ADVISORY 

CIFtcULAR 

DEPARTMENT OF TRANSPORTATION 

FEDERAL AVIATION ADMINISTRATION 

SUBJECT: FATIGUE EVALUATIONOFROTORCRAFTSTKIClsJRE 

1 0 PURPOSE. This advisory circular sets forth acceptable means of 

compliance with the provisions of Federal Aviation Regulations, 

Sections 27.571 and 29.571 dealing with the fatigue evaluation 

of rotorcraft structure. 

2 l CANCELLATION. Appendix A, Main Rotor Service Life Determination, 

January 15, 1963, to CAM 6 is cancelled. 

3 0 HOW TO GET THIS PUBLICATION. 

a. Order copies of this publication from: 

Department of Transportation 

Federal Aviation Administration 

Distribution Unit, TAD 443.1 

Washington, D.C. 20591 

b . Identify the publication in your order as: Fatigue Evaluation 

of Rotorcraft Structure, AC No. 2w5. 

C. This advisory circular will be distributed free of charge. 

‘1 I 

L 

L i 

( 7g. ‘; 7 6’ t,;r ’ 

t J. A. FERRARESE 

‘Acting Director, Flight Standards Service 

Initiated by: A&120 
AC 20-95 

SECTION lo ZNTRODUCTION 1 

- . . 

1 0 Background . 1 

SECTION I? 0 

tXCTION 3. 

XXTION. 5. EXtI,-SAFE ST~EMGTH EVALUATION 10 

5/=/76 

TABLE OF CONTENTS 

2 l General 2 

3 l Instrumentation 3 

4 l Parts to be strain-gaged 3 

5 l Flight regimes and conditions to be investigated 3 

l7XQUENCY OF LOADING 

6 l <Type of operation 

7 0 Loading spectrum 

FATIGUE STR.ENGTf! EVALUATION 

8 l General 

9 0 Analytical method 

10 0 Testing methods 

11 l Combination replacement time and fail-safe 

evaluation 

12. Extension of replacement time 

13 

14: 

General 

Identification of critical portions of primary 

structure 

15 0 Jlxtcnt of fail-safe damage 

16 l Determination of probable failure locations 

17 0 Fail-safe clemonstration 

18 0 Inspection 

Percent of occurrence 

1 0 Power On - Rotor R.P.M. Airspeed Envelope 

2 0 Power Off - Rotor R.P.M. Airspeed Envelope 

3 

4: 

Oscillatory Stress vs. Steady Stress 

Stress vs. Number of Cycles 

4 

4 

5 

5 

5 

6 

7 

9 

9 

10 

11 

11. 

1' L 

12 

13 

14 

16 

16 

16 

16 
SECTION 1. INTRODUCTION 

1 l BACFRWND. The fatigue evaluation procedures outlined in this 

advisory circular are for guidance purposes only and are not mandatory 

nor regulatory in nature. Although a uniform approach to -.fatigue 

evaluation is desirable, it is recognized that in such a complex problem, 

new design features and methods of fabrication, new approaches to 

fatigue evaluation, and new configurations may require variations 

and deviations from the procedures described herein. Engineering 

judgment should therefore be exercised for each particular application. 

a. The flight structure of the rotorcraft is subject to vibratory 

stresses in practically every regime of flight. In addition, 

since it is a highly maneuverable aircraft that is capable of 

forward, rearward, sideward, vertical, and rotational flight, 

operating limitations due to fatigue are possible in practically 

all flight situations. For these reasons, it is required that 

special attention be focused on the fatigue evaluation of the 

flight structure of the rotorcraft. 

b. Fatigue evaluation of the flight structure is intended to verify 

structural reliability. Assurance of s true tural rel iabili ty 

starts with design, including choice of materials for resistance 

to crack initiation and/or propagation, detail design to minimize 

stress concentration, and specification of surface finishes, fits, 

etc. Design analysis should include estimation of expected 

flight loads, and estimation of resistance to fatigue. Fatigue 

strength should be based on past full scale fatigue tests and/or 

materials fatigue data with appropriate reductions for the variability in fatigue strength, size, shape, surface finish, and 

environments of the structure. In addition, design for fatigue 

should consider mode-of-failure analysis, areas susceptible to 

fatigue cracking, and methods to assure detectability of fatigue 

cracks. The residual strength of a cracked structure is an important 

consideration of fail-safe design. 

c. Assurance of structural reliability also includes manufacture and 

fabrication in accordance with design requirements and specifications, quality control to monitor compliance, and effective 

service inspection procedures. 

Par 1 Page 1 
AC 20-95 5/18/w 

d l Fatigue evaluation of the structure, . measurement of flight loads 

and stresses, and evaluation of fatigue strength and/or fatigue 

crack propagation are the subjects of this advisory circular. 

There is some question whether a cotipletely reliable method for the 

prediction of time to fatigue crack initiation and fracture 

exists. Nevertheless, one engineering approach to the subject is 

to use the "Linear Cumulative Damage Hypothesis." This hypothesis 

states that every cycle of stress above an "endurance limit" 

produces fatigue damage proportional to the ratio of cycles 

accumulated at the stress to fatigue "life" at that stress, 

e. Laboratory tests of this hypothesis indicate that it is reasonably 

valid when the loading spectrum consists of stresses which are, 

in effect, random6 Despite the lack of an adequate theory connecting this hypothesis with more basic properties of materials, 

it has been successfully used in a number of applications. 

f 0 In addition, fatigue evaluation generally requires a method of 

accounting for the effect of steady loads and stresses on fatigue. 

Where the manufacturer does not provide other substantiating data, 

a Goodman diagram may be used to account for these effects. 

F:* In any rational fatigue evaluation, the following factors should 

be considered: 

(1) Identification of the structure to be considered in the 

fatigue evaluation. 

(2) The stresses associated with steady and maneuverinp operating I 

conditions expected in service. 

(3) The frequency of occurrences associated with various flight 

conditions and the corresponding spectrum of loadings and 

stresses. 

(4) The fatigue strength, fatigue crack propagation characteristics of the structure, and the residual strength of the 

cracked structure. 

SECTION 2. FLIGHT STRAIN MEASUREMENT PROGWU4 - 

l l 2 0 GENERAL. Subsequent to design analysis, in which aircraft loads and 

associated stresses are derived, the stress level and/or loads are to 

be verified by a carefully controlled flight strain measurement 

program. 

Page 2 Par 1 
5/18/N AC X1-95 

3 0 INSTRUMENTATION. 

a. The instrumentat-Len system used in the flight strain measurement 

program should accurately measure and record the critical strains 

under test conditions associated with normal operaticn and 

specific maneuvers. The location and distribution of the strain 

gages should be based on a rational evaluation of the critical 

stress areas. This may be accomplished by appropriate analytical 

means, supplemented, when deemed necessary, by strain sensitive 

coatings or photoelastic methods. The distribution and number of 

strain ege S IlKl st def 

part es sent i .a1 to the 

ine 

sa 

the load p spectrum ade qua tely for each 

fe operat ion of th e ro tor craft. 

p 0 The corresponding flight parameters (airspeed, rotor r.p.m., 

center of gravity accelerations, etc.) should also be recorded 

simultaneously by appropriate methods. This is necessary in 

order to correlate the loads and stresses with the maneuver or 

operating condition at which they occurred. 

c. The instrumentation system should be adequately calibrated and 

checked periodically throughout the flight strain measurement 

program in order to ensure consistent and accurate results. 

4 a PARTS To BE STRAl+GAGED. Fatigue critical portions of the rotor 

systems, control systems, fuselage, and supporting structure for 

rotors, transmission, and engine are to be strain-gaged. For 

rotorcraft of unusual or unique design, special consideration might 

be necessary s to insure that all of the essential parts are evaluated. 

5. FLIGHT REGIMES AND CONDITIONS TO BE INVXSTIGATED. 

il. Typical flight and ground conditions to be investigated in the 

flipht.strain measurement program are given in Table 1. Flight 

regimes that should be investigated for Dower-on and power-off 

operation are shown in Figures 1 and 2. -For clarity, the 

parameters which define these regimes are included in these 

figures. As noted on Figure 1, complete coverage at 111 percent 

VNE should be demonstrated for power-on operation. However, for 

power-off operation, Figure 2, complete coverage at 111 percent 

VI\IE for maximum and minimum design r.p.m. need not be obtained if 

points are obtained at Vm at both maximum and minimum design 

r.p.m. and at 111 percent V~;E at both maximum and minimum placarded r.p.m. as indicated in the figure. 

Par 3 Page 3 
AC 20-95 

b 0 The determination of flight conditions to be investigated in the 

'flight strain measurement program should be based on the antici- 

' . pa&d use of the helicopter, and, if available, on past service 

records for similar designs. In any event, the flight conditions 

considered appropriate for the design and application should be 

representative of the actual operation in accordance with the 

rotorcraft flight manual. In the case of multiengine helicopters, 

the. flight conditions concerning partial engine-out operation 

should be considered in addition to complete power-off operation. 

The flight conditions to be investigated should be submitted in 

connection with the flight evaluation program. Suggested flight ., 

conditions for single-engine helicopters used in normal operation 

are shown in Table 1. 

c. The severity of the maneuvers investigated during the flight 

strain survey should be such that it is unlikely that service use 

will be more severe. 

d l All flight conditions considered appropriate for the- particular 

design are to be investigated over the complete rotor speed, 

airspeed, center of gravity, altitude, and weight ranges in order 

to determine the mst critical stress levels associated with each 

flight condition. In order to account for data scatter and to 

determine the stress levels present, a sufficient amount of data , 

points should be obtained at each flight condition. In some 

instances, the critical weight, center of gravity, and altitude 

ranges for the various maneuvers can be based on past experience 

with similar designs. This procedure is acceptable where adequate 

flight tests are performed to substantiate such selections. The 

combinations of flight parameters that produce the most critical 

stress levels should be included in the fatigue evaluation. 

SECTION 3. FREQUENCY OF LOADIIJC 

6 l TYPE OF OPERATION. The probable types of operation (transport, 

utility, etc.) for the rotorcraft should be established. The type 

of operation can have a major influence on the loading environment. 

13ormally, the rotorcraft should be substantiated for the most 

critical general type of operation with consideration of special 

occasional types of operation. 

Page 4 , Par 5 
AC 20-95 

7 0 LQkmGa The spectrum allocating percentages of time or 

frequencies of occurrence to flight conditions or maneuvers is to 

be baked on the expected usage of the rotorcraft. This spectrum 

is tosbe such that it is unlikly that actual usage will subject 

the structure to damage beyond that associated with the spectrum. 

Considerations to be included in developing this spectrum should 

include prior knowledge based on flight history recorder data, 

design limitations established in compliance with FAR’s 27,309 

or 29.309 and recommended operating conditions and limitations 

specified in the.rotorcraft flight manual. The distribution of 

times at various forward flight speeds should reflect not only 

the relation of these speeds to VNE but also the recommended 

operating conditions in the rotorcraft flight manual which govern 

VC or cruise speed.. Where possible, it is desirable to conduct 

the flight strain-gage program by simulating the usage as determined 

above, with continuous recording of stresses and loads, thus 

obtaining directly the stress/load spectra for structural elements. 

Table 1 contains typical percent of occurrences for the various 

.flight condi 

x 

ions for single-engine piston helicopters used in 

utility oper tions. This table should be used only as a guide and 

should be modified as necessary for each particular rotorcraft. 

m FATIGUE STRENGTH EVALUATION 

8 0 GmERAL. Information to guide fatigue evaluation based on safe-life 

considerations leading to recommended replacement times is provided 

in this section. Although there is a large quantity of information 

available on the fatigue strength characteristics of material 

specimens, built-up specimens and parts, the prediction of strength 

of parts of new designs based on this information is less reliable 

than testing the actual part. Consequently, additional conservatism should be used with this method. However, in many cases the 

differences between past test specimens and the actual part (which 

involve such factors as stress concentration, size, and fretting) 

cannot be accounted for with a reasonable degree of accuracy. 

Therefore, it is usually necessary. that the structural components 

be subjected to repeated load tests using information determined 

in the flight strain measurement program. Special operational or 

functional characteristics which could affect the fatigue strength 

should also be considered in the service life evaluation. Such 

factors as high blade operating temperatures due to tip jets or 

Par 7 Page 5 
AC 20-95 5/18/N 

turbine exhaust impingement on the tail rotor should be considered 

as well as other special operating conditions. In addition, 

effects of special purpose use such as hoist and sling operation, 

spraying, surveying, etc. v s.hould be considered if appropriate to 

the particular type. The fatigue strength should be evaluated by 

either of the methods outlined below, but fYull scale testing is 

considered more accurate. 

9 0 ANALY'TICPFL METHOD,. It is recognized that if allowab,ti bzess 

levels are established by acceptable means* and the stresses 

measured in flight are lower than these established levels, no 

fatigue testing is necessary. 

a. Simplified method. The following techniwes, based on the use 

of the Goodman diagram, are considered acceptable for establishing 

this allowable stress level: 

Estimate the mean endurance limit of the part from test 

results of specimens with similar stress concentrations. The 

test specimen material should be representative of the actual 

part and sufficient test data should be available to 

substantiate the mean endurance limit. The estimate should 

account for surface conditions, fabrication methods, fretting, 

size and shape effects, as well as differences in stress 

concentrations between the test specimen and the actual part. 

Referring to figure 3, the mean endurance limit may be represented by a straight line drawn through the yield stress 

(point A on the horizontal. tis) and the maximm oscillatory 

stress which the aver e specimen can withstan at a give 

steady stress (Point B 7 without failure for 10 ff 9 to 5 x 10 

cycles depending on the material. 

A factor of safety of 3 should then be applied to the mean 

endurance limit so that the slope of line AC would be l/3 

of line AB. A smaller factor is acceptable when substantiated 

by a sufficient number of tests on similar parts in similar 

applications. 

If the flight strain measurements indicate that al of apeating stresses fall below the operating boundary line (AC), 

no fatigue testing is necessary. When the measured stresses 

.are above the operating boundary line, however, fatigue 

testing of the actual parts should be conducted. 

Caution should be exercised in the application of the analytic& 

method above, particularly when the following items are 

involved: 

Page 6 

Par 8 
5/18/7h AC m-95 

( a 1 

b) 

( C 1 

(a) 

( e > 

. 

Large parts in proportion to the laboratory specimens. 

Irregularly shaped parts containing numerous or supeimposed fillets, holes, threads, or lugs. - . 

Parts of unique design for which no past service experience 

is available. 

Parts subject to fretting. 

Bolted or pinned connections. 

b. Rational methods. Methods may be used which do not involve full 

scale testing but which apply the variables of fatigue strength 

with a calculation of retirement times in a manner that provides 

equivalent reliability to the fatigue testing and simplified 

methods and is acceptable to the Administrator. 

10 . TESTINGMETH~DS. The fatigue strength of the flight structure may 

be determined in appropriate laboratory tests and evaluated in 

terms of a loading spectrum. The strength indicated by the test 

results should be reduced by a factor such that a replacement time 

based directly on this reduced strength level and the loading 

spectrum of paragraph 7 will assure that the probability of 

failure is extremely remote. This reduction factor should be 

based on consideration of the number of specimens tested, the 

variability' of the fatigue results, the effects of service use, 

and, where available, previous test data for the same material or 

similar components as well as service experience. 

The test methods outlined below are considered acceptable. 

a, S-N brves. 

(1) Fatigue tests should be conducted over a range of 

oscillatory stresses or loads to define the S-N curves. 

Fatigue tests should be perfonned at steady stresses or 

loads representative of those occurring in flight. 

(2) In order to determine the mean fatigue strength and the 

variability in fatigue strength, it is necessary to test 

a number of specimens in establishing SIN curves. In 

order to account for the variability in fatigue strength, 

a reduction factor should be applied to the mean curve 

in arriving at a working SIN curve. This factor should 

include consideration of the number of specimens tested, 

Par 9 Page 7 
AC 20-W 

the variability of the fatigue results, and, where 

w available, previous test data on the same material or 

similar components, as well as service experiencE. 

Where new materials or designs are being evaluated, 

it is recommended that a larger reduction factor be 

used until such time as additional test data justifying 

a change are available. The mean and reduced S-N curve 

should reflect the curve shapes of typical published S-F 

data on notched or unnotched specimens, as applicable. 

The reduced S+ curve and the loading spectrum of 

paragraph 7 should be used in determining replacement 

. times. Figure 4 represents the method of constructing 

. .a typical S-N curve from the fatigue test data. 

1 0 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: 

(1) Definition of the test loading spectra based on either: 

; (a) Analysis, ,supported by extrapolation of available 

load history data or prior knowledge where 

available, or 

(b) Stress histories based on flight test data obtained 

. for flight and ground conditions and maneuvers 

cons*idered'appropriate for the particular rotorcraft, 

and a spectrum allocating percentages of time or 

frequencies of occurrence to these flight and 

ground conditions and maneuvers. 

(2) Fatigue tests in which the loading spectra are applied 

such&at effective randomization of loadings is 

obtained, 

((3) Unless performed prior to step '(l), determining by flight 

test the stress levels associated with each flight condition 

and maneuver considered appropriate for tlhe particular 

rotorcraft. 

(4) Assignment of replacement times. The fatigue test results 

should be eval.uated in terms of the loading spectrum of 

.paragraph 7 (if different than test spectr&) and reduced 

Pal: IO 
5/18/76 ’ AC 20-95 

by a factor considering the number of specimens, the 

fatigue strength variability, and applicable prior 

data, in arriving at a replacement time. -_l 

C. 'Major system tests. Another method of determining the replacement 

times is to perform a fatigue test or tests of the m3jor 

systems, Examples of such testing are whirl tests, tiedown 

tests, and bench .tests. The test results should be evaluated 

in terms of the loading spectrum of paragraph 7 (if different 

from test loading) and reduced by a factor considering number 

of specimens tested, variability in results, and applicable 

prior test data in arriving at a replacement time. 

11 . COMBINATION OF REPLACEXENT TIME AND FAILSAFE EVAUJATION. It 

may be possiLh to extend the replacement time of safe-life 

components which exhibit limited fail-safe capability by using a 

combination of the safe-life and fail-safe concepts. This is 

accomplished by evaluating both the fatigue strength and failsafe characteristics as described elsewhere in this circular and 

by assigning both a replacemen .t time and inspection period to 

these components. The replacement time may then he based on the 

combined probability of not initiating a fatigue crack at or 

before the replacement time and the probability that the crack if 

initiated will be detected prior to catastrophic failure or loss 

of limit load (or maximum attainable load, whichever is less) 

carryiq'capability. The probability of detection should be based 

on consideration of the inspection effectiveness, the inspection 

interval, and the fatigue life remaining after an obvious partial 

failure. A lower strength reduction factor, commensurate with 

this probability of detection, may then be used in the determination 

of the replacement time. 

12 l EXTENSION OF REPLACEXIENTIME. Parts should be replaced or retired 

'at the established service period unless additional data indicate 

that an extension of the service period is justified. Important 

factors; in the consideration of such extension would be: 

a. Recorded load 'data, Recording load data entails instrumenting 

aircraft in service to obtain a representative sampling of ' 

actual loads experienced. The data measured should include. 

airspeed, altitude, and rotor speed versus time, or the airspeed, altitude, and strain ranges versus time, or similar 

data. The data obtained by instrumenting aircraft in service 

should provide a basis for correlating the estimated loading 

spectrum with the actual service experience. 

Par 10 Page 9 
AC 20-W 5/18/76 

b . Additional analyses and tests. If test data and analyses based 

on repeated load tests of additional specimens are obtained, 

a reevaluation of the initial strength reduction or scatter w. . factor may be made. 

c. Tests of parts removed from service. Where conservatism was 

used in initial calculation of replacement times because of 

lack of knowledge of service environment, repeated * .?d tests 

of replaced parts may be utilized to reevaluate the rni tial 

scatter factor selected. The tests should closely simulate 

service loading conditions. 

d l Rework of the structure. In some cases, rework of the structure 

may result in an increase in replacement time. 

SECTIOK 5'. FAIL-SAFE STRENGTH EVALUATION 

13 l G-L. The fail-safe strength evaluation of the flight structure 

is intended to insure 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. On components predominantly loaded by centrifugal force, 

care should be taken in selecting limit load to assure that it is 

the maximum expected in service. Design features which may be used 

in attaining a fail-safe structure are: 

a. Selection of materials and stress levels that provide a controlled 

slow rate of crack propagation combined with high residual 

strength after injtiation of cracks. 

b 0 Design to permit detection of cracks including the use of crack 

detection systems, in all critical structural elements before 

&he cracks can become dangerous or result in appreciable strength 

loss, and to permit replacement or repair. 

c -0. Use of multipath construction and the provision of crack stoppers 

to limit the growth of cracks. 

d l Use of composite duplicate structures so that a fatigue crack or 

failure.occurring in one element of the composite member will be 

PaEe 10 Par 12 
5/=/76 AC x1-95 

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 0 IDENTIFICATION OF CRITICAL PORTIONS OF FLIGHT STRUCMRE. Those por- w- tions 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 0 Rotor heads, including hubs, hinges, dampers. 

c. Control system components, including control rods, servos, 

awashplates. 

d l . R6cor . . supporting structure. 

e. Fuselage, including stabilizers and auxiliary lifting surfaces. 

15 0 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 6f more than one element, or only a partiaLfailure 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 OY cutouts 

which can be readily detected by visual inspection of the area. 

b 0 A circumferential or longitudinal skin crack In the basic fuselage 

structure of such a length thatit can be readily detected by a 

visual 'inspection of the surface area. 

c. Complete severance of interior frame elements or stiffeners in 

addition to a visually detectable crack in the adjacent skin. 

Par 13 Page 11 
AC 20-95 

d . Failure of one element where dual construction is utilized in 

components. 

e. Failure of primary attachments, including control hinges and 

fittings. 

16 0 DETERMINATION OF PROBABLE CRACK LOCAe~ONS, 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. V%en 

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 

low margin of safety. 

maximum stress and 

b l Conducting strain surveys on undamaged structure to establish 

points of high stress concentration as well as the magnitude 

of such concentration. 

c. Dcamining static test results to determine locations where 

excessive deformation occurred. 

d a Determiningfrom 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 0 Selecting partial fracture locations in an element wherein high 

stress concentrations are present in the residual structure. 

g a Assessing design detail areas which are prone to fatigue damage 

based on service experience records of similarly designed 

components. 

17 l FAILSAFE DEMONSTRATIO!L 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 
5/33/74 AC 20-95 

failure is extremely remote. Tha loading spectrum should include 

at least one application of limit load. In test 

demonstrations, the damage may be lnit%ated or simulated by cuts 

made with a fine saw, sharp blade,or guillotiae In those-"cases where 

it is not necessary and not practical to produce fatigue cracks by 

tests. In those cases where damage is airmrlated at joints or fittings, 

bolts may be removed to simulate failure if this condition would be 

representative of an actual failure. In 8ome instances, the failsafe characteristics may be shown analytically. The aaalyelcal 

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 0 TNS~AECTION. 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 l GROUND CONDITIONS . . -1.5 

(a) Racid increase of r.p.m. on-ground to q@cUzwG clutcbO.5 

(b) Taxi'ing with frill cyclic control 0-v 0 5 . 

(c) Jump takeoff --am- (1) 

(d) Ground-air-ground cycle (1) . 

II l HOVERING -- a-w- "-" a---n -2.0 - 

(a) Steady hovering -'~~~~-n~~n-vnNov-~~~-~---- 0 5 

(b) Lateral reversal---------- v-n--o- --o.-- 0 5 

(c) Longitudinal reversal------ ~~o~wov~~--- 0 5 

(d) Rudder reversal- ' -MD-_- ow~-~-vuu 5 

III. FORWARD FLIGHT-POWER ON----- noo~~~~~~--~- '87 0 S . 

(a) Level flight - 20% VNE-- . vuv~~nu~~~vn~~o~~~-~~~ 1.0 

(b) Level flight - 40% V~------------------------3.0 

(c) Level flight - 60% VNE----------------------------- 18.0 . 

(d) Level flight - 80% VW--- ----------..25.0 

(e) Maximum level flight (but not greater than V&--------15.0 l 

. 

(f) VNE-~~~~-~~-~~-~-~ ~-no~------------~-~~3~0 

w 

09 

(0 

(3 ) 

(k) 

(1) 

( m 1 

( n 1 

( 0 1 

(P) 

111% VNE ~~----n~--n a’~~~-~~~~c----~~~- 5 

Right turns n 30, 60, 90% ~--an-n---- (2) 3:o 

Left t&s - 30, 60, 90% vm------------- ----o----(2) 3.0 - 

. 

Climb (Takeoff power)--------- n--m --- 2.0 

Climb (Max. continuous power) --~~~n~~~~.~~~- 4.0 

Change to autorotation from power-on-flight - 30, 60, 90% V&5 

Partial power descent (including condition of zero flow 

through rotor)------- --~--___- ~~~n-~~~-I, 2.0 l 

Cyclic and collective pull-up from level flight------o(2) 

. Pushovers -~~~~~~~-~o~vo~~~~~~~~~---~-- (2) 

(q) Quick stops ~-~-~-~~~~-~~~~---n- aomn--nn,- (2) 

(r) Flares ~-~~~-n-~nw~n-~w-nn ---00-vnv- (2) e 

(s) Lateral reversals 'at VH---------- -on-n--v- 0 5 

. 

(t) Longitudina1 reversals at VH----------- ---so-- 0 5 

. 

(u) Rudder reversals at VH -v-nnvmn '-n-WV----- 5 

(v) Landing approach--- -~ovo~-no-~-~nnoov~--- 3'0 0 

(w) Sideward flight----. v~v~n~-n--~- -n-n-- 0 5 

. (x) Rearward flight v ~--n-n~n~nv Pm-- v 0 5 

IV AIJTOROTATION -POWER OFF---------- * -. 

-vvonnn--- . 9.0 

(a) Steady forward flight---------------- -m.--- 2.0 

(b) Rapid power recovery from autorotational flight---- 5 

(c) Right turns - 30, 60, 90% vm-- -n--n--o-- (2) 1:o 

(d) Left turns - 30, 60, 90% VNE------------ -n--v- (2) 1.0 

Page 14 
S/18/76 AC XL% 

(e) Lateral reversals-~----~nn--n----- 0 5 

(f) Longitudinal reversals- m-J-n,- --n--o" 0 3 

(g) Rudder reversals------ ye--- ~N--v-y' 0 5 

(h) Cyclic and collective pull-ups -rrrnr~-~rr-~rr---- (2) 

(i) landing approach---------- -ma-am-mwmmn-- 20 

(1) We flight every $0 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 specif led per cent of 

occurrence. 

Page 15 
AC -95 S/18/76 

NOTE: DASHED LINES IN FIGURES 1 & 2 INDICATE 

TEST BOUNDARIES. CROSS-HATCHED AREAS 

INDICATE OPERATING REGIMES. 

VNE 27.1606-, 

MAX. DESIGN OR 

DEMONSTRATED RPM 27.33,27,309 I/ 

11% VNE 

27.33, 27.309 

AIRSPEED 

FIGURE 1. POWER ONROTOR RPM Al RSPEED ENVELOPE 

I VNE 27.1696 + 

MAX. DESIGN OR 

I 

DEMONSTRATED RPM 27.33,27.309 

~~---~-~~~ --w-w 

M 

z 

a 

a 

0 

IO 

a 

MIN. DESIGN OR 

--~~~~~---- 

DEMONSTRATED RPM 

AIRSPEED 

STEADY STRESS 

FlGURE3. Oscillatory Stress 

vs. Steady Stress 

STEADY STRESS LEVEL-A 

I TEST DATA CURVE 

NO FAILURE 

AT REQUIRED 

NUMBER OF 

CYCLES (6x107) 

REDUCED CURVE 

1 

1 

I 

I 

u N 

I 

FiGURE 2 POWER OFFROTOR R.P.M. AIRSPEED ENVELOPE 

FIGUREA Stress vs. 

Number of Cycles 

Page 16 

_ . . A -, 

-*- --.- .-.- 

-9 F