National TIchnicallnfonnation DEPARTMENT OF COMMERCE · Mobility Research anc Develo,ment...
Transcript of National TIchnicallnfonnation DEPARTMENT OF COMMERCE · Mobility Research anc Develo,ment...
AD-772949
CH-47A DESIGN AND OPERATIONAL FLiGHTLOADS STUDY
A. Herskovitz, et al
Boeing Vertol Company
Prepared for:
ArillY Air Mobility Research and DevelopmentLaboratory
November 1973
DISTRIBUTED BY:
National TIchnicallnfonnation SeniteU. S. DEPARTMENT OF COMMERCE5285 Port Royal Road, Springfield Va. 22151
BestAvailable
Copy
r¥=.4tQ!j@i1tPJ2£%~j- ~,;4£me3Q3@..,t.4._~¥,!
<R:;>
~
-"'-""""~
DEPARTMENT C:f' THE ARMYu. S ARMY A!1t ,,"OBILITY RESEARC~ " OEVELO~ENT LABORATORY
EUSTIS DIRECTORATEFORT EUSTIS. VIRGINIA 236~
This program was conducted under Contract DAAJ02-72-C-0087 withThe Boeing Company, Vertol Division.
The inforruation presented herein is the result of an analy~ical effortto derive improved structural design criteria for transport-typehelicopters based upon flight parametecs measured on transporthelicopters operating in Southeast Asia. This is one of four similarefforts being conducted concurrently to develop improved criteria forutility, crane, and observation as well as transport-type helicopters.
The report has been reviewed by the Eustis Directorate, U.S. Army AirMobility Research anc Develo,ment Laboratory alld is considered t:> betechnically 30und. It is published for the exchange of information andthe stioulation of future research.
This program was conducted under the technical management of Mr. HermanI. MacDonald, Jr., Technology Applications Division.
Unclassified
I
••_&c•• DO~_ len. I ~_N. _C" ••O_L.T..._ "'..... us..DD,~••1473
Unclassified A]) - /]1,1 CjLjCJ5Kurily a ....ir..,.li_
DOCUMENT CONTROL DAT 4 • R.. D .{s.curl'r cr...,lIc..,_.1 ""e.~ .,....-._1~"'1t4_'0"_ ......, w __.... o".,en _,. cI....UH
t OJlIGINATING ACT.v,T ... !c.".,.......} aa.. ___0." SCCu•• TY CLA...",CATtON
Boeing Vertol Company UnclassifiedA Division of The Boeing Company u. ....0\1"
~
Philadelphia, PennsylvaniaJ -EPO"" TITLE
CII-47A DESIGN A..\O OPERATIONAL FtIG:rr LOADS SIDDY
• DlE.C'''''1".VlE "OTlE.~ .,_,__,..., ......)
Final Reports AU THolles, (,.,., ..... 8l~. ,.,,,.,.....,~)
A. HersKovitzH. SteiJUllann
e.....C~O.. T OATC 7.. TOTAL ItO. OF ..aca. 1''' NO. 0.; "....November 1973 .~ -_.erg
... CONTRaCT 011 GIIANT NO• i- 0"'"'''''' TO"'. R."O"1" ..........lal
.. Plb~Qh",~-C-0087US~IRDL Technical Report 73-40
c. Task 1F162204AA8201 M. OTM£" __"T NOla' ("'"____.. _ ........_)Co 0210-10579-1
to. OtSTtlta"TtON STATCM£..T
Approved for public release; distribution unlimited.11. SU."LC"CNTA"Y NOT£S 12'.. .-o-.SO"'NG MtLITA.Y ACTIVITY
Eustis DirectcrateU. S. Army Air Mobility R&D Lab"ratoryFort Eustis, Virginia
u A.,T"aCT •ll&e purpose of this study was to evaluate the adequacy of current structural designcriteria for future cargo- and transport-t)~e helicopters based on the design,development, and operational use of the CIf-47A Olinook helicopter. It was &:on-eluded that current structural design criteria are adequate to insure structuralsafet)·. Specifications for procureaent of new ~elicopters should be modified toprovide the most realistic mission description possible for fatigue design, withthe objective of sicplifying the design task.
"''bile analyzing CII-47A operational data, several deficiencies were identified inthe data acquisition and analysis process. The deficiencies can be overcoze infuture field survey worl by cooperative advanced planning between the cognizantArJAy agency, the helicopter manufacturer, and the contractor responsible for dataacquisition and 3nalysis. Current state-of-the-art recording systems and automateddata reduction and analysis techniques are reco~cnded for future surveys.
". . ~ ~ ~ ",-,~
,,:a," I.(J\l.,. ':" ~"''''-J._& _ .... ;-- l '!' ~:-- - .
··0
'-_ .
~ -,F- ~~.-... • ( " -'=--'-'~ ~
~.~ -~ .T-. 0 ".; .. "'.
r""l'-"""""-........·----""'----~>J<-#--~"'~ ~..,..=~.-~~- ... """""1
f~ .~
~f!'
i!l
Unclassified
ty •• ea Oft
'4 LINK" LIfr.,( • LINk Ck~Y woaos -a." L IE. WT "OLE WT .t")LE "T
Cli-47'\ helicopterFlight 1 ..~~ stud)·Str~ctura~ design ~riteria
Cargo-type helicoptersTrans~rt-~;pc helicopters
I
I.
I
II
I
r!!!lM)gg;-M!i~-'_""il!@!_'M"""--' '"""'-".~~,., -j".t=c""<"'--'''"--''~''''
i~
i11"
Task lr162204AAB?~1
Contract DAP~02-72-C-0087
USAAMRDL Techrical Report 73-40November 1973
CH-47A DESIGN AND OPERATIONALFLIGHT LOADS STUDY
By
A. HerskovitzH. Steinmann
Prepared by
Boeing Vertol CompanyfA Division of The Boeing Company}
for
EUSTIS DIRECTORATEU.S •.Z\RMY .'HR MOBILITY RESEARCH AND DEVELCPl4ENT LABORATORY
FORT EUSTIS, VIRGINIA
Approved foz public release;distribution unlimited •..,',f,
--~
,_~E£k;;,~."},~c q~;:osv~,"="",,",_,_,,,__!,~,,,,,,,__~-,,,y.,-=,,-..c=,~-,c",c=" ......,,..,
1,,ABSTRAC1'
The purpcse of this study was to evaluate the adequacy ofcurrent structural design criteria for future cargo- andtransport-type helicopters based on the design, development,and operational use of the CH-47A Chinook helicopter. Itwas concluded that current structural 5esign criteria areadequate to insure structural safety. Specifications forprocurement of new helicopters should be moaified to providethe most realistiL mission description possible for fatiguedesign, with the objective of simplifying the design tc~k.
While analyzing CH-47A operational data, several deficiencieswere identified in the data acqu~sition and analysis process.The deficiencies can be overcome in future field survey workby cooperative advanced planning between the cognizapt Armyagency, the helicopter manufacturer. and the contractor responsible for data acquisition and analysis. Current stateof-the-art recording systems and automated data reductivn andanalysis techniques are reco~~ended for future surveys.
iii
TABLE OF CONTENTS
ABSTRACT • • • • • • .
LIST OF ILLUSTRATIONS
LIST OF TABLES • • • •
LIST OF ABBREVIATIONS AND SYMBOLS
· .· . . .
• iii
vi
vii
viii
1. INTRODUCTION • • · . . . . . . . . . . . . 1
2. .'HSSION PROFILE · · · · · . . · · · · 3
3. FATIGUE DAMAGE · · · · 13
4. AIRCRAFT CONFIGURATION . . . . . · · · · · . .. · · 24
5. OPERATING LIMITATIONS · · · 32
6. CONCLUSIONS AND REC01-lMENDAT101 is · · · · 39
7. LITERATURE CITED · . . . · · · · · · · 42
DISTRIBUTION . . · . . . . . · · · · · 43
Preceding page blankv
1
2
Percentage of Vne Occurrence Shows Importanceof Extracting Data in the Fo~ to be Usedin Analysis • . • . . • • . • • • • • •
Comparison of r-tis~;ion Profile£ ;md Operation.llUse of the CH-47A Helicopter • • • • • . •
7
• • 12
5 CH-47A Aft. Rotor Blade Spar Fatigue I.oau Ex-ceedance for Four Mission Profiles • • • 22
Normal Load Factor Exceedance for OperationalData and Missio~ Profiles
3
Elements of Fatigue Lif~ Evaluation
14
. 16
6 CH-47A Aft Pivoting Actuator Fatigue LoadExceedance for Four Mission Profiles • • 22
7 Structural Airspeed Limitations for CH-47 Helicopters at 6,000 Feet Density Altitude •••••• 31
8 Percent of Steady-State Time for 98.2 Hours(USA) and 135.9 Hours (SEA) of CH-47A OperationalData at Gros~ Weights Below 28,000 Pounds •••• 34
9 Percent of Steady-State Time for 9.4 Hours(USA) and 18.9 Hours (SEA) of CH-47A OperationalData at Gross Weights Above 28,000 Pounds •••• 35
10 CH-47A ~aneuvering Load Factors Relative toDesign Envelope . • • • • . • • • • • • • • • 37
11 Operational Load Factors Extrapolated to HighFlight Hours •. • • • • • • • • • • • • • • • • • 37
5
2
Page
. . .
LIST OF TABLES
Criteria Used for Determination of FlightSegment from Measurements in References 1and 2 •••••.•• • • • ~ • • • •
Summary of Significant Characteristics ofA, B, and C Models of the Chinook Helicopter
I
II
Table
r--""'''''~----~=''~~'~~'~''~'~~ ~"'~,-..~~--"Ii..~~~
I
IX S:.m'u'nary of CH-47 Cha:lges to Improve Fatiguednd Ultimate Strength of Components ••••••• 27
Summary of 207 CH-47 Changes by HelicopterSystem and Reason for Change • • • • • • • • • •• 25
Load Factor <nz } Peak Frequency as DeterminedFeom CH-47A Flight Strain Survey • A • • • • 10
8Redistribution of Time for Mission Segments
Fatig~e Damage Rate for Aft Fotor Blade Sparand Aft Pivoting Actuator by Flight Conditionand Mission Profile •••••••.••• 20
Comparison of Various Fatigue Mission Profilesby Operating Condition and by Mission Segment 11
CH-47A Part Lives for Various Mission Profiles •• 18
VIII
III
IV
V
~
~"VI
,-VII
N~~~~-~~~--'---- --~~
-~
"]r-h
•
iI~
IAccel.AIRcgDesECPEJSgG-A-GmolHdlASISAnzPPDSRDTOVHVneWtG
~
LIST OF ABBREVIATIONS AND SYMBOLS
AccelerationAutorotationCenter of gravityDescentEngineering change proposalEngineering job sheetAcceleration due to gravity, ft/sec 2Ground-air-groundGross weight, lbDensity altitude, ftIndicated airspeedInternational standard atmo~pheLe
Load factor in helicopter z-axis (vertical)Partial power descentService revealed difficultyTakeoffAirspeed limit, used interchangeably with VneAirspeed limit, used interchangeably with VHWeightStandard statistical deviationImplies
viii
1. INTRODUCTION
rr _".~,""~~,,,~.=,'C_ ..,~~ ~--~~~-~=-~~~~-~-~-~ --. -~,~- _-__ c- ~
r:[
l~~~~~
1~~
"
BACKGROUND
CH-i.7A operational data \<lere collected by the U. ~. Army inboth simulated and actual combct conditions as reported inReferences 1 and 2. The combat data were separately analyzedby the Boeing Company under u.s. Army contract and reportedin Reference 3.
The twin tandem CH-47A Chinook helicopter was designed to meetU.S. Army medium lift requirements as a personnel transportand cargo carrier. Cargo can be transp0Lted either internallyor externally. The CH-41n a~d CH-47C models of the Chinookwere developed to meet increased payload and range requirements,with particuloL concern for operations in hot climates.Table I surr~arizes some signiiicant characteristics of thethree models.
This design and operational flight loads study is centeredabout the CH-47A model of the Chinook because the operationaldata were obtained on the A model. Some of the lessonslearned from the CH-47A field data resulted in design improvements incorporated in the Band C models. CH-47B and CH-47Cdesign features which are baseG o~ CH-47A experience arenoted in £ection 4.
DESCRIPTION OF TASKS
This report follow~ the basic sequence of the separate studiesconducted under the contract:
Section 2MISSIO~ PROFI~E
Section 3FATIGUE Dl'..MAGE
Section 4AIRCRAFTCONFIGURATION
Section 5OPERATINGLIMITATIONS
Compares two independent analyses ofcombat data, constructs a new missionprofile based on field data, and comparesvarious mission profiles.
Evaluates load spectrum, fatigue damagerate, and calculated fatigue life forsix components and various mission profiles.
Correlates configuration changes in theCH-47 history with mission profile changes.
Investigates the factors which limited theCH-47A ~perations in simulated anr actualcombat conditions.
1
"·"'.....,..rr~h1t5Bt5--itW. 7T5S.WmettW~~--nm=~T77e7 rntr?Wf~-tWnir.'WWwrnsF--5"i5wnrss·iiiI
TABLE I. SUMMARY OF SIGNIFICAN~ CHARACTERISTICSOF A, B, AND C MODELS OF THE CHINOOKHELICOPTER
MODELDESCRIPTION
CH-47A CH-47B CH-47C
Number of blades per rotor 3 3 3Airfoil Symmetrical Cambered CamberedRotor dIameter, ft 59 60 60Constant chord, in. 23 25.25 25.25 <!:Rotor speed,rpm-Power on 230 225-230 235-245 1
-Power off 261 261 261Gross weight,lb-Design 28,550 33,000 33,00°0
-Maximum 33,000 40,000 46,000Load factor, nz,at design -0.5 to -0.5 to -0.5 to
GW 2.67 +3 +3Limit true airspeed, kt, atdesign GW 130 170 170
Payload, Ib, at SL/STD,50-mile radius 12,300 17,400 20,950
NOTE: @ CH-47C capability at 50,000 pounds gross weightand 250 rpm has been substantiated by the con-tractor and evaluated by the U.S. Army.
2
Recommends adequacy and/or deficienciesin structural design criteria for cargoand transport-type helicopters. Additi~nal recommendations are made forsta'~-of-the-art improvements in fielddata acquisition and analysis techniques.
2. MISSION PROFILE
This report section comprises a comparison of mission profilesas used to structurally design and evaluate the CH-47A withthe actual helicopter usage as reported in References 1. 2,and 3.
M~ssion profiles are used for various evaluations in helicopterdesign and analysi.;, such as pcyload-range, duty cycle, lowcycle and high cycle part fatigue, and part wear. The philosophy of mission profile construction for the various evaluations is beyond the scope of this study, bu~ the philosophymust ~ecessarily vary with the purpose and risks associatedwith a particular evaluation.
This section is dir~cted at the fatigue mission profile:
Fatiguc_Mis3ion Pro£il~ is us~d for fatigue design and toc?lculate the safe-life retirement interval for criticalcomponents. It is also used to design wear-criticalcomponents and to calculate required overhaul intervals.
DATA SOURCE EVALUATION
An understanding of the content and 11mitations of the fielddata sources is essential to a rational interpre~~tion of theinformation.
The total accuracy of a measurement should be considered,but cannot readily'be established from the information available.The "Quality Control Values" tabulated in R.aferences 1 and2 indicate that reasonable precedures were used to minimizedata reducti0n errors for the measured values. However, theaccuracies of sensors, amplifiers, and powe~ ~~~rces were notincluded, and procedures for establishing calibration factorswere not identified. The lack (·f i! :ormation on accurccy doesnot preclude an analysis of the cat~, but the possible implications of errors in the data should be considered in formingconclusions.
~,
~'
~
ci:-
_ -_~'"*-<E-~ :~~~~"':.~~r-~~~--~ ."""_".......,,-~_ d'J-~ _ -r-.-"'=--,<~=-;.---~~_O-r-=-.-=..-~,;.~-~ ~_""-~~~"""_-t"~:~~~~"" ~"--~.F~:::E__ "'~_"'r~.d'~~r=.[;=-::'.=-_=-r_=- ~~~ ......-~~~~ __%-_ .._~>~"':~~--=..:a>=:..~;?~~~~---""'--,...====~? ...
1
The source of gross wei¥ht information was not specified. Itis assumed that this in ormat ion comes from a log maintainE·d byflight crews.
No information on center of gravity distribution is provided.
Flight Segments (steady state, ascent, descent, and maneuver)were identified by the data characteristics of longitudinalcontrol displacement, collective control displacement, airspeed,altitude, and normal acceleration as sho~~ in Table II. Obsp-rvethat yaw and roll displacements cannot be identified from thedata.
Acceleration Peaks were not evaluated between 0.8 and 1.2g inthe data reduction process. This threshold limitation would exclude, for example, a steady 330 banked turn. Acceleration peaksless than 0.8g and greater than 1.2g were subdivided into maneu~- or gust-induced accelerations. An acceleration peak wasclassified as maneuver induced if either or both of the longitudinal and collective control positions were displaced justprior to the observed acceleration. Using the same example asbefore, the load factor developed in a turn could be clas~ified
as either maneuver or gust induced depending on the longitudinaland/or collective control exercised in conjunction with theturn.
Finally, th~ uata samples for simulated and actual combat operations each re?resent the activity of one operational unit in alimited locale for a time period much less than a year. Conclusions drawn from this data with respect to aircraft utilization should consider the possible effects of aifferent missionsand environments.
DATA ANALYSIS PROCEDURE
The operational data obt;tined in actual combat conditions wereseparately an~lyzed and reported in References 2 and 3. Eachanalysis used the same reduced data, but apparently with different objectives. Reference 2, prouuced for USAAMRDL by Technology, Inc., appears to be tailored for easy comparison to operational data gathered on other types of helicopters. Reference3, prepared for USAAVSCOM by the Boeing Company, was compiledfor the speci~ic purpose of fatigue mission profile evaluation.
Except for the parameter increrrents chosen, the documents werefo~nd to be in good agreement with respect to gross weight, density altitude, rotor sFe'~d, aI\d accel{.'ration pE:3k occurrences.Th\ R~ference 2 analy~~s included a 3~-minute flight at approxim?~ely 37,000 pounds gross weight in a "greater than 32,000ponnd" increment. The gross weight for this flight was more accurately defined in Reference 3.
TABLE II. CRITERIA USED FOR DETERMINATION OFFLIGHT SEGMENT FROM MEASUREMENTS INREFERENCES 1 AND 2
Flight segment Transient SteadyMeasurement - Ascent Descent Maneuver state
Longitl~dinal Not Not Not Rel"\tivelycontrol steady steady steady steadyposition
Collective Not Not Not Relativelycontrol steady steady steady steadyposition
Airspeed Frequent Frequent Frequent Steady orchanges changes changes smooth change
Altitude Freouent Frequent Frequent Steady orchanges changes changes smooth change(increas- (decreas-ing) ing)
Normal No No URually Noacceleration, criteria criteria very criterianz active
5
Evaluation of airspeed was significantly different in the twodocuments. In Reference 2, indicated airspeed (lAS) is shownas a time increment within designated boundaries of gross weight,altitude, and indicated airspeed. In Reference 3, the analystwas interested in airspeed as a per.centage of Vne (never-exceedairspeed = flight manual limitation). Since Vne varies withgross weight, altitude, and rotor speed, the percent of Vne wasca1cu1~ted at each discrete time increment of data reduction.The results were presented as a histogram of percent of Vneincreinents.
An attempt was made to express the lAS data of Peference 2 inpercent of Vne • The results, shown in Figure 1, are less thansatisfactory. The wide band of possible solutions arises fromthe 1argp variation of Vne within the gross weight, altitude,and airspeed boundaries which define an lAS occurrence. Similarproblems would be encountered in attempting to determine lASdistr1bution from the percent of Vne histograms.
MISSION ?ROFILE DERIVATION
Mission Segment Analysis
The data analysis procedure employed in References land 2separated transient flight conditions into three flight segments identified as ascent, d~scent, and maneuver. A fourthsegment, ste~dy state, inc1udec climbing, :eve1, and descendingsteady flight conditions. The ascent se~ment also included"the takeoff and climb to the initial steady-flight altitude~,
and the descent segment included "the unsteady part of flareand landing". Table III shows that the ra~io of maneuveringtime from level flight conditions to the time spent in steadylevel flight was significantly less than the maneuver to steadytime ratios obtained in ascent and descent conditions. The acceleration tc climb airspeed and flare to landing can be accepted as maneuvers peculiar to the ascent and descent segmentsrespectively. However, other maneu~ers such as turns, pull-ups,pushovers, and control reversals sLould be nearly as common toconstant altitude operations as to climbing and descendingflight. In addition, in Reference 2, normal load factor peaksoccur 2.8 and 4.1 times more frequ~ntly ip. the maneuver segmentthan in the ascent and descent seg.ents respectively.
The observations of the preceding -aragraph led to the conclusion that a significant amount of tle time in the ascent and descent mission segments of Referenc!s 1 and 2 was, in reality,ste&dy ascent and descent. Accorc.ng1y, the flight time wasre-allocated as shown in Table III.
6
;~
1
100
- -- - --~-----=--~-~- ~---~ --~-~_..... ---,,"""~7"">'~';;"'-"""'~- ~- -. --
5f
SAMPLE HAS SOLUTIONS BETWEEN90 AND 152% Vne FOR 3.1% OFTIME
10
60 80 100
INDICATED AIRSPLED - KT
5,000
REF 2 SAMPLE
GW S 28,000 LB285.5 MIN~.1% OF DATA
10, 000 ;"--;~r------"T---iiF~-
DENSITYALTITUDEFT
~ = Figure 1. Percentage of Vne Occurrence ShowsImportance of Extracting Data inthe Form to be Used in Analysis.
7
l'lj!i~;m~~I!1i~~li~~'~1i1Ir,'!;!'1Ifi~~~,
iI"iI
,r:r,~~'i~L1T"
I,II~
it:Wl1
1''r''
ll;«I'
r.,:I
llr;tt
, "...."
i~,t\
,'1·""ri~\f1'·~i:~r~,t:'~~~'llr)
H~;'1'i
:trr..,
"lIf",~,;,,'H
;1'11i;il1iHii~iltf,;~!!';,":'':I'''''I;
"i"ii
~~r,;i~
Pl'Fil"l'lI'r;\r,;rN~f,n~'JI'l;;',")i~n~"lI;n
~lt~;,~r;rr"w~r
"l',f
ln~il
1ll'i'
i'NI'~
,n}lii
f"~l
~"l
l~~r
~'U'
"1Jf
."J}
!l""
il1'
~rr,
'~I'
,,'i
f,j'
i~~(
ljJI
j'r,
~~'~
ll~'
Rl'~
i'11
~~~!
'..."
"""'
!!,~;I
'~~
1 11 ;~ :'~ ~j 1<
,!l,
co
TABL
EII
I.R
EDIS
TRIB
lJ'L
'ION
OF
TIM
EF
onM
IS$!
ON
SEG
MEN
TS
Fli
qh
tS
eqm
ent
Perc
en
tO
ccu
rren
ceT
ran
sien
tD
eriv
edS
tead
yM
issi
on
Asc
ent
Des
cen
tM
a,:!
uver
Sta
teP
rofi
leD
esc
rrp
tTO
-nR
efR
efR
efR
efF,
efR
efR
efR
ef1
21
21
21
2
Val
ues
inR
ef1
&2
11
.80
12
.50
16
.81
19
.97
6.4
81
.87
64
.91
65
.66
-R
epo
rts
Red
istr
ibu
tio
nu
sed
for
mis
sio
np
rofi
led
eri
vati
on
Ste
ady
lev
el
(±.5
00fp
mR
IC)
63
.48
59
.40
60
Ste
ady
asc
en
t(>
500
fpm
RiC
)7
.11
8.2
80
.97
4.3
611
Ste
ady
desc
en
t«-
500
fpm
RiC
)1
1.7
11
5.6
00
.46
1.9
016
Man
euv
er-f
rom
19
vel
flig
ht
6.4
81
.87
Man
euve
rin
Asc
ent
4.6
94
.22
13
Man
euve
rin
Des
cen
t5
.10
4.3
7IT
OTA
I,:;
';)0
Met
hod
of
red
istr
ibu
tio
no
fti
me
seg
men
ts--
I.S
tead
y-s
tatese~
brea
kdow
nis
bas
edon
rate
-of-
c1
imb
data
rep
ort
ed
inR
efer
ence
sI
an•
2.
Asc
ent
and
desc
en
tse
gm
ent
brea
kdow
nis
bas
edo
nth
efo
llo
win
glo
gic
:A
.Thirt~
seco
nd
sp
er
flIg
ht
isas
sum
edfo
raccele
rati
on
tocl
imb
air
speed
inasc
en
tse
gm
ent,
and
thir
tyse
con
ds
per
flig
ht
isas
sum
edfo
rla
nd
ing
flare
ind
esc
en
tse
gm
ent.
B.
The
rati
os
of
tim
esfo
rasc
en
tm
aneu
ver
sto
stead
yasc
en
tan
dd
esc
en
tm
aneu
-v
ers
tost
ead
yd
esc
en
tw
ere
assu
med
eq
ual
toth
era
tio
of
tim
esfo
rle
vel
maneuv~rs
tost
ead
yle
vel
flig
ht.
The
rati
o,
det
erm
ined
from
lev
el
co
nd
i-ti
on
s,w
asap
pli
ed
toall
asc
en
tan
dd
esc
en
tti
me
exce
pt
for
the
accele
ra-
tio
nan
dfl
are
tim
eid
en
tifi
ed
in(1
\)ab
ov
e.
:1, i:l Ii I' " :\ " 'j' ::~ 'rl,~ :i' ~~ ,<,I :~1!. I,' :i~ 'I" 'it" ;/, ; " ;~ )~, \1' ~ Ill' jl '1,1: i I' tj ':j IIII I <,.
i q ,l I ~ i 1 I I 1 I~''l 'I'
."t,
!,
Derived Mission Profile
Normalization of Mission Profiles
The normalization of mission profiles was accomplished by thefollowing process:
profile derived from the operational dataand 3 is shown in Table V. The distribumission segments was based on the analysis
The basis of other distribution factors
Gross Weight Cumulative % Time at(lb) or Below
Derived Ref RefProfile 1 2 & 3
25,000 66 85 6428,550 92 92.5 9033,000 100 99.5 99.8
9
Increments selected to be consistent withavailable CH-47A flight strain survey testdata.
1. Gross weight
3. The fractional times were proportionately normalized tounity and expressed in percent of the total missionprofile.
2. Each operating condition in the mission profile was expressed as a fraction of a common time base {I hour wasused). The s~ of these fractions then exceeded unity.
1. Th:~ time required to conduct a maneuver was determinedfrom Cn-47A flight strain survey testing, as shown inTable TV.
The derived mission profile i~ compared to the operational dataand to the original design, current, and AR-56 mission profilesin Figure 2.
The nonoalized mission profiles are displayed in rable V forReference 4 (current) and Reference 5 (AR-56). The designmission profile from Reference 6 and the mission profile derived in the Derived Mission Profile section are also sho~~
in Table V.
The fatigue mission profiles defined in References 4 and 5include some operating conditions with occurrence factors expressed in number of events per unit time. For comparison ofmi~sion profiles, it was desirable to express all mission segments in percent of the total mission profile.
The fatigue missionof References 1, 2,tion of time into 4shown in Table III.follows:
1i~1
~~II
J":'
lft[
~~~m
r~~'
;~~~
11~N;J
Il'jl:
IIII'.
!;1iM
"!:T't<
'llWl'I
'I~'~l
l'~~il
tt;JI;
1~i'n~
'I'Mr
II,t!"
1.'rfl
,rn7J
';'IJi
'J~~
T~'Fi.
fli:1<
:flljH
~~'iTi
l1np"~
:r~ln~
liFllf
~m,rfT
llim:~
1i[1tF
',11l"
r';ll'
lI'1i
"tl'li
'l~rr~
l"if<i
~~i'~J
:m'Uk1
t!~',l
!}<"Ii'
)lf.~,
~I::NVA
~(li::
~f,iIN"q,~~1t~fM~~j';tn~l~~~~"';J'tIl:r,i>~!'''''h~,
~~I\q
;JM~""nVi~M~1l1'''tIl~:~li9l
t~I,~t
iIJI"l
Ij;!il
!ijfll
;JI'l~
1!1~rl
't~~(J
r;'~'i
%,'tiJ
~':'i1
jiN'f'
i~~"'j
;:Pf~~
,\!!l;
~li,i"
"'iltl
i'i'i
I,:, l ~.~ '1
j I iJ /, I ) 1 t 1 ,l ,; l I ~ I ! I' II
.... o
'I'AU
T.!::
IV.
LOA
DF'
AC
'I'O
U(n~)
PI~1\1<
!-'U
F.Q
lJF:
NC
Y1\
Sm
::'I'
F.I
lMIN
lmFR
Ot-!
CII
-47
/\1·'
1.ta
ll'!
'S
'l'H
AJN
!>lIH
VI-:
Y_
__.
-_.
."-'-
'.'
..=
'.
..
..-
_.._
--"
.....
._._-+I-:~li9ht
Hel
;()r
c!s----.
_L
one!
Fil
et'o
r(n
zl
P(,l
lk~/
j1ou
rCD
Nil'
I.en
gth
----
-:..-
----
----
-1~I
iIl
ui"
nP
rof
i1t'
Sec
,.
-+r.
I.on
d0
pN
'nti
nq
Con
<li
l'it>
r\G
)CD
I.,O
ilel
Fn
ctO
lH
ilng
c..
.:.O
.():
>F
acto
r0
.20
.70
.80
.91
.11
.21
.31
.41
.75
Ste
ad
yC
onc!
.ili
on
ll!
!lo
ver
.6
12
0.7
17
4.0
'l'r
un
sit
ion
174
P.'
J7
4.2
O,2
VII
92
6.5
11
6.0
0.7
VII
'0
.8"1
1'
O,C
)VII
123
5.6
40
4.0
VH
5I!
>•
6i
36
7.0
1.
)VII
51
5.2
I?
3oJ
.01
..t5
VII
12
I3
9.8
Un.
.°1
.2V
"7
I2
4.
[)1
46
.0C
lim
bS
11
2.0
46
.0P~rtia1
pcw
erJU
8cen
tJ?
11
9.0
~O.9
60
.6Autorot~tion
171
16
.03
1.0
31
.06
1.Q
Q2
.83
1.0
M~ncuvcr
Co
nd
itio
ns
Sid
ew
ard
Cit
,re
arw
ard
flig
ht
92
6.5
1J6
.CAcccler~tion
tocli
mb
air
sp
eed
13
0.0
12
0.0
12
0.0
rull
-up
,1.
4Cl
(,A
/Pr.
,7
.10
';0
7.0
50
7.0
PU
l.l-
Up
,1
.'7
5g
3I
7.L
OI
50
7.0
PU
!lh
ov
ert
7.1
05
07
.0L
eft
turn
55
3.0
67
.Q
nig
ht
turn
56
4.7
55
.6L
on
git
ud
inal
co
ntr
ol
rev
ers
al
57.~0
48
0.0
48
0.0
r.~tcr~1
co
ntl
'ol
rov
crs
al
5I
l>.O
O4
50
.04
50
.0D
irecti
on
n1
co
ntr
ol
rovers~l
~I
6.4
05
62
.0S
62
.0L
an
din
gappro~ch
and
Fl~ro
41
0.0
12
0.0
12
0.0
N01'
I~S:
CDn
zp
eak
s/h
r..
(No
.o
fn
zp
allk
so
bse
r.v
edin
lon
dfa
cto
rrn
ng
e.;.
rcco
r<'
lon
qth
inse
(.:o
nd
s)x
,360
0(]
)R
eco
rdll
'nq
thfo
rfl
ten
dy
co
nd
itio
n!>
..su
mo
fall
flig
ht
reco
rclf
l~
Hec
ord
len
gth
for
man
eUV
erco
nd
itio
ns
..le
ng
tho
flo
ng
cll
t(l
igh
tre
co
rd
.~:~ ,,' j, I~ III 1:~ "IliI .~I I'! '!~! \' I,! 1~l ~I} j,:1 I~~
1,-
" :~:i'
,
,':1
;1:1
',
i I, l~ I" 1:(' ';1 Ii' 1 1
:'1'
1 :1 1'
Ij
cd J,""'
I",,,
,,,j",
..t.t,,"~!~(
0 .. :
..v •
4.1
:. '.
1.1I.:
'.;'
".1
~. '-1.6
0.2
... - ~. ,
-.¢
: ..
... :
~.:
n .. 2*\.:- ,
:.il: .~
: . .-.
- ---- -- --,,--,-..,...----.----~::-r,-.
~.
h
,-._---
]
i. -
: ..... ",
':'~h:-! ...... ~~f'N.PA~~S-~; nt' v; R%"')i:~ ~'A':'h~ r ~tS<'~(\·; pj )rI:'''~ ~y
"'rfY..A'"::~;C ~••,~,,:,-:~; 1"'-::> 8Y ~:~~:("~: $t~!.""::-
c- -~~.~-'O~~='---~==""======-=~_~-=~=
--~ ~-------------- ----~--- ------".. ~ ..
:...-... ~~ : F·:"":... ~.~~~ : ..~-..,.:
- ~ ·h- ". - '-'
". !"':!1~··z ':'.! ..... ~i:
-.::.;'".,,,"'.!_. -'u
... ? ::-,..~ :.':'-=- t-~~:--r ~ .... .:... .:. -:; ::. ::--r:... ;' ... 'S:"; r "'.:~
=-~=__=..;;-=====__=::=_====_==_=_==_==_ -==="d
11
~~E>.)~: 1(·-...........$~.::";i~ r ..:l~-i:--"
.;::-;4""'~: .j,~ ~~ ....-~
'!:.!4-~~£.....~ _
-.r:i'-- -~~---------- ~~ ----- . ~- -~ ... - ---
~ ~ _~ - -----'=>0'--
I .'~- ..l';': .::u;;c ....~· s· ::""_::;:~ .::'~:'-! - V...l.X ,:.:;. ~e'::- ...":-
I - ~ ><r_' •• ?~_~~
~;.)...~"~' .is':-,~:
LEGEND
IDERIVEg FROr-1.~
I OPERATIONAL DATAMISSION PROFILES
DESIGN CURRENT AR-S6
S - STEADY LEVEL FLIGHTA - STEADY ASCENT
.-5- ,.-S ... D - S'I'EADY DESCENT1-1 - HANEUVER
-S-
r-S--s- -S-
- 1-1-~D-'-M- '-D- f-M-
~1-1_I-M- f-D- ~A=
A f-A->-M
~A_ f-D-I-D- f-A-o
80
~
60
~ t:J~I-f
~ 8"
8 40zt:.l
~
U0::t:.l0..
=20-
~
Figure 2. Comparison of Mission Profiles and OperationalUse of the Ch-47A Helicopter.
~-
12
13
3. FATIGUE DAMAGE
Ref2
3.8
Ref1
3.9
Occurrence, %
4.0
DerivedProfi::"e
Occurrence selected to obtain normalload factor (nzl distri~~tion consistentwi th References 1 and 2.
Occurrence based or. Table III analysisof References 1 and 2.
Distribution between partial powerdescent ar'd autorotatio>1 based onReference 5 distribution (AR-56).
Distribution based on R~ference 3analysis of time at pe~cent of Vne(never-exceed airspeed).
.;.,.,.....;e-:-.;;"...;..-..c~1._~",-I_-_=I-n~~_I,~~I,,",_I_-I__IE__'_I-_,!g!lK~~II-_~.&._.~2~2~~._.L~~···_·i2-----1
~
Airspeed
Acceleration toclimb airspeedand landing flare
Steady descent
Symmetric maneuvers, tl.rns, andcontrol reversals
3.
2.
5.
It was intended to include a comparison of design values tothose resulting from operational use of the CH-47A helicopter.However, such a comparison does not provide useful informationfor future helicopter design for the following reasons:
In this s~ction the effect of mission profiles on fatigue lifefor six critical components is evaluated.
4.
a. Fatigue loads measured during developmental testing of treCH-47A exceeded predicted values, particularly in the control system. Subsequent redesign of components was basedon trade-offs considering related costs and structuralperformance.
The incremental normal load factor exceedance for combined gustand maneuver load factor peaks is shown in Figure 3. Operational data is shown as a O.lg band representing the range ofvalues reported in References 1 and 2. Mission profileexceedance was determined from CH-47A flight strain surveydata, Table IV, in conjunction with the operating conditionoccurrence from Table V. The derived mission profile isshown to reasonably represent the operational experience withrespect to the frequency and magnitude of load factor peaks.The operational load factor peak occurrence frequency is significantly exceeded in both the current and AR-56 mission profiles. The maximum load factor to be considered for fatigueanalysis is evaluated in Section 3.
10-2 t---+--+----f--+--+---H-l---4--.J----4--~
14
10- 3 '--_-L._-..L__L.-_.J-_--L__l.-_.J-_-L.__l.----J-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0
LOAD FACTORt>
Figure 3. Normal Load Factor {nz } Exceedance forOper1tiona1 Data and Mission Profiles.
T
0 10r:.:~
u~t:J
"'" 0::0
::::~- U
< 1~0::
0~ Eo<-
Ul0::::>g
10-1
r-*2¥_~'_¥!--'-'h,","'_""""'--h""'-""'~
i 103
I~ It. OPERl~TIONAL DATAEc REF 1 Po.ND 2 COMBINEDI 102 t--_.+~r-1~~-+-_MA_N+-E~U"...V..:::E:=.RJ-A_N-=D~G_U_ST~.__..J;;.;::;:i:J-4_~
Fatigue Strength
Any discussion of fatigue evaluation requires some backgroundknowledge of the method used to assess damage. Figure 4displays the relationship of the elements required to calculate a componE :: life. The Boeing Vertol Company currentlyemploys the following criteria:
• Full-scale flight hardware is bench fatigue tested tofailure. A minimum of six failed specimens is desired.
---~
since the CH-47A wasthe current use ofthree standard deviaused for CH-47A
S-N curve shape is determined f~om coupon tests andmakes maximum use of available sources such as MII.HBK-5B.
Fatigue methodology has been changeddesigned. Of particular importance,statistically significant mean minustion (M-30) endurance limits was notdesign.
•
b.
• Endurance limit at the assumed stress asymptote isevaluated statis~ically assuming a log-normal distribution (the number of cycles at which the fatigue strengthis assumed asymptotic varies with th0- material). Thelowest endurance limit of the following is used for lifecalculations:
a. bottom of d3ta scatter
b. 80% of the mean value
c. three standard deviations below the mean value(mean -3G).
Fatigue Loads
• Calculated or measured loads at flight conditions consistent with the expected use of the telicopter (designmission profile) are used.
• Top-of-scatter fatigue loads are assumed to exist ~on
tinuously for steady flight conditions such as levelflight, climb, and descent.
• Cycle counted fatigue loads are used for transient flightconditions such as turns, pull-ups, and landing flares.
Calculation Method
• Minor's cumulative damage theory is usea.
15
-~ --- ~~-
16
CALCULATECOMPONENT LIFE
LIFE CA~CuLATION
METHODOLOGY
Figure 4. Elements of Fatigue Life Evaluation.
~-
Mission Profile
• Initial fatigue evaluation is made utilizing the designmission profile.
• The design mission profile is reevaluated and modifiedas necessary based on field experience. The field experience could be fatigue failures, observer reports,or operational load surveys (as in this case).
It should be noted that some of the previously discussed fatigue evaluation criteria were not in use at the time of theReference 6 component life evaluation. In particular and ofmost importance is that bottom of scatter endurance limitswere used in the Reference 6 document.
COMPONENT LIFE
Calculated part lives for six critical aft rotor componentsfor each of four mission profiles, along with the endurancelimits used, are presented ~n Table VI. The wide range ofpart lives requires some explanation.
The basic mission profiles defined in Ta~l~ v are not completely descriptive of all the factors which influence partlife. The following in~ormation supports the missiondescriptions:
Gross Weight
The design mission profile, Reference 6, was evaluatedusing the worst loading from any flight data up to 33,COOpounds.
The current, Reference 4, AR-56, Reference 5, and derived mission profiles used flight test data at weightsconsistept with Table V.
Center of Gravity
The design mission profil~ assumed a cg distribution of25% maximum forward, 50% miu/ and 25% maximum aft cg forall flight conditions.
All other mission profiles were evaluated with level flighttime distributed equally between extreme forward and aftcg, and maneuvers were assumed tp occur at. the cg whichproduced the highest load on the part.
"~
I~I
\"
~,J
,I'~I
I"
r[d
ltl
',,)
I,'
",,
1'1
,,',
',I,
I~'~I"
Ifl"
l,I"
..,~
~jIH
,I~,I
I,\i!
,1'I.,
hj
I~,I
II"
1,'I"
)II
I'~!
"II
I~~
:,1l
~i~~
~".
tj,I
l~,'
:Ii~
,'II
I,r.,
I,m,~tN$mf.\Ufii1,W:W:tnf:lti:lllJlF::1T.llf;I"~?mWll":mmrJ;~~~~~/l.P"O'l'~~1W~"T!l~~ijt,'t11~\~~,n;r,"2l'i;'n'~fir'"'i,i
, ~ "
I, I,,,
..... CO
..TA
BLE
VI.
CH
-47A
PAR
TL
IVE
SFO
RV
AR
IOU
SM
ISSI
ON
PRO
FIL
ES
Fa
tig
ue
L.i
fe(H
ours
)M
issi
on
Pro
fi.1
eE
nd
ura
nce
Lim
it
Com
pone
nt<D
curr
ent.
(2)
Pa
rtN
umbe
rD
esig
nC
urr
ent
AR
-56
Der
ived
Oes
ig>
lA
R-5
6D
eriv
ed--
Aft
Bla
de
Sp
ar5
0,1
00
4,9
20
8,9
30
52
,55
0±
25
,00
0±
23
,20
01
14
RI0
42
-2p
si.
psi
Aft
Bla
de
So
cket
5,9
40
3,8
20
5,8
60
11
,69
0±
48
,20
0±
47
,90
011
4R10
43-4
in-
1bin
-1b
A'f
tR
otor
Sh
aft
9,4
40
9,6
00
20
,20
03
3,3
40
±2
16
,00
0±
23
3,0
000>
11
40
30
02
-7,
-8in
-lb
in-
1bA
ftR
ota
tin
gSwa~np1ate
17
,80
01
,56
54
,81
04
,49
0±
1,0
35
±88
111
4R33
24-1
1b1b
Aft
nta
tio
na
ryS
was
hp
1ate
42
3,0
00
23
,40
04
0,4
00
67
,20
0±
1,6
00
±1
,25
011
4R33
50-8
1bIb
Aft
Piv
ot
Act
ua
tor
1,4
00
1,3
40
1,4
70
1,3
50
±1
,61
0'!"
1,1
93
114H
4000
-20
1b1b
_.(J
)E
nd
ura
nce
lim
itb
ase
don
bo
tto
mo
fsc
att
er
(lJE
nd
ura
nce
lim
its
~a~ed
onle
sser
of
bo
tto
m0
_.
sca
tter,
0.8
mea
n,
or
mea
n-
30
Q)
Red
esig
ned
sha
ft(s
ho
tp
een
ed)
-,
-"
..
; !I , :~ I: 'f ,~ ~! :~ :1 ,~l I~f
I~1'1
,Ill:
'j'1, ,\~i l; jl\ ,'il )~ II ~10' ,~, 'I j" l, ,II: I:~' "I '~l II" I "~ 'I , i'j' :"'. i;1 I
11 '1 I'i I, Ii I' J,i
I1,'
1,1,
~ M*¥.&L¥£3::£~~::ggi-=~~~~~3L"!-"&S::iil¥€",~2Jh2'&"~~.TM~~=-~~W::~~~~~::'~-·~'::'-::~-::':_'-~--~--~~-=-:~-.~ ~--i-:;~':-_:----=:_-::~'~'::."-"'-;--~~------"-=--~~....;=...r---~,~--~~
-.;i't
Altitude
No altitude distributions were used in any of the lifecalculations. The effect of an altitude split on the current part lives was reported in Reference 7.
~ Speed
The power-on rotor speed was 230 rpm and the power-offrotor speed was 261 rpm for each mission profile.
Part lives from the design mission profile, Reference 6,are influenced by too many factors other than the missionprofile to draw any conclusions as to the effect of theprofile.
Because of fatigue design criteria changes whi~h occurredsin~e the initial evaluation of fatigue lives in Reference6, the component lives of Table VI are not directly comparable. Specifi.cally, the endurance limits used for design mission profile calculations only considered bottom ofscatter strength, whereas the other calculations additionally considered statistically significant values.
The current mission profile, Reference 4, produces extrern~ly conservative fatigue lives relative to those obtained from the derived operational mission profile on fiveof the six components. The aft pivoting actuator life isidentical for both of these profiles. Table VII shows thatnearly half of the damage to the actuator occurs at 120%Vne for the derived profile, while the maximum airspeed inthe current profile is 110% V •ne
The AR-56 mission profile fatwo times those of t~e curreche six parts, but a similarpivoting actuator.
lives are approximately-ion profile for five ofs indicated for the aft
The strength of the aft rotor shaft appears to change in thewrong ~irection in Table VI. This is because the higher endurar.ce limit is based on specimens which we~e shot peenedto improve fatigue strength.
Fatigue live3 for the aft rotor shaft assume that 5% of theoperating time is with retracted longitudinal cyclic trimfor each mission profile. Other components evaluated arenot adversely affected by retracted cyclic trim.
To gain further visibility into the effect of mission profile on fatigue life, two components were selected for evaluation of fatigue load spectra and fatigue damage rate. Theselected components were the aft rotor blade spar and theaft pivoting actuator.
19
inIJ1l~~~Jl:11:~iI';,~li
,~"
~li:
r1il
~jT'
F;I'
II,"
,'l
li~~
~'~"
'I';
h~R~t~l1"*1l:i
,':i~~j~l~
'V;!
I~"~
;!ti
if~~
ilt~
i~"
t"!Ii
'1i1.'
t";I}
:'ff'l:
i>',~~
rr;,r;
"r;~mR
i~'Wrt
Jl':nl
1t'1l'
J,.jI~
li':lr
r,rt;;
~·fm~i
~"!..,~
;J!l:~'i~
I~r~lb;r;i~ili'~'~,f
~!""~",~'lJii
ri:JI-
~l"':;
1[j'«);~"~~~~h}l~1I,1'?1;~'~~'~'IIt-I~~~l~~~I"'I'\~
v,iil'!l1;~r~11Qli~5')~'Vi'I~;~
"'1~~ "),; ',
~[l'
~ II tj~ ~S il~ 11 i: l' I 'I :~ i: 'Il !','~ Il'i' ~~ "~ f j 11: ~
~-i; i; '1t li '"
"._I
~...._....
_--
'Il1l
1,1I
11E
SP
IIH
"NO
IM
Il;r
,IO
NI'H
OP
11.1
: ......_.......
_...,....
._.._
..1
I'ro
fi
I.,
-_..._-
_..,,-
,\((
..';(;
I>('
riv
('d
_. .__
....'"-"
._"-
--' -
----
---
(lccll
r-IM
I"'l'lc
Occ
llr-
llll
mll
gcr,
'ncc
'rc
mcp
.,
~
",
..-.
A,'1
30II
I'S2,S~.0
III'
,..
...-
0.0
41
.50
.02
5'i
,J
0.0
01
0.1
0.0
20
.7...
.......
~..
.,.
'/f:!i
'/io"
'6'1j~'}-
1.7
(,4
1.6
0.'1
21
7.2
O.
'61.
40
.1(,
l,c;
0.3
00
1(,
.70
.1(,
r,.'
iO
.OI)A
lo6
0.01
l2.
"'0.
<)(,
9.
I0.
01l
2.2
0.4
01I.
20
.04
06
.60
.1~.
3.4
-¥.....
,......
-~~-~.
--
1,4'
1011
1'1
,3~,O
Ill'
.,~.,-
,~
..~_¥
..
-._~
....._.
._..-
......
.....-
"~
..~'".~
._..~
........_..
.,_.-
-'__.~~._1.~
0..
•0.
0'1
A.l
0.2
00
311.
00
.04
2.0
I0
.02
51
.21'
••11
11,I.
?0
.02
LO
0.2
2'1
.20
.40
01
5.6
0.2
20
.10
.40
00
.20.
0:>
0.11
0.1
30
O.
J0
,10
0.6
0.1
:10
0.3
,hll.
3..
•!l
,.L
. '-'o:"s
if6-'
''To'
:-r-
1.7
(;3A
.20
,72
15
.(,
0.3
63
.10
.10
01
9.7
0.3
67
.A0
.16
1.'
i0
.30
02
.50
.16
0.3
0.0
06
0.0
(LO
ll0
.20
.96
8.3
O.o
n0
.7n
.d
On
.,
0.9
1.0
7.0
O.
I
0.7
0.3
8.0
10
.6
4,'1
20III
'
4.i
li)"
af.
i
0.2
0
0.1
00
.0'i
O••
,~,
0.7'
1
0.2!
1
0.0
5
·1:'ll
O'
.•?
A:''''
0.2
0
21
.4
7A.(
,
10
0.
0.5
0.5
10
.0
33
,00
0
1.1
,00
0
1,1
fll
..I
'i0
,10
0II
I'"'"
--"1"
--2
8.r,s
c;
-t
hovC
II'
lonc
!dO
llcc
mt
'-'-"---miJ:"ij'ji"~-I:;NI'I<
:111
~IW
'II<
:1~
HII
'I'I:
mil
/If"
"IH
yr(
M'T
1'1
V(Y
I'W
GII
C'l'
lIM
'OH
IlV~'I,
1<:1
1'1'
CON
I>I'
I'W
N/11
11
II!
t1
l1.1
do
.S
Pil
t:
MiI
Illl
"n11
<'8IlI
n--·T
-··
'(:~;;~:;;;"_-
0IH
'r.,
UnC
JO
ccll
r---
iO-;
:-;\
lr-
Con
elIt
IClIl
ren
cc'
1MIn
IIIII
,'r~
'llc
clO
nlM
IlO_
__
_'I.
\,
1.
!'1I1
1-lI
pl.
oIq
I'll
I1
-up
I.7
',./
•••
II(H
.1'1
I11-
u1'
"1',,\
1\14
II.l
on
I.Il
V"
Irl
iqh
t-
0.2V
IIl.
ov,,
1fi
ll)h
t-
O.o
IVII
I:Id
c.w
'lI'd
f1iq
llt
HO
ilrw
lll'c
!fI
i'll
lt',c
c,,·
1to
cli
mb
IIlr
llIH
lt·c
!l'
ull
-lIp
1.4
9',
/Hp
ull
-up
"III
'n..
ho
vo
r('
on
tl'o
lIlC
lv
Co
ntl
'ol
Il<'v
"o
ntr
c)l
!lc,'V
-I I
.•.
,
11ft
,I'.
i:~~_~
~~.~.
~~":~t
or.._
...._
.•.•r..~~~.~
~·...~~......I.~'
~.O~..~~'
._•..
·_1..1.
~3"~..
I,I.~_
...~(1Vol
qig
ht
-h
I_~
'!00
0":,•
•.•_
,,_
_._
."._
._••
.__
"".
I.n
vo
lr
il/h
t-
1.1
VII
0.4
01
.1J.
.:vc·
1fl
igh
t-
1.1'
Wil
'1.
\lv,
,1f
!i<
lht
..1
.2V
lII'
uil-l
lp1
.4/,
Io
.or,
6.~
l'u
ll-l
Ip1.
7">1
/1I/~
pu
ll-I
I))
I28
,%0
'tll
rn-
loft
Ill'
01T
llrn
-rl'lllt
Inv
l'l
IT
urn
-Il
'ft
dOll
ccnt
IC
on
tro
lr,1
V..
Inlo
rr'
Co
n"I
'oI
rew
-d
Iro
cl..
on.I
IC
on
t.ro
lrQ
v-
d<ll
g:qn
t'r
rlllllli
tIo
nI,
evu
lfl
lC}
ht
..O
.2V
lII..
wo
lfl
igh
t-
0.4V
III,'
JVU
1f
1i<
Jht
-VI
ILo~nl
flig
ht
-1.
15V
III,
llvo
lfl
lC/M
,-
1.2V
lllJ
iclo
w.l
rdf1
illh
t/!'
.'III
:'war
clr1
iI/l
it1
\',c
\ll
t.c
0;:1
\mil
l\Ir
np
ocd
I'I,
II-u
p1.
4<1
III/
!P\
I.~
I-li
p'r
'Jrn
-h
ov
<!r
Co
ntr
ol
rHV
-h
ov
or
Co
ntr
ol
I'OV
-II
l.'u
ct:i
on
l1lI
Ij
0.2
00
.2I
...
._~~~!
~~.;~.
....,:
~:,-::
~._c:~
~~:~:.
!~:".:
=.,~...::
:.::::
.-=·",
~=l",-
":=:':
='~'''
',,.,.
,,=~.
.~.,,.
'''''-
:-'--:
::::-:
-~-:::
:~;.~~
:":=::
.,,,-~
:';=L
.--.--------..
Ii
._-=
---,.:.
.:._._
_.
IV o
~/~,u.
ii~
~,,;
I
FATIGUE LOAD SPECTRA
Fatigue loadings for the CH-47A aft rotor bl~de spar and aftpivoting control actuator are shown in Figures 5 and 6 respectively. The loads are displayed in percent of endurancelimit, and occurrence is expressed cumulatively in percent oftime to reach or exceed. Only the high load, low occurrenceportion of the curves is shown in order to improve dataseparation.
Looking first at the blade spar, there appears to be a closerelationship in loads and lives for the current and AR-56 mission profiles. Similarly, the design and derived mission profiles have comparable shapes and nearly identical lives. Thefact that resulting part Jives differ by an order of magnitudeis significant, and will lJe explored further in the FatigueDamage Rate section.
The occurrence of higher loads on the blade spar with thedesign mission profile requires explanation. An extensivereview of fatigue methodology was conducted at Boeing Verto]subsequent to the life calculations of Reference 6. Theculmination of this review was the current life calculationsfor the CH-47A presented in Reference 4. The review included a reevaluation of measured flight loads, and errorsfound in the original data reduction were corrected. Thecorrections included both increases and decreases in loadvalues, and in the case of the aft blade spar, the maximumfatigue load was lower than originally reported.
The actuator fatigue load spectrum, Figure 6,exhibits a widevariation of loads above 110 percent of the endurance limitfor the various mission profiles, but the resulting part livesare surprisingly similar. The Fatigue Damage Rate sectionprovides further visibility into the specific conditions whichcause fatigue damage. The load variation is attributed to thefollowing:
• Design mission profile load spectrum is normalized to ahigher endurance limit than the other mission profiles, asshown in Table VI.
• AR-56 mission profile loads at 115% VH are greater thanthose of the current mission profile.
• Derived mission profile maximum loads occur at 120% VH andin the 1.75 g pull-up maneuver conditions which are p~culiar
to this profile.
The actuator fatigue life is approximately 1,400 hours as calculated, using four different mission profiles. The fatigueload spectra resulting from the four mission profiles are
21
HRl1Q:
1,4001,3401,4701,350
LEGEND
LIFE,PROFILE SPAR
Q DESIGN 50,1001 -f------:::II~----4 I:!J. CURRENT 4,920
OAR-56 8,930o DERIVED 52,550
10 ....---lJ6---+----_....._
PERCENTCUMULATIVEEXCEEDANCE
10-1+-----.f---¥ilr__1------
10-2 ....----+-----180 100 120
PERCENT ENDURANCE LIMIT
Figure 5. CH-47A Aft Rotor Blade Spar Fatigue Load Exceedancefor Four Missions Profiles.
10~---4-l-3~-I- +- -+- -+_
1+-----i---_+_~
PERCENTCUMULATIVE
EXCEEDANCE 1----~_---_Ir_-\_--..J~~oQ~_4-10-1 +
10-2-I-----+----~.....---.......IIIft'-_+-80 100 120 160
PERCENT ENDURANCE
Figure 6. CH-47A Aft Pivoting Actuator Fatigue LoadExceedance for Four Mission Profiles.
22
Table VII summarizes the fatigue damage for the blade spar andcontrol actuator for each mission profile by flight condition.Fatigue damage is expressed in percent of total fatigue damagefor each mission profile element, and the corresponding occurrence in percent from Table V is shown for reference (grossweight for AR-S6 i~ distributed the same as current missionprofile) • ThE: calculated part life is also shOW:l for reference.
r.-, A ::;;;:e::;o;;:t-¥9-*±¥A-;<-jdA"!!"!f""<·-- -rtf!.'t!";1'! en::: ::; :;--. .
~ nearly identical at 110% endurance limit and exhibit rather~ large dispersion at greater load values. These observationsf-.
t emphasize the importance of designing for fatigue strength@ greater than the loads expected in frequently encounteredi flight conditions.
~ FATIGUE DAMAGE RATE~~
iiE
As previously discussed, fatigue damage for the design missionprofile was calculated using differen~ methodology than theother mission profiles. In addition, the limiting airspeedenvelope for the design profile calculations was approximately10% greater than the current airspeed restriction for the CH47A. It is concluded that th€ design profile is a poor basisfor comparison in this study.
It is not surprising that components with low fatigue livesexperience fatigue damage in steady state flight conditions.Elimination of fatigue damage in transition and level flightfrom Table VII would increase all calculated lives tc atleast 4200 hourE.
A 1.75g pull up was included in the derived mission profile atan occurrence consistent with the load factor frequency experienced in th~ operational load surveys. Figure 3 shows theload factor spectrum of the derived mission profile relativeto the operational data. Because of the very low occurrence,the effect of the 1.75g pull up on fatigue life is insignificant as shown in Table VII.
MISSION PROFILE RELATIONSHIP TO FATIGUE LIFE
& _ AT
From the analyses in this section,several cor-elusions can bemade with respect to missior- profiles and component fatiguelives:
1. Design fatigue strength for safe-life should insure thatno fatigue damage is incurred in steady flight conditionsor in transient conditions which are frequently encountered.Based on CH-47A loads and S-N curves for common metals,transient conditions of 10 seconds or less which occurless than once in 10 flight hours need not be consideredin fatigue design. Calculated design fatigue loads shouldLe amplified by a credibility factor for the load analysis
23
A review of the developmental histcry of the CH-47 helicopterwas conducted with the objective of determining the extent towh1~h missiop profile changes affected configuration changesof dyl1amic components.
4. AIRCRAFT CONFIGURATION
method. Procedures for establishing a credibility factorare required.
Fatigue damage tracking of critical safe-life componentsmay be an effective method of increasinq the utilizationof components.
CH-47A component retirement lives are conservative basedon the operational data available.
3.
2.
r,.zs:w;:.,,4A4!4¥i!AlitJ1W-•••V, JII "-"'-,%dif1!!<'#-"""",,,-,,,," _c' L. ,_A..., c" " _-,-
:{
f~~~
~!j
~
~~
i~
IThe study was initiated by reviewing all Engineering Job Sheets(EJS) and Engineering Change Proposals (ECP). The EJS was thepredecessor of the ECP, and these combined documents defineall of the major changes to the CH-47A,B, and C helicopters asformally proposed to the Army by the Boeing Vertol Company.
Not all EJS and ECP are accepted by the Army. This study waslimited to the company proposals which the Army approved. Allchanges involving rotor blades, hubs, drive, and control system components were included in the study.
Since the interest in this study is the effect of aircraftusage on dynamic component changes, the Service Revealed Difficulty (SRD) documentation was also reviewed. All changes ofinterest resulting ~rom investigation of 85 SRD's were imple-mented by either EJS or ECP action. Therefore, the SRD reviewproduced no new information relative to this study.
Minor changes to the aircraft can be accomplished throughdrawing changes without EJS Gr ECP action. In order to avoidan oversig~t of significant changes, a review of the historyof assembly drawing changes was conducted. As with th~ SRDreview, no new useful information relative to the study wasobtained from the drawing review.
EJS/ECP &~ALYSIS
A total of 207 approved EJS and ECP changes in the dynamicsystem were identified. The changes were classified intoeight categories chosen to provide a general description ofthe reason for the change. Table VIII quantitatively summarizes the analysis of the changes.
24
-~
TABLE VIII. SUMMARY OF 2~7 CH-47 CHANGES BY HELICOPTEkSYSTEM AND REASON FOR CHANGE
~Number of Changes
Reason for Change Blades Hubs Drive Controls
1- t-1aterial, marp.1facturing and 4 2 4 3maintenance improvements
2. Environmental protection 1 2 - -(sand, dirt, corrosion)
3. Lubrication and cooling - 1 19 3improvements
4. Joint retention - - 12 2improvements
5. Reduce friction, improve - - - 33sensitiv~ty, improvedamping, eliminateinterference
6. Wear improvements - - I 4 31 5bearings and gearteeth
7. Fatigue strength 8 10 25 24improvements
8. Ultimate streagth 3 1 3 4improvements
Totals 19 20 94 74 i
25
Classification of the changes was difficult in many cases,since subjective judgements were involved. It was desirablefor analysis p~rposes to assign only one reason for eachchange. Obviously, some changes could be identified with morethan one of the classifications shown in Table VIII. A controlinterference, for example, could possibly affect strength andbe classifi.ed as a fatigue strength improvement, ultimatestrength improvement, or both fatigue and ultimate strengthimprovements. In this analysis the classification selectedwas the one which appeared to be the primary reason for initiating the change proposal.
The changes made for fatigue and ultimate strength improvementare further summarized in Table IX. Two changes which separately define fatigue strength improvement by shot peening ofthe forward and aft rotor shafts are clearly related to serviceusage. The changes were made following analysis of the operational data reported in References 2 and 3. Two operationalfactors contributed to the reduction in fa~iaue life and resulted in the need for citange: flight at airspeeds in excessof Operator's Manual limitations, and flight with longitudinalcyclic trim retracted (failed trim actuators). The airspeedexceedance was identified from the operational data. Thefailed trim actuator condition became apparent based on pilotreports, field service engineer reports, and a study of actuator malfunction frequency.
Twelve of the identified fatigu~ strength changes resultedfrom planned growth of the CH-47 helicopters. The structuralflight limitations for the CH-47A, CH-47B, and CH-47C at 6,000feet density altitude are compared in Figure 7. The need forgreater payload than the CH-47A could produce became a?parentduring the combat operations in Southeast Asia. Structuralchanges associated with model changes were based on operationalneed. This should not be confused with changes resulting fromoperational use.
The balance of strength improvemel1ts in Table IX resulted fromeither developmental tests or service failures. Developmentaltests include static and dynamic benc~ tests as well as flighttests, and bear no relationship to operational use of the helicopter. Ch.:;.nges \Plhich result from service failures could beattributed to operational use, but not necessarily so. Inmanycases the distinction between design deficiency, manufacturingdefect, operatiol.al use, and operational mis-use is hard todefine, and qaite often is influenced by the perspective ofthe one making judgement. There were ~nsufficient facts available in most of the serVice-related changes to determine theinfluence of operational use on the service failure.
26
, ii, j 'J \ I I I II I ~ I, ,"
tv -...J
TA
BL
l:;
IX.
~:Y
OF
CH
-47
CHA
NG
ESTO
IMPR
OV
EFA
'l'IG
UE
AND
UL'
!'IM
ATE
STRE
NG
THO
FCO
MPO
NEN
TS-
Sy
stem
Rea
son
Num
ber
of
Cha
nge
Resu
lts
Cha
nge
Desc
rip
tio
nfo
rC
hang
eC
han
ges
Fro
m
Bla
ne
sock
et
incid
en
ce
pin
ho
le4
Serv
ice
fail
ure
co
rrecte
dfo
rm
anu
fact
uri
ng
de-
(bla
de
loss
)fe
et
(Lu
rr),
imp
rov
edco
rro
sio
np
rote
cti
on
,re
-ori
en
ted
tore
-F
ati
gu
ed
uce
load
sst
ren
gth
Dev
elo
pm
enta
lB
lad
etr
ail
inq
edg
ere
desi
gn
ed
Ro
tor
3b
lad
ete
sts
toim
pro
ve
fati
gu
est
ren
gth
1A
ircra
ftg
row
thN
ewb
lad
ed
esi
gn
for
CH
-47B
and
CH
-47C
-U
li:.
imat
eD
evel
op
men
tal
Ro
tor
bla
de
tip
co
ver
and
bal-
stre
ng
th3
tests
ance
wei
gh
tre
ten
tio
nim
pro
ve-
men
ts
Dev
elo
pm
enta
lD
roop
sto
pch
ang
esto
min
imiz
eF
ati
gu
e8
tests
dro
op
sto
pco
nta
ct
and
incre
ase
stre
ng
thlo
cal
fati
gu
est
ren
gth
-V
ert
ical
pin
join
tan
dla
gda
m-
Ro
tor
2A
ircra
ftg
row
thp
ers
red
esi
gn
ed
for
CH
-47B
and
hub
CH
-47C
Serv
ice
fail
ure
Dro
opst
op
red
esi
gn
tore
du
ceU
ltim
ate
(bla
de
stru
ck
aft
bla
de
sta
tic
dro
op
an
gle
stre
ng
th1
fuse
lag
ed
urL
1g
hig
hw
ind
shu
t-dO
'..,n)
I
~~: ,n: J!\j
I "; Ii! III I',
'; ,)Jj'~
)'i~i·I"JI':~;'"')~:
1·'~~I;~·it"i~m~)~~~~"r
lii'
r~\l
~~~~
~~1f
if1I
~'T"
~~~I
~1:i
1I~'
~liI
i,1l
rtir
·~(l
,~~·
~p~f
t!~:
~~~~
n<';
1f1l
iWrm
n~ij
·~,~
~Jr!
li;i
ir'!
1'-r
nro'
~~"1
l!!~
!f'l
i~n~
~Ji1
1~4N
~~"~
'~~~
~II.
,J'r
')l1
;flt
"f,l
~~t,
't]I
~~:I
I'jo
~ifi
";~ih7l~1'ji;Jl'mIY~~~Y~~~'.I'~~tI¥~r,ni~I~~''I',>f-;\r~;'''?'''~ij;n;r~~'I'~·;;YI~·(1IJ·!ll,11W.'''~f;,TfrH'~~''
II~;; ~ I~·,
I J ~' ~, !~ I~ .~ ,I :1 ~
,
~~ I: ~,
tv ex>
TAB
LEIX
.Continu(;~d
Sy
stem
Rea
son
Num
ber
of
Cha
nge
Resu
lts
for
Cha
nge
Ch
ang
esF
rom
Cha
nge
Desc
rip
tio
n..
-3
Dev
elo
pm
enta
lT
ran
smis
sio
np
inio
ng
ears
sho
tte
sts
pee
ned
toim
pro
ve
fati
gu
est
ren
gth
4D
evel
op
men
t'll
Tr3
nsm
issi
on
pla
net
beari
ng
re-
tests
tain
errede~igned
toim
pro
ve
fati
gu
est
ren
gth
Dri
ve
Fatigu~
4D
evel
op
men
tal
Tra
nsm
issi
on
pin
ion
gears
red
e-
stre
ng
thte
sts
sig
ned
toeli
min
ate
r~sonance
2S
erv
ice
fail
ure
Th
rust
beari
ng
reta
iner
red
c-
(cra
cks)
sign~d
toim
pro
ve
fati
gu
est
ren
gth
"
1S
erv
ice
fail
ure
En
gin
eq
uil
lsh
aft
red
esi
gn
ed
to
--~
(pow
erlo
ss)
imp
rov
efa
tig
ue
stre
ng
th
E~ginetransmis~ion
sp
iral
bev
el
1S
erv
ice
fail
ure
(pow
erlo
ss)
r1n
gg
ear
red
esi
gn
ed
toim
pro
ve
fati
gu
est
ren
gth
-.
2S
erv
ice
fail
ure
Syn~hronizin~
shaft
ad
ap
ter
re-
(cra
cks)
des
1g
ned
to1m
prov
efa
tig
ue
,...
-st
ren
gth
2S
erv
ice
usa
ge
Ro
tor
shaft
ssh
ot
pee
ned
toim
-p
rov
efa
tig
ue
stre
ng
th
6A
ircra
ftg
row
thT
ran
smis
sio
ns
and
shaft
sre
de-
sig
ned
for
CH
-47C
J:~ li~ I!~ '/;ll:
,~ ~~ ~,~ )' ~i ~ t)~i ;~ ,I 1~1' i' i ~ {~1 11 I~~' 1~,.. 'l ~~ :1 ' ~ ., I, ~'i , I ,~ ~! ,,·1 ~ Jr'l
"
"I
r'1i
1PII
,,~~
~,,:
r~~:
~lfi
;r~'
fi1i
f"![
l1I'
!'m~
'1'P
:ii~
r;':
I;tr
r;H~
;::'
ll't
'M;:
~If~
';r;'.
....F:'r~~~'if.r.~:w":iI!'1i~~~1i}Ill:~~i11!"rf~'iJ11n:p,~'l:~1"i'f'!'r~j"~i'lil~'il1':~~!JjPt~~fi',~)7i
~~n"%\"I,!,'~\l(.I,~:,''',~g'l
~"'~""'ilf'"',\wlN""',r':r'I,I4',F"i'11\"""'111',litl:*",'Itltf,it,l'j"'!ii/,l4W"'''I~,'p,it,I'
\""1,1
1,"'",'1
,i'"Or
l',',1l'~""\W'lrli1'H:,im,p:"I~C:'il":~i
i; ~! j"
j ",
tv '"
.TA
BLE
IX.
Co
nti
nu
ed
Sy
stem
Rea
son
Num
ber
of
Ch
ange
Resu
lts
for
Ch
ange
Ch
ange
sF
rom
Ch
ange
Desc
rip
tio
n
1D
evel
op
men
tal
En
gin
ed
riv
esh
aft
bala
nce
wei
gh
tD
riv
ete
sts
rete
nti
on
stre
ng
then
ed-
1D
evel
op
men
tal
Aft
tra
nsm
issi
on
ho
usi
ng
stren
gth
-te
sts
ened
torea
ct
un
pch
edu
led
ro
tor
lock
eng
ag
emen
t
Ult
ima
te1
Se.
r.vi
cefa
ilu
re
Au
xil
iary
gea
rbo
xq
uil
lsh
aft
stren
gth
(gro
un
dpO
Wer
loss
)st
ren
gth
ened
torea
ct
lo~ds
du~
toclu
tch
ma
lfu
nct
ion
3D
evel
op
men
tal
Lo
ng
itu
din
al
cy
cli
ctL
imsc
hed
ule
tests
cha
ng
edto
red
uce
hu
ban
dsh
aft
loa
ds
J.D
evel
opm
el1t
a1SA
Sa
uth
orit
ych
an
ged
tore
du
ce~ests
co
ntr
ol
syst
emfa
tig
ue
loa
ds
Fa
tig
ue
.1.1
Dev
elo
pm
enta
lI
Fix
edco
ntr
ol
syst
emco
mp
on
ents
stren
gth
tests
red
esig
ned
toim
pro
vefa
tig
ue
stren
gth
Co
ntr
ol
2D
evel
op
men
tal
Ro
tati
ng
co
ntr
ol
syst
emco
mp
on
ents
tests
red
esi
gn
ed
toim
pro
ve
fati
gu
est
ren
gth
'--
3A
ircra
ftg
row
thC
on
tro
lsre
des
ign
edfo
rC
H-4
7Ban
dC
H-4
7C
-..
I
,~:. ~I II ~ !\ :~ ;~ :t~t ~~: " i :;-~ l!~ Jf~ " l ,I II j, * I;
j; t:~ ~: 'i' ~ ~ ~~ ~! 1~1~: ;:~: " I, (~j :~~, $ 1'~ t" :,'t· i( f '~I, ~! j il I ,i i J" I'
,I'0
0'
r.
1~:r:'W'ffll"'i'1lli~"Wll.1i~n::v~;;rrl,~~~~'tPjjil'~Tnm~IWIWf~-ln;)I'I'[I1'l,'
m~~I~,1'1j1li;r.~r~r,~~lff,I+~1;1l?i~i~I~~%ifilj!'i~I~li~i;~~~
li'r
'l'~
i1~
:I~i
'~il
:';f
'T~'
1IJr
M~fT
q'jl
~Pi~
,~i~
W'~~
~'ff
~tr~
i4i~
~lil
'NNI~!'~I~~~'lfil'illi,~T'fij'~I~i\I.~~illllii~~li'~;If~~1Iif;tj'
~'~"
1f!t
1li;r,
!';~'
i;j~
;';~
'f.:
'IHm
n'I,
rf*~
;nr.
l'~r
:·li
r.~.
]'~,
ilio
'~1l
.'jl
~~~.
W~J
~1·~r'''~''~L~I;rMir~~~,1~~I':lj'~rji;l:~~:ii.'~t~')'I'4'1r'''';l'
~f"rl
il~I;~
i~._ '"~1 ,[ !i ~ :I'~
Vol
o
TABL
EIX
.C
onti
nu
p.d
Sy
stem
Hea
son
Num
ber
of
Ch
ange
Res
ult
sfo
rC
han
geC
han
ges
Fro
mC
han
geD
esc
rip
tio
n._
-
Fa
ti9
ue
Serv
ice
fail
ure
Dri
ve
co
lla
r(3
)an
dp
itch
lin
kC
on
tro
lS
tren
gth
4(c
rack
s)p
ro
tecti
ve
cov
er(1
)re
des
ign
edto
imp
rove
fati
gu
est
ren
gth
3D
evel
op
men
tal
Low
erco
ntr
ols
red
esig
ned
torea
ct
Ult
ima
tete
sts
par
ked
bla
de
loa
ds
stre
n<
;th
Co
llecti
ve
pit
ch
co
ntr
ol
ma
gn
etic
1S
erv
ice
fail
ure
bra
ke
(trj
md
evic
e)re
des
ign
edto
rea
ct
pil
ot
loa
ds
,-
:k ~~ I, .1: ;1 ~~" ~I ;~ i~ t;. ~ 1,1 f'1~1
;' , ," ~~l III :1; "~ I II 1,1 i' !! J:~0'i~!
"I~
tI
50,000
GROSS WEIGHTPOUNDS
25,000
CH-47C
CH-47B
CH-47A
o+----+-----+-Aoo..__...........o..-
o 50 100 150
Figure 7.
TRUE AIRSPEED - KNOTS
Structural Airspeed Limitations for CH-47Helicopters at 6,000 Feet Density Altitude.
31
The effect of aircraft usage on component fatigue lives can bemade by comparing the calculated lives documented in References 4 and 6. A sample comparison is shown in Table VI,and it is evident that significant reductions in part lifeoccurr.ed. As discussed in Section 3, the current mission profile of Reference 4 wa~ not the only factor affecting reduction i~ part life. The most importaut difference between thetwo mission profiles is the high speed level flight condition.The design mission profile assumed that the airspeed limitations of the Operator's Manual would not be exceeded, whilethe CH-47A operational data shows that this airspeed was freque~tly exceeded. Flight with failed longi~udinal cyclic trimactuators affects the fatigue loads on some components, inparticular, the rotor sqafts.
DISCUSSION OF RESULTS
Op~rational use of the CH-47A as compared to the design missionprofile led directly to the rework of the forward and aft rotorshafts an~ contributed to reduced reti~ement lives for severaldynamic system components. Exceeo~nce of the airspeed limitations of the Operator's Manual is cited as the principal causeof the reduced fatigue lives. The failure mode of a controltrim device also contributed to the rotor shaft rework requirement.
The experience gained in evaluating the operational use of theCH-47A helicopter was used in the development of the CH-47Band CH-47C models. The structural performance of the growthmodels shows a great improvement over the CH-47A, at least apart of which should be attributed to the CH-47A operationalexperience. No operational survey ha~ been conducted on theBand C models, however, so the adequacy of their mission profiles cannot be evaluated.
5. OPE~.TING LIMITATIONS
An attempt is made in this section to determine the factorswhich limited operational use of the CH-47A, and to projectthe effect of various limita~ions on the use of future cargoand transport-type helicopters.
Steady state mi~sion time from the Reference 1 data gatheredin ~e United States (USA) and from the Reference 2 datagathere(1 in Southeast Asia (SEA) was analyzed with respect togross weight, airspeed, and altitude. The operational measurements were compared to the Operator's Manual limitations, powerlimits, and vibrations.
Flight duration and the rotor start-stop cycle were brieflyexamined.
32
Finally, the load factors experienced in the USA and SEA wereexamined to evaluate the suitability of the design load factor.
Two factors which may be significant to this study cannot beev~luated:
1. . ..: effect that external cargo stability may have had onairspeed.
2. The influence of a monitor system on crew performance.This has always been a nagging problem in field surveywork, although the data analyzed herein did not appear tobe biased to hide airspeed exceedance of operating manuallimitations.
Power limited airspeeds were calculated for both normal andmilitary power. However, it was only necessary to includenormal power values in the analysis. The power values weredetermined at Interna~:onal Standard Atmosphere elSA} conditions, which approximate the USA data atmosphere, and at ISA+20oC, which more nearly represents the SEA operations.
The limitations of toe TM55-1520-209-10 Operator's Manual werecopied directly from the Manual. These limitations were established from the fatigue load to strength relationship asdetermined in contractor developmental bench and fJight tests.
The vibration environment in the cockpit was estimated interms of a pilot comfort index. Values of the index are shownon the plots where this index is used. Vibration levels aredisplayed as a scatter band varying with airspeed but independent of gross weight and altitude.
STEADY-STATE OPERATIONS
Th~ steady-state segments of USA and SEA data are presented inFigures 8 and 9 for gross weights below and above 28,OOOpoundsrespectively. In the figures, the flight times are expressedas percent of total steady-state flight time in the weightcategory. The upper portion of each figure breaks up the datainto altitude and airspeed blocks which are compared to normalpower and to the limits of the Operator's Manual. The lowerportion of Figures 8 and 9 have a pilot comfort index shownrelative to the operational data sample with all altitudesgrouped together.
The first observation to be Made on Figures 8 and 9 is thatthe CH-47A limits in the Operator's Manual are less than thoseattainable with normal power (twin engine) for nearly all operating conditions. The relationship of power limits to otherlimits in the operator's Manual may be important in projectingaircraft usage, and analysis of operational data should consider the relationship.
33
O.G =+ <. 0.05%
15,000
DENSIT'ALTITUDERANGE
FEET
NORMAL POWER- - - - GW 20,000 LB (ISA)--- GW 28,000 LB (ISA +20oc)
I .... jUSA
l~-- ....SEA 11,900 i"-
t-. I "I !oo". 0.1 0.0 "I II10,000 i
USA .............~ ~'" ffiSEA i'--. ..
0.5 1.8 10.6 7.0 6.8 4.5 0.7 0.0Vne I I5,000 GW ~ 28,550 LB I I
"- 1~.1USA 0.1 0.1 2.2 3.2 4.2 4.9 2.3 \- " "SEA 7.6 4.4 25.9 9.0 7.7 8.2 1.7 0.2 0.0T
2,000 :,USA 4.3 0.5 6.1 7.8 7.8 4.8 3.3 1.2 0.1 ,SEA 1.2 0.1 0.4 0.3 0.6 0.5 0.1 \ \ 1
!~r
1,000 I. 1USA 8.9 1.0 8.1 11.1 8.5 4.5 2.6 0.9 0.1
SEA 0.1 0.0 0.0 0.0 10.0 1\ ~;v
c 40 60 PO 90 100 110 120 130 140 150 160
I~DICATED AIRSPEED - KN01'S
10 --- SHORT-TERr-t LIMIT
3.4 0.3
0.2 0.0
--- LONG-TER."t LIMIT
--~~----ACCEPTABLE
PILOTCO~tFORT 5INDEX
~
,USA 13.3 1.6 16.4 22.1 20.
ALL 0SEA
ALTITUDES 9.4 6.3 37.0 16.3 15.
Figure 8. Percent of SteadY-State Time for 98.2 Hours (USA)and 135.9 Hours (SEA) of CH-47A Operatipnal Dataat Gross Weights Below 28,000 Pounds.
34
35
80 90 100 110 120 130 140 150 160
NORMAL POWER______ GW 28,000 LB (ISA)___ GW 33,000 LB (ISA +20oC)
INDICATED AIRSPEED - KNOTS
6040
USA9,200 ,,...--, -...... ... -..."....I 6,200 ' ~-.
SEA 0.3"1......
~4.1 ~2 .........0-:2 ·12.7, 1.0 0.6 ,
5,000 I i iI ••
" '. ..'\,USA 0.1 0.8 2.1 1.6 1.5 1.5 0.3 0.0
SEA. ",5.1 5.9 ~2.0 15. 9.4 8.5 1: 9 0.1 .,
GW 28,550 LB •V t,
2,000 ne GN 33,000 LB- ••USA 1.2 0.3 7.0 117. ] 126. C110.2 2.1 0.9 0.1
SEA, , 1\1.6 0.3 0.6 \ I.
I ;1,000 I•
USA 5.5 0.3 4.6 4.2 3.0 5.5 3.0 1.1 0.0 :;;SEA 0.1 0.1 0.1 0.1
,\ ~ I
Figure 9. Percent of Steady-State Time for 9.4 Hours (USA)and 18.9 Hours (Sea) of CH-47A Operational Dataat Gross Weights Above 28,000 Pounds.
10,000
FEET
o
DENSITYALTITUDI:RANGE
10 SHORT-TERM LIMl~
PILOTCOr-IFORTINDEX
5 ACCEPTABLE
LONG-TEJU.1 LIMIT
USA 6.8 1.4 13.7 5.4 2.0 0.1ALLALTITUDES SEA 7.1 6.5 45.4 2.5 0.1
0.0 =}< 0.05%
Airspeed occurrence fell off rapidly above 110 knots, nearlyindependent of altitude and gross weight. Above 110 knots thepilot comfort index s~ggests that vibrat10ns can reach pilotfatiguing levels.
=----- .~""""~~"=,.,-~~ ,""",,~~~~~-'""-~~--~-'"'1r_§.!@1F~Ca:;'~J",";LJ2<E09J2P!R;;g:t<'&§3S%li"'iL -.- ,- - ~:
~ ~
~ ~
~ ~
~ The limits of the Operator's Manual were frequently exceeded ~i in both the USA and SEA data sets, while all of the data could l~ be within normal power limits. The power observation does not i~ imply that the pilots never exceed normal power, but simply jf that high power settings were not usually required for steady "t level flight. ~
~~
~
•~I>
The frequent exceedance of the airspeed limitations of theOperator's Manual is a cause for concern, since componentfatigue lives are generally reduced by this excessive airspeed.Since it is not logical that pilots would willfully violatemanual limits, it must be assumed that the limits are eithernot understood or are too difficult to comply with in combatoperations.
FLIGHT DURATION AND THE ROTOR START-STOP CYCLE
A re~iew of the rotor start frequency and number of flightsshowed that the average flight was very short. The shortflight duratior. was consistent with the high occurrence oftime at low air3peeds. The flight and rotor start data areshown below:
No. of No. of No. of loVerage Avg startflights fIt hrs rotor starts flight cycle
USA 769 165 230 12.9 min 43.0 minSEA 1081 235 395 13.0 min 35.7 min
l-1AXIMUM LOAD FACTOR
36
The maximum load factors experienced in the operational dataare presented in Figure 10. The USA and SEA data samples werevery similar, and each gives the impression that the CH-47Acould have been designed tc a much lower load factor. ~he
problem with arriving at such a conclusion is the uncertaintyof the adequacy of the data sample to represent the long-termexposure of a full fleet of aircraft. In Figure 11 the dataof Figure 3 is replotted with an exponential extrapolation ofthe mean of the data. The contractor's demonstrated loadfactors are shown fo~ reference. The extrapolation, which isonly one of several possibilities, suggests that exceedance ofthe positive design load factor would occur around 10 milliontotal flight hours. The CH-47 fleet has logged over 1.2 million flight hours as of October 1972. Other extrapolations ofthe data could lead to the conclusion that the CH-47 shouldhave been designed to a higher load factor.
Figure 11. Operational Load Factors Extrapolated toHigh Flight Hours.
x2
DATA
35
DESIGN ENVELOPE
-1 0 1
6 LOAD FACTOR <nz)
= O.594e-9 • 3 -6
= O. 061e l l. 4x_l
I ""t------If---'l,.---- OPERATIONAL ~4------IDATA I
j
-2
GROSS WEIGHT - LB X 10- 3
37
l~-----+---
y -----..,.-------------~----......,\1...41---1 DE140NSTRATED n z --4---#-~
\ DESIGN n I~ f106 1------.;:.".,--+---- --+--"':'---A4---#--=-~
\ EXTRAPOLATED
Figure 10. CH-47A Maneuvering Load FactorsRelative to Design Envelope.
HOURS TOREACHOR EXCEED
102 1-------1---:
3
2NORMAL LOADFACTOR 1,nz)
0
-1
~ I I I20 25 30
SUMMARY OF OPERATING LIMITATIONS
From the data presented in this section, the following conclusions may be extracted for use in future cargo- and transport-type helicopter design:
rJ!W-~t"'*""""" .4-"&%""'"""'" -.... , --.-",""-...",",,,~.,_.,,,,,,,",.,.,,,.. -"",","""""~"~-'-""-- ~'""""!
~ The design load factor for the CH-47A was based on the thrust~ capability of the rotors. The demonstrated load factors on~ the CH-47 models were the extremes that the pilot could attain, for the required demonstrations. Therefore, if the required~ demonstrations were properly chosen, it is not possible for a[ fleet pilot to exceed the demonstrated values. This reasoning~ leads to the conclusion that the extrapolated load factors of~ Figure 11 should become asymptotic to the demonstrated load~ factors.
i1. Frequent flights of 10 to 15 minutes duration can be
expected.
2. Rotor start-stop cycles may average up to 2.0 per flighthour.
3. Structural lil!litations should exceed power capability forall primary missions. Improved methods of providing acockpit display of structural limitations are desirablein order to avoid overdesigning of rotor systems.
4. Vibration attenuation should improve pilot comfort andconfidence in the aircraft, and should not be a factor inlimiting operations in future designs.
5. Helicopters should be designed to sustain flight loadswithin the transient capability of the rotor system. Theinherent versatility of helicopters to perform ~ widerange of missions could be compromised if the aircraft aredesigned to less severe criteria.
38
6. CONCLUSIONS AND RECOMMENDATIONS
Analysis of the separate studies of Sections 2 through 5 leadsto the following conclusions and recommendations for structuraldesign criteria for cargo- and transport-type helicopters.Because of limitations identified in the CH-47A usaqe daTa,recommendations are also submitted for future aequ~sition andanalysis of operational data.
STRUCTURAL DESIGN CRITERIA
1. Design load factors should be based on the maximum transient thrust capability of the rotor system. As shown inSection 5, the operational data does not support an argument for reducing the design load factor to a lesservalue. No historical evidence was found to suggest thatgreater strength is required. With this design criteria,the maximum design load factor would be determined at theminimum operational gross weight.
2. The minimum load factor of 2.0, as specified in MIL-S8698, appears conservative when compared to the operational data. If the entire helicopter is designed toreact the transient thrust capability of the rotor system, as recommended ~n the preceding paragraph, no minimum positive design load factor is required. Structuralsafety will be maintained for any operations within theperformance capability of the helicopter.
3. Fatigue mission profiles should be simplified for helicopter design. Attaining part lives in excess of SpOGOhours is not likely if fatigue damage is incurred in anysteady flight condition, as shown in Section 3. Transientconditions of less than 10 seconds duration which occurless than one time in 10 flight hours usually need not beconsidered in the fatigue design of most metallic dynamiccomponents. The requirements for fatigue design shouldinclude the following:
• Listing of steady flight conditions required to performthe design mission(s). Include steady bank angle andsideslip requirements.
•"
•"
The frequency, duration, and severity of transient flightconditions required to perform the c2s~Jn missions.
The payloads, airspeed range, and design altitude forsteady and transient flight conditions.
39
4.
•
•
Fatigue mission profilp~ for life calculations should bebased on design flight conditions and reevaluated by conducting surveys of actual helicopter operations. The CH47A experience clearly indicates the need for reevaluatingmission profiles, as evidenced by the reduction in partlives due to exceedance of Operating Manual limitations.Concepts for control of fatigue damage other than ope~a
tional surveys include:
Cockpit display of fatigue loads. The Cruise Guide Indicator (CGI) is a suitable device, but it has no capabilityto assess fatigue damage. The nGed for such a display isapparent in Sections 4 and 5.
Direct monitor of fatigue damage to critical components.
• Fail-safe design which provides a secondary loa~ path andsuitable detection and warning of a primar/ load pathfailure.
5. Flight times and rotor start-stop cycles should influencethe f~tique mission profiles. The CH-47A operational dataindicate t~at frequent flights of 10-15 minutes durationcan be expected, and 1.5 to 2.0 rotor start-stop cyclesper flight hour should be considered (Section 5).
6. Vibration criteria shoulo insure that operations are notrestric~ed due to crew or passenger comfort. It was concluded ~n Section 5 that pilot comfort had a strong influence on the maximum airspeeds attained in operationaluse. However, improved vibration attenuation techniquesdeveloped after the CH-47A should eliminate this constrainton future helicopters.
OPERATIONAL DATA ACQUISITION fu~D ANALYSIS
During the review of operational data, several shortcomingswere found in the types of measurements and analysis techniques for use in assessment of fatigue exposure. The following recomrrlendations are made for future acquisition and analysis of opera~ional data:
1. The data sample should include a complete cross section ofoperating conditions, such as training and operationalsquaorons in variors parts of the world.
2. The measurement list should be tailored to meet t.he needsof fatigue assessment. Direct measurement of loads infatigue critical components should be considered.
3. Recording systems compatible with automated data reductionand analysis are desirable.
40
41
-- ~ -~- ----~-
4. The data should be reduced in a format appropriate to theintended analysis to be performed.
1.
2.
Braun, J.F., et aI, CH-47A Chinook Flight Loads Investiga~ion Program; USAAVLABS Technical Report 66-68, U.S. ArmyAviation Materiel Laboratories, Fort Eustis, Virgina, July1966.
Geissler, F.J., et aI, Flight Loads Investigation cf Cargoand Transport CH-47A Helicopters Opc:"ating in SoutheastAsia; USAAVLABS Technical Report £8-)" U.S. Army AviationMateriel Laboratories, Fort Eustis! 'iirginia, April 1968.
3. Investigation 0:: Impact of RVN Flight Survey Data onFatigue Critical Components, Boeing Report 114-SS-025,July 1968.
4. Dynamic Component Fatigue Life Evaluation, CH-47A Helicopter, Boeing Report 114-SS-001, March 1966.
5. Structural Design Requirements (Helicopter), Naval AirSystems Command Aeronautical Requirement (AR)-56, 17February 1970.
6. Boeing-Vertol CH-47A Helicopter Dynamic System StrengthS~bstantiation Bench Test Data and Life Calculations,Boeing Report 114-T-112.2, January 1965.
7. CH-47A Helicopter Dynamic Component Life Evaluation forFlight Envelope, Boeing Report 114-SS-002, December 1965.
42