Pressure Altitude Air data basis for Flight Speed establishing power Free Air Temperature plant inlet pressure (stagnation) and temperature.
Torque Direct measurement of Power Turbine Speed power output.
Gas Generator Speed(s) Primary thermodynamic Measure Gas Temperature and limiting parameters Fuel Flow Secondary trend monitoring and potential limiting parameter
(ii) The overall power check accuracy can be assessed on a suitable stat istical basis using equations that link the measured parameters and inserting system accuracy distributions for each value. This approach will provide an overall assessment of power check accuracy and will highlight major contributors to error. The accuracy assessment at each parameter should include the following elements:
Sensor error Indicator error System error Reading error
(iii) This assessment might show that while conventional instrument displays of air data are acceptable, servo driven digital displays are desired for engine parameters. Further, displays that provide a “snapshot” of engine readings at a given moment may be useful in avoiding variation in power level during the finite period needed to manually read and log the set of parameters.
(2) Installation Loss Definition.
(i) Installation loss definition is an extremely important aspect of any form of rotorcraft engine performance. Engines are certificated and sold with uninstalled performance guarantees and estimates as to the power output capabilities. Installation of the engine in the rotorcraft imposes power output penalties that must be accounted for in any sort of power assurance check procedure. Normal practice dictates that the engine manufacturer provides a computer program that accurately predicts the engine power output capability throughout the approved flight envelope. This computer program has the capability to correct the power output for the losses incurred by the rotorcraft installation.
(ii) Losses that can reduce engine power available are as follows:
(A) Air intake total pressure loss Page MG 9 - 6
(B) Air intake total temperature rise (C) Exhaust back pressure (D) Accessory power extraction (E) Compressor bleed air extraction (F) Off-optimum power turbine output speed effects
(iii) The above items and methods of dealing with them are clearly defined in SAE Aerospace Recommended Practice (ARP) 1702. Typically, these losses will not be a fixed percentage but will vary with engine operating conditions and environment.
(iv) Any calculations involving power assurance data should use the approved engine performance program, and the rotorcraft losses should be input on a discrete basis so that the interaction between losses and their independent variability is properly considered. This approach is clearly defined in ARP 1702. Accurate consideration of the losses should produce a Power Assurance Check that will preclude premature removal of acceptable engines or continued operation of inadequate power plants.
g. Rotorcraft Flight Manual (RFM).
(1) The Power Assurance Check data for the installed engine (engine data adjusted for inlet losses, exhaust losses, bleed extraction, power extraction, and off-optimum output shaft speed operation) should be presented in the RFM in an easily useable format. The data format may consist of charts of engine torque (at constant power turbine shaft speed) versus allowable values of gas generator speed and gas path temperature covering the range of ambient conditions for takeoff operations.
Associated limitations for the rotorcraft transmission and the engine should be noted.
(2) The RFM should also address the following:
(i) Include succinct statements of the reason for the Power Assurance Check and what must be done if the Power Assurance Check results are not acceptable.
(ii) Clearly state that Power Assurance Check either is a pre-takeoff or in-flight procedure, as required by operations, specifications and/or other approval authority documents.
(iii) Be kept simple, easy to use, and identify equipment operation limitations and requirements.
Page MG 9 - 7
NEEDLE VALVE AIRCRAFT FUEL TANK
FUEL FLOW TRANSMITTER BY-PASS TEE FUEL PRESSURE TRANSMITTER FUEL PRESSURE GAGE INSTALLED ENGINE FUEL FLOW CAPACITY CHECK
FUEL NOZZLE COCKPIT FUEL FLOW GAGE
FUEL CONTROL FUEL PUMP FIGURE AC 29.MG 9-1 Page MG 9 - 8
CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 10 ADVISORY MATERIAL FOR SUBSTANTIATION OF EMERGENCY
FLOTATION SYSTEM
a. Reference. FAR sections 29.521, .563(b), .751, .753(a)(1), (a)(2), .801(b), (d), .807(d).
b. Explanation .
(1) This section pertains to emergency flotation systems used to provide buoyancy for rotorcraft not specifically certificated for ditching but performing over-water operations. According to paragraph AC 29.801, ditching may be defined as an emergency landing on the water deliberately executed with the intent of abandoning the rotorcraft as soon as practical. Currently, ditching certification is not required by FAR 29; however, certification requirements are prescribed for applicants requesting ditching certification approval. If a rotorcraft operates over water during a Part 135 operation, the rotorcraft must comply with FAR 135.183, which may require floats.
(2) There are no airworthiness rules specifying the minimum standards for emergency flotation systems on rotorcraft not certificated for ditching. Equipment presented for evaluation must perform its intended function and not create a hazard for the rotorcraft or occupants. The objective in evaluating emergency flotation systems is safe flight and evacuation of the rotorcraft in emergency situations. Adequate emergency flotation systems would aid in keeping rotorcraft sufficiently upright and in adequate trim to permit safe and orderly evacuation in an emergency water landing.
c. Procedures. The following guidance criteria is based on past certification policy and experience for emergency flotation systems. Demonstration of compliance to other criteria may produce acceptable results if adequately justified by rational analysis.
Model tests of the appropriate emergency water landing configuration may be conducted to demonstrate satisfactory flotation and trim characteristics where satisfactory correlation between model testing and flight testing has been established.
Model tests and other data from rotorcraft of similar configurations may be used to satisfy the water requirements where appropriate.
(1) Flotation Systems.
(i) Normally inflated. The flotation systems which are normally inflated and intended for emergency use only, should be evaluated for:
Page MG 10 - 1
(A) Structural integrity when subjected to:
(1) Air loads throughout the approved flight envelope with floats installed,
(2) Water loads during water entry, and
(3) Water loads after water entry at speeds likely to be experienced after water impact.
(B) Rotorcraft handling qualities throughout the approved flight envelope with floats installed.
(ii) Normally deflated . Emergency flotation systems which are normally stowed in a deflated condition and inflated either in flight or after water contact during an emergency water landing should be evaluated for:
(A) Inflation.
(1) Proper Inflation. The inflation system design should minimize the probability of the floats not inflating properly or inflating asymmetrically. This may be accomplished by use of a single inflation agent container or multiple container system interconnected together. Redundant inflation activation systems will also normally be required. If the primary actuation system is electrical, a mechanical backup actuation system will usually provide the necessary reliability. A secondary electrical actuation system may also be acceptable if adequate electrical system independence and reliability can be documented.
(2) Inadvertent actuation. The inflation system should be safeguarded against spontaneous or inadvertent actuation for all flight conditions. It should be demonstrated that float inflation at any flight condition within the approved operating envelope will not result in a hazardous condition unless the safeguarding system can be shown to be reliable. Limitations to the approved envelope can be established so inadvertent actuation does not impose a hazard at the new envelope.
(3) Float actuation. The float activation means may be fully automatic or manual with a means to verify primary actuation system prior to each flight. If manually inflated, the float activation switch should be located on one of the primary flight controls. These activation means should be safeguarded against spontaneous or inadvertent actuation for all flight conditions.
(4) Flight Limitations. Maximum airspeeds for intentional in- flight actuation of the float system and for flight with the floats inflated should be established as limitations in the Rotorcraft Flight Manual (RFM) unless in-flight actuation is prohibited by the RFM.
Page MG 10 - 2
(5) Inflation time. For floats inflated automatically by water contact, inflation time from actuation to neutral buoyancy should be short enough to prevent the rotorcraft from becoming submerged to the point where egress is impeded.
(6) Pressure checking. A means should be provided for checking the pressure of the gas storage cylinders prior to each flight. A table or device showing acceptable gas cylinder pressure variation with ambient temperature and altitude (if applicable) should be provided.
(7) Over inflation. A means should be provided to minimize the possibility of over inflation of float bags under any reasonably probable actuation conditions.
(8) No puncture inflation. The ability of the floats to inflate without puncture when subjected to actual water pressure should be substantiated. A full scale rotorcraft immersion demonstration in a calm body of water is one acceptable method of substantiation. Other methods of substantiation may be acceptable depending upon the particular design of the flotation system.
(9) Flotation bag containment. Float installations should be evaluated to ascertain that emergency exits are not blocked by the inflated floats when the float bags are inflated to their maximum inflation pressure or their most adverse inflation pressure for emergency exits and the rotorcraft at its most critical weight and center of gravity configuration.
(B) Struc tural Integrity. The flotation bags should be evaluated for loads resulting from:
(1) Airloads during inflation and fully inflated during the most critical flight conditions and water loads with fully inflated floats during water impact for the rotorcraft desiring float deployment before water entry; or
(2) Water loads during inflation after water entry.
(C) Handling qualities. Rotorcraft handling qualities should be verified by test or analysis to comply with the applicable regulations throughout the approved operating envelopes for:
(1) Deflated and stowed condition,
(2) In-flight inflation condition,
(3) Fully inflated condition, and
(4) Partially inflated condition, assuming the most critical float compartment fails to inflate.
Page MG 10 - 3
(2) The float system attachment hardware should be shown to be structurally adequate to withstand critical air loads and water loads during water entry when both deflated and stowed and fully inflated (unless in-flight inflation is prohibited). The appropriate vertical loads and drag loads determined from water entry conditions (or as limited by flight manual procedures) should be addressed. The effects of the vertical loads and the drag loads may be considered separately for the analysis.
(3) Flotation and Trim should be investigated for a range of sea states from zero to the maximum selected by the applicant and should be satisfactory in waves having height/length ratios of 1:12.5 Category A rotorcraft, 1:10 Category B rotorcraft with Category A engine isolation, and 1:8 for Category B rotorcraft.
(i) Demonstrated to be satisfactory to at least sea state 4 water conditions.
(ii) Flotation tests should be investigated at the most critical rotorcraft loading condition.
(iii) Flo tation time and trim requirements should be evaluated with a simulated, ruptured deflation of the most critical float compartment. Flotation characteristics should be satisfactory in this degraded mode to at least sea state 2 water conditions.
(iv) Probable rotorcraft door/window open or closed configurations and probable damage to the airframe/hull (i.e., failure of doors, windows, skin, etc.) should be considered when demonstrating compliance with the flotation and trim requirements.
(4) Float System Reliability. Reliability should be considered in the basic design to ensure approximately equal inflation of the floats to preclude excessive yaw, roll, or pitch in flight or in the water.
(i) Maintenance procedures should not degrade the flotation system (such as introducing contaminants which could affect normal operation, etc.).
(ii) The flotation system design should preclude inadvertent damage due to normal personnel traffic flow and excessive wear and tear. Protection covers should be evaluated for function and reliability.
(5) Buoyancy requirements for emergency flotation systems should be a minimum of 25 percent excess buoyancy at maximum internal gross weight. The weight of fresh water (density 62.42 lb/ft3) displaced by fully submerged float or floats should be a minimum of 25 percent greater than the maximum certificated internal gross weight of the rotorcraft. Analysis may be used for buoyancy verification.
Page MG 10 - 4
(6) Sufficient watertight compartments should provide an acceptable margin of positive stability with any single main float compartment flooded or deflated. The location of the floats, the most critical compartment, the rotorcraft weight, mass moment of inertia, and center of gravity location are also important considerations for stability. Analyses, tests, or a combination thereof may be used to substantiate a positive margin of stability with the most critical compartment flooded or deflated.
(7) The inflatable bag type floats should be designed for the maximum pressure differential developed at the maximum design altitude. That is, the resulting pressure difference between an operational altitude and a take-off site elevation should be established and substantiated. This resulting pressure differential may become an operating limitation.
(8) The float landing loads may be determined from the drop test of the float landing gear or the loads may be derived from landing gear drop test or loads may be determined from model or full scale water entry tests. The vertical loads are distributed over three fourths of the bag's projected area. Bag floats are not subject to the side loads. Rigid floats are to be designed for vertical, horizontal, and side loads distributed along the length of the float.
(9) Design and/or support of the forward part of bag type floats should be evaluated for maximum design speeds to prevent collapse or significant distortion of the bag while in flight.
(10) Resistance to puncture and abrasion at attach/wear points is an important design consideration. Girt or attachment design loads should be sufficient to withstand the maximum imposed design loads.
(11) Occupant Egress and Survival. Each practicable design measure should be taken to minimize the probability that the behavior of the rotorcraft would cause immediate injury to the occupants or prevent evacuation of the rotorcraft after an emergency landing on water. Emergency exits should be located such that they are above the waterline and will not be blocked by the inflated floats or partially inflated floats, impeding evacuation of the rotorcraft. The flotation time and trim of the rotorcraft should allow the occupants to evacuate the rotorcraft. i.e., the rotorcraft should remain sufficiently upright and in adequate trim to permit safe and orderly evacuation of all personnel. For configurations which are considered to have critical occupant egress capabilities due to float proximity, an actual demonstration of egress may be required. When a demonstration is required, it may be conducted on a full-scale rotorcraft actually immersed in a calm body of water or using any other rig/ground test facility shown to be representative. The demonstration should show that floats do not impede a satisfactory evacuation.
(12) Rotorcraft Flight Manual. The Rotorcraft Flight Manual should contain the information pertaining to the emergency flotation system. This material should include:
Page MG 10 - 5
(i) The information pertinent to the limitations applicable to the emergency float system and operating limitations for the emergency float system,
(ii) Procedures and limitations for flotation device inflation,
(iii) Procedures for use of emergency flotation equipment, and
(iv) Procedures for emergency water landing occupant evacuation.
Page MG 10 - 6
CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 11 FATIGUE TOLERANCE EVALUATION OF TRANSPORT CATEGORY
ROTORCRAFT METALLIC STRUCTURE
a. Purpose. This advisory material provides an acceptable means of compliance with the provisions of § 29.571, Amendment 28, of the Federal Aviation Regulations (FAR) dealing with the fatigue tolerance evaluation of transport category rotorcraft metallic structure. For safe- life evaluations, AC 27-1B, MG 11 (Fatigue Evaluation of Rotorcraft Structure) provides in depth guidance on acceptable means of compliance.
The fatigue evaluation procedures outlined in paragraph AC 29 MG 11 are for guidance purposes only and are neither mandatory nor regulatory in nature. Specific issues related to substantiation of composite rotorcraft structures are addressed in AC 29 MG 8. Although a uniform approach to fatigue evaluation is desirable, it is recognized that in such a complex problem, new design features and methods of fabrication, new approaches to fatigue evaluation, and new configurations may require variations and deviations from the procedures described herein. It is recommended that major deviations from the procedures be coordinated with the certifying regulatory authority to assure compliance with the regulatory requirements.
b. Special Considerations . The structure of rotorcraft is subject to cyclic stresses in practically every regime of flight. In addition, since rotorcraft are highly maneuverable and capable of forward, rearward, sideward, vertical, and rotational flight, operating limitations due to fatigue are possible in practically all flight situations. Corrosion and other environmental damage are also common in rotorcraft operations. For these reasons, special attention should be focused on the fatigue evaluation of rotorcraft structure.
Background.
(1) Fatigue substantiation of rotorcraft dynamic components was first implemented in the 1950’s by means of Safe-Life Methodology. Many advances in design and analytical methods have been made in the state-of-the-art and in industry practices since that time. To date, Safe- Life Methodology, as described in AC 27-1B, MG 11, is considered successful in providing a high level of reliability, but could be improved by taking into account the strength-reducing effects of damage likely to occur in manufacturing, maintenance, and in service, including corrosion, accidental damage, or manufacturing/maintenance flaws. The introduction of composites led the manufacturers/certifying authorities to take into account the specific static and fatigue strength-reducing effects of aging, temperature, moisture absorption, impact damage, and recognition of a minimum quality standard. In parallel, crack growth methodology in Page MG 11 - 1
29-2C, Chg 1 2/12/03 metal structure has been successfully used for solving short-term airworthiness problems in rotorcraft, and as the certification basis for civil and military transport aircraft applications.
(2) Recognizing that advances in state-of-the-art and industry practice warranted changes to the existing fatigue requirements in Part 29, the regulatory requirements of § 29.571 were substantially revised. The revision to § 29.571 requires new guidance material containing compliance provisions related to the changes.
AC 29 MG 11 provides this material with respect to the flaw tolerance requirements for metallic structure and is supplemented by AC 27-1B MG 11. Guidance material specifically addressing composite rotorcraft structure is provided in Chapter 3, MG 8 of this AC (1) and is supplemented by AC 20-107A.
d. Introduction .
(1) Definitions. These definitions are provided to define the terms used in applying the requirements specified in FAR 29.571, Amendment 28, and may differ from other definitions.
Fatigue Tolerance. The capability of structure to continue functioning without catastrophic failure after being subjected to fatigue (repeated) loads expected during operation of the rotorcraft. Fatigue tolerance should be achieved by flaw tolerance design, or if impractical, safe-life design, or a combination.
(ii) Safe Life. The capability of as -manufactured structure as shown by tests, or analysis based on tests, not to initiate fatigue cracks during the service life of the rotorcraft or before an established replacement time.
(iii) Flaw Tolerance. The capability of rotorcraft structure to achieve fatigue tolerance accounting for the presence of flaws and damage which may occur in manufacturing and service use. Flaw tolerance can be achieved by either flaw tolerance safe-life or fail-safe designs. The term ‘Damage Tolerance’ is frequently used to describe the ability of a structure to tolerate the effects of flaws and damage;
however, the terminology of FAR 29.571, Amendment 28, is used in this AC to maintain consistency.
(iv) Flaw tolerant Safe Life. The capability of as -manufactured structure with expected flaws as shown by tests, or analysis based on tests, not to initiate fatigue cracks during the service life of the rotorcraft or before an established replacement time.
(v) Fail-Safe. The capability of structure remaining after a partial failure to withstand design limit loads without catastrophic failure within an inspection period.
(vi) Multiple Load Path. Structure providing two or more separate and distinct paths of structure that will carry limit load after complete failure of one of the members.
Page MG 11 - 2
(vii) Active Multiple Load Path. Structure providing two or more load paths that are all loaded during operation to a similar load spectrum.
(viii) Passive Multiple Load Path. Structure providing load paths with one or more of the members (or areas of a member) relatively unloaded until failure of the other member or members.
(ix) Crack Arrest Feature. Structure that does not provide completely separate and distinct load paths but does provide features of design such as bonded and/or riveted straps, changes in geometry, or special processing techniques such as rolling or coining to retard or arrest crack growth.
(x) Slow Crack Growth Feature. Structure (single element or multiple element) that provides for slow crack growth by material selection, material processing, limitation of stress levels, geometrical design features, or by other methods.
(xi) Flaw. Intrinsic imperfec tions such as inclusions, forging laps, or porosity, and discrete damage such as gouges, scratches, nicks, corrosion, fretting, wear, or impact, that could be expected during manufacture or operation.
(xii) Design Limit Loads. The maximum loads to be expected in service, as defined by § 29.301(a).
(xiii) As Manufactured. Product and/or component that has passed the applicable quality control process and has been found to conform to the approved design within the allowable tolerances.
(xiv) Residual Strength. The strength retained for some period of unrepaired use after a failure or partial failure due to fatigue, corrosion, and accidental or discrete source of damage.
(xv) Principal Structural Element (PSE). A structural element that contributes significantly to the carrying of flight or ground loads and whose failure due to fatigue can lead to catastrophic failure of the rotorcraft.
(2) Rotorcraft Fatigue Tolerance. Fatigue tolerant design as substantiated by fail-safe flaw growth or flaw tolerant safe-life means outlined in § 29.571 and paragraph AC 29 MG 11g is required for all PSE’s, unless it entails such complications that an effective flaw tolerant structure cannot be achieved within the limitations of geometry, inspectability, or good design practice. Good design practice includes consideration of component complexity, component weight, methods of production, and component life cycle cost. Under these circumstances, a design that complies with safe-life criteria should be used. Typical examples of structure that might not be conducive to flaw tolerance designs are swashplates, main rotor shafts, push rods, small rotor head components (i.e., devices, bolts, etc.), landing gear, and gearbox Page MG 11 - 3