AC 29-2C, Chg 4 2/27/2014 (C) “Provisions Only” Considerations. In some instances systems are approved that only include provisions for a supply system consisting of the high pressure cylinders, regulators, and their associated lines and fittings. In these instances, a placard should be provided that refers to a supply system that is considered satisfactory for the remainder of the installation. An example of an acceptable placard for this situation is:
Oxygen Supply System must be in accordance with the requirements given in STC SH _________. Deviations to the configuration specified must be evaluated and approved by the Manager (include reference to the appropriate FAA ACO).
(11) Medical Communication Equipment. This equipment is provided to allow for communication between the rotorcraft and ground medical personnel. It includes voice communication and may also include telemetry equipment for the transmission of graphic data. It should be demonstrated that this equipment functions properly and the range at which this determination was made recorded in the project file. The functional demonstration should include a 360° turn (clockwise and counterclockwise) to ensure no significant sections of signal blanking exist. The remainder of the emphasis on this equipment should be to ensure that operation of this equipment does not interfere with normal operation of any rotorcraft systems whose installation is required for safe operation of the rotorcraft.
(12) Cabin Lighting. EMS interiors normally include higher intensity cabin lighting than other interiors. This lighting capability should be carefully evaluated to ensure it does not interfere with operation of the rotorcraft. In some installations a special curtain is required to separate the cockpit from any interference by the lighting. The FAA project file should document the approach of how this evaluation was conducted. See paragraph b.(6) for other curtain considerations.
*(13) FDMS. If required under an oper ating regulation, an FDMS (not to be confused with a flight data recorder (FDR) certificated under § 29.1459) may be comprised of a system or combination of systems that record a helicopter’s flight performance and operational data. An FDR certificated under § 29.1459 and the appropriate operating rules would be acceptable to meet this requirement; however, an FDMS would not be adequate to meet the § 29.1459 requirement for an FDR. The FDMS should be capable of capturing 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 Page MG 6 - 16
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 in stalled 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 Organizati on of 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 Systems.”
Equipment testing should be conducted to the categories most applicable to the aircraft type, and the location of the equipment to be installed. The equi pment 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 ma y 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 aircraft parameter data. The applicant determines and maintain s 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 la ter 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 fo r the promotion of operational safety within a FOQA program is directly correlated to the recording capabilities of the system installed.
Page MG 6 - 17
AC 29-2C, Chg 4 2/27/2014
(ii) Recording. The FDMS shoul d 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.
(iii) 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 instruction for continued airworthiness.
*(14) Radio Altimeter (RAD ALT). RAD ALTs insta llation 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 mini mum performance requirements for an FAA approved RAD ALT system can be found in TSO-C87.
(15) Other EMS Equipment. These items of equipm ent 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 so meone knowledgeable Page MG 6 - 18
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 co llection of several STCs and may also include some FAA field approvals. Fo r 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.
(ii) Electrical Load Analysis. An electrical load analysis should be conducted, and additional guidance is available in AC 29-2 MG 1. If the analysis indicates the generator(s) 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 t he amount of overload is relatively small and the rotorcraft is not an IFR-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.
(iii) 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 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 s hould 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, etc.
(vi) Equipment in Cargo and Baggage Compartments. When components are added to the compartment, revisions should be made to protect the system components Page MG 6 - 19
AC 29-2C, Chg 4 2/27/2014 due to shifting cargo. In addition, w hen 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 compartment weight limitations placard should be changed. AC 29-2, section 29.787 pertains to cargo and baggage compartments.
(vii) Safety Assessment. When installing any new equipm ent 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 no rmally 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. Ot her differences would not be considered significant. A specific seat and litter design is approved as a part of the EMS configuration. Substitutions may be approved in accordance with the standards.
c. Related Regulations and References.
(1) Regulatory Sections. 14 CFR 29.337, 29.471, 29.561, 29.773, 29.783,
29.785, 29.803, 29.805, 29.809, 29.811, 29.813, 29.815, 29.831, 29.853, 29.1301,
29.1309, 29.1353, 29.1357, 29.1411, 29.1431, 29.1557(d), 29.1583(d), 29.1585,
29.1589, part 91, and part 135.
(2) Other References Refer to the current version of each document.
(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.
Page MG 6 - 20
(vii) MIL-D-19326G, Design and Installation of Liquid Oxygen Systems in Aircraft, General Specification for Military Specification.
Page MG 6 - 21
AC 29-2C, Chg 4 2/27/2014
BUILDUPIVENT VALVE IVA INTEGRAL ASSEMBLY (Most Cases) TO GAUGING SYSTEM LIQUID CONTAINER PRESSURE RELIEF PRES CLOSING VALVE PRES. BUILDUP COIL SHUTOFF VALVE
EVAPO- RATING ASSEMBLY
OVERBOARD VENT LINE RELIEF FLOW REGULATOR CONTROL
TO PATIENT FIGURE AC 29.MG 6-1 TYPICAL LIQUID OXYGEN SYSTEM Page MG 6 - 22
FILL, BUILDUP AND VENT VALVE NTEGRAL ASSEMBLY (Most Cases) TO GAUGING Z SVSTEM LIQUID CONTÂINER PRESSURE RELIEF PRES. CLOSING VALVE PRES BUILDUP COIL EMERGENCY SHUTOFF VALVE
EVAPO- RATING COIL ASSEMBLY OVERBOARD VENT LINE PRESSURE RELIEF • REGULATOR FIGURE AC 29.MG 6-2 TYPICAL LIQUID OXYGEN SYSTEM - USING
COMBINATION VALVE Page MG 6 - 23
CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 7. STRUCTURAL CONDITION INDICATORS.
a. Related Sections. § 29.301 - Loads; § 29.305 - Strength and Deformation;
§ 29.571 - Fatigue Evaluation of Flight Structure; § 29.1301 - Function and Installation;
§ 29.1309 - Equipment, Systems and Installations; § 29.1321 - Arrangement and Visibility; § 29.1322 - Warning, Caution, and Advisory Lights; § 29.1355 - Distribution System; § 29.1503 - Airspeed Limitations: General; and § 29.1529 - Instructions for Continued Airworthiness.
b. Background.
(1) Structural condition indicators have been used on rotorcraft for several years in two main programs: as part of the basic type design and as part of airworthiness directive (AD) action. When approved as part of the basic type design, only limited “credit” has been given for the installation of structural condition indicators;
i.e., components provided with a structural condition indication system were required to be designed to § 29.571 “safe-life” criteria considering the structural condition indicator system inoperative. So- called “nonhazard” approvals were granted. When used as part of the mandatory actions of ADs, structural condition indicators have had a degree of “credit” recognized, primarily in the recognition of “fail-safety” provided by the indicator system.
(2) Since structural condition indicators have been used during both original type design and AD issuance, and since there is movement toward increased damage tolerance in rotorcraft design, policy concerning condition indicator use is considered appropriate.
c. At present, the use of structural condition indicators alone on new type designs is not considered an acceptable substitute for providing the necessary safe life for each component. However, areas which may be considered when approving these indicators for fail-safety credit are delineated in the following paragraphs.
d. What, how, when, where, and who of structural condition indicators.
(1) Indication of what?
(i) Previous structural condition indicators have primarily been used for crack detection. Several types of through- the-thickness crack detection systems are currently in use. Two types which detect changes in pressure in an instrumented chamber due to gas movement through a cracked wall are known as the blade Page MG 7 - 1
inspection method (BIM) system and the integral spar inspection system (ISIS). These systems can only detect full-depth cracks which are large enough to allow loss (of gain) of pressure from the instrumented chamber. This presents a limitation since full -depth cracks may be fast growing before detection. Another through-t he-thickness crack method is a pressurized, dyed fluid or oil system to detect through cracks in specially designed bolts (NASA patent), spindles, pins, or other closed chamber mechanical equipment.
(ii) Surface cracks can be found by systems such as surface-mounted crack detection wires. These systems would allow a greater safe crack growth period for assuring safe landing after detection than the through-crack-detection systems, but they have been used little in operations because of significant limitations; e.g., complexity of installation, durability problems, limited areas of coverage, and strain level limitations.
(iii) Some aircraft have had mast moment indicators or other load indicators to help prevent the pilot from inadvertently applying a high load to the instrumented system or to help the pilot reduce the load by control movements. These load indicators only indirectly give indications of structural condition; therefore, only limited “credit” is allowed for this use. “Credit” is limite d in that the fatigue life substantiations of § 29.571 should consider a reasonable number of excursions into the higher ranges established for the load indicator, and special inspections, rework, or replacement instructions should be provided for any strength degradation associated with high range excursions.
(2) How indicated?
(i) Current BIM systems use two types of indicators. The visual blade inspection method (VBIM) uses a gauge mounted on the blade which must be read visually by maintenance personnel while the aircraft is parked. The cockpit blade inspection method (CBIM) uses lights mounted in the cockpit which may be monitored by the crew. Other pressurized chambers have used dyes or oils to improve visual inspection effectiveness. Mast moment indicators and other load indicators use instruments with marked ranges and needles.
(ii) No specific types of load indicators are required by the FAA/AUTHORITY but the type used should be evaluated for accuracy, readability, and overall effectiveness. Paragraphs AC 29 MG 7(e) and (f) cover, in more detail, the use of structural condition indicators.
(3) When indicated? Structural condition indicators are used before flight, during flight, and for normal maintenance inspections. Paragraphs AC 29 MG 7 (e) and
(f) contain guidance for cockpit-mounted instruments which are monitored during flight.
Indicators used for normal maintenance inspections are the preferred type since they can be scheduled to allow the most effective use of available maintenance personnel of well-equipped maintenance facilities and of parts available.
Page MG 7 - 2
(4) Where indicated? Indications on the component are provided by VBIM systems and by systems utilizing dye or colored oil leakage. Cockpit -mounted lights and gauges may be used for certain critical structures which require frequent, but simple, checks. Maintenance panel locations (cabin, equipment bay, etc.) are the preferred locations for use in routine maintenance.
(5) Who reads indicators? The flightcrew, of necessity, monitors indicators mounted in the cockpit for use during flight. Gauges with ranges of values representing mast bending moments or other structural loads are monitored by the flightcrew, as necessary, to reduce or to prevent control operations from imposing excessive loads or to prevent too many high load applications. Maintenance personnel are generally responsible for reading component-mounted indicators and for monitoring indicators which are mounted on maintenance panels. The before- flight checks may be conducted by maintenance personnel or by flightcrew in certain cases (i.e., cockpit-mounted gauges or “push-to-test” checks).
e. Actions required by indicators .
(1) On-ground indications. Indications noted on the ground should be followed by a functional check of the indication system as provided for by its design. If indications persist after the system has been checked and found to be functional, further inspection of the affected component(s) should be conducted for damage assessment.
Any damage found as a result of the detailed inspections should be repaired or replaced as appropriate.
(2) In-flight indications.
(i) Indications used for in-flight monitoring have in the past been used for two main reasons: to provide a structural load display (such as mast bending moment) and to help resolve a service problem (CBIM systems have been used to supplement conventional inspection methods in blind areas).
(ii) Structural load display systems should not be used instead of correcting deficient designs. Structural load display systems are appropriate for use in locating control positions, such as the cyclic stick, under transient conditions such as slope landings and hover in sidewinds, but structural load display systems are not considered appropriate for routine operations such as climbout or cruise with constant attention required by the flightcrew. If the load indicator provides a needed tool to the pilot in limited types of operations and does not significantly add to pilot workload otherwise, its use can be considered.
(iii) In the past, certain service problems have been solved by adding in-flight indicators such as CBIM systems. When retrofit of the affected structure is impossible or impractical, and when conventional inspection techniques are shown to be inadequate by themselves, CBIM or similar systems may be the only practical Page MG 7 - 3
solution, despite the increase in pilot workload and the potential for problems caused by overreaction by the pilot to a structural fault indication. When used for correction of service difficulties, the structural condition indicator system should be accompanied by clear, concise crew directions to prevent possible catastrophic overreaction. Load reduction measures such as rotor speed changes, airspeed reductions, altitude changes, etc., should be clearly provided, if needed. Crack propagation time from indication should be sufficient to allow continued safe flight to a safe landing area. For new designs, CBIM or similar systems which add to the pilot’s workload are considered inappropriate. Proper redesign to provide the needed safe life, fail -safety, and inspectability is considered the appropriate action.
f. Complementary considerations of structural condition indicator use.
(1) Two basic programs are commonly used for approval of structural condition indicators. Basic type certification procedures are used for mast moment indicators and similar systems, and AD’s (with appropriate type design changes) are used for CBIM systems which require pilot attention and corrective action when an indication of a structural fault is detected.
(2) The fatigue substantiation required by §§ 27.571 and 29.571 should consider a conservative number of excursions into the high load range monitored by a structural condition indicator such as a mast moment indicator. Static strength should not be adversely affected by a single excursion into the high load range monitored by the indicator.
(3) Complementary design provisions should accompany the use of a structural condition indicator system. Redundancy of load paths and inspection systems and indicator system failure analyses should be provided, as necessary, to meet the requirements of § 29.1309. The life remaining after the indicator system detec ts a structural failure should be calculated (with test verification), and compatible inspection and/or overhaul programs should be provided.
(4) The FAA/AUTHORITY approval of a structural condition indicator system requires evaluation by the airframe, systems and equipment, and flight test specialists.
The airframe specialist has the responsibility to review effects of structural condition indicator system use on aircraft loads, strength and deformation, and structural fatigue evaluation as well as the instructions for continued airworthiness. The systems and equipment specialist needs to evaluate the system for function and installation as well as the reliability requirements of § 29.1309. Flight test evaluation of the instruments’ arrangement and visibility, effect on crew workload, and possible changes for RFM is also needed. Care should be exercised to assure that responsibilities are not given to the flightcrew which would be more appropriately handled by a redesign or by the maintenance personnel. Early coordination between all specialists is necessary to prevent delays from last minute design changes.
Page MG 7 - 4