confined to that element and the remaining structure will still possess appreciable load-carrying ability.
e. Use of backup structure wherein one member carries all the load, with a second member available and capable of assuming the extra load if the primary member fails.
14. IDENTIFICATION OF CRITICAL PORTIONS OF PLIGHT STRUCIURE. Those portions of the flight structure which may be critical in fatigue should be identified. Typical portions of the structure are:
a. Rotor blades and attachment fittings.
b. Rotor heads, including hubs, hinges, dampers.
c. Control system components, including control rods, servos, swashplates.
d. Rotor supporting structure.
e. Fuselage, including stabilizers and auxiliary lifting surfaces.
15. LILV OF FALL-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 which should be considered are outlined below:
a. Cracks emanating from the edge of structural openings or cutouts which can be readily detected by visual inspection of the area.
b. 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.
Complete severance of interior frame elements or stiffeners in addition to a visually detectable crack in the adjacent skin.
Par 13 Page 11
d. Failure of one element where dual construction is utilized in components.
e. failure of primary attachments, including control hinges and fittings.
16. DETEFMINATION 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 loadings 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 judgment should he used and a variety of factors such as the following taken into account:
a. Conducting an analysis to locate areas of maximum stress and low margin of safety.
b. Conducting strain surveys on undamaged structure to establish points of high stress concentration as well as the magnitude of such concentration.
c. Examining static test results to determine locations where excessive deformation occurred.
d. Determining from fatigue analysis where cracks may initiate.
e. Selecting locations in an element where the stresses in adjacent elements would be the maximum with that element failed.
f. Selecting partial fracture locations in an element wherein high stress concentrations are present in the residual structure.
g. Assessing design detail areas which are prone to fatigue damage based on service experience records of similarly designed components.
17. FAIL-SAFE DEMONSTRATION. It is to be demonstrated by analysis, tests, or both, that the structure with the partial failures as defined in paragraphs 15 and 16 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 spectrum of paragraph 7 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 Page 12 Par 15
failure is extremely remote. The loading spectrum should include at least one application of limit load.
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 failsafe 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: (1) it can be shown that the failure would be detected considerably before the critical crack length is approached;
(2) the margins of safety resulting from the analysis are well in excess of the fail-safe residual static strength level; and (3) 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.
18. INSPECTION. Detection of fatigue cracks before they become dangerous is the ultimate control in insuring 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.
: Par 17 Page 13
TABLE 1 Percent occurrence I. GROUND CONDITIONS- 1.5
(a) Rapid increase of r.p.m. on ground to quickly engage clutch-0.5 (Б) Taxiing with full cyclic control- .(I) •5
(c) Jump takeoff- (1)
(d) Ground-air-ground cycle- HOVERING- -2.0
(a) Steady hovering 5
(b) Lateral reversal 5 (с) Longitudinal reversal- • 5
(d) Rudder reversal • 5 III. FORWARD FLIGHT-POWER ON- -87.5
(a) Level flight - 20% VNE- L. O
(b) Level flight - 40% VNE- 3.0 (с) Level flight - 60% VNE 18.0
(d) Level flight - 80% VNE- 25
(e) Maximum level flight (but not greater than VNE)- (£) VNE
(g) 111% VNE-
(h) Right turns - 30, 60, 90% VNE (2)
(1) Left turns - 30, 60, 90% VNE- • (2) 3.0 Climb (Takeoff power)-
(k) Climb (Max. continuous power)
(1) Change to autorotation from power-on-flight - 30, 60, 90% VNE1•5
(m) Partial power descent (including condition of zero flow through rotor) - 2.0
(n) Cyclic and collective pull-up from level flight- •(2)
(0) Pushovers• •(2) (₽) Gusts-+ (2)
(q) Quick stops- (2) (I) Flares (2)
(s) Lateral reversals at VH
(t) Longitudinal reversals at VH (и) Rudder reversals at VH- • 5
(v) Landing approach- 3.0
(w) Sideward flight- • 5
(x) Rearward flight-- • 5 IV. AUTOROTATION - POWER OFF-- -9.0
(a) Steady forward flight- 2.0
(b) Rapid power recovery from autorotational flight • 5
(c) Right turns - 30, 60, 90% VNE- (2) 1.0
(d) Left turns - 30, 60, 90% VNE- •(2) 1.0 Page 14
Lateral reversals-- • 5 Longitudinal reversals-- • 5 Rudder reversals- •5 Cyclic and collective pull-ups-- - (2) Landing approach--- 2.0 flares-- • (2)
100.0
(1) One flight every 10 minutes with less frequent rotor stops.
(2) A vertical load factor frequency curve should be developed that is representative of the more critical types of operation. The time spent in each turn should be adjusted to give the specified per cent of occurrence.
NOTE: DASHED LINES IN FIGURES 1 & 2 INDICATE TEST BOUNDARIES. CROSS-HATCHED AREAS INDICATE OPERATING REGIMES.
VNE 27.1505
ROTOR RPM AIRSPEED FIGURE 3 Oscillatory Stress FIGURE 1. POWER ON- vs. Steady Stress ROTOR R.P.M. AIRSPEED ENVELOPE
STEADY STRESS LEVEL-A VNE 27.1505
ROTOR RPM FIGURE 2 POWER OFF- FIGURE 4 Stress vs.
ROTOR R.P.M. AIRSPEED ENVELOPE Number of Cycles Page 16
_ . . A -, -*- --.- .-.- -9 F