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