AC 27-1B, Chg 4 2/27/2014 *(13) EDMS. If required under an operating regulation, an FDMS (not to be confused with a flight data recorder (FDR) certificated under § 27.1459) may be comprised of a system or combination of systems that record a helicopter's flight performance and operational data. An FDR certificated under § 27.1459 and the appropriate operating rules would be acceptable to meet this requirement; however, an FDMS would not be adequate to meet the § 27.1459 requirement for an FDR. The FDMS should record digital or analog raw data, images, cockpit voice or ambient audio recordings, or any combinations thereof, according to a broadly defined set of parameters including information pertaining to the aircraft's state, condition, and system performance. This data can be used to perform post flight analysis and provide critical information to investigators in the event of an incident or accident as well as to promote operational safety. When used in conjunction with an FAA-approved flight operations quality assurance (FOQA) program, part 135 certificate holders would be required to collect flight performance and operational data that characterizes the state of the helicopter and its subsystems that the certificate holder determines is pertinent to its safety program. FDMS data should be recorded and stored on digital media, and when selecting a location to install the hardware device used for storing the data, consideration should be given to the potential for survival in the event of a crash. The system should receive electrical power from the helicopter's bus that provides the maximum reliability without jeopardizing service to essential or emergency loads, and capable of being operated continuously from the time power is applied to the aircraft until power is removed from the aircraft.
(i) Safety. The FDMS equipment should not, under normal or fault conditions, adversely affect the airworthiness of the systems to which it is interfaced or of other aircraft systems. The equipment should be installed in accordance with all applicable safety regulations. The equipment should be tested under the standards of RTCA DO-160F, "Environmental Conditions and Test Procedures for Airborne Equipment," or subsequent issue. The European Organization for Civil Aviation Equipment (EUROCAE) specification ED-14 may be used in lieu of RTCA DO-160. Additional crashworthiness testing may be conducted according to EUROCAE specification ED-155, "Minimum Operational Performance Specification for Lightweight Flight Recording Systemsi Equipment testing should be conducted to the categories most applicable to the aircraft type, and the location of the equipment to be installed. The equipment manufacturer typically defines the test class within each environmental category. The objective of this level of testing is to ensure that the equipment does not present a hazard to the aircraft, and can survive and continue to operate under the environmental conditions to which it will be subjected throughout its life. Specific testing may be required to demonstrate that the equipment performs its intended function when operated over the full environmental conditions to which it has been declared to comply.
Consideration should be given to the extremes at which it may be subjected during an incident or accident. These tests can be undertaken during the specified RTCA DO- 160F (or later revision) tests or separately. Analysis may be substituted for a test where its use can be shown to produce equivalent evidence of compliance. The system should be capable of recording up to 2 hours of image or acoustical data and 6 hours of Page MG 6 - 16
aircraft parameter data. The applicant determines and maintains the data stream format and parameter documentation, including which parameters are recorded, how often the parameters are recorded, the bit resolution of each parameter, the operational range of each parameter, and the conversion algorithm from decimal units to engineering units. The Design Assurance Level (DAL) for an FDMS that is required by an operating regulation is DAL "D." RTCA DO-178B (or later revision) provides acceptable software development standards, which in this case would be for DAL "D" software. RTCA DO-254 (or later revision) provides acceptable airborne electronic hardware (AEH) development standards, which in this case would be for DAL "D" AEH.
Note: The duration between data downloads for the promotion of operational safety within a FOQA program is directly correlated to the recording capabilities of the system installed.
(ii) Recording. The FDMS should be capable of capturing and recording any combination of the following parameters in order to monitor the aircraft's state, condition, and system performance:
• Positioning system time.
Positioning system latitude.
Positioning system longitude.
Positioning system altitude.
Positioning system error.
Altitude.
Heading.
Pitch attitude.
Pitch rate.
Roll attitude.
• Roll rate.
• Yaw rate.
• Air speed.
• Ground speed.
• Ambient acoustic data.
• Engine parameters.
• Main rotor revolutions per minute.
• Transmission ambient audio.
• Any other parameters deemed appropriate by the operator.
Notes: Parameters may be recorded directly or deduced from recorded data from the FDMS. Additional guidance on parameters can be found in EUROCAE specification ED-155.
Recording individual pilots, using hot microphones, on separate pilot audio channels can provide useful information in the investigation of incidents and accidents.
Page MG 6 - 17
AC 27-1B, Chg 4 2/27/2014
(ili) Maintenance. The maintenance requirements to ensure the serviceability and continued airworthiness of the FDMS are typically established by the equipment manufacturer and installer. These maintenance instructions should be included in the applicable helicopter model instructions for continued airworthiness.
*(14) Radio Altimeter (RAD ALT). RAD ALTs installation is required. Its information display must be in the pilot's primary field of view in all helicopters operating under a part 135 certificate. The minimum performance requirements for an FAA approved RAD ALT system can be found in TSO-C87.
(15) Other EMS Equipment. These items of equipment installed for the EMS mission are considered optional equipment and should be operated to ensure they function properly. This evaluation would normally be done by someone knowledgeable about the particular type of equipment, since correct operation of the equipment is essential to a valid determination that the required rotorcraft systems are not being interfered with. This includes all removable pieces of medical equipment that are used for patient care. The primary purpose of the evaluation of this equipment is to emphasize the possibility of any interference between operation of the EMS equipment and the systems whose installation is required for safe operation of the aircraft, the adequacy of the installation provisions, and assurance that failure modes will not result in a hazardous condition for the rotorcraft.
(16) Miscellaneous. The following areas are not peculiar to EMS installations;
however, their significance is enhanced by the complexity of an EMS installation.
(i) Compatibility. Many EMS installations are a collection of several STCs and may also include some FAA field approvals. For this situation, it should be shown that the overall installation provides for safe operation of the aircraft. Operation of a search light, if included, should be addressed since in using this system it can be difficult to keep light from interfering with the pilot view.
(il) Electrical Load Analysis. An electrical load analysis should be conducted, and additional guidance is available in AC 27-1 MG 1. If the analysis indicates the generators) can be overloaded, appropriate measures should be taken to account for the problem. In some instances (e.g., in a visual flight rules (VFR) approved rotorcraft), a placard that specifies certain operating limitations may be satisfactory, while in other instances (e.g., in an instrument flight rules (IFR) approved rotorcraft), an electrical interlock may be in order. In general, if the amount of overload is relatively small and the rotorcraft is not an IF-approved rotorcraft, the placard solution will probably be satisfactory, whereas if the amount of possible overload is significant, it is more likely that an interlock scheme will be necessary.
(ili) Aircraft Grounding. It should be emphasized in an appropriate place in the STC data (e.g., RFM, maintenance information) that any time the EMS systems are Page MG 6 - 18
being operated or serviced (e.g., oxygen) on the ground, the rotorcraft itself must be grounded.
(iv) Electrical Outlets. All electrical outlets provided in the cabin should be the three-prong grounded type. When not in use, these outlets should be suitably protected against the entry of fluids.
(v) Placards. All medical outlets (e.g., air, oxygen, vacuum) should be placarded. Electrical power outlets should be placarded for type of voltage and amperage capacity. A placard stating "No Smoking When Oxygen Is In Use" should be included. Other placards would include information appropriate to the oxygen system, operation of special controls, and so forth.
(vi) Equipment in Cargo and Baggage Compartments. When components are added to the baggage compartments, provisions should be made to protect the system components due to shifting cargo. In addition, when oxygen components are installed, the compartment should be placarded against the storage of oil or hydrocarbons. A smoke detector is recommended for a compartment if oxygen cylinders are installed in a closed, non-accessible compartment. Also, the cargo weight limitations placard should be changed. AC 27-1, section 27.787 pertains to cargo and baggage compartments.
(vil) Safety Assessment. When installing any new equipment or modifying existing equipment, a safety assessment must be made to assure the FAA that all possible failure conditions that could occur from these changes have been adequately addressed to show compliance to the regulations.
(17) Equipment Substitution. The EMS modification that is presented for approval will contain specific items of equipment, and the approval will make reference to this equipment. If other equipment (e.g., new model, manufacturer) is to be substituted, then an evaluation should be made to ensure the substitute equipment is also satisfactory. This evaluation would normally consist of comparing the attachment means, design features, failure modes, specifications, and operation of the two units.
The purpose of the evaluation is to ensure there are no differences that have an adverse effect on the airworthiness of the installation. Other differences would not be considered significant. Specific seats and litters are generally approved as a part of the EMS configuration. Substitution may be approved in accordance with the standards.
c. Related Regulations and References.
(1) Regulatory Sections. 14 CFR 27.337, 27.471, 27.561, 27.773, 27.783, 27.785,
27.807, 27.831, 27.853, 27.1301, 27.1309, 27.1353, 27.1357, 27.1365, 27.1367,
27.1411, 27.1413, 27.1431, 27.1557(d), 27.1561, 27.1581(a)(2), 27.1583(d), 27.1585,
27.1589, part 91, and part 135.
(2) Other References. Refer to the current version of each document.
Page MG 6 - 19
AC 27-1B, Chg 4 2/27/2014
(i) Helicopter Association International, Emergency Medical Services Recommended Guidelines.
(ii) National Highway Traffic Safety Administration, Air Ambulance Guidelines.
(iii) AC 23-2, Flammability Tests.
(iv) AC 43.13-2, Acceptable Methods, Techniques, and Practices-Aircraft Alterations.
(v) AC 135-14, Emergency Medical Services/Helicopter (EMS/H).
(vi) Oxygen Equipment for Aircraft, Society of Automotive Engineers Aerospace Information Report No. 825.
(vii) MIL-D-19326G, Design and Installation of Liquid Oxygen Systems in Aircraft, General Specification for Military Specification.
Page MG 6 - 20
INTEGRAL ASSEMELY 2/27/2014 (Most Cases) FIGURE AC 27.MG 6-1 TYPICAL LIQUID OXYGEN SYSTEM OVERBOARD
VENT LINE PRESSURE BUILDUP/VENT
RELIEF VALVE TO
GAUGENE C SYSTEM
PRESSURE
RELIEF CLOSING PRES, LIQUID VALVE
CONTAINER TO
PATIENT - CON TROI VALVE REGULATOR EVAPO PRES.
RATING COIL BUILDUP COIL CHECK I FILL VALVE
ASSEMBLY VALVE Page MG 6 - 21 EMERGENCY AC 27-1B, Chg 4
SHUTOFF VALVE
Page MG 6 - 22 INTEGRAL ASSEMBLY AC 27-1B, Chg 4 (Most Cases) FIGURE AC 27.MG 6-2 TYPICAL LIQUID OXYGEN SYSTEM - USING OVERBOARD
VENT LINE
PRESSURE
RELIEF FILL, BUILDUP
AND VENT VALVE TO
COMBINATION VALVE PRESSURE PRES
RELIEF CLOSING LIQUID
VALVE CONTAINER FLOW
CONTROL REGULATOR
WALVE PRES
EVAPO- BUILDUP
RATING COIL
COIL OHECK
ASSEMBLY VALVE EMERGENCY 2/27/2014
SHUTOFF VALVE
CHAPTER 3
AIRWORTHINESS STANDARDS
NORMAL CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 27 MG 7. HOUR METERS OR TIME-IN-SERVICE RECORDING DEVICES IN ROTORCRAFI.
a. Explanation.
(1) Time in service is a required maintenance record according to § 91.173(a)(2)(i). Manual recording (rather than automatic) of the rotorcraft time in service is required whether or not the rotorcraft is equipped with a time accumulating device or hour meter. Time recording devices (hour meters) are optional and not required by the maintenance or operating rules or as a part of the aircraft type design.
(2) Time in service is defined in FAR Part 1 as follows: " ...with respect to maintenance time records, means the time from the moment an aircraft leaves the surface of the earth until it touches it at the next point of landing."
(3) The allowable total time in service, that is, service life or retirement time, has been typically specified for critical rotorcraft components that were subject to fatigue. The life was determined under § 27.571 or its predecessor standards.
(4) Hour meters may be mechanical and a part of a recording tachometer.
They may be electrical and activated by the main battery switch, engine oil pressure, or some other source. These devices may not record "real time" when an engine or rotor is at idle speed or when deactivated by an oil pressure sensor switch. This real time omission is not considered significant in the life of a rotorcraft if the design and installation of the sensor minimizes this omission. However, for an individual aircraft that is used extensively, the omission of real time may become significant, and the time in service should be adjusted prior to recording the time in service in the aircraft records.
b. Procedures.
(1) Time in service must be recorded in the aircraft maintenance records.
(2) Hour meters are used to accumulate operating time. The meters may be activated by various sensors/switches. The type design data should include hour meter installation data including the sensor/switch location and any adjustment or "rigging" information to ensure proper installation.
(3) One acceptable installation of a sensor/switch would activate the hour meter whenever the collective pitch control was moved from the lowest pitch stop position.
Page MG 7 - 1
That is, real time in service was recorded whenever the collective control was moved from the "full down" collective position.
The location of the sensor is important to ensure time in service is measured when collective lift or thrust of the main rotor is desired.
Time in service will not be accumulated in flight whenever full down (or near full down) collective is used, such as in some phases of autorotation. However, collective controls are used to maintain proper rotor speed during autorotation.
(iii) Time in service should be corrected (increased) for individual aircraft whenever "full down" collective is used a significant amount.
(4) If the hour meter installation is significantly deficient in recording real time in service, corrections must be made or the installation changed to eliminate any significant deficiency in the correct time in service.
Page MG 7 - 2
CHAPTER 3
MISCELLANEOUS GUIDANCE (MG) AC 27 MG 8. (Amendment 27-30) SUBSTANTIATION OF COMPOSITE
ROTORCRAFT STRUCTURE
a. Reference. 14 CFR §§ 27.305, .307, .571, .603, .605, .609, .610, .611, .613, .629, .923, .927, .931, .1529 and Appendix A.
b. Purpose. These substantiation procedures provide a more specialized supplement to the general procedures outlined by AC 20-107A, "Composite Aircraft Structure." These procedures address substantiation requirements for composite material system constituents, composite material systems, and composite structures common to rotorcraft. A uniform approach to composite structural substantiation is desirable, but it is recognized that in a continually developing technical area which has diverse industrial roots, both in aerospace and in other industries, some variations and deviations from the procedures described herein will be both necessary and acceptable.
Significant deviations from this material should be coordinated in advance with the FAA/AUTHORITY.
c. Special Considerations. Since rotorcraft structure is configured uniquely and is inherently subjected to severe cyclic stresses, special consideration is required for the substantiation of all rotorcraft structure, including composites. This special consideration is necessary to ensure that the level of safety intended by the current regulations is attained during the type certification process for all structure with special emphasis on composite structure because of its unique structural characteristics, manufacturing quality and operational considerations, and failure mechanisms.
d. Background.
(1) Historically, rotorcraft have required unique, conservative structural substantiation because of unique configuration effects, unique loading considerations, severe fatigue spectrum effects, and the specialized comprehensive fatigue testing required by these effects. Rotorcraft structural static strength substantiation for both metal and composite structure is essentially identical to that for fixed wing structure once basic loads have been determined. However, rotorcraft structural fatigue substantiation for composites is significantly different from fixed wing fatigue substantiation. Since AC 20-107A, as developed, applies to both fixed wing aircraft and rotorcraft; it, of necessity, was finalized in a broad generic form. Accordingly, a need to supplement AC 20-107A for rotorcraft was recognized during type certification programs. One significant difference in traditional rotorcraft fatigue substantiation programs and fixed wing fatigue programs is the use of multiple component fatigue tests for rotorcraft programs rather than just one full-scale test. Also, constant amplitude, accelerated load tests are typically used rather than spectrum tests because of the high frequency loads common to rotorcraft operations. These rotorcraft fatigue tests have Page MG 8 - 1
AC 27-1B, Chg 4 3/31/2014 traditionally involved the generation of stress versus life or cycle (S-N) curves for each critical part (most of which are subjected to the cyclic loading of the main or tail rotor system) using a monotonic (sinusoidal) fatigue spectrum based on maximum and minimum service stress values. Unless configuration differences or flight usage data dictate otherwise, the monotonic fatigue spectrum's period is typically based on six ground-air-ground (GAG) cycles for each flight hour of operation. The S-N curves for the substantiation of each detailed part are typically generated by plotting a curved line through three data points (reference AC 27-1B, Chg 1, MG 11, "Fatigue Evaluation of Rotorcraft Structure"). The three data points selected are a short specimen life (low-cycle fatigue), an intermediate specimen life, and a long specimen life (high-cycle fatigue). Each raw data point is generated by monotonically fatigue testing at least two full-scale specimens (parts) to failure or run out for each data point on the S-N curve.
The raw data point values are then reduced by an acceptable statistical method to a single value for plotting to ensure proper reliability of the associated S-N curve.
Order 8110.9, "Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and Other Power Transmission Systemsi and AC 27-1B, Chg 1, MG 11, "Fatigue Evaluation of Rotorcraft Structure," contain comprehensive discussions of the S-N curve generation process. The rotorcraft S-N curve process contrasts sharply with the fixed wing process of using a single full-scale fatigue article (usually an entire wing or airframe, which constitutes a single full-scale assembly data point), generic material or full-scale assembly S-N data (e.g., MIL-HDBK-5 for metals, MIL-HDBK-17 for composites, or AFS-120-73-2 for full-scale assemblies), a non-monotonic spectrum and relatively large scatter factors to verify or determine the design fatigue life of the
(2) Also, rotorcraft have employed and mass-produced composite designs in primary structure (typically main and tail rotor blades) since the early 1950's. This was 10 or more years before composites were type certificated for primary fixed-wing structure in either military or civil aircraft applications (with some notable limited production exceptions, such as the Windecker fixed wing aircraft). In any case, the early 1950 period was well before a clear, detailed understanding of composite structural behavior (especially in the areas of macroscopic and microscopic failure mechanisms and modes) was relatively common and readily available in a usable format for the average engineer working in this field. It also predated the initial issuance of AC 20-107. Currently, much composite design information is proprietary, either to government, industry, or both, and many data gathering methods have not been completely standardized. Consequently, a significant variation from laboratory to laboratory in material property value determination methods and results can exist. The early rotor blade designs (as well as current designs) are by nature relatively low strain, tension structure designs. Also, by nature, these designs are not damage or flaw critical. Thus by circumstance as much as design, early composite rotor blade and other composite rotorcraft designs incorporated an acceptable fatigue tolerance level of safety. In the 1980's, more test data, analytical knowledge, and analytical methodology became available to more completely substantiate a composite design. Current 14 CFR parts 27 and 29 contain many sections (reference paragraph a above) to be considered in substantiating composite rotorcraft structure, but this advisory material is needed to Page MG 8 - 2
supplement the general guidance of AC 20-107A by providing specific rotorcraft guidance for obtaining consistent compliance with 14 CFR sections applicable to rotorcratt.
e. Definitions. The following basic definitions are provided as a convenient reading reference. MIL-HDBK-17, and other sources, contain more complete glossaries of definitions.
(1) AUTOCLAVE. A closed apparatus usually equipped with variable conditions of vacuum, pressure and temperature. Used for bonding, compressing or curing materials.
(2) ALLOWABLES. Both A- basis and B- basis values statistically derived and used for a particular composite design.
(3) BALANCED LAMINATE. A composite laminate in which all laminae at angles other than 0° occur only in + pairs (not necessarily adjacent).
(4) A-BASIS ALLOWABLE. The "A" mechanical property value is the value above which at least 99 percent of the population of values is expected to fall, with a confidence of 95 percent.
(5) B-BASIS ALLOWABLE. The "B" mechanical property value is the value above which at least 90 percent of the population of values is expected to fall, with a confidence of 95 percent.
(6) BOND. The adhesion of one surface to another, with or without the use of an adhesive as a bonding agent.
(7) COCURE. The process of curing several different materials in a single step. Examples include the curing of various compatible resin system pre-pregs, using the same cure cycle, to produce hybrid composite structure or the curing of compatible composite materials and structural adhesives, using the same cure cycle, to produce sandwich structure or skins with integrally molded fittings.
(8) CURE. To change the properties of a thermosetting resin irreversibly by chemical reaction; i.e., condensation, ring closure, or addition. Cure may be accomplished by addition of curing (crosslinking) agents, with or without catalyst, and with or without heat.
(9) DELAMINATION. The separation of the layers of material in a laminate.
(10) DISBOND. A lack of proper adhesion in a bonded joint. This may be local or may cover a majority of the bond area. It may occur at any time in the cure or subsequent life of the bond area and may arise from a wide variety of causes.
Page MG 8 - 3