AC 27-1B, Chg 1 2/12/03 database of historical test data of components with similar characteristics. Care should be exercised when pooling such data to be sure that difference in failure modes and curve shapes are considered. Whatever reduction factor is selected, a rationale should be provided to substantiate it. The reduced S/N curve and the loading spectrum developed per paragraph c. and d. should be used in determining replacement times, see paragraph e(4).
(II) Spectrum tests. The establishment of replacement times based on fatigue tests in which each specimen is subjected to a spectrum of loading is to include the following considerations:
(A) Definition of the test loading spectra based on either:
(1) Load histories based on flight test data obtained for flight and ground conditions and maneuvers considered appropriate for the particular rotorcraft, and a spectrum allocating percentages of time or frequencies of occurrence to these flight and ground conditions and maneuvers, or
(2) Analysis supported by extrapolation of available load history data or prior knowledge where available.
(B) The effects of high infrequent load cycles on the test result particularly when such cycles may occur only rarely in service.
(C) The effects of omitting low load (high frequency) cycles to reduce test time should be fully established and supported by test experience to be adequately accounted for.
(D) Fatigue tests in which the loading spectra are applied such that effective randomization of loading is obtained.
(E) Assignment of replacement times. The fatigue test results should be evaluated in terms of the loading spectrum of paragraph d. if different to the test spectrum, and reduced by factors for strength and life based on similar approach to those derived for the constant amplitude tests above.
(4) Safe-Life Calculation Methodology. The key procedures for safe life determination, as shown hereafter, are based on the three basic elements of strength, loads, and usage as established in the preceding sections. These elements are reiterated below and combined according to Figure AC 27.MG 11-6 to calculate the retirement life: 1) The conservative working S/N curve developed from the mean S/N strength curve of the component using reduction factors based on the material and manufacturing variability and test parameters. 2) Loads and usage combined in an individual fatigue loads spectrum for the component determined conservatively through test and analysis. The loads may be processed by conservative methods using maximum load for the duration of each condition or by suitable cycle counting methods.
Page MG 11 - 12
GAG cycles, once per maneuver cycles, and other high load cycle events should be accounted for as the loads are combined with the established usage spectrum. The service life of the component may then be determined using Miner's linear cumulative damage rule, considering both high and low cycle fatigue damage. The calculated service life obtained for a total damage equal to one is then the maximum allowable replacement time for the component.
f. Fail-Safe Evaluation.
(1) General. The fail-safe evaluation of the flight structure is intended to ensure that, should fatigue cracks initiate, the remaining structure will withstand service loads without failure until the cracks are detected. The fail-safe evaluation generally encompasses establishing the components which are fail-safe, defining the loading conditions and extent of damage for which the structure is to be designed, conducting structural tests and analysis to substantiate that the design objective has been achieved, and establishing inspection programs to assure detection of fatigue damage.
Design features that may be used in attaining a fail-safe structure are:
Selection of materials and stress levels that provide a controlled slow rate of crack propagation combined with high residual strength after initiation of cracks.
(i¡> Design to permit detection of cracks, including the use of crack detection systems, before the cracks result in an appreciable loss of residual strength.
(#) Use of multi-path construction and the provision of crack stoppers to limit the growth of cracks.
(iv) Use of composite duplicate structures so that a fatigue crack or failure occurring in one element of the composite member will be confined to that element and the remaining structure will still possess limit load-carrying ability. It may be necessary to employ the design techniques of f(ii) above to assure effectiveness of these features.
(v) Use of backup structure wherein one member carries the entire load, with a second member available and capable of assuming the load if the primary member fails.
(2) Extent of Fail-Safe Damage. The extent of the partial failure is to be such that it would be readily detectable during the specified inspection. It may involve complete failure of a principal element, failure of more than one element, or only a partial failure of an element, depending on the rate of crack propagation, the ease of detection, and the inspection interval. Damage in inaccessible areas should extend into inspectable areas.
Typical examples of the fatigue damage that should be considered are outlined below:
AC Page MG 11 - 13
AC 27-1B, Chg 1 2/12/03 Cracks emanating from the edge of structural openings or cutouts which can be readily detected by visual inspection of the area.
(ii) A circumferential or longitudinal skin crack in the basic fuselage structure of such a length that it can be readily detected by a visual inspection of the surface area.
(ili) Complete severance of interior frame elements or stiffeners in addition to a visually detectable crack in the adjacent skin.
(iv) Failure of one element in a multiple load path design.
(v) Failure of primary attachments, including control hinges and fittings.
(3) Determination of Probable Crack Locations. The probable crack locations are to be determined by tests, analysis, or both. In cases of unusually critical or complex components or when initial fatigue loading may affect the rate or mode of cracking, the probable crack locations should be determined by fatigue test. When determination is made by analysis, sound engineering judgement should be used and a variety of factors such as the following should be taken into account:
(i) Conducting an analysis to locate areas of maximum stress and low margin of safety.
(ii) Conducting strain surveys on undamaged structure to establish points of high stress concentration as well as the magnitude of such concentration.
(iii) Examining static test results to determine locations where excessive deformation occurred.
(iv) Determining from fatigue analysis where cracks may initiate.
(iv) Selecting locations in an element where the stresses in adjacent elements would be the maximum with that element failed.
(v) Selecting partial fracture locations in an element wherein high stress concentrations are present in the residual structure.
(vi) Assessing design detail areas which are prone to fatigue damage such as joints, holes, and other features as based on service and test experience of similarly designed components.
(4) Fail-Safe Demonstration. It is to be demonstrated by analysis, tests, or both, that the structure with the partial failures as defined in paragraphs f(2) and f(3) can withstand the maximum load and the repeated loads expected in service during the period prior to detection. The repeated loads should be as defined in the loading Page MG 11 - 14
spectrum of paragraph d(2) and the structure should be capable of supporting this loading after a partial failure for a sufficient time with respect to the inspection interval to assure that catastrophic failure is extremely remote. In test demonstrations, the damage may be initiated or simulated by cuts made with a fine saw, sharp blade, or guillotine in those cases where it is not necessary and not practical to produce fatigue cracks by tests. In those cases where damage is simulated at joints or fittings, bolts may be removed to simulate failure if this condition would be representative of an actual failure. In some instances, the fail-sate characteristics may be shown analytically. The analytical approach may be used when the structural configuration involved is essentially similar to one already verified by fail-safe tests, whether on a previously approved type design, or on other similar areas of the design currently being evaluated.
The analytical approach may also be used when:
(i> It can be shown that the failure would be detected considerably before the critical crack length is reached;
(ii) The margins of safety resulting from the analysis are well in excess of the fail-sate residual static strength level; and, (II) The stress levels in the partially failed structure and the design are such as to assure adequate crack propagation time relative to the inspection interval.
(5) Inspection. Detection of fatigue cracks before they become dangerous is the ultimate control in ensuring the fail-safe characteristics of the structure. Therefore, the manufacturer should provide sufficient guidance information to assist operators in establishing the frequency and extent of the repeated inspections of the critical structure.
g. Further Considerations.
(1) Control of Fatigue Sensitive Parts. Control of the part in manufacture, operational service, and maintenance is vital to ensure the full benefits of the fatigue life substantiation process. Any part that has been selected for fatigue assessment should be considered using the following guidance. This is particularly important for safe life parts with no damage tolerance capability. The details of the manufacturing procedures and processes for fatigue sensitive parts, including material manufacture and source, forging procedures, machining operations and sequence, and inspection techniques and acceptance and rejection criteria should be established. Sensitivity should be established on the basis of identifying the processes, which if incorrectly completed could significantly affect the fatigue life. The tested components should be produced in accordance with the above manufacturing procedures. For life-limited and fatigue sensitive parts, the design and manufacturing standards should be frozen subject to further evaluation by the design authority. Parts produced in whole or in part under subcontracting or partnership arrangements should be subject to the same procedures.
Life-limited and fatigue sensitive parts should be marked with a serial number and records relating to the marking maintained, such that it is possible to establish the AC Page MG 11 - 15
AC 27-1B, Chg 1 2/12/03 relevant manufacturing modification and service history of the individual parts. Special instructions should be provided, as necessary, to ensure the part is handled in an appropriate manner, particularly during maintenance. Processes for determining the disposition of parts having manufacturing errors or material flaws should be established.
In a similar manner, processes controlling changes to the design or manufacture of the component or to its operating environment or loading spectrum are required. For any such changes, their effects on the fatigue evaluation of the part should be established.
This evaluation should involve further fatigue testing, unless it can be shown that testing is not necessary.
Page MG 11 - 16
4 ROTOR RPM VNE 27.1505 MAX. DESIGN OR DEMONSTRATED RPM 27.33, 27.309
NORMAL OPERATING RPM MIN. DESIGN OR DEMONSTRATED RPM 27.33, 27.309 111% VNE 27.309
AIRSPEED FIGURE AC 27.MG 11-2: Flight Regime to be Investigated Power-On Operation.
(Note: Dashed lines in these figures indicate test boundaries. Shaded areas indicate operating regimes.)
ROTOR RPM VNE 27.1505 MAX. DESIGN OR DEMONSTRATED RPM 27.33, 27.309 MAX. PLACARD RPM 27.33, 27.1509 (95% MAX.DESIGN) MIN. PLACARD RPM 27.33, 27.1509 (105% MIN.DESIGN) 111% VNE MIN. DESIGN OR DEMONSTRATED RPM 27.33, 27.309 27.309
AIRSPEED FIGURE AC 27.MG 11-3: Flight Regime to be Investigated for Power-Off Operation AC Page MG 11 - 17
AC 27-1B, Chg 1 2/12/03 Estimated Mean Endurance Limit Adjusted Specimen Endurance Limit Vibratory Stress Operating Boundary (C=1/ЗB) C Yield Stress D Steady Stress FIGURE AC 27.MG 11-4: Simplified analytical method for safe life evaluation Vibratory Load or Stress Reduction Life - Test Data Curve Strength Statistically Reduced — * Reduction Working Curve Cycles, N FIGURE AC 27.MG 11-5: Typical S/N curve for safe life evaluation.
Page MG 11 - 18
Measured Loads Component Strength Usage (Flight Load Survey) (S/N Curve) (Flight Spectrum) Miner's Rule Life Determination FIGURE AC 27.MG 11-6: Elements of Safe-Life Determination AC Page MG 11 - 19
AC 27-1B, Chg 1 2/12/03 FIGURE AC 27.MG 11-7
FLIGHT LOAD MEASUREMENT PROGRAM CONDITIONS TO BE INVESTIGATED
1. GROUND CONDITIONS.
a. Normal start.
b. Rapid increases of RPM on ground to maximum power-on RPM of main rotor.
c. Taxiing with max allowable full cyclic control.
d. Landing run (if applicable).
e. Braking (if applicable).
f. Normal shutdown.
g. Special ground checks (if applicable).
2. HOVERING IN AND OUT (1) Steady with rotor at OF GROUND-EFFECT maximum side of RPM tolerance.
(2) Steady with rotor at minimum side of RPM tolerance.
(3) 90-degree right turn
(4) 90-degree left turn
(5) Control reversals
i) Longitudinal.
ii) Lateral.
ili) Rudder.
(6) Sideward flight
i) Left.
ii) Right.
(7) Rearward flight
3. MANEUVERING IN (1) Jump takeoff.
GROUND EFFECT (2) Normal takeoff (*) and accelerate to climb airspeed.
(4) Full autorotational landing. (3) Normal Landing (*) i) Multiengine. ii) One-engine-inoperative.
Page MG 11 - 20
FIGURE AC 27.MG 11-7 (continued)
4. FORWARD FLIGHT-POWER (1) 40 percent VH i) Minimum side of main rotor ON. RPM tolerance (RPM -)
a. Level flight. (**) ii) Maximum side of main rotor RPM tolerance (RPM +)
(2) 60 percent VH i) (RPM +) ii) (RPM -)
(3) 80 percent VH i) (RPM +) ii) (RPM -)
(4) VH i) (RPM +) ii) (RPM -)
(5) VNE i) (RPM +) ii) (RPM -)
b. Maneuvers. (1) Full power climbs. (**) i) All engines operative.
ii) One-engine-inoperative.
(2) Cyclic pull-ups. i) 60 percent VH.
ii) 90 percent VH.
(3) Normal acceleration from climb airspeed to 90 percent VH.
(4) Turns, right and left over a range of bank angles and speeds up to the lesser of VH or VNE and including:
i) Right at 60 percent VH and 90 percent VH.
ii) Left at 60 percent VH and 90 percent VH.
(5) Control reversals at 90 percent
i) Longitudinal.
ii) Lateral.
iii) Rudder.
(6) Deceleration from 90 percent Vu to descent airspeed.
(1) Partial power descent. (*)
(2) Normal approach
(3) Steep Approaches (to landing) i) All engines.
or Flare FIGURE AC 27.MG 11-7 ii) One engine out.
AC Page MG 11 - 21
AC 27-1B, Chg 1 2/12/03 (continued)
5. POWER TRANSITIONS.
a. All engines operating to one (1) In full power climb.
engine out. (2) At 90 percent VH.
b. One engine out to all engines operating in powered descent.
c. All engines operating to (1) At 60 percent VH.
autorotation. (2) At maximum forward transition speed.
d. Stabilized autorotation to all engines operating at normal autorotation airspeed.
6. AUTOROTATION.
a. Stabilized. (1) At 70 percent VNE•
(2) At VNE•
b. Turns at 70 percent and 100 (1) Right.
percent VNE• (2) Left.
c. Cyclic pull-up.
d. Control reversals. (1) Longitudinal.
(2) Lateral.
(3) Rudder.
(**) side slip conditions should be considered (*) max slope angle and aircraft headings should be considered Page MG 11 - 22
FIGURE AC 27.MG 11-8
LIFTING HELICOPTER LOAD MEASUREMENT in addition to normal conditions CONDITION %RATED LOAD % VNE
ROTOR START VERTICAL LIFT 100 87 HOVER including spot turns 100 sideward and rearward flight and control reversal.
87 CRUISE with load 100 100 87 100 100 90 87 90 CRUISE no load 100 90 CLIMB max. rate DECELERATIONS max. rate
GROUND IDLE
SHUTDOWN AC Page MG 11 - 23