviations are acceptable.
(iii) Assuring that maintenance data is provided for functional checks and calibration of instruments and devices which are used to monitor or protect critical turbine rotor limitations. Preflight checks for automatic limiter devices may be appropriate.
(iv) Verifying that instrument markings are clear and relatively simple, that corresponding flight manual instructions and descriptions are straightforward and complete, and instruments are located and orientated to minimize the probability of reading error.
398A. $27.903 (Amendment 27-23) ENGINES.
a. Explanation. Amendment 27-23 adds a requirement to § 27.903(a) that requires reciprocating engines used in rotorcraft to be certified in accordance with the special rotorcraft engine test requirements in § 33.49(d). This change was needed to ensure that certification requirements are not overlooked when reciprocating engines are installed in rotorcraft to be certified under Part 27 requirements. Section (b) was revised to prescribe tests and qualifications for powerplant area cooling fans. This rule 506 Par 398
change requires cooling fans to be designed and installed to enable safe landing of the rotorcraft following a fan blade failure. Compliance with the previous requirements could result in hazards to the rotorcraft with the loss of cooling air to critical powerplant components. A new paragraph was also added to the rule for cooling fans, which are part of the powerplant installation. It should be determined that no cooling fan blade resonant conditions exist within the operating limits of the rotorcraft unless a fatigue evaluation is conducted. These requirements will ensure that correct qualification procedures are used for rotorcraft engines and that all powerplant cooling fans are properly tested.
b. Procedures.
(1) Engine type certification. All engines installed in rotorcraft should have a type certificate. The specific certification requirements for reciprocating engines when installed in rotorcraft are found in the paragraph listed in Part 33. Engines certificated under other approved certification rules (CAR Part 13 and FAR § 21.29, for imported engines) are also eligible. If a component, system, or arrangement is certified under Part 33 or other requirement, the applicant is not relieved of the necessity to comply with the requirements of Part 27. If the component, system, or arrangement, supplied as a part of a certificated engine, meets the Part 33 and Part 27 requirements, subsequent changes to these components, systems, or arrangements could negate compliance with Part 27.
(2) The applicant may select §§ 27.903(b)(1)(i), (b)(1)(ii), or (b)(1) (III) to show compliance with this section.
For compliance with § 27.903(b)(1)(i), a demonstration should be conducted to show that at the expected maximum fan speed, a failed blade will be contained within a housing or shroud that is included in the proposed type design and designated as the containment shield. The maximum fan rotational speed may be related to an overspeed limiting device or to the expected maximum transient speed from analysis or test of the engine, system, or component which drives the fan. For fans driven directly by the engine, output shaft disconnect and the subsequent terminal speed of the engine may establish the maximum fan speed for the test condition. To conduct an overspeed blade failure containment demonstration, applicants have found it convenient to progressively weaken a blade to induce tailure at or above the required demonstration speed. Blade failure may be expected to subsequently occur on some or all of the remaining blades. This condition, provided all blades are contained, is acceptable for showing compliance with this rule. However, the corresponding loss of cooling may be unacceptable if it causes the loss of any function essential to continued safe flight and landing.
(ii) For § 27.903(b)(1)(ii) compliance, if containment protection is not installed, damage to any component or structure within the trajectory of the failed fan rotor should not cause the loss of any function essential to a controlled landing.
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(ili) For § 27.903(b)(1) (ili) compliance, a spin test should be conducted.
For fans driven directly by the engine, the test should be conducted at 122.5 percent of the terminal engine rotational speed that will occur under uncontrolled conditions, or at
122.5 percent of the maximum engine rotational speed that would be controlled by a reliable, approved engine overspeed limiting device. For fans driven by the rotor drive system, the test should be conducted at 122.5 percent of the maximum rotor drive system rotational speed expected in service, including transients.
(Note: Capability to withstand the ultimate load of 1.5 times the centrifugal force means that no failure should occur and distortion should not result in fan element contact with housings or other adjacent components during the 122.5 percent spin test which equates to 150 percent centrifugal force).
(3) Fatigue. If the cooling fan is not included in the fatigue evaluation under § 27.571, it should be shown that the cooling fan blades are not operating at resonant conditions within the normal operating limits of the rotorcraft.
399. § 27.907 ENGINE VIBRATION.
a. Explanation. Section 27.907 is intended to require the design of the rotor drive system, including the engine, to be free from harmful vibration. A vibration investigation is required.
b. Procedures. Review Order 8110.9, Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and other Power Transmission Systems. Note that the mechanical coupling of the engines to the rotor drive system creates, for torsional vibration considerations, one, rather complicated, drive system which responds to any forced or resonant frequency. Antinodes or nodes and frequencies may exist in the engine shaft which are absent when the engine is operated on a test stand; therefore, the vibration investigation conducted under Part 33 is not conclusive with respect to torsionals. As noted in Order 8110.9, the engine manufacturers' assistance is necessary to find compliance. Section 27.571 was amended by Amendment 27-12 to include "rotor drive systems between the engines and the rotor hubs" as part of the flight structure. This rule supplements § 27.907 and requires coordination with the structures certification engineer to avoid duplication of effort by the rotorcraft manufacturer. Advisory Circular 20-95, which provides acceptable methods of compliance with § 27.571, may also be used to find compliance with § 27.907.
In addition to basic drive system components such as main and auxiliary rotor drive shafts, the vibratory evaluation should include couplings, gear teeth, gear cases and splines, and should consider, where appropriate, low cycle fatigue associated with ground-air-ground cycles.
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400.-420. RESERVED.
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SECTION 25. ROTOR DRIVE SYSTEM
421. § 27.917 (through Amendment 27-11) DESIGN.
a. § 27.917 (a):
(1) Explanation. This paragraph requires the design of the drive system to include a means to automatically disengage the engine(s) from the rotor drive system in order to prevent excessive drag from an inoperative engine from adversely affecting the pertormance of the rotor system.
(2) Procedures. The design objective usually is met by installing a freewheeling or overrunning clutch in the drive shaft between the engine and the first part of rotor drive system. If lubrication for these clutches is required, it should be provided by a means that continues to function after an engine is made inoperative except that for single-engine rotorcraft, clutch lubrication need only be provided for autorotation descent with the engine inoperative. A 15-minute demonstration of freewheeling or overrunning operation is usually acceptable.
b. § 27,917(b):
(1) Explanation. This paragraph requires that control rotors (tail rotors, for example) will continue to be driven by the main rotors when the rotorcraft is in autorotation.
(2) Procedures. Provide hard mechanical interconnect shafting between the rotors such that the main rotor will drive the control rotor (tail rotor). Note that this requirement must be met with all engines inoperative, thus, the driving force for the tail rotors must depend on the autorotative driving forces inherent in the main rotor(s).
c. § 27,917 (c):
(1) Explanation. This paragraph pertains to any device or feature designed into the rotor drive system intended to prevent damage in the event of excessive torque in the rotor drive system from high engine power or mechanical interference with normal rotation of the rotor drive system. The rule prohibits location of these devices in any part of the rotor drive system that is required to continue functioning to provide control of the rotorcraft.
(2) Procedures. Review the arrangement of the rotor drive system to determine that any intentionally designed weak links in the system, such as shear sections or slip clutches installed to relieve high torsional loads, are located so that their function will not compromise the interconnect mechanism between the main and auxiliary rotors of the rotorcraft.
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d. § 27.917 (d):
(1) Explanation. This paragraph sets forth a definition of the rotor drive system and its associated components.
(2) Procedures. Coordinate with other certification personnel to ensure that other rules pertaining to rotor drive systems are properly addressed.
422. $ 27.921 ROTOR BRAKE.
a. Background. Rotor brake safety requirements are intended not only to prevent adverse effects on aircraft performance due to brake drag but also to minimize the possibility of fire. These fires, caused by friction from a dragging rotor brake, have occurred both in flight and during ground operation with extremely hazardous consequences.
b. General. This rule requires (1) that any limitations on the use of the rotor brake must be established, and (2) that the control for the brake must be guarded to prevent inadvertent operation.
c. Limitations.
(1) The limitations on the use of the rotor brake should first be defined by the applicant and will normally consist of merely the maximum speed eligible for application of the brake. In some installations, other limitations associated with engine operation may be specified.
(2) Control guard mechanisms to prevent inadvertent operation may be conventional. A cockpit evaluation of the guard should be conducted by flight test personnel to affirm the function of the guard, that markings, if any, are adequate, and that both latched and unlatched positions of the guard do not interfere with other cockpit functions.
d. Other rules require both generalized and specific rotor brake qualification tests.
However, some significant aspects of brake safety tests are listed below for reference.
(1) Routine application of the brake at shutdown during the endurance test of § 27.923 and during the function and reliability tests of § 21.35.
(2) Torsional vibration loads in the rotor drive system and oscillatory loads in the brake components during a critical brake engagement procedure should be determined with appropriate consideration in the fatigue evaluation for these components. Brake engagements should be conducted with and without collective control displacement as authorized by the flight manual or a training manual.
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