ad-773 551 - dtic · 1 rcport titl1 ch-54a design and operational flicht loads study 4. oeicmlptlvt...

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AD-773 551

CH-54A DESIGN AND OPERATIONAL FLIGHT LOADS STUDY

A. L. Mongillo, Jr., et al

United Aircraft Corporation

Prepa red for:

Army Air Mobility Research and Development La boratory

November 1973

DISTRIBUTED BY:

urn National Technical Information Service U. S. DEPARTMENT OF COMMERCE 5285 Port Royal Road, Springfield Va. 22151

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iicnoi-AVfit.; DISCLAIMERS

The.findings In thin report are not to be construed as an official DeÄrtment oi? the Army position unless so designated by other authorized doiumnts. I

When Government drawings, specifications, or other data are used for any purpose other than In connection with a definitely related Government procurement operation, the United States Government thereby incurs no responsibility nor any obligation whatsoever; and the fact that the Government may have formulated, furnished, or in any way supplied the said drawings, specifications, or other data Is not to be regarded by implication or otherwise as in any manner licensing the holder or any other person or corporation, or conveying any rights or permission, to manufacture, use, or sell any patented invention that may in any way be related thereto.

Trade names cited in this report do not constitute an official endorse- ment or approval of the use of such commercial hardware or software.

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Destroy this report when no longer needed. Do not return it to the originator.

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I odiorN« TIN« ACTIVITY (Cotporml* mulhor)

Sikorsky Aircraft Division Unitcil Aircraft Corporation Stratford, Connecticut

2M, NCPOMT SKCUNITV CL AtllFIC A TION

Unclassified

1 RCPORT TITL1

CH-54A DESIGN AND OPERATIONAL FLICHT LOADS STUDY

4. OEICMlPTlvt NOTCt fTVp« of r*porr and/ncfua/» <<•»<>

Final Report 9 AuTHOHlt) f^iraf n«aM. mtddlm Inlllml, tmalnmmi)

A. L. Mongillo, Jr. S. M. Johnson

«. »cponr DATt

November 1973 •a. CONTRACT OH GRANT NO.

DAAJO2-72-C-0060 ?tUlJ ». nOJICT NO.

«Task 1FI(>22()4AA8201

7«. TOTAL NO- OP PAOCS

M. onioiNA

IO- OP PAOCS

TOR'* RIPORT

7». NO. OP RCPt

13 NUMVCRtt)

USAAMRDL Teclinical Report 73-39

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SER-64370 . 10. DISTRIBUTION tTATKMCNT

Approved for public release; distribution unlimited.

II. SUPPUCMCNTARV NOTES 12. SPONSORING MILITARY ACTIVITY

Eustii Directorate, U. S. Army Air Mobility Rescvrcli and Development Laboratory, Fort Eustii, Virginia

If. ABSTRACT

An analytical and correlation study ol predicted fatigue design data and operational fliglit loads data for crane-type helicopters was conducted to compare operational mission profiles with a design mission profile and to provide data for use in establishing structural design criteria for future Army helicopters.

Flight loads and usage data for CH-54A helicopters operating in Southeast Asia were compared with CH-54A design data. The effects of gross weight and altitude on true airspeed were determined. Fatigue spectra were developed for six dynamic components, and fatigue lives were calculated lor thesi components. These fatigue lives were compared with lives predicted during CH-S4A design. Service histories lor these components were reviewed, and it was found that none ol the changes made in these components resulted from load conditions. Peak operational load parameters were compared with limit static design values. Recommendations were then developed to assist in establishing future crane helicopter fatigue design criteria.

NA ! i' iNAI 11 i'I INI' M INI uKMAlldii Si K'Vii !

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DD IIM»VM1473 OSMLBTB rea «MIY usi. UNCLASSIFIED ■•curlly ClHiineallM

UNCLASSIFIED ■•curilT Claatlflcatlon

KMt WOKOt NOLI WT

CH-54A Helicopter Flight Spectra Helicopter Operations Aircraft Structures

UNCLASSIFIED

f Ay Security CUttlflcttlon

DEPARTMENT OF THE ARMY U S ARMV AIR MOBILITY RESEARCH t DEVELOPMENT LABORATORY

IUSTIS DIRECTORATE FORT EUSTIS, VIRGINIA 23804

This program was conducted under Contract DAAJO2-72-C-O060 with Sikorsky Aircraft Division, United Aircraft Corporation.

The Information presented herein Is the result of an analytical effort to derive Improved structural design criteria f.tr crano-type helicopters based upon flight parameters measured on crane helicopters operating In Southeast Asia. This is one of four similar efforts being conducted concurrently to develop Improved criteria for observation, gunship, and transport as well as crane-type helicopters.

The report has been reviewed by the Eustis Directorate, U.S. Army Air Mobility Research and Development Laboratory and is considered to be technically sound. It Is published for the exchange of Information and the stimulacluu of future research.

This program was conducted under the technical management of Mr, Herman I. MacDonald, Jr., rechnology Applications Division.

/Ä/

Task lFl6220i+AA8201 Contract DAAJ02-72-C-0060

USAAMRDL Technical Report 73-39 November 1973

CH-5UA DESIGN AND OPERATIONAL FLIGHT LOADS STUDY

Final Report

Sikorsky Aircraft Report SER-6it370

By

A. L, Mongillo, Jr. S. M. Johnson

Prepared By

Sikorsky Aircraft Division United Aircraft Corporation

Stratford, Connecticut

for

EUSTIS DIRECTORATE U. S. ARMY AIR MOBILITY RESEARCH AND DEVELOPMENT LABORATORY

FORT EUSTIS, VIRGINIA

Approved for public release; distribution unlimited.

(/

ABSTRACT

Sikorsky Aircraft has conducted an analysis and correlation study of predicted fatigue design data and operational flight loads data for crane- type helicopters. The purpose of the study was to compare operational mission profiles with a design mission profile and to provide data for use in establishing structural design criteria for future Army helicopters,

In this stud^, flight loads and usage data from USAAVLABS Technical Report 70-73, "Flight Loads Investigation of CH-5Â Helicopters Operating in Southeast Asia," were compared with Ctt-^hk design data. The effects of gross weight and altitude on true airspeed were determined. Fatigue spectra were developed for six dynamic components, and fatigue lives were calculated for these components. These fatigue lives were compared with lives predicted during CH-5Â design. Service histories for these components were reviewed, and it was found that none of the changes made in these components resulted from load conditions. Peak operational load parameters were compared with limit static design values. Recommendations were then developed to assist in establishing future crane helicopter fatigue design criteria.

Comparison of CH-5Â operational mission profiles with the design mission profile indicated that crane operating conditions in a combat environment were generally less severe than predicted. The fatigue substantiation of the six selected components confirmed this. Extended or "unlimited" replacement times resulted for all six components.

Airspeeds above 90 knots were rarely associated with an external payload configuration. Most aircraft flight time occurred in a density altitude range of 2000 to 5000 feet. Approximately 91% of the measured load factor peaks occurred at gross weights at or below 29,000 pounds.

Future Army helicopter designs will benefit from improved data collection and editing techniques. Better definition of discrete ground and flight regimes is required to develop accurate mission profiles. Consideration should be given to development of a composite operational spectrum based upon a combat environment and on peace-time operation. Knowledge of peak loads and specific load parameters, such as main rotor head moment or main and tail rotor flapping angles, would yield more accurate fatigue load prediction.

iii

FOREWORD

This report covers the work performed by Sikorsky Aircraft to analyze and correlate CH-5^A design and operational fatigue spectrums. The program was authorized by the U.S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, under Contract DAAJ02-72-C-0060, Task IFI6220UAA82OI. The project monitor for the Army was Mr. H. I. MacDonald, Jr.

The authors express appreciation to Messrs. M. Chrissanthis, G. Chuga, A. Clarke, R. Frye, W. Jackson, E. McLaud, and L. Vacca of Sikorsky Aircraft for their contributions to this report.

Preceding page blank

TABLE OF CONTENTS " Page

ABSTRACT iii

FOREWORD v

LIST OF ILLUSTRATIONS viii

LIST OF TABLES x

LIST OF SYMBOLS xi

INTRODUCTION 1

OPERATIONAL MISSION PROFILES 3

Assumptions for Developing Operational Mission Profiles ... 3 Gross Weight, Altitude, and Airspeed Distribution T Comparison of Crane Mission Profiles 13 Effects of Altitude and Gross Weight on Airspeed 15

FATIGUE ANALYSIS 20

Selection of Components 20 Selection of Flight Loads 20 Replacement Time Calculations 30 Component Evaluation 30

CH-5HA STRUCTURAL COMPONENT HISTORY hk

Introduction hk Main Rotor Pitch Control Horn hk Main Rotor Shaft 1*5 Tail Rotor Spindle 1*6 Main Rotor Hub 1*7 Main Rotor Head Spacer Assembly 1*7

COMPARISON OF OPERATIONAL AND DESIGN LOADS 52

Introduction 52 Gross Weight and Airspeed Comparisons 53 Rotor Speed Comparisons 5l* Maneuver and Gust Load Factor Comparisons 55 Prediction of Load Occurrence Frequency 56

CONCLUSIONS 6l

RECOMMENDATIONS 62

LITERATURE CITED 61*

DISTRIBUTION 65

vii Preceding page blank

LIST OF ILLUSTRATIONS

Figure

1 CH^i+A Three-View Drawing, Basic Data, and Design Limits . 2

2 Grouping of Data from TR 70-73 8

3 Parametric Distribution - Cruise 9

h Parametric Distribution - Ascent 10

5 Parametric Distribution - Descent 11

6 True Airspeed Vs. Cumulative Percentage of Time for Gross Weights Less Than 29,000 Pounds 16

7 True Airspeed Vs. Cumulative Percentage of Time for Gross Weights Greater Than or Equal to 29,000 Pounds but Less Than 37,000 Pounds 16

8 True Airspeed Vs. Cumulative Percentage of Time for Gross Weights Greater Than or Equal to 37,000 Pounds . . 17

9 True Airspeed Vs. Cumulative Percentage of Time for Gross Weight Ranges and a Mean Density Altitude of 2,500 Feet 17

10 Main Rotor Pitch Control Horn Fatigue Substantiation Data 31

11 Main Rotor Shaft Fatigue Substantiation Data 32

12 Tail Rotor Spindle Fatigue Substantiation Data .... 35

13 Main Rotor Hub Fatigue Substantiation Data 36

I** Magnesium Main Rotor Spacer Fatigue Substantiation Data . 39

15 Aluminum Main Rotor Spacer Fatigue Substantiation Data . hi

16 Main Rotor Pitch Control Horn Assembly kk

17 Main Rotor Shaft Assembly 1*5

18 Tall Rotor Spindle Assembly 1+6

19 Main Rotor Head Assembly 1|8

20 Load Factor Vs. Gross Weight Envelope 53

21 Maneuver V-N Diagram 5^

vill

Figure Page

22 Gust V-N Diagram 55 z

23 Prediction of Maneuver Load Factors Above 200 Hours . . 57

2k Prediction of Gust Load Factors Above 200 Hours ... 58

ix

LIST OF TABLES

Table Page

I Comparison of Mission Profiles 1+

II Time Ratios for Parametric Distribution 12

III Indicated Airspeed Corrected to True Airspeed .... 18

IV Operational Fatigue Spectra 21

V Main Rotor Shaft Replacement Time Calculation .... 33

VI Main Rotor Hub Replacement Time Calculation 37

VII Magnesium Main Rotor Spacer Replacement Time Calculation 1+0

VIII Aluminum Main Rotor Spacer Replacement Time Calculation . 1+2

IX Comparison of Component Replacement Times 1+3

X Component Histories 1+9

XI Comparison of Static Design Loads With Operational Flight Loads 52

SYMBOLS

C™, Nondimensional blade loading

D Fatigue damage per 100 hours for entire mission profile

d. Fatigue damage per 100 hours for a particular segment of the mission profile

F_ Bending stress on main rotor shaft at strain gage location, G

psi

H H , H , Time ratio for distribution of altitude in each mission u' J segment, t, . -t H^ ^ hi/

h. Density altitude, ft

KIAS Indicated airspeed, kn

L Distance from centerline of hub to strain gage location on main rotor shaft, in.

M Bending moment on main rotor shaft at strain gage location due to blade flapping, in.-lb

M Net bending moment on main rotor shaft at strain gage location, in.-lb

NL. Bending moment at centerline of main rotor hub, in.-lb n

Mq Tail rotor spindle moment, in.-lb

n Number of segments in the mission profile contributing to fatigue damage

N Vertical load factor at aircraft center of gravity, g

P Main rotor vibratory pushrod load, lb

R Time ratio for ascent portion of each mission segment, tA/T

R^ Time ratio for cruise portion of each mission segment, (tc - W^

R Time ratio for descent portion of each mission segment, tD/T

S/N Curve indicating cycles to crack detection

XI

T Total time for a particular data sample, min

t Time for each mission segment, min

TAS True airspeed, kn

T. Total ascent time, percent

T Total cruis^ time, percent

T Total descent time, percent

t Time in each mission segment for climb rates > +300 ft/min, min

t Time in each mission segment for climb rates > -300 ft/min but <+300 ft/min, min

t Time in each mission segment for climb rates < -300 ft/min, rain

t Tirae in each mission segment for a particular altitude group, rain

tj. Time in each mission segment for y <0.05 at climb rates > -300 ft/min but < +300 ft/min, min

T._ Main rotor axial thrust, lb MR

t . Time in each mission segment for a particular speed group, rain

t . Time in each mission segment for a particular gross weight wl

group, mm

V_T Limit dive speed, KTAS Db

Vu Limit operating speed, KTAS n

V , V , V , Time ratio for distribution of airspeed in each mission

v^, v5 se^en^ Ki/*

W , W , W Time ratio for distribution of gross weight in each mission segment, t . ,t

wi/

3 Z Main rotor shaft section modulus at strain gage location, in.

ß Main rotor flapping f-ngle, deg

xii

\i Advance ratio, JL QR

3 p Air density, slug/ft

p Air density, sea level standard, slug/ft

<|) Aircraft bank angle, deg

S2R Rotor tip speed, ft/sec

xm

INTRODUCTION

The CH-5UA crane helicopter was used extensively in Southeast Asia to assist in U.S. Army airmobile operations. During regular missions, several CH^Âs were instrumented to record flight data for an operational flight loads report. The resulting report, USAAVLABS Technical Report 70-73, "Flight Loads Investigation of CH-5ÜA Helicopters Operating in Southeast Asia" (Reference l), was used in the present study to establish a rational approach for development of future crane helicopters and to improve fatigue criteria for dynamic component design. Figure 1 shows a three-view drawing, basic data, and design limits for the CH-5Â.

In the present study, fatigue spectra are developed for six components: the main rotor pitch control horn, main rotor shaft, main rotor hub, tail rotor spindle, magnesium main rotor spacer, and aluminum main rotor spacer. The main rotor pitch control horn analysis employs a spectrum of main rotor vibratory pushrod loads, and the tail rotor spindle fatigue analysis employs a spectrum of tail rotor spindle moments. The main rotor shaft, hub, and spacer fatigue analyses employ a spectrum of main rotor head hub moments. Fatigue lives are then compared with current replacement times to assess the impact of the operational flight loads spectrum on life- limited components.

A historical summary tracing changes in component fatigue lives or configurations is presented to identify any situation in which the need for design improvement was indicated.

Comparison of reported maximum one-time occurrences of selected load parameters with structural limit design values indicates that service usage of the CH^Â was predominantly within the structural design limits of the aircraft. However, in certain instances the aircraft were operated above the design gross weight and the design speed. A rational basis is pro- posed for predicting the frequency of occurrence of maneuver and gust load factors for future crane helicopters.

Finally, recommendations are made to facilitate the derivation of future mission profiles, and conclusions are offered based on the results of this study.

■ -^7% 4 JUT' -i

BASIC DATA

Main Rotor Dia. M.R. Solidity M.R. Blade Chord No. of M.R. Blades Engines (two) Max. Design Gr Wt Empty Weight

72 ft 0.0861+9 1.629 ft 6 . P&W JFTD-12A-14A 1+2000 lb 19231+ lb

DESIGN LIMITS

Normal Rated Power *l+000 shp Transmission Limit Power 6600 hp Max. Horiz. Fit. Airspeed 110 kn Limit Dive Speed 125.5 kn Limit Load Factor, N 2.26g

* One Engine

Figure 1. CH-5I+A Three-View Drawing, Basic Data, and Design Limits.

OPERATIONAL MISSION PROFILES

ASSUMPTIONS FOR DEVELOPING OPERATIONAL MISSION PROFILES

Operatioral mission profiles were derived from the data in TR 70-73 for samples 1 and 2 (see Table I). Because the data as presented in TR 70-73 did not conform to standard Sikorsky flight regime definitions and since fatigue load levels on main and tail rotor head components are highly- dependent on flight regime, a procedure was established to reclassify the field data into a more detailed mission profile. The four-segment breakdown of TR 70-73, namely, ascent, maneuver, descent, and steady state, was restructured into specific flight regimes consisting of ground, hover, sideward and rearward flight, ascent, descent, cruise, takeoff maneuvers, maneuvers, hoist, and gust. This detailed mission profile enabled a more accurate selection of flight conditions to be made from Sikorsky flight test records.

The reclassification of the TR 70-73 four-segment mission profile to the ten-segment profile used to develop fatigue loading spectra was derived from Tables VI (sample l) and XLII (sample 2) by the following procedure:

From Tables VI and XLII

Ascent

Descent

Steady State

Maneuver

— Ascent

- Descent

r- Ground-

.Steady State- •Hover•

— Cruise

Ascent

Descent

■ Ground

—Steady Hover & Turns

— Sideward & Rearward

mm Takeoff Maneuvers

—Hoist

f Cruise

Maneuvers

Gust

Three basic assumptions were made to derive the ascent, descent, and steady-state flight regimes from Table VI and Table XLII.

1) Ascent was assumed to be all positive climb rates.

2) Descent was assumed to be all negative climb rates except for the -300 ft/min climb category.

TABLE I. COMPARISON OF MISSION PROFILES

Percentage of

Sample 1 Sample 2 of TR 70-73 of TR 70-

Aircraft Time Cll~5kA Design Mission

-73 Profile

CH(Crane) Mission Profile of AR-56

Ground 0.295 0.186 2.110 1.000

Hover, Steady & Turns 5.880 7.015 21.385 16.000

Sideward & Rearward 0.653 0.779 2.175 1.500

Ascent 19.373 17.0U 8.1+22 5.670

Descent 20.»450 18.798 5.270 5.558

Cruise 52.790 55.753 53.21+0 61.231+

Takeoff Maneuvers 0.159 0.100 0.020 1.330

Maneuvers 0.303 0.21+8 7.328 7.708

Hoist 0.01+9 0.01*5 - -

Gust 0.0J+8 0.032 - -

Total 100.000 100.000 100.000 100.000

3) Steady state was assumed to be all time at -300 ft/min climb rate, since this category includes all time between -300 ft/min and +300 ft/min.

The detailed procedures by which the ascent, descent, and steady-state regimes were expanded to the ten-segment mission profile are explained in the following eleven steps:

1) Using Tables VI (sample l) and XLII (sample 2) of TR 70-73, all time at -300 feet/min climb rate (all climb rates - -300 to < +300 ft/min) was considered to be cruise, hover, and ground operation regimes.

2) Data from the same tables were employed, end the ground operation regime was assumed to be the sum of the following:

a) All time for -300 ft/min climb rate for C /o = less and W = 0.0, plus

b) 23% of the time for -300 ft/min climb rate at C /o = less and U = less.

3) Data from the same tables were employed, and hover time, including sideward and rearward flight, was assumed to consist of

a) All time for -300 ft/min climb ate and

C^/a = .06, u = less and y = 0.0 and

C /o = .09, y = less and y = 0.0, plus

b) 13% of the time for -300 ft/min climb rate, C /o = less and y = less.

k) The time for the cruise regime (including maneuver and gust N ) was found by subtracting results of 2 plus 3 from 1 above.

5) Data from Tables VI and XLII were employed, and the ascent regime was established as all time at climb rates greater than +300 ft/min.

6) Data from the same tables were employed, and the descent regime was established as all time at climb rates less than -300 ft/min.

7) The maneuver Nz occurrences were categorized as listed below in a, b, and c. The pulse times used in conjunction with these maneuvers are given in parentheses. These pulse times were developed from a study of Sikorsky flight test data and were conservatively assumed to bp a rectangular pulse of peak Nz versus time.

a) Hoist Maneuvers (3.0 sec) - The number of Nz occurrences for the hoist mission segment, as given in Tables XXVI (sample l) and LXI (sample 2). Hoist Nz occurrences were recorded during hover while picking up or releasing pay loads.

b) Takeoff Maneuvers (12 sec) - Based on Tables XXVI (sample l) and LXI (sample 2), all Nz occurrences at y - 0.0 for the ascent, descent, maneuver, and steady-state mission segments.

c) Maneuvers During Cruise - All load factor (Nz) occurrences in Tables XXVI (sample l) and LXII (sample 2) at y > 0.0 for the ascent, descent, maneuver, and steady-state mission segments. These maneuvers were classified as transitions, pull-ups, and turns using the following procedure:

i. For each mission segment (ascent, descent» steady state, and maneuver), all occurrences for Nz - .8 were considered negative A Nz transition maneuvers.

ii. All remaining maneuvers (Nz - 1.2) were distributed as follows:

a. The number of positive ANZ transition maneuvers equaled 50% of the number of negative ANZ transition maneuvers for all mission segments except steady state, where 23% was used.

b. The remaining positive ANZ occurrences were distributed between pull-ups (3 sec) and turns for the ascent, descent, and maneuver mission segments. All remaining positive ANZ occurrences for the steady-state mission segment were assumed to be turns.

iii. The total number of transition maneuvers obtained above were further subdivided into landing approaches (6 sec), lateral reversals (3 sec), and longitudinal reversals (3 sec). Similarly, the turn maneuvers were broken up into 90° turns (6 sec) and 180° turns (12 sec).

8) The maneuver times calculated in Ta and 7b above were subtracted from the hover time found in 3 above.

9) Ninety percent of the time obtained in 8 above was assumed to be steady hover including hover turns, and the remaining ten percent was assumed to be sideward and rearward flight.

10) For all mission segments, the gust Nz peaks in Tables XXV (sample l) and LXI (sample 2) were assumed to occur during the cruise regime.

11) Net cruise time was calculated by subtracting the maneuver times of 7c and 10 from the time computed in h above.

GROSS WEIGHT. ALTITUDE. AND AIRSPEED DISTRIBUTION

Further breakdown of the operational mission profiles in Table I by gross weight, altitude, and airspeed was required. Since sample 1 and sample 2 were similar (see Table l), sample 1 was chosen for a detailed usage spectrum. To reduce the number of data points that would result from this breakdown, the full ranges of gross weight, altitude, and airspeed were grouped into several broader categories (see Figure 2).

Based on the histograms in TR 70-73, the percentage of time was determined for each of the three gross weight, four density altitude, and five airspeed categories. Using the assumptions discussed earlier for developing the mission profiles from TR 70-73, the total aircraft life (203.^ hours for sample l) was divided into descent, cruise, and ascent by rate of climb (see Figure 2).

Figures 3, *+, and 5 illustrate the procedure used to calculate the time for each element of the detailed usage spectrum. Time ratios, noted by symbols in each of these three figures, were determined using the gross weight, altitude, and airspeed breakdowns from Tables IV and XL and the mission segments from Tables VI anO XLII. All maneuver mission segment time, however, was combined into the cruise regime (RQ), since it represented only .k6% of sample 1. (Maneuvers are discussed later.) These time ratios were multiplied together as indicated by the arrows in Figures 3, *+. and 5, and the results for each mission segment were summed to obtain the total time in each element of the detailed usage spectrum (see Table IV, presented in the "Fatigue Analysis" section).

For example, the total cruise time (T^) for 25,000 pounds (VL ), sea level (H1), and 70 knots (V ) was

Tc = 100 {(Rc x W1 x ^ x V ) for ascent

+ (R^ x W1 x IL x V ) for descent

+ (R_ x W1 x H.. x V_) for steady state

+ (R,, x W1 x IL x V )} for maneuver Equation (l)

(Rc, W , H , and V are defined in Figure 3 and Table II)

Realistic mean ascent and descent airspeeds of 60 and 70 knots, respectively, were used throughout these regimes. Using this procedure, gross weight and altitude distributions were the only considerations for the ascent and descent regimes. The procedure was similar to that for the cruise regime, and the resulting calculations are reflected in Table IV.

Parameter

Gross Weight (lb)

Data Category (TR 70-73)

^ 29000

15*29000 but ^37000

^ 37000

Value Used In Study

25000

33000

39000

Density Altitude (ft)

^rlOOO

1000 but ^ 2000

2000 but -i 5000

^5000

Sea Level

1000

2000

5000

1+0

I ^ l+o but ^ 60 Airspeed

(KIAS) / ^60 but ^80

/ ^80 but ^95

\ ^95

/ ^-300

Rate of Climb (ft/min)

) ^300 but ^+300

+300

20

140

70

85

100

Descent

Cruise, Hover, Maneuver

Ascent

Figure 2 . Grouping of Data from TR 70-73

GROSS DENSITY AIRSPEED WEIGHT ALTITUDE (KIAS ) (POUNDS) (FEET)

ASCENT PORTION OF EACH MISSION SEGMENT

CRUISE PORTION OF EACH MISSION SEGMENT

IRr

DESCENT PORTION OF EACH MISSION SEGMENT

(M

Figure 3. Parametric Distribution - Cruise (See Equation l).

GROSS DENSITY WEIGHT ALTITUDE (POUNDS) (FEET)


(RA)


(Rc)


(RD)

Figure h. Parametric Distribution - Ascent (See Equation 2).

10


(Ri


IRC)


GROSS DENSITY WEIGHT ALTITUDE (POUNDS) (FEET)

IR,

Figure 5. Parametric Distribution - Descent (See Equation 3),

11

TABLE II. TIME RATIOS FOR PARAMETRIC DISTRIBUTION |

Time Ratio j Symbol

Mission S 2Kment (From TR 70-73) i Ascent Descent Steady State Maneuver

RA .1766 .OOl+U .0127 0 i

RD .0030 .1880 .0135 0

Rc .0682 .0836 .371+9 .001+6

Wl .5125 .631+3 .61+28 1.000 1

W2 .0799 .0979 .0972 0 1

W3 .1+076 .2678 .2601 0

Hl .0097 .0115 .0070 0

H2 .1021+ .1267 .01+75 0 1

H3 .78I4U .71+91 .6090 .6361+

h .1036 .1127 .3365 .3636 1

Vl .0873 • 0855 .0097 0 1

V2 .21+37 .11+78 .01+51+ 0

V3 .1+667 .31+91 .31+69 .1031+ 1

Vk .3065 .1+521+ .6988 j

u5 .0318 .1112 .11+56 .1979

12

For example, the total ascent time (T ) for 39,000 (W ) and 2,000 feet (H3) was

A i

T. = 100 l(R, x W0 x H_) for ascent A ' A 3 3

+ (R x W x H ) for descent

+ (R x W x H )| for steady state Equation (2) A j j

(R , W , and H are defined in Figure h and Table II.)

The total descent time (T ) for 33,000 (W ) and 5,000 feet (H,) was

T = 100 {(R x W x H. ) for ascent

+ (R x W x H, ) for descent

+ (R x W x HI )) for steady state Equation (3)

(R , W , and H, are defined in Figure 5 and Table II.)

COMPARISON OF CRME MISSION PROFILES

Samples 1 and 2 of TR 70-73, the CH-5I+A design mission profile, and the OH (Crane) mission profile of AR-56 (Reference 2) were tabulated for comparative analysis (see Table I). Sample 1 correlated well with sample 2. There were significant discrepancies between several regimes of samples 1 and 2 and the CE-^kA design mission profile and similarly between samples 1 and 2 and the CH (Crane) mission profile of AR-56. The CH (Crane) mission profile of AR-56 correlated well with the CH-51+A design mission profile. The Crane profile is believed to be derived from the CH-5'+A design mission profile.

Ground Time

The ground time derived from TR 70-73 for the operational mission profiles (samples 1 and 2) was small compared with the CH-5^A design spectrum and the CH (Crane) mission profile or AR-56. This time constituted 0.295^ and 0.186^, respectively, of the total data sample times compared with 2.11% for the CH-51+A design spectrum and 1.0^ for the AR-56 spectrum. TR 70-73 indicated that the pilots used a switch in the cockpit to start the recording system that recorded in-flight data. The apparent ground time was probably recorded Just prior to takeoff, whereas the design spectrum accounted for full warm-up.

Hover

Time in hover experienced under combat conditions in Southeast Asia differed significantly from CH-5^A design hover time. These results are understand- able. Under combat conditions, extended hovering is avoided to reduce vul- nerability to enemy fire. The design spectrum, on the other hand, accounted for tasks involving extended periods of hover, including pilot training.

13

Ascent and Descent

Comparatively high ascent and descent times for the operational mission profile were influenced by the requirements of the combat environment. Field reports indicated that a typical ascent or descent maneuver was executed in a spiral fashion. Apparently, the pilot could make his approach to a drop zone at cruise altitude, staying out of rifle range as long as possible. He could then slowly make his descent while in relative safety from small-arms fire throughout his approach, since he would be close to an assumed friendly position. The same pattern would be used for ascent. Sikorsky field reports from Southeast Asia also indicated that most missions involved legs of 30 nautical miles or less. This would also tend to increase ascent and descent time.

Maneuvers

The low frequency of maneuvers during operation compared to design raised the question of whether raw data editing had eliminated large-radius, low-load-factor turns as maneuvers and, instead, categorized these turns as other mission segments, such as ascent and descent. As indicated in TR 70-73, only those load factors - .80g and - 1.20g were classified as Nz

occurrences. This implied that turns at bank angles up to 33.5° were not included as maneuvers, since bank angle is a function of load factor. That is,

$ = cos"1 l

= cos-1

= 33.5°

N z

1.2

where $ is aircraft bank angle.

Addition of these turns to the maneuver regime derived from operational data would produce a significant increase in maneuver time and bring it more in line with the CH-5^A design mission profile.

Ih

EFFECTS OF ALTITUDE AND GROSS WEIGHT ON AIRSPEED

Gross weight, density altitude, and airspeed are important considerations in the fatigue design of helicopter dynamic components. These parameters have a diret t influence on the magnitude of control loads and flapping and power requirements which, in turn, affect dynamic component rejlace- ment times. With this in mind, it becomes apparent that accurate knowledge of gross weight, altitude, and airspeed combinations would improve design mission profile development.

Data from TR 70-73 were compiled for sample 1, relating gross weight and density altitude to airspeed. Time was tabulated for airspeeds greater than or equal to 1+0 knots (IAS), neglecting any lesser speeds. This accounted for 87.2!?. For this particular study, only 1000-foot, 2000- foot, and 5000-foot density altitudes were considered, accounting for 98.W of the time at airspeeds greater than or equal to hO knots. Gross weight groupings described in Figure 2 were also aiplicable to this study.

As a procedural example, the time for the 1000-foot density altitude was compiled for all gross weights less than 29 »000 pounds and for the entire speed range - ho knots. The percentage of time for a particular speed within the overall speed range was then calculated. These individual percentages, totaling 100?» of the time at airspeeds - hO knots, were accumu- lated beginning with the highest speed. The results were then plotted acainst true airspeed (Figures 6, 7> and 8), indicating the percentage of time that any given speed was equaled or exceeded for a particular density altitude and weight range. In this way, typical operating procedures could be shown graphically. In a similar manner. Figure 9 was developed to illustrate the percentage of time that any airspeed was equaled or exceeded for a particular weight range.

As noted in Figures 6, 7> 8, and 9» the indicated airspeeds of TR 70-73 were corrected to true airspeed, accounting for instrument error and density altitude (Table III). The results appearing in Figure 9 were further adjusted, using a weighted average, to combine the data for all altitudes into one mean altitude (2500 ft). This approach lent itself to c more meaningful comparison of gross weight versus airspeed.

Reduction of Southeast Asia flight loads data indicated that certain operating procedures were consistently practiced by U.S. Army pilots. One of the more predominant characteristics of crane helicopter operation was the airspeed-gross weight relationship. At aircraft gross weights greater than or equal to 37»000 pounds, the aircraft were flown very conservatively. High-gross-weight missions generally involved transport of bulky items by the single-point hoist system, making low-speed flying essential for maintaining safe control of the slung load.

Figure 9 confirms that heavy loads were flown at relatively low speeds, especially compared with gross weights below 29,000 pounds. For this weight range, the aircraft would not have been carrying any external cargo.

15

120-

§100 z

a it!

< UJ

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2 3%-) 40 6 % (

17 8 %

5000 FT , MAX RECORDED SPEED = 132 2000FT , MAX RECORDED SPEED = 118 1000 FT , MAX RECORDED SPEED = 106

* TAS - TRUE AIRSPEED (SEE TABLE IH) ** PERCENTAGE OF TIME AT DENSITY ALTITUDE FOR GROSS WEIGHTS < 29000

AND SPEEDS > 40 KNOTS

40 _i_ -L. J_ -L. -1- -1- J 2 3 4 5 6 7 8 9 10 20 30

PERCENTAGE OF TIME AIRSPEED WAS EQUALED OR EXCEEDED

40 50 60 7D 809000

Figure 6. True Airspeed Vs. Cumulative Percentage of Time for Gross Weights Less Than 29,000 Pounds.

5000 FT, MAX RECORDED SPEED = 113 2000 FT , MAX RECORDED SPEED = 97 IOOOFT , MAX RECORDED SPEED = 101

TAS1

100

a

<

80-

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40L-

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J_ 5 6 7 8 9 K) 20 30 40 50 60 7D 8090100


Figure 7. True Airspeed Vs. Cumulative Percentage of Time for Gross Weights Greater Than or Equal to 29,000 Pounds but Less Than 37,000 Pounds.

16

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40

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PERCENTAGE OF TIME

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20

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EXCEEDED

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Figure 8. True Airspeed Vs. Cumulative Percentage of Time for Gross Weights Greater Than or Equal to 37,000 Pounds.

120

in

o 100

en 80

a. 60-

40'r

GROSS WEIGHTS ■ 29000 POUNDS GROSS WEIGHTS ■ 29000 BUT ■ 37000 POUNDS GROSS WEIGHTS • 37000 POUNDS

NOTE TRUE AIRSPEEDS OBTAINED BY WEIGHTED AVERAGES OF DENSITY ALTITUDE CORRECTIONS (SEE TABLE ID )

^L -I _J_ _L. O- 2 345678910 20 3040


J l_i 50 60 70 809000

Figure 9. True Airspeed Vs. Cumulative Percentage of Time for Gross Weight Ranges and a Mean Density Altitude of 2,500 Feet.

17

TABLE III. INDICATED AIRSPEED CORRECTED TO TRUE AIRSPEED

Indicated Calibrated • True Airspeed (kn) *»

Airspeed Airspeed (kn) (kn) 1000 Ft 2000 Ft 5000 Ft

ho ^5 1+6 1+6 1*8 60 62.5 63.5 61+ 67 65 67 69 69 72 70 71 72 73 76 75 75.5 77 78 81 80 80 81 82 86 85 81*.5 86 87 91 90 89 90 92 97 95 9^ 95.5 97 101

100 99.5 101 103 108 105 10lt.5 106 108 113 110 110 112 113 118 115 115 117 118 121+ 120 123 125 127 132

•Reference 3

**True Airspe Calibrated Airspeed

and 1000 Ft, .971 § 2000 Ft Where V$o = .985 i .928 @ 5000 Ft

18

but would have been loaded to maximum fuel capacity or less. Adjusted for true airspeed corrections at a mean altitude of 2,500 feet, the curves indicate that, for a given speed, considerably more time was spent at gross w ights less than 29,000 pounds than at weights involving a payload. As an example, for true airspeeds of 80 knots, 80% of the time spent below 29,000 pounds was flown at or above this speed compared to only 15^ for weights at or above 37,000 pounds. Examining the curves from another viewpoint, 50% of the time at gross weights less than 29,000 pounds was spent at or above airspeeds of 91 knots (TAS), while the equivalent time at gross weights greater than or equal to 37,000 pounds equaled or exceeded only 69 knots (TAS).

These facts are important considerations in the development of a speed spectrum for cruise and maneuvers . There is a tendency to be overly conservative in establishing an airspeed vs. gross weight distribution; however, the data compiled for this report may help promote a more accurate approach to this problem.

Another important aspect of mission profile development is the relationship of density altitude to aircraft spned. It is common design practice to develop fatigue load spectra at various density altitudes above sea level standard for speeds up to the limit d've speed at the design gross weight. From the standpoint of control loaas, this approach could be excessively conservative for crane helicopter design. Analyzing control loads for high-speed flight at high density altitudes creates high loading conditions. If a more rational approach could be formulated for deter- mining practical density altitude, gross weight, and airspeed combinations, component design improvements could be realized.

Figures 6, 7> and 8 help to illustrate the relationships among altitude, gross weight, and airspeed. True airspeed vs. cumulative percentage of time for each altitude as a funci ion of the entire forward speed range has also been provided in each figure. With this information, the percentage of time that an airspeed was equaled or exceeded can be calculated for any combination of altitude and gross weight.

For example, to determine the percentage of time spent at or above 80 knots for gross weights greater than 37,000 pounds and a density altitude of 5»000 feet or more. Figure 8 would be used. From the figure, 32^ of the time for 5,000 feet and 37,000 pounds was spent at or above 80 knots. Relating this percentage to the total aircraf life spent in forward flight, the result is .32 x h.1% = 1.3%. This approach provides a useful tool for future mission profile development.

19

FATIGUE ANALYSIS

SELECTION OF COMPONENTS

Initially, the main rotor pitch control horn, main rotor shaft, and tail rotor spindle were selected for fatigue analysis. The intention was to compare replacement times based on the operational mission profile with those established during design. However, the analysis showed that the replacement time based on the operational mission profile for both the main rotor pitch control horn and tail rotor spindle was "unlimited," as was the case under the existing Army replacement time schedule. Subse- quently, three additional components were evaluated, using the operational fatigue spectrum: the main rotor hub, magnesium main rotor head spacer (currently being phased out for reasons explained later in the text), and the newly incorporated aluminum main rotor head spacer.

SELECTION OF FLIGHT LOADS

Flight test results on the CH-5^A were available from four Sikorsky engi- neering reports (References 3, k, 5, and 6). Flight loads were selected by matching specific flight regimes of the operational mission profile with actual flight demonstrated maneuvers. When operational flight regimes could not be matched exactly by flight test conditions, conservative alternative conditions were used. When two or more loads were available from flight test data for a particular regime, the highest load was used. When a peak load appeared in the flight test records for conditions of low intensity (low airspeed, altitude, or gross weight conditions), this same load was used for similar conditions of high intensity (high airspeed, altitude, or gross weight conditions). In other words, for reasons such as gust response or high transient pilot inputs, there were cases when peak loads for low-intensity conditions exceeded the highest recorded loads for high-intensity conditions. Table IV summarized the detailed usage spectra and corresponding critical flight loads that were used for component evaluation. Mean and working S/N curves for each component were used to determine damaging flight load levels. Gross weight, center of gravity, altitude, and airspeed were considered in compiling conditions for each component. Only those conditions with corresponding flight test loads exceeding the working strength of the component at IQÖ cycles were considered for the operational fatigue analysis.

No hoist conditions were simulated during the CH-5^A flight test program to determine rotor loads. Therefore, damaging loads experienced during hover regimes were applied to hoist maneuvers.

20

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29

REPLACEMENT TIME CALCULATIONS

Appropriate fatigue design load parameters were selected from CH-5ltA flight test results and paired with corresponding flight conditions of the operational mission profile. The related percentages of time, which were derived from Southeast Asia field data, were used together with these parameters to calculate replacement times.

Replacement Time = 100 D

i=n where D = Ed.

i=l 1

and , % Time d. = - i Allowable Cycles

Cycles/Hour

The allowable cycles at each load level were obtained from appropriate working S/N curves. Cycles per hour were calculated, assuming a constant rotor speed of 185 rpm and one load cycle per rotor revolution.

COMPONENT EVALUATION

(a) Main Rotor Pitch Control Horn

The mean and working S/N curves for the main rotor pitch control horn (Figure 10) were obtained from Reference 7.

As indicated, the maximum vibratory pushrod load encountered during any flight regime of the operational mission profile does not intersect the working S/N curve. Therefore, no fatigue damage is incurred, resulting in an "unlimited" replacement time for the control horn.

Because this does not represent any change from the existing krray replacement time, it was decided to evaluate an alternative component. The main rotor hub was chosen, since it has a relatively low replacement time of 750 hours. This evaluation is made in section (d) below.

(b) Main Rotor Shaft

The mean and working S/N curves for the main rotor shaft (Figure 11) were obtained from Reference 8. Table V presents the replacement time calculations.

Head moments in Table IV were calculated as shown on page 33, using the upper shaft bending stresses and the blade flapping angles recorded in the flight test reports.

30

During laboratory fatigue testing, the main rotor pitch control horn exhibited no

fatigue failure.

LABORATORY TEST FAILURE MODES

>

5 +

o

>: 2

1 +

3 + MAXIMUM MEASURED FLIGHT LOAD t 980 LB,

TOASTS

i I i m-

FAHGUE TEST DATA FROM REF 7

CURVE SHAPE

ALUMINUM WITH CHAFING

MAIN ROTOR PITCH CONTROL HORN S/N CURVE

-\—< I I HIH H t-h ■ft MH I 11 ill

10° 10» I0T

CYCLES TO CRACK DETECTION 10«

Figure 10. Main Rotor Pitch Control Horn Fatigue Substantiation Data.

31

34.977 IN.

I-6.810H IN.

10

z 8 -

6

9 Ul z

2 -

IFATIGUE TEST FAILURE MODEl

Failure Mode ■ Cracking occurred across the lower portion of the upper splines, extending almost completely around the circumference. Cracks also extended longitudinally. Fretting and galling of the upper splines were also evident, concentrated mainly at the ends.

Upper Splinrs

FATIGUE TEST DATA FROM REF CURVE SHAPE

STEEL WITH CHAFING

MAIN ROTOR SHAFT S/N CURVE

ol—H-H-WJ 1—i t 11 ml 1—i—h'-H-wJ 1—i i j Hiti 10» 10« 10 T 10»

CYCLES TO CRACK DETECTION

Figure 11. Main Rotor Shaft Fatigue Substantiation Data.

32

1 TABLE V . MAIN ROTOR SHAFT REPLACEMENT TIME CALCULATION j

^ 6 (in. -lb x 10 ) 1 Ref. Table IV

I % Time Ref. Table IV

Allowt.ble (r Cycles x 10~ ' Ref. Fig. 11

Allowable Hours *

Damage per 100 Hours I

.1+23

.U38

! .391

.698

.395

1.7808

.1958

5.508

.0205

.00082

14.8

3.8

1U.0

.1*25

11.8

14 32

3^

1261

38

1063

.001+12 |

.00057

.001+37 j

.00053

0

I Total Damage/lOO Hours .00959 |

Replacement Vimr - 1^0 -in 1+00 Hours * Timc " .00959 " 10'

* Assumes a main rotor speed of 185 rpm and one cycle

revolution.

per |

33

and G " Z

MG = MH + MB

= M, + T.„ x L x h MR 57.3

where F = Bending stress on main rotor shaft at strain gage location, psi

M = Net bending moment on main rotor shaft at strain gage location, in.-lb

Z = Main rotor shaft section modulus at strain gage location, in.3

M = Bending moment at centerline of main rotor hub, in.-lb

M = Bending moment on main rotor shaft at strain gage location due to blade flapping, in.-lb

T,^ = Main rotor axial thrust, lb MR L = Distance from centerline of hub to strain gage

location on main rotor shaft, in. g = Main rotor flapping angle, deg

The calculated replacement time of 10,U00 hours is compared with existing replacement times in Table IX.

(c) Tail Rotor Spindle

The working S/N curves for the four designs of the tail rotor spindle used at one time or another on the CH^'+A aircraft were obtained from References 9 and 10 and are presented in Figure 12. It should be noted that neither the -Ohl nor the -Ohh spindle is in service at the present time.

As shown in Figure 12, the maximum vibratory spindle moment encountered during the operational mission profile does not intersect the working S/N curve for either the -0U6 or -0^7 tail rotor spindle (see Table X for part descriptions). As a result, no fatigue damage occurs, and an "unlimited" replacement time is obtained. Since this does not represent any change from the design replacement time shown in Table IX, it was decided to evaluate the fatigue life of the main rotor spacers.

(d) Main Rotor Hub

The mean and working S/N curves for the main rotor hub, shown in Figure 13, were taken from Reference 8. The replacement time calculation, based upon the operational mission profile, is shown in Table VI. As can be seen from Table IX, the calculated replacement time of 7 »TOO hours is more than 10 times the FAA-approved replacement time.

3U

5.140 * IN. ♦ 2.570^

IN.

13.125IN

T 2.25 IN. 1

< (9

O z a.

30

LABORATORY TEST FAI LURE MODES

FAILURE MODE: (1) Shank origin (O.D.) - Fatigue cracks originated at the inboard end of shank.

(D-

^r (i) (2) Thread root origin - Fatigue cracks originated in the root of the threads at the outboard end of the shank.

-(2)

MAXIMUM MEASURED MOMENT ♦ 16750 IN-LB, TAKEOFF AND CLIMB CONDITION, REF 6

CURVE CURVE SHAPE DASH NO REFERENCE! TITANIUM WITH- OUT CHAFING

-041,-044 9

TITANIUM WITH CHAFING

046 9

TITANIUM WITH CHAFING

EAR MODE

-047 10 [

TAIL ROTOR SPINDLE S/N CURVE


Figure 12. Tail Rotor Spindle Fatigue Substantiation Data.

35


FAILURE MODE: (1) Hole origin - cracks originat'.-d around bolt holes damaged by heavy fretting.

437 IN.

(2) Cone origin - heavy fretting and galling occurred on cone and flange.

(3) Flange-to-wall radius origin - cracks originated at flange radii extending through the hub wail.

10- r o M

z

9 UJ r 4..

i L

FATIGUE TEST DATA FROM REF 8 CURVE SHAPE

STEEL WITH CHAFING

MAIN ROTOR HUB S/N CURVE

I Mill \-—i—i i I MMJ 1—i i li ml 1—i i j i HJ 10« 10» I07


Figure 13. Main Rotor Hub Fatigue Substantiation Data.

36

TABLE VI. MAIN ROTOR HUB REPLACEMENT TIME CALCULATION

"" 6 (in.-lb x 10 ) Ref. Table IV

^ % Time Ref. Table IV

Allowable g Cycles x 10-

Ref. Fig. 13 Allowable

Hours * Damage per 100 Hours

.1+23 1.781 3.8 31+2 .00520

.1*38 .1958 3.0 270 .00073

.391 5.508 9.1+ 81+6 .00651

.698 .0205 .39 35 .00058

.357 .0280 100.0 9000 0

.395 .00082 8.3 71+8 0

Total Damage/100 Hours .01302

Replacement m.- _ 100 _ -7700 U^,v.„ 1XJ,K- " .01302

* Assumes a main rotor speed of 185 fV™ and one cycle per

revolution.

37

(e) Magnesium Main Rotor Spacer

The original main rotor spacer was made of magnesium and had an Army replacement time of ^,175 hours. Inspection of in-service parts revealed fatigue cracks initiating at corrosion pits in the central hole of the spacer. As a result, the material for the spacer was changed to aluminum. The two parts have identical dimensions. The magnesium configuration is currently being replaced at overhaul. Since it will be some time before the conversion is completed, an evaluation of the magnesium design was made. The aluminum spacer is analyzed in (f) below.

Mean and working S/N curves for the magnesium main rotor spacer, shown in Figure Ik, were taken from Reference 8. Table VII gives the replacement time calculation based upon sample 1 of the operational mission profile.

(f) Aluminum Main Rotor Spacer

Mean and working S/N curves for the aluminum spacer, sh^wn ii; Figure 15, were taken from Reference 8. Replacement time calculations using the operational mission profile are given in Table VIII. As summarized in Table IX, the aluminum spacer has a replacement time approximately twice that of the magnesium spacer.

38

LABORATORY TEST FAILURE MODES]

2.625 IN

2.625 INI _1_

O]

o o

o

O 0 o ■ Y

o ■■;■ o

o

o 3.0 IN.

Fatigue cracks generally occurred on a 45 diagonal on one or both sides of the web boss. Crack initiation was associated with corrosion pitting around the hole.

^A

7.257 IN,

, i of o

Ul

a < ui z

s

6-

fi 2' > S 2-- MAGNESIUM SPACER S/N CURVE

FATIGUE TEST DATA FROM REF 8 CURVE SHAPE:

MAGNESIUM WITHOUT CHAFING

MEAN CURVE

oi i | iiiii 1—i i 11 ml 1—i- i 11 ml—-i—i M i ml 10s 10» I07 10«


Figure ih. Magnesium Main Rotor Spacer Fatigue Substantiation Data.

39

TABLE VII. MAGNESIUM MAIN ROTOR SPACER REPLACEMENT TIME CALCULATION

(in.-lb x 10 ) 1% Time Allowable ,- Cycles x 10 Allowable Damage per

Ref. Table IV Ref. Table IV Ref. Fig. ll+ Hours * 100 Hours

.1+23 1.781 3.3 297 .00599

.1+38 .1958 2.6 231+ .00081+

.391 5.508 8.0 720 .00765

.698 .0205 .205 18 .00111

.357 .0280 62.0 5585 0

.395 .00082 7-2 61+8 0

Total Damage/I00 Hours .01559

Replacement Tii 100 6I1 20 Hours

- " .01559

* Assumes a main rotor speed of 185 rpm and one cycle per

revoluti on.

1 1

1+0

2.625 IN ; o O 0 o o ^>T .v::r"J""I"-:•--":-'• 3.0 IN. O GO. O O^V _i_


During laboratory fatigue testing, the aluminum main rotor spacer exhibited no fatigue failure.

-t -8.500 IN

3.500 IN.

t

2.625 IN./ /^

^rv

7.257 IN.

i

10--

8 ■

6-



2--

FATIGUE TEST DATA FROM REF 8 CURVE SHAPE

ALUMINUM WITH CHAFING

€AN CURVE

ALUMINUM SPACER VN CURVE

I mil 1—i—i I nitl 1—i i i i ml ^—i i I i m I05

I06 10' CYCLES TO CRACK DETECTION

10»

Figure 15. Aluminum Main Rotor Spacer Fatigue Substantiation Data.

hi

TABLE VIII. ALUMINUM MAIN ROTOR SPACER REPLACEMENT TIME CALCULATION

MH (in.-lb x 10"6] Ref. Table IV

1$ Time Ref. Table IV

Allowable . Cycles x 10~6

Ref. Fig. 15 Allowable

Cycles * Damage per

100 Hours

.k23 1.781 6.3 568 .00311+

.1+38 .1958 5.1 1*59 .0001*3

.391 5.508 12.8 1153 .001+78

.698 .0205 .53 1*8 .00043

.357 .0280 38.0 31+23 .00001

.395 .00082 11.8 1063 0

Total Damage/lOO Hours ,00879

Replacement Time = g = 11380 Hours

* Assumes a main rotor speed of 185 mm and one cycle per

revolution.

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CH-Â STRUCTURAL COMPONENT HISTORY

INTRODUCTION

Since the development of the CH-Ûk, many design improvements have been incorporated into its dynamic component systems. In selecting components for a fatigue study, special attention was given to the main rotor head and tail rotor head assemblies. Six components were analyzed to determine replacement times based on the operational mission profile. Table X des- cribes the parts, their functions in the system, design changes during their service life, and corrective measures taken to prevent problems. Other information includes a summary of part numbers , dash numbers, and aircraft effectivities.

MAIN ROTOR PITCH CONTROL HORN

The CH-5Â main rotor pitch control horn (Figure l6), designed and used on the CH-3T» functioned as an integral part of the main rotor servos, swash- plates, and pushrods. The design never changed during the U.S. Army procurement program except for elimination of blade fold capability. This change saved weight and simplified assembly procedures.

Figure 16. Main Rotor Pitch Control Horn Assembly.

hh

MAIN ROTOR SHAFT

The CH-5UA main rotor shaft (Figure IT) was designed to adapt to the new CH-5UA main rotor gearbox and to interface with the existing CH-37 main rotor head. The upper end of the shaft was necked down to fit the CH-3T main rotor hub, since the design of the crane helicopter was developed around the CH-3T main rotor system.

Modifications were made in the upper and lower bearing seats during the development stage by incorporating liners to tighten the bearing fit. For the production aircraft, however, wall thicknesses in the bearing seats were increased during manufacturing, thus eliminating the need for the liners. An additional nonstructural improvement was incorporated into the shaft assembly by installing a pressed cork plug in the previously open upper end of main .-otor shaft. The plug prevented rain, snow, dirt, and tools from contaminating the interior of the shaft and main gearbox oil sump.

UPPER BEARING '///),

Figure IT. Main Rotor Shaft Assembly.

^5

TAIL ROTOR SPINDLE

As has been stated, the Sikorsky crane was developed with a CH-3T tail rotor system. The first three aircraft were built as a company-funded project, using off-the-shelf tail rotor components, despite certain performance limitations. When the U.S. Army indicated an interest in a crane helicopter, Sikorsky took advantage of the CH-53A program and incorporated its new tail rotor design into the CH-5H. The CH-53A tail rotor provided excellent performance characteristics and appreciable growth potential. That tail rotor was used for the duration of the CH-5^A program.

Although fatigue-related structural problems on the tail rotor spindle (Figure 18) were never experienced, design improvements were incorporated to improve fatigue life further (Table X).

Figure l8. Tail Rotor Spindle Assembly.

he

MAIN ROTOR HUB

The Ch-SUPi main rotor hub (Figure 19) was designed originally for the CH-37 and later adapted for the crane helicopter as part of the off-the-shelf CH-5Â main rotor system. During service on the CH-37» improvements made in the hub increased corrosion resistance and made manufacturing easier. Additional bolt holes were added to the hub flanges during the CH-5Â design program to provide added gross weight potential (Table X). No changes in the hub design were ever required to eliminate fatigue damage problems.

MAIN ROTOR HEAD SPACER ASSEMBLY

The main rotor head spacer assemblies (Figure 19) are bolted between each pair of arms of the CH-5'*A upper and lower main rotor plates . The combination of the upper and lower plates with the six spacers forms an I-beam that reacts vertical bending moments and vertical shear resulting from blade loadings. The spacer itself, acting as a web, reacts shear forces.

Such an assembly was incorporated previously on the five-bladed CH-3T. When the same spacer forging was later considered for the CH-5Â, lag damper clearance considerations required that a hole be drilled in the spacer web. During whirl stand fatigue testing, cracks developed at the hole, and the spacer had to be redesigned with a boss around the hole. This appeared to solve the problem, but further cracking eventually occurred.

The spacers originally had been machined from magnesium castings. Because magnesium is highly corrosive, the spacers were pitting in normal service environments. High stress concentration around the pit holes caused the magnesium spacers to crack well before their designed replacement time. To solve this problem, the spacer material was changed to aluminum, pro- viding fatigue strength and corrosion resistance superior to those of magnesium.

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51

COMPARISON OF OPERATIONAL AND DESIGN LOADS

INTRODUCTION

TR 70-73 provided a means for comparing peak operational flight loads with static design loads. Except for a few specific peak loads, only approxi- mate peak values could be derived. TR 70-73 reported that most of the recorded data points represented ranges of loads, making it impossible to establish the exact peak value within the range. Table XI compares gross weight, airspeed, main rotor speeds, and maneuver and gust load factors for operational and design conditions. Additional qualifying parameters are included to relate the data effectively.

TABLE XI. COMPARISON OF STATIC DESIGN LOADS

WITH OPERATIONAL FLIGHT LOADS

Load Parameter

Operational Flight Loads (Ref. l) Static Design

Gross Weight (lb)

Uk009 a 142000

Airspeed (kn)

120 (KIAS) a

132 (TAS) (G.W.= 27000 lb)

VD = 126.5 (TAS)

(VH = 110 (TAS))

Main Rotor Speed (rpm)

198 § 60 (KIAS) a 203 (Hover) b

201+ (Power on) 215 (Power off)

Load Factor (Nz)

1.88 (Hoist g 140100 lb) a 1.5 (Maneuvers) max. a

.5 (Maneuvers) min. a

2.26 max. j -0.5 min.

Gust Load Factor

(Nz)

j25000 lb 1.3 max. <100 (KIAS) a

1103 (TAS) (25000 lb

0.6 min. <100 (KIAS) b (l03 (TAS)

193 (U2000 lb@ 110 TAS)

a Sample No. 1

b Sample No. 2

52

GROSS WEIGHT AND AIRSPEED COMPARISONS

The CH-5ltA was designed for a basic gross weight of 1*2,000 pounds combined with a load factor of 2.26g and a limit dive speed of 126.5 knots. The basic design gross weight also represented the maximum operating weight, since there were no alternate design gross weight provisions for the CH-5^A.

TR 70-73 reported that a peak gro^s weight of l+l+,009 pcunds had been recorded at a speed of 60 knots. Field studies conducted by Sikorsky engineers stationed in Southeast Asia in 1968 indicated that significant numbers of missions were being flown at or above k2,000 pounds. This practice, although possible on the basis of static strength, may have produced damaging loads on fatigue-related components . Overload gross weights probably were flown rarely at speeds higher than 60 to 70 knots or at load factors greater than 1.0 ±0.2g. But even in these low speed and load factor ranges, flapping, power, or control requirements could have approached damaging levels and should have been indicated either by the cruise guide or torque meters. Figure 20 compares load factor vs , gross weight enve- lopes for design and operational data.

2.4

N Z

2.2-

20-

1.8-

1.0-

0.8

CH-54A STATIC DESIGN REF. 11

OPERATIONAL FLIGHT LOADS REF. I

-L _L J_ J_ _L J_ X _L 20 22 24 26 28 30 32 34 36 38 40 42 44

GROSS WEIGHT, LB x I0"3

Figure 20. Load Factor Vs. Gross Weight Envelope.

Since more than 30% of the sorties were flown at gross weights in excess of ^2,000 pounds, as indicated by Sikorsky field studies, considerations of excessive ride discomfort apparently did not discourage this practice. It is possible that any high cockpit vibration experienced, such as during transition into flare, did not persist long enough to warrant any change in flight practices.

53

The CH-5UA was designed for a limit dive speed (VD) of 126.5 knots at the design gross weight of ^2,000 pounds. The maximum indicated airspeed recorded in TR 70-73 was approximately 120 knots for gross weights of 25tO00 and 27,000 pounds. For a density altitude of 5.000 feet, at which this airspeed appeared to be measured, the equivalent true airspeed would be 132 knots. Again, this indicated that structural design limits were exceeded during Southeast Asia operations. In addition to structural limits, a limit operating speed (VH) of 110 knots was established, because severe nose-dc»m attitudes were experienced at low gross weights for high speeds. Apparently, these redlines were also exceeded (Figure 21).

30

a < o

•1.0

CH-54A STATIC DESIGN- G.W.» 42000 LB

REF II

OPERATIONAL- FLIGHT LOADS

(ALL GROSS WEIGHTS) REF I

_L 20 40 60 80

TRUE AIRSPEED, KN 100 120 130

Figure 21. Maneuver V-N Diagram.

ROTOR SPEED COMPARISONS

The CR-^kA was designed for a main rotor speed of 20l* rpm, or 110^ NR (power on), and 215 rpm» or 116^ Np (power off), at li2,000 pounds. The recommended limit operating rotor speed is 19^ rpm, or 10555 NR. Accord- ing to Sikorsky pilots, operation at rotor speeds greater than 105^ produces distinct tail rotor whining, well above ordinary noise levels. Pilots normally accept his signal as a warning to cut back rotor rpm.

In Southeast Asia, flight load recorders measured peak main rotor speeds of 198 rpm and 203 rpm, the former occurring in forward flight at 60 knots and 1+0,3^ pounds, and the latter during hover at 25,009 pounds.

These results raised some question. In autorotative flare, it is possible to attain a rotor speed of 198 rpm, but a hover rotor speed of 203 rpm does not sound feasible. This reading indicates either a data recording error or the possibility of poor ground maintenance rigging procedures between the pilot's collective controls and the engines.

51*

MANEUVER AND GUST LOAD FACTOR COMPARISONS

As mentioned above in discussing the CH^A design gross weight, the design limit load factor (Nz) is 2.26g. The minimum design load factor as specified in Reference 11 is -0.5g- During the data recording program in Southeast Asia, the cranes operated well within the design load factor limits, as shown in the V-Nz diagram (Figure 21). Approximately 91% of the peak load factors recorded during maneuvers were experienced at gross weights at or below 29,000 pounds, indicating that no payload was being carried at the time. None of the Nz occurrences recorded at gross weights greater than or equal to 37,000 pounds exceeded 1.2g in forward flight. The minimum measured load factor was 0.5g in forward flight, well above the design minimum.

Because the basic use of the crane was to pick up and release heavy loads with the single-point winching system, load factor measurements were taken during what was termed the "hoist" mission segment. This type of operation produced a peak Nz of 1.88g, again below the maximum limit load factor.

2.0

1.8

1.6 N

z *

s 1.4 1- o 5 1.2

o < 1.0 o -1

H .8 (/) 3 <D

.6

.4

.2

CH-54A STATIC DESIGN 7 G.W. =42000 LB —

REF.

OPERATIONAL ^ FLIGHT LOADS —'

(ALL GROSS WEIGHTS) REF I

40 60 80 TRUE AIRSPEED, KN

100 120

Figure 22. Gust V-N Diagram, z

55

During data editing, gust load factors were distinguished from maneuver and hoist load factors. As indicated by the gust V-Nz diagram (Figure 22), maximum gust peaks were well below the maximum design gust load factor of 1.93g. In addition, the 1.93g design point was applied at a 142,000-pound gross weight and 110 knots, well above the 25,000-pound, 100-knot condition for which a peak gust of 1.3g was measured.

PREDICTION OF LOAD OCCURRENCE FREQUENCY

There are seven helicopter design load parameters for which the frequency of occurrence of peak values must be predicted: maneuver load factor, gust load factor, airspeed, gross weight, rotor speed, horsepower, and main rotor blade flapping. Each parameter is affected by aircraft configuration (weight, drag characteristics, center of gravity, and installed power), payload configuration (drag characteristics, size, single-cable slung loads, or four-point payloads), and aircraft flight profile. These effects govern the magaitude and frequency of occurrence of peak values. Data from TR 70-73 have been employed to establish typical operating limits.

Maneuver Load Factor

Extrapolation of load factor exceedance curves presented in TR 70-73 provides a practical approach for extending the load factor data beyond 200 hours. As a valid sample extrapolation, 5000 hours is selected, in line with 5000-hour replacement times for major dynamic components. Figure 23 indicates that, for a composite load factor spectrum, a peak load factor increment of l.lg, or a total load factor of 2.1g, would be experienced within 5000 hours.

A similar composite load factor exceedance curve, presented in TR 66-58 (Reference 12), provides similar data. In that study, data were collected during 110.38 hours of military operations in the Fort Banning, Georgia, area. Extrapolation of these data to 5000 hours results in a peak load factor of 1.97-

These two extrapolations of technical report sample data support the assump- tion that a peak load factor of 2.0g, in combination with the design gross weight configuration, would provide a conservative approach to fatigue design. In addition to extrapolating to 5000 hours, the use of a composite spectrum, which includes all gross weights and the hoist regime load factor occurrences, provides additional conservatism. For airspeeds above UO knots (IAS), load factor recordings for payload configurations did not exceed 1.1+g to 1.5g. The highest load factor recorded was 1.88g, during the hoist regime. However, these hoist load factor pulses are normally of short duration.

A realistic approach to establishing maneuver load factors for fatigue should depend largely on the type of payload. A moderate peak load factor of 1.8g would be employed for configurations involving single-point slung loads. A peak load factor of 2.0g would be applicable to four-point, rigidly attached payloads, such as a pod. When no payload is carried, a

56

10000

5000

1000

500

z < a a H 100

50

x o <

DATA FROM TECHNICAL REPORT 70-73 (REF I

O POSITIVE ANZ SAMPLE NO. I • NEGATIVE ^N z SAMPLE NO I

a POSITIVE AN z SAMPLE NO 2

■ NEGATIVE AN z SAMPLE NO 2

RECORDED DATA

EXTRAPOLATED DATA

4 6 8 10

MANEUVER LOAD FACTOR, AN2

12

Figure 23- Prediction of Maneuver Load Factors Above 200 Hours.

peak load factor of 2.2g would be applicable. A peak load factor of 2.0g would be employed during pickup and release of payloads.

Gust Load Factor

Gust load factor data from the 200-hour samples of TR 70-73 can be extrapolated validly to 5000 hours. Weather phenomena are assumed to follow a straight-line log-normal distribution. Figure 2h indicates that, for a composite load factor spectrum, a peak gust load factor increment of .hkg, or a total load factor of l.UUg, would be experienced within 5000 hours. A similar composite exceedance curve, presented in TR 66-58, provides similar results when extrapolated to 5000 hours. A peak gust load factor of l.h5g is derived from these data. These results suggest that the probability is extremely remote that applied loads from gust response would ever exceed maneuver loa ^.

57

10000

»00

1000

500

o UJ ulOO u X UJ

50

x 8

o (-

CE D

DATA FROM TECHNICAL REPORT 70-73 (REF I )

O POSITIVE ANZ

• NEGATIVE AN;

□ POSITIVE ANZ

■ NEGATIVE UN2

SAMPLE NO. I

SAMPLE NO. I

SAMPLE NO. 2

SAMPLE NO 2

RECORDED DATA EXTRAPOLATED DATA

0 2 4 6 B 10 1.2 GUST NORMAL LOAD FACTOR, ANZ

Figure 2h. Prediction of Gust Load Factors Above 200 Hours.

Airspeed

Data from TR 70-73 support a realistic airspeed of 100 knots for aircraft configurations involving single-point slung loads. Figures 8 and 9 indicate that 99% of the time in which aircraft gross weights were greater than or equal to 37,000 pounds, airspeeds were 90 knots (TAS) or less. Future crane helicopter fatigue designs should incorporate these facts.

Peak airspeeds for future fatigue designs should realistically depend on the handling quality characteristics of specific externally carried payloads. Payload classes should include a single-point or multipoint carried vehicle, artillery piece, container, or netted cargo, a rigidly attached container or medical pod, and no payload at all.

58

For example, an approach to development of a speed spectrum for single- point slung loads would be as follows:

a) 99% of the time for slung load configurations is distributed at or below 100 knots. (Figure 9 presents a typical distribution.)

b) The remaining 1^ is distributed T5$-25$ between 110 knots and 120 knots, respectively.

In considering four-point, rigidly attached loads, handling quality characteristics of the particular payload would be combined with drag and pow'ir characteristics to establish a peak airspeed for fatigue analysis. Configurations without payload would be restricted to the limit dive speed established by static design.

Gross Weight

For a fatigue design effort, a three-segment breakdown is sufficient for gross weight distribution. Using the gross weight categories established in the "Operational Mission Profiles" section of this report, the following time distribution was derived from TR 70-73:

% Time Gross Weight Sample 1 Sample 2

<: 29000 lb 60,8 60.h

g 29000 lb but < 37000 9.3 10.3

g 37000 lb 29.9 29.3

These results provide reasonable support for a 60% - 10% - 30% distribution of time for aircraft weights of 70^, 85^, and 100^ of the design gross weight, respectively.

Main Rotor Speed

Since normal variations of main rotor speed are well within the endurance limit of rotor speed related components, 100^ rotor speed for 100^ of the time is reasonable for development of fatigue spectra. For such components as main rotor and main gearbox bearings, however, for which designs are affected by a rotor speed spectrum, the following distribution is offered, based on data from TR 70-73:

% Rotor Speed % Time

08-1025? 85%

103-1065? lU.9^

106-11055 0.15?

59

Horsepower

There axe three essentials for the development of horsepower spectra: mission requirements, engine capability, and rotor performance. Also, rotor performance is dependent on main rotor blade characteristics and aircraft flat plate drag. The variability of these parameters from one aircraft design to another makes prediction of frequency of occurrence inapplicable for horsepower. However, fatigue design capabilities of power-related components often can be assessed quickly by visual inspection of the mission requirements. The extent of hover time, sideward flight time, and high-speed time, coupled with preliminary power data, can aid in establishing peak power requirements in the early stages of fatigue design.

Main Rotor Blade Flapping

A main rotor blade flapping spectrum impacts significantly on the fatigue design of main rotor head dynamic components and main rotor shaft. A peak main rotor blade flapping angle(p) is established, for example, by control limits. A straight-line log-normal distribution of flapping for maneuvers between the peak value and the level flight value is validated by correlation with data from the CH-53A 25-hour flight test substantiation program (Reference 13). This assumes a .OOOljS probability of occurrence of the peak flapping angle. Level-flight blade flapping values of 3 to k are established by Sikorsky flight test.

60

CONCLUSIONS

1. Hover time during combat operations is approximately one-third to one-half the time currently used for crane helicopter design spectrums.

2. Ascent and descent times during combat operations are approximately four times greater than used for current crane helicopter design spectrums.

3. The level of command from which crane helicopters are currently deployed impacts on normal crane operating gross weight categories . Crane helicopters are rarely employed for missions for which lower gross weight transport helicopters can be substituted.

U. Maximian operating gross weight limits are often deliberately exceeded when it is necessary to transport special items in the Army inventories; for example, trucks, armored vehicles, and artillery pieces.

5. Airspeeds above 90 knots are rarely associated with an externaJ. payload configuration.

6. Most aircraft flight time occurs in a density altitude range of 2000 to 5000 feet.

7. In forward flight, load factors rarely exceed 1.3g for external payload configurations.

8. During the hoist regime, load factor peaks are of significant magnitude compared with typical maneuver load factors.

9. Future Army helicopter designs will benefit from improved data collection and editing techniques.

61

RECOMMENDATIONS

INTRODUCTION

Future Army helicopter designs will benefit from improved data collection and editing techniques. Better definition of discrete ground and flight regimes is required to develop accurate mission profiles. Consideration should be given to development of a composite operational spectrum based upon a combat environment and on peace-time operation. Knowledge of peak loads and specific load parameters, such as main rotor head moment or main and tail rotor flapping angles, would yield more accurate fatigue load prediction.

DATA COLLECTION

Thorough data collection is required to establish discrete flight conditions from the recorded data. The four-segment breakdown (ascent, descent, steady state, maneuver) should be replaced. The current procedure does not provide enough information to derive specific conditions: hover, hover turns, sideward and rearward flight, ascent, descent, etc.

DATA EDITING

The first phase of data editing should begin in the field. Review of raw data after each fli^-rt, accompanied by pilot debriefing, would promote accurate correlation of recorded flight loads with specific flight conditions. An effort should be made to define the low-speed range of conditions, such as hover, hover turns, sideward flight, and transitions. Current methods combine data below ^0 knots into one category. Assump- tions must then be formulated to reduce the data to the desired format.

The method for editing maneuver load factors should be revised to include time during banked turns. The low number of maneuver occurrences reported in TR 70-73 indicates the absence of low-load-factor turns below the 1.2g editing limit.

DATA PRESENTATION

Mission profile analysis would benefit significantly from revised procedures for the grouping and presentation of operational flight loads data. Meaningful combinations of recorded data would be valuable in developing mission profiles and related fatigue spectra.

For example, one specific load parameter, such as rate of climb, should be cross-plotted against time for each important load parameter:

Load Factor (N ) Airspeed Main Rotor Speed Gross Weight Altitude Horsepower

62

A distribution of each of these parameters could be developed individually as well as collectively, since each would have a common reference. The rate-of-climb parameter helps to establish a general flight regime: ascent, descent, or steady state. A detailed cross-plot with airspeed would establish the precise flight condition within these three flight regimes. Besides load factor exceedance curves, similar plots for airspeed and main rotor speed would be helpful. The airspeed plots would provide a general distribution of airspeeds for cruise and maneuvers, and the rotor speed plots would provide valuable information for rotor and gearbox bearing design. Information on load attachment methods would be helpful, whether single-cable slung loads, hard-point connections, shade roller connections, or no payload.

The availability of specific operational hardware load parameters would prove beneficial in establishing operational trends. Valuable information would be gained for future main and tail rotor fatigue designs if data were available for main rotor blade flapping, tail rotor flapping, tail rotor collective, and tuil rotor thrust. Ideally, these data would be presented along with rate of climb and airspeed.

FUTURE STUDIES

An operational mission profile, such as that developed from TR 70-73, does not provide an absolute design tool. Certain limiting conditions, which were prevalent during Southeast Asia operations, affected the results. For example, hover time was not typical for universal crane operations. The level of command from which aircraft were deployed restricted gross weight operating ranges. The evasive flight tactics adopted for combat operation may have impacted on flight procedures.

Additional analytical studies are recommended to examine the need for a weighted composite design spectrum that would encompass the full use of the CH-5^A in U. S. Army military opera xons. It would add accuracy and realism to future crane helicopter designs.

63

LITERATURE CITED

1. Giessler, F. Joseph, et al., FLIGHT LOADS INVESTIGATION OF CH-51+A HELICOPTERS OPERATING IN SOUTHEAST ASIA, Technology Incorporated; USAAVLABS Technical Report 70-73, Eustis Directorate, U. S. Army Air Mobility Research and Development Laboratory, Fort Eustis, Virginia, January 1971, AD 881238.

2. Naval Air Systems Command, STRUCTURAL DESIGN REQUIREMENTS (HELICOPTERS), Department of the Navy; AR-56, February 17, 1970.

3. Sikorsky Aircraft, FLIGHT TESTS FOR STRUCTURAL SUBSTANTIATION OF THE S-61+ HELICOPTER INCLUDING PERFORMANCE AND VIBRATION MEASURE- MENTS, SER-61*095, April 9, 1961*.

k. Sikorsky Aircraft, ADDITIONAL FLIGHT TESTS FOR FAA CERTIFICATION OF THE S-6UA HELICOPTER, SER-6i*137, April 15, 1961+.

5. Sikorsky Aircraft, STRUCTURAL SUBSTANTIATION OF THE S-61*E MAIN ROTOR HEAD AND CONTROL SYSTEM FOR FAA CERTIFICATION, SER-6l4l*70, May 7, 1969.

6. Sikorsky Aircraft, EXTENSION OF FAA V CERTIFICATION OF THE S-6UE, SER-6I+I471, May 19, 1969. ne

7. Sikorsky Aircraft, FATIGUE SUBSTANTIATION OF THE YCH-5l+A/S-61*E MAIN ROTOR HEAD CONTROL COMPONENTS BASED ON LABORATORY AND FLIGHT TEST DATA, SER-61*10li, March 12, 1965.

8. Sikorsky Aircraft, S-6I4 MAIN ROTOR WHIRL/FATIGUE TEST RESULTS, SER-61*561+, July 6, 1971.

9. Sikorsky Aircraft, FATIGUE SUBSTANTIATION C S-61iA/S-61»E AUXILIARY ROTOR ASSMBLY, SER-61»103, March 29, 1965.

10. Sikorsky Aircraft, FATIGUE SUBSTANTIATION OF CH-53A ROTARY RUDDER OUTPUT SHAFT, HUB, AND BLADE RETENTION COMPONENTS, Test Report, SER-65076, November 26, 1965.

11. Sikorsky Aircraft, STRUCTURAL DESIGN CRITERIA REPORT - MODEL S-6I4A HELICOPTER, SER-6I1O9I-I, May 16, 1966.

12. Braun, Joseph F., and Giessler, F. Joseph, CH-5I+A SKYCRANE HELICOPTER FLIGHT LOADS INVESTIGATION PROGRAM, Technology Incorporated; USAAVLABS Technical Report 66-58, U. S. Army Aviation Materiel Laboratory, Fort Eustis, Virginia, June 1966, AD 63836U.

13. Sikorsky Aircraft, AIRCRAFT MISSION SIMULATION AND FLIGHT LOAD MEASUREME.'T PROGRAM, Test Report, SER-65I+I7, May 2k$ 1968.

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