783. 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?
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 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-the-thickness crack method is a pressurized, dyed fluid or oil system to detect through cracks in specially 1230 Par 783
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.
(™) 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 limited 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?
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 783(e) and (t) 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 783(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.
(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 Par 783 1231
АС 29-2B 7/30/97 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.
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.
(її) 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.
(lI) 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 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 1232 Par 783
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 detects 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.
Par 783 1233
784. EULL AUTHORITY DIGITAL ELECTRONIC CONTROLS (FADEC) FOR
INSTALLATIONS WITH CATEGORY A ENGINE ISOLATION
a. Explanation. The advent of "microprocessor technology" has resulted in rotorcraft engine controls being implemented by digital process control rather than by conventional means. These digital, processor-based full authority engine controls offer many performance advantages (such as isochronous governing) which were not feasible with conventional technology pneumatic or hydromechanical controls.
Because of the incorporation of this advanced technology, some additional considerations must be made of the engine installation to ensure regulatory compliance.
b. Procedures. The following is a discussion of some special attention areas for a Part 29 Category A FADEC engine installation. Paragraph 621(b)(3)(ii)(D) of this AC contains a general definition of what constitutes a "full authority" control.
(1) Software Qualifications.
Paragraph 621(f) contains a general discussion of the use of the recommended RTCA document DO-178A which is utilized for the approval of system software. When utilizing that document, one might arrive at the conclusion that the engine control as a required function is essential and thus level 2 software would be appropriate for the control. However, for this level 2 category software, errors are presumed to exist, and since a software error in a full authority control could result in simultaneous unacceptable malfunctions in all engines, the provisions of § 29.1309(b)(2)(i) for continued safe flight and landing and the engine isolation rule, § 29.903(b), would generally preclude the use of this classification.
System designs which provide redundant distinctive software or an alternate technology control which is automatically selected and meets all of the minimum regulatory requirements would reduce the impact of software errors and may allow the level 2; i.e., essential software classification. At level 1, it is accepted that the software is sufficiently error free that the software does not require further verification in the installation evaluation.
(2) Lightning Strike Protection. A complete discussion of an acceptable method of demonstrating that the FADEC, as installed, is adequately protected against the catastrophic effects of lightning is contained in Paragraph 621b(3) of this AC.
(3) Electrical Power System Considerations.
Normal Operation. The system should be evaluated with all power sources operating normally. If additional power source capability is being provided that is above the minimum required for certification, a certain portion of the evaluation 1234 Par 784
7/30/97 АС 29-2B should be conducted while operating in the minimum configuration. The minimum power source configuration should consider the provisions of § 29.903(b).
(її) Malfunction Conditions. Beginning with the minimum configuration that is required for certification, electrical power system malfunctions should be introduced and the impact on continued FADEC operation determined.
(iii) Circuit Protection Location. The circuit protective devices for the FADEC should be located in the cockpit such that they can be readily reset or replaced in flight. The operation of the FADEC system is considered to be essential to safety in flight. Reference § 29.1357(d).
(iv) System Separation. On multiengine applications, each system should be separated from the other system to the maximum extent practical. Wiring should be routed separately. Power should be taken from independent busses and grounds, and system components should be independent of one another.
(V) Periodic Checks. Where periodic checks are appropriate, they should be made at reasonable intervals. This would normally range from preflight checks for certain items of greater concern to a tie-in with normal aircraft maintenance intervals for other items. If a crew check is specified, it should be evaluated to ensure it is a reasonable check. If items to be checked are located in an area that can be covered by interior upholstery, for example, a crew check would not be considered reasonable, and further design considerations may be in order.
(4) Powerplant Insta lation Considerations.
A demonstration of compliance with § 29.901(c) would generally include a failure mode and effects analysis (FMEA) of the powerplant systems as installed. When a FADEC is utilized, the analysis would consider the control's failure modes, the installed engine reaction, the effect on the aircraft, and the crew response to the situation. Combinations of undetected failures should be considered. Engine failures which may be escalated in severity by the FADEC's response to the initial failure should be analyzed. Potentially hazardous failures should be evaluated during flight testing. The requirements of §§ 29.903(b)(2) and 29.1309(b)(2)(i) should be reviewed in determining acceptability of failures.
(ії) Section 29.903(b)(2), Category A engine isolation, is intended to ensure that a failure will not prevent the continued safe operation of the remaining engine(s) or require immediate action of the crew to ensure continued safe operation.
The FADEC's of the individual engines should be independent. Where communication between FADEC's is required (for example, for torque sharing), care should be exercised to ensure that failures which may occur will not result in a power loss to the extent that total power available is less than would be available under OEI conditions.
The no-required immediate-crew-action provision would preclude credit for manually Par 784 1235
selected or operated backup systems in meeting the § 29.903(b) rule. These unrequired backup systems, which may offer the advantage of get-home multiengine capability rather than forced OE operation, would be evaluated on a no hazard basis.
Section 29.939, turbine engine operating characteristics, intends a flight investigation to ensure that no adverse characteristics are present to a hazardous degree during normal and emergency operation in the allowed flight envelope. The evaluation should include assessment of the minimum FADEC system certification configuration; i.e., the minimum proposed by the applicant to meet Part 29 requirements. Reduced capabilities (e.g., restrictions on normal collective movements, limited aircraft maneuvers, etc.) may be acceptable for degraded FADEC modes or backup systems not required to meet Part 29 requirements if those degraded capabilities are reasonable and not hazardous as determined by flight evaluation. The restrictions should be specified in the flight manual.
(iv) The rotorcraft with FADEC engines must of course meet all of the Part 29 requirements, but the areas described herein are those which deserve special attention.
785. AGRICULTURAL DISPENSING EQUIPMENT INSTALLATION.
NOTE: This paragraph has been extensively revised and expanded to clarify the restricted category certification of agricultural dispensing equipment installations on rotorcraft.
a. Explanation. In the early development of the rotorcraft one of its primary usages was agricultural operation. The FAA recognized that the existing requirements, which were designed primarily to establish an appropriate level of safety for passenger-carrying aircraft, imposed an unnecessary economic burden and were unduly restrictive for the manufacture and operation of aircraft intended only for use in rural, sparsely settled areas. Therefore, a special document that established new standards for agricultural dispensing equipment and other special purposes was developed. Restricted Category CAM 8 became effective October 11, 1950.
(1) During the recodification of 1965, CAR 8 ceased to exist as a regulatory basis and selected portions addressing certification were incorporated into FAR 21.
While the specific standards in CAR 8 were not changed substantially when adopted into FAR 21, the less restrictive philosophy of CAM 8 and the policy material that was stated in the preamble to CAM 8 were not clearly conveyed.
(2) Advisory material published in 1965 and revised in 1975, summarized the information contained in the advisory portions of CAM 8. This new advisory material indicated that the CAM advisory material would be applicable to the related FAR's.
Unfortunately, this document specified that CAM 8 could be used in conjunction with 1236 Par 784
certain FAR's for restricted category certification of small agricultural airplanes only.
Rotorcraft were omitted.
(3) A survey of restricted category rotorcraft projects related to agricultural modifications indicates that the CAM 8 philosophy was interpreted to allow the use of AC 43.13-2A structural criteria for most STC's issued through the early 1980's. Since then more restrictive guidance based on CAR 6 and FAR 27 requirements has been applied by some ACO's to several STC applications. Since the more restrictive guidance imposed a significant economic burden on the industry, the HAI requested a meeting with the FAA during the 1990 annual convention in Dallas. As a result of the meeting, an Action Notice to clarify the interpretation of FAR 21.25(a)(1) for restricted category aircraft has been issued.
(4) The following advisory material is a result of a reassessment of past and present policy.
b. Procedures. The certification basis for agricultural dispensing equipment in the restricted category is FAR 21.25(a)(1) as interpreted by Action Notice 8110.22. The accountable Directorate guidance for the substantiation requirements for rotorcraft is as follows:
(1) Substantiation of the agricultural dispensing system hoppers or spray tanks to the load factors provided in Figure 785-1 provides for proof of structure. The load factors of Figure 785-1 address the critical structural load conditions of dispensing equipment mounted in or near the fuselage and provide adequate margins of safety.
FIGURE 785-1
ACCEPTABLE ULTIMATE LOAD FACTOR FOR
AGRICULTURAL DISPENSING EQUIPMENT DESIGN
UP DOWN SIDE FORWARD AFT Tanks & Equipment Mounted 1.5g 4.0g 2.0g 4.0g In Or Near The Fuselage Note 1 Spray Booms 1.5g 2.5g ... Note 1 2.5g Note 2 Note 1: An ultimate load factor of 2 G's is acceptable for externally side or under fuselage mounted tank and forward mounted spray booms where failure in a minor crash landing will not create a hazard to occupants or prevent exit from the rotorcraft.
Par 785 1237
Note 2: The aft loads for spray booms may be developed by the applicant based on the 111 percent of VNE for which certification is requested or the load factors of Figure 785-1, whichever is greater.
(2) The applicant may elect to substantiate his/her product by either static or dynamic testing, by analysis, or any combination thereof.
(3) Lower load factors may be used only when justified by manufacturer's data, rational analysis, or actual rotorcraft flight and ground load demonstrations.
(4) Tank pressure test, while not mandated, is recommended for safety reasons. An acceptable procedure is included in Paragraph (c)(4).
(5) Dispensing equipment installation attach points. If attach points exist which are an integral part of the rotorcraft and these attach points have been certified to the standard category requirements no further substantiation of the attach point is required if an analysis indicates the dispensing system does not impose loads which exceed those for standard category certification.
(6) Ground clearance for dispensing equipment installation. A 5-inch ground clearance has typically been used for skid gear equipped rotorcraft which incorporate belly mounted supply tanks/hoppers or systems which have dual side mounted supply tanks/hoppers and the design incorporates cross tubes or other system components which are located beneath the bottom of the fuselage when these components are rigidly attached to the airframe structure. The 5-inch dimension is measured vertically from the ground to the lowest point of the installed system, with the rotorcraft in its operational configuration and gross weight (including disposable load) and while resting on a smooth, level asphalt surface. For rotorcraft equipped with wheels and/or landing gear struts, the maximum system deflections should be considered when determining the 5 inches of acceptable static ground clearance. The 5-inch ground clearance would only apply to original configuration of newly manufactured rotorcraft. However, a 3-inch ground clearance has been found acceptable and may be approved for skid gear equipped rotorcraft to account for the in-service permanent set allowed for skid gear members, (l.e., cross tube deflections allowed per the maintenance manual). Cable supported systems, (i.e., cargo hook installations) or dispensing systems utilizing flexible ducts (certain types of dry material dispensing equipment which may or may not be retractable) have been approved even though portions of the system may contact the surface during a normal landing.
(7) A number of rotorcraft are approved for external cargo operations that allow a gross weight higher than the approved internal gross weight limit. This difference is usually due to the allowable weight limit restriction of the landing gear.
(The gear is not approved for the higher weight.) Those types of dispensing equipment, that can be loaded in flight to a weight that exceeds the allowable limit of the landing gear should incorporate a reliable means that rapidly reduces the total aircraft gross 1238 Par 785
7/30/97 АС 29-2B weight to within allowable landing gear limits. In most cases, this will involve jettison of the disposable load. The time interval for this operation should be demonstrated, and should not exceed a recommended 3 seconds from a level flight condition.
(8) A flight check or demonstration of the agricultural dispensing equipment installation is normally conducted. This flight check should also qualitatively determine that no hazardous deflection or resonance in the rotorcraft or dispensing system exists.
This flight check should be conducted in accordance with the requirements of FAR 133.41.
(9) For rotorcraft certificated in dual categories, the inspection requirements of FAR 21.187(b) must be observed when converting from restricted to normal category.
c. Acceptable Means of Compliance.
(1) Analysis Method. Structural analysis (static) may be used if the structure is of a configuration for which experience has shown the method to be reliable. Structural substantiation of tanks that are designed to contain liquid materials may be accomplished by pressure testing. For tanks or hoppers designed to contain dry material, (e.g., dust or fertilizer) static load tests may be used to verify structural integrity. The tank/hopper, mounting hardware, and support structure should all be substantiated to the load conditions specified by this paragraph considering the effects of internal fluid pressures when applicable.
(2) Static Tests. Static tests of tank/hoppers, mounting hardware, and support structure for each critical load condition may be accomplished using conventional techniques; such as, dead weight loading, whiffletree systems, and hydraulic rams. If tests of the tank and its mounting hardware are conducted using a test fixture representing the rotorcraft, the rotorcraft support structure may be substantiated independently by means of test and/or analysis. Static test loads should be applied in combination with associated internal fluid pressure loadings. The ultimate loads specified in Paragraph 785 should be sustained for at least 3 seconds without failure.
(3) Dynamic Tests.
If the applicant elects to test to the load factors noted herein, the maneuvering and gust loadings will be considered to be adequately substantiated. For each condition, the critical volume and density of fluid should be used.
(i¡) The tank and mounting hardware should support ultimate loads without detrimental permanent set or failure, respectively. The rotorcraft support structure may be included in the dynamic tests, or it may be substantiated separately via static test and/or analysis for each condition specified by this paragraph.
Par 785 1239
(4) Pressure Testing. Internal pressure loads may be applied using the water standpipe technique. Standpipe water height should be accurately computed for each critical spray tank static test loading. Pressure testing of spray tanks is not absolutely essential but is recommended for safety reasons. This testing will also determine whether the joints and connections are tight and will not leak in addition to determining any weak spots in the construction. Where spraying is done with highly volatile and flammable liquids, or where the tank has a return line, such as in an engine oil tank where the fluid is pumped back into the tank, it is recommended that the tank be tested for a pressure of 5 pounds per square inch. For other liquids, and where no fluid return line is used, testing to 3½ pounds per square inch should be satisfactory. There are many ways of pressure testing a tank, however, it is believed that the simplest and easiest method is to fill the tank with water and use a standpipe filled with water. A 1 1/8-inch pipe can be connected to the venting tube or one adapted to the filler opening. In either case the height of the pipe would be the same. For a 3 ½ PSI test of the tank the height of the water in the pipe would only need to be 8 feet and for a 5 PSI test only an 11 ½-foot height of water will be needed.
1240 Par 785
Pipe or 11 ½ TANK Figure 785-2. Sketch of Tank Pressure Test Par 785 1241